Study of an advanced General Aviation Turbine Engine (GATE)

TABLE of CONTENTS Section,: Title Pa e , ...

Summar'y .. 1 . . 3 II Iatroduat? on. .

IzI General.. Approach...

Market Analysis 6 Broad Scope Trade-off Studies . . . _. 6 Evaluation "of a Common Core Concept . . . . . 6 , 6 .'

Technology Program Flan.....'

Discussion of Results._ 7 IV ,. ..

4a Market Analysis . ... . . . 7 Approach. ..: . 7.

Market Projections. • . . . . . .

. 12 GATE Potential Categories_ . . . . . . . . . . 13 Engine Production Forecast.. . . . . . . 20 Market. Anal.ysis..Conclusions.. . 20 4b Broad Scope Trade-off Studies 24 Technology Forecast far Market Analysis 24 ( Effect of Engine Size on Weight and Pei•- formance. . . . . . . . . . 24 Effect of Engine and Production.

Size Quantity on Cost. . . . . . . . . . 27 Turboprop Engine Trends for Analysis. . . . 27 ^- Turbashaft Engine Trends. . . . . . . . . 28.

Cost Performance Trade Studies. . . . . . . . . 31 Approach .. . . . . 31 31 Mission Requirements. . . . . . . 32 ^.'

Airframe Characteristics. . . . . . . . . . 32 . . . . . . . .

Economic Model.. . . 34 E ngine Cost . . . . . . . . . . . . . . . . 36 Engine Maintenance. . . . . . . . . . . . 44 Current Technology (Baseline) Engine.

Study Engines . . . . . . . . . . . . . . . 51 Dual- and Single-Stage Compressor Engine i .

Matrix. . . . .. . . . 53 Performance and Gast Comparison of Dual and Single Stage. . . . . . . . . . . . 54 Mission Evaluation -f Dual Stage. . . . . .

Mission Evaluation of Single Stage. . . . . 63 Candidate Engines . . . . . . . . . . . . . 66 } . . 66 Engine Component Comparisons. . . . . .

Candidate Engine Comparison . . . . . . . 69 ATE Versus CTE Study Results. . 73 i Turbine Engine versus Piston Engine . . . . .

Trade-Offs . . . . . . . . . . 78 Technology Elements . . . . . . . . . . 81 Sensitivity Studies . . . . . . . . . . . . 89 Environmental Considerations. . . . . . . .

- Noise . . . . . . . . . . . . . . . . . . . 96 ^ Emissions 103 i Summary 105 k i Sect ion Title page .

4c Evaluation of .a Common Core Concept . . .

. . 106.

t € Approach.

106.

Engine Configuration.

I05 Power Range .

. .,... ... . . . . . . . . 108 4d Technology Program Plan . . . . . . .

Y v Summary of Results. .. .

Market Analysis ... I13 ', ..:.

Broad. Scope Trade-off Studies . . . . . . .

Evaluation of a Comon . Core . Concept :. , iii Technology Program Plan Appendix . . . . . . . . . . .

. .

' GATE.: Cost i Analysis Computer `Program. . .

. . . . . .

'. 119'' r; LJ i

ii

Interre:lationshi-oof vc:icle.system parameters and cost of ownership. ....

2 GATE Task I - methodology.

• • `• . 7 4 3 Aircraft comparison. ... . . . . . . . 14 GATE-category V--high-performance, single-engine; retract-.` . 15 able. gear., aircraft . .

.

.

5 GATE cate.go^ry.VI--utility, single-engine aircraft.

b GATE category VII ^-singX a engine aircraft aerial .appli-- . .. .,.

G a - '..: ', '. , '.. .. cation • • '. • • ^ a 7 GATE category VIII--single--eroine aircraft .aerial . appli- cation . . . . .

. • . -. . . 17 .• . " . ", . - ."

•. Single-engine, fixed-wing aircraft production. forecast

summary.

, 17 • .. ..

9 GATE category . X---twin-engine, cabin class, unpressurized } aircraft . . . . . . ..

. . ... . . , .. . ... ._ . . . . . . 18 GATE category Xl--twin-engine, pressurized aircraft under 4082. kg ( 9000 lbm) . a :.:.. , ..

1Z GATE category XII--twin-engine aircraft over 4082 kg (.9000 lbm) Z2, . Twin-engine, • . fixed--wing aircraft production forecast summary.

GATE rotary wing aircraft category I--ultralight (recip) -! 14 Rotary wing aircraft production forecast summary : . . . . . 21

Turbine engine market forecast---1988 . . . . . . . 22

16 Turboshaft and turboprop engines . sfc sensitivity to scale size relative trend at slss T.O. .

. . 25 Turboshaft and turboprop engines specific mass sensitivity to scale size relative trend at slss T.O.. . . . . . . . 25 Turboshaft and turboprop engines specific cost sensitivity J to scale size relative trend at slss T.O.. . . . . . . .

Turboshaft and turboprop engines production volume power size impact on cost relative . trends. . . . .

. . . . 27 Turboprop sfc trends . . . . . .

. . . . . . . . . . . . .

Turboprop mass trends . . . . . . . . . . . . .

. . . . . . . 29 Turboprop cost trends . . . . . . .

. . . . . . . . . . . . .

Turboshaft sfc trends .

. . . . . . . . . . . . . . 30 24 Turboshaft mass trends . . . . . . ... . . . .

. . . . . . . 30 Turboshaft cost trends . . . .

. . . . . . . . . . . . . . . 31 fixed-wing missions . . . .

27 .

. . . Rotary-wing missions . . . . ... . . . . . . . 33 Log-log linear learning curve. .

. . . . . . . . . . . 39 Comparative average manufacturing --CTE and ATE type costs { ( turbine engines . .

"s . . . .

TBO and PRR relationships--100 A/C, 15--yr forecast 31.

GATE parametric engine study---two-stage. centrifugal com- ressor (engine sfc) . . . .

. . . . . . . . . . . . . . 54 p GATE parametric engine study--two-stage centrifugal com- pressor (engine mass) . . .

. . . . . . . . . . . . . . 55 GATE parametric engine study--two-stage centrifugal com- pressor (engine cost) . . . . . . . . . . . . . . . . 55

Figure

Title Page

GATE parametric engine study---on e -stage centrifugal com-

.

pressor (engine sfc) .'". .'. . :. . .....:. . . . . . . 56

35 GATE parametric engine study--one-stage centrifugal com- ,..

..

:pressor (engine.mass)s . . ..... . .. 56 36 GATE parametric engine study -one stage'c entrifugal.com- pressor (engine cost). ".' . . :: . . . . 57 ' Dual-stage centrifugal compressor'engine matrix--gross mass,

shaft-power and- total :. cost of.. and cash flow : is

ownersiiip: ..

requirement ` trends q (unpressurized twin). 58.

.".. I

.38 Dual-stage.. centxifugal..compress.or . engine matrix=-direct

operating cost' total cost of ownership, and cash flow ;.

requirement trends -(unpressurized twin). 59 g . `.' `.

39 Dual -stage centrifugal compressor engine matrix--gross mass, L'

total .

aircraft .cost, and . total cost of ownership trends (light twin) . . . . . ... . . . . . . .

. . . 59 G Dual-stage centrifugal compressor"engine matrix---gross mass, La total. aircraft cost, and total cost of ownership trends , . 6 ., ..

. .

(heavy twin) . .: . . `. ... ..

Dual-stage centrifugal compressor engine matrix--gross mass, total aircraft cost, and total cost of ownership trends (light agricultural) . . . . . . . . . . . 61 . . . . .

Dual-stage centrifugal compressor engine matrix-=gross mass, total aircraft cost, and total cost of ownership trends . .

(helicopter-light single). . . . . . . . . . . . 61 L 43 Dual-stage compressor engine matrix--gross mass, total air- craft cost, and total cost of ownership (helicopter- light twin) . . . . . . . . . . . .

. . . . . . . . . 62 44 Single-stage centrifugal compressor engine matrix---grass mass and total cost of ownership trends (unpressurized trend). . . . . . . . . . . . . . . . . . . . . .

. . 64 45 Single-stage centrifugal compressor engine matrix--gross mass and total cost of ownership trends (heavy twin) 65 46 Single-stage centrifugal compressor engine matrix--gross mass and total cost of ownership trends (helicopter- light . . . . . . . . . . . .

twin) . . . . . 65 .• 47 GATE candidate engines—gross mass trends (unpressurized . . .

twin) . . . . . . . . . . . .. . . . . . . 69 48 GATE candidate engines--total cost of ownership trends (Unpressurized twin) . . . . . . . . . . . . . . 70 GATE candidate engines--gross mass trends (heavy twin) . . . 70 50 GATE candidate engines--total cost of ownership trends .

(unpressurized twin) . . . . . . . . . . . . . . . . . 71 51 GATE candidate engines--gross mass trends (helicopter- light twin) . M . . . . . . 71 . . . . . . . . . . . . .

. 52 GATE candidate engines--total cost of ownership trends (he.licogter-light twin)... . . . . . . . . . . . 72 Gross mass comparision--best advanced technology engines versus current technology engines. . . . . . . . . . . . 74 - 54 Total cost of ownership comparison--best advanced technology engines versus current technology engines. 75 55 Direct operating cost comparison---best advanced technology engines versus current technology engines. . . . . . . . 75 iv T Figure Title Page Total aircraft cost comparison--best advanced technology - engines versus current technology engines.

. . . . . . . 76 Mission fuel comparison--best advanced technology engines versus current technology engines. . . . . . . .

. . . . 76 Cross mass and total aircraft cost comparison---advanced gas turbine versus piston (unpressurix.ad twin) . . . . . 79 Gross mass and total cost of ownership com;,Ar•ison--advanced gas turbine versus piston (helicopter--light twin).

. . . 80 T,.;tal aircraft-- cost versus ATE to NAP engine price ratio (unpressurized . . . . . . .

twin) . . . . . . . . . . . 81 Engine sfc, mass, and cost sensitivity data---CTE (unpres- surized .

twin) . . . . . . . . . . . . . . . . . . .

Engine sfc, mass, and cost sensitivity data---CTE (heavy twin) . . . . . . . .

. . . 90 Engine sfc, mass, and cost sensitivity data---CTE (hell:_ copter-twin) . . . . . . . . . . . .

. . . . . . .

Engine sfc, mass, and cost sensitivity data--ATE (unpres- surized twin) . . . . . . . . . . .

. . . . . . . . . . .

Engine sfc, mass, and cost sensitivity data---ATE (heavy twin). . . . . . . .. . . .

. . . . . . .

Engine sfc, mass, and cos; sensitivity data--ATE (heli-- copter-twin) . . . . .

. . . . . . . . . . . . .

. . . . 92 Total cost of ownership sensitivity to use and fuel cost (unpressurized twin) .

. . . . . . . . . . . . . . . . . 93 Total cost of ownership sensitivity to use and fuel cost (heavy twin) . . . . . . . . . . . . .

. . . . . . . . . 93 Total cost of ownership sensitivity to use and fuel cost (helicopter-twins) . . . . . . . .

. . . . . . . . . . . 94 Component efficiency, cooling air, and leakage sensitivity data (unpressurized twin) . . . . . . . . . . .

. . . . . 95 Component efficiency, cooling air, and leakage sensitivity data (heavy twin) . . . . . . . . . .

Component efficiency, cooling air, and leakage sensitivity data (helicopter-twin) . . . . .

. . . . . . . . . . . . 96 Comparison of initial and current production engines at takeoff power . . . . . . . . . . . . .

. . . . . . . . . 97 Turboshaft engine exhaust noise as a function of sfc . . . . 98 Inducer tip Mach number for single- and dual--stage centri- compressors.

fugal . . . . . . . . . . . . . . . .

. 99 A weighted sound power level for single- and two-stage centrifugal compressors . . . . . . . . . . . .

. . . . . 99 Narrow band frequency spectrum for the flyover noise of a typical single-engine, general aviation aircraft .

. . . 100 Influence of propeller section properties on the observed noise..

. . . . 101 Certification noise levels for small propeller-driven aircraft .

.. . .

Passible raise Iimit for type certification of helicopters 103 Impact of EPA emissions rule making on GATE study. . .

. . . 105 Effort of rated rotor inlet temperature. . . . . . . . . . . 109 83 Task IV---technology plan . . . . . . . . . . . . . . .

. . . 111 84 Specific .

cost trends . . . . . . . . . . . . . . . . . . .

Small gas turbine technology lag . . . . . . .

. . . . . . . 117 v Fi gure Title _e GATE turbashaft and turboprop engine scaling ratios for engine-s p ecific and maintenance costs. . . . . . . . . . 122 87 Cost curve —' iixed-wing aircraft. • . . . • . . . . . . . • . 122 88 Helicopter acquisition coat . . . . . . . . . . . . . . . . . 123 vi LIST OF TABLES Table Title 21&e I Forecast GATE Categories —Fixed Wing - Propeller Driven. . . . 8 1I GATE Forecast Categories —Fixed Wing - Thrust Driven . . . . . 9 III . . . . . . 9 Forecast GATE Rotary Wing Categories . . . . . .

IV Estimated Active General Aviation Aircraft by Type of Ai rcraft . . . . . . . . . . . . . . . . . . . . . . . . 10 V Eng . . . . . . . . . . . . . . . . . . . 11 ine Domain . . . . .

C^ VI Single - Engine Fixed Wing Aircraft U.S. Production Quantities (1975 to 1988) . . . . . . . . ..

Twin-Engine Twi-Engine Fixed Wing Aircraft U.S. Production Quantities . . . . . . 13 (.1975 to 1988) . . . . . . . . . . . . . . .

VIII Rotary Wing Aircraft U.S. Production Quantities ( 1975 to 1988) . . . . . . . . . . . . . . . . 13 Ix - - General Aviation 1988 Market Forecast Turbine Engine. . . . 22 X Missions Identified in Market Survey . . . . . . . . . . . . 23 XI . . . . .

Description of Reference Aircraft . . . . . . . . . 34 XII Economics of Aircraft Ownership . . . . . . . . . . . . . . . 35 .

XIII Costing . . . . . . . . . . . . . . . . . . . . . . . . . . 35 XIV Material Indices for MIF Technique . . . . . . . . . . .

. . 36 XV T63 Engine Cost History. . . . . . . . . . . . . . . . . . . 38 KVI Candidate Engine Data units)--Unity Size . . . . . . . .

( SI 41 XVII Candidate Engine Data ( custo m ary units)--Unity Size. . . . . 42 XVIII "Optimum" Engines. . . . . . . .. . . . . . . . . . . . . . . 45 XIX Turboshaft Maintenance Standards . . . . . . . . . 45 . . . . .

XX Turboprop Maintenance Standards . . . . . . . . . . . . . . .

M Schedule Overhaul Recommendations. . . . . . . . . . . . . . 48 XXII Projected Premature Removal Rates . . . . . . . . . . . . . .

XXIII Maintenance Plan . . . . . . . . . . . . . . . . . . . . . . 49 XXIV . .

NASA GATE--Baseline Maintenance Input. . . . . . . . . . 50 XXV Price Structure. . . . . . . . . . . . . . . . . . . . . . . 50 XXVI Current Technology Engine Data ( SI units) . . . . . . . . . . 51 XXVII Current Technology Engine Data (customary units) . . . . . . 52 XXVIII Overview of Matrix Engines ( turboshaft and turboprop) Unity Size: Approx 615 kW ( 825 hp) . . . . . . . . . . . . . 53 .

XXIX Summary of Cycle Selections . . . . . . . . . . . . . . . . . 62 XXX Current Technology Engines Versus Advanced Technology Engines Gross Mass Comparison . . . . . . . . . . . . . . 63 XXXI 66 Cycle Selection One-Stage Centrifugal Compressor Matrix.

XXXII Axial versus Radial Inflow EPT Configuration Study Results 67 XXXIII Cooled versus Uncooled Turbine Study Results . . . . . . . . 68 XXXIV Single - Stage versus Dual-Stage Centrifugal Compressor Study Results . . . . . . . . . . . . . . . . . . . . . . . . .

XXXV Installation Effect on Shaft Specific Fuel Consumption . . . . . . . . . . . . . . . 72 (unity size engine) . . . .

XLXXVI Engine Comparisons ( Turboprop) . . . . . . . . . . . . . . . 73 XXXVII Engine Comparisons 73 ( Turboshaft) . . . . . . . . . . . . . . .

XMOV I I I Percent GM Reduction ( CTE as reference gross mass) . . . . . 74 XXXIX Percent TCO Reduction ( CTE with production base as re- ference . . . . . . . . . . . . . . . . . . . . . . . . . 74 XL Unpressurized Twin Summary--Use = 600 h / yr and Fuel . . . .

$0.22/ ($0.83/gal) . . . . . . . . . . . . . . 77 vi i i Li Table Title Page XLI Unpressurizad Twin Summary--Continued DOC Breakdown-- 78 $/fl h (util ° 600 h/yr and fuel = $0.22/ ($0.83/gal)).

XLII 79 Engine Comparisons---Turboprop versus Piston .. . . . . . . .

82 XLIII Engine Comparisons--Turboprop. . . . . . . . . . . .

XLIV 83 Dual Property Turbine Evaluation ( , .npressurized twin). . . .

XLV 84 Engine Comparisons--Turboprop . . . . . . . . . . . . . . . .

84 XLVI Ceramic Turbine Stator Evaluation (unpressurized twin) . . .

XLM 85 Engine Comparisons--Turboprop. . . . . . . . . . . . . .

86 XLVIII Advanced Composite Gearbox Evaluation (unpressurized twin) .

IL GATE Turboshaft and Turboprop Engines--Unity Size 373 kW (500 hp) Candidate Engines--Design Point Data. . . . . .

L 112 GATE Technology Plan . . . . . . . . . . . . . . . . . . . .

LI 114 Task II "Optimum!' Engines . . . . . . . . . . . . . . . . . .

LIZ Relative Performance and Cost---Advanced vs Current Tech- 115 nology 373 kW (500 hp) Gas Turbine Engines . . . . . . .

L Cost Parameters . . . . . . . . . . . . . . . . . . . . . . .

YITITI7 121 LIAircraft Specific Costs. • . . . . . . . . . . . . • . . . .

131 LVSample Problem Input . . . . . . . . . . . . . . . . . . . .

132 LVIComputer Output Order . . . . . . . . . . . . . . . . . . . .

r viii I. SUMMARY The objective 'of the General Aviation Turbine Study (GATE) was to iden best technology program for a small., economically Vable gas turbine engine applicable to the general aviation helicopter and aircraft market for 1985-1990. The study considered turboshaft and turboprop engines in the 112 to 746 kW (150 to 1000 hp) range and turbofan en gines up to 6672 N 050C lbf) CD thrust.

The scope of the effort encompassed five tasks: Task I Market Analysis Task II Broad Scope Trade-off Studies Task III Evaluation of a Common Core Concept Task IV Technology Program Plan Task V Reporting Based on the 4% annual growth rate anticipated within the next 10 years in general aviation aircraft production, a good market for new turbine engines was predicted for 1988 providing aircraft are designed to capitalize on the advantages of the turbine engine particularly its light weight and compact size. The 1988 turbine engine market was found to be essentially equal in terms of dollar volume in all power classes; however, the greatest impact on the general aviation field for a GATE engine exist under 447 kW (600 hp) be- cause no new technology engines are currently planned in this class. Penetra- tion of the fixed-wing aircraft market is required for attractive engine quan- tities. Turbine engines can achieve greater penetration into the general avi- ation market through improvements in cost, preformance, and TBO (time between overhaul). Any successful new design must carefully select- a balance of these parameters to best meet the market needs. No significant market for turbofan e_Rgines under 6672 N (1500 lbf) thrust was found.

Parametric engine families were defined in terms of design and off-design per- formance, mass, and cost. These were evaluated in aircraft design missions selected to represent important market segments for fixed- and rotary-wing ap- plications. Payoff parameters influenced by engine cycle and configuration changes were aircraft gross mass, acquisition cast, total cost of ownership, and cash flow. Significant advantage over a current technology, small gas turbine engines was found especially in cost of ownership and fuel economy for airframes incorporating an air-cooled high-pressure ratio engine. Gross mass reductions of 10 to 20% for the same capability was indicated. Mass savings were reflected in 8 to 20% lower ownership costs. A power class of 373 kW (500 hp) was recommended as the next frontier for technology advance where large improvements in fuel economy and engine mass appear possible through component research and development. The advanced turbine engine appears com- petitive with the piston engine in terms of installed performance and has a large advantage in engine mass; however, a large premium in price for the turbine engine appears inevitable based on manufacturing experience and fore- casts through the late 1980's.

f The technology plan recommended gy p programs in component research applicable to small turbine engine compressors, combustors, turbines, seals, controls, gear- ing, and shafting, including on-going efforts in materials and engine airframe integration. Toward this end, 24 specific program plans Caere described and provided to NASA. it was also recommended that N49A serve as the catalyst to encourage introduction of useful new technology into engine designs through application studies and possibly core demonstrator programs as component R&D matures.

a II. INTRODUCTION The GATE (General Aviation Turbine Engine) study was sponsored by NASA to de- termine possible benefits to general aviation through development of the small gas turbine engine.

The gas turbine engine has made tremendous inroads in certain segments of gen- eral aviation today. Over 75% of new helicopters manufactured in the free world each year are powered by turbine engines, and this percentage is fore- cast to increase steadily as more turbine-powered machines reach the operating fleet.

Turbine engine use in fixed-wing general aviation aircraft is also growing, although the piston engine dominates under 298 kW (400 hp) shaft power. rn the larger multiengine aircraft used for third-level carrier (commuter air- line), corporate/executive, business/utility, and personal flying, turbine engine power is used nearly exclusively. Approximately 800 small turbine en- gines were produced in 1976 in North America to supply the demand for domestic fixed-wing aircraft production. An additional 500 were produced for domestic commercial helicopter production. In addition, nearly equal numbers of .2n- gines were produced to satisfy the requirements of foreign airframe manufac- turers.

General Motors Corporation, Detroit Diesel Allison (DDA) Division, as the worlds leading supplier of small gas turbine engines for helicopters, shared NASA's interest in the project, and was one of four successful contractors selected to perform the GATE study. DDA's experience in this field includes the production of over 10,000 Model 250 engines at its Indianapolis facility.

These engines ranged from 186 to 485 kW (250 to 650 hp) shaft power and accum- ulated over 10 million hours of operation.

The overall objective of the GATE study was to define the most effective pro- gram to develop and demonstrate advanced technologies for small-sized turbine engines for the 1985--1990 general aviation market. This objective is con- sidered essential and timely in view of the expanding market in general avia- tion, the need for energy conservation, the demand for more stringent environ- mental controls, and the desire to keep the U.S. aircraft industry strong and internationally competitive.

One problem addressed in the study was that of achieving a high performance level in small gas turbine engines. In transferring technology demonstrated in the more efficient large engines to engines with much smaller flow paths, the rising importance of minimal tip clearance for the rotating blading and improved seal efficiency are well known. The increased cycle efficiency ob- tainable with high temperature is more difficult to obtain in the small engine because of aerodynamic constraints that tend to allow insufficient passage area for turbine blade cooling air thus limiting maximum cycle temperature.

Engine cost, another prime drYV._r in small engine design was addressed. Air- craft gas turbine engine materials are relatively expensive and difficult to work, resulting in a costly engine. As aircraft size diminishes, powerplant costs become an increasing percentage of the total cost, until at some point, a less costly engine type may be selected. Nor can one excessively over em- phasize cost to the detriment of performance because the additional weight of the engine and fuel required to perform a given mission will react on the air- frame design and reflect in a less efficient and more costly total vehicle.

Trade studies conducted by DDA were designed to identify the most effective cost/performance balance.

Inherently related to engine cost is commonality of core, of components, and of parts. Maximum commonality as a means of spreading costs over a larger production base is sought as a means of reducing engine price. Commonality of the engine core for application as a turboprop, turboshaft, or turbofan engine is considered in the design concept.

Noise and emission requirements for eneral aviation aircraft reflect the eco- nomic impact on the industry and the need far protecting the public from ex- cessive noise and pollution. The need for considering environmental con- straints in a new en gine technology plan was evaluated.

Turbine engines can achieve greater penetration into the general aviation mar- ket if improvements can be made in cost, performance, and TBO (time between overhaul). Any successful new design must carefully select a balance of these parameters to best meet the market needs. It is the purpose of the GATE study to arrive at a conceptual engine design that does this--to identify the tech- nology elements necessary and to describe a technology program plan for gov- ernment sponsorship that will best support the effort.

Complex relationships are involved in properly selecting an engine for general aviation. Engine characteristics, i.e., specific fuel consumption (sfc), power-to-mass ratio (P/M), and power-to-volume ratio (P/V), impact the total vehicle cost of ownership (COO) sometimes more so than the cost of the engine, by affecting the gross mass (GM) of the vehicle for a given mission as well as the operating expenses once usage begins. If the vehicle has high use, oper- ating expenses dominate and engine sfc and TBO considerations are paramount.

With low use, capital costs control and engine initial cost takes on added significance. Noise and emission standards can adversely affect engine and propulsion, performance and cost. Cost criticality of .materials required by the design is another prime driver on engine cost. Fabrication technique is also a strong determinant of cost as is production rate. The mission require- ments impact the sensitivity of the vehicle to engine characteristics; for instance, vehicles with high GM in relation to payload (i.e., long--range and/or high-power aircraft) tend to be the most sensitive to sfc and P/M ratio. Figure 1 shows the interactions of the market needs, the engine and air vehicle characteristics, and the cost of ownership.

The study attempts to analyze all of these factors in a systematic broad scope manner to determine the general, aviation market needs and to arrive at a con- ceptual engine configuration best satisfying these needs. Both of these ob- jectives are supportive to the main objective of determining the most effec- tive technology program plan for government sponsorship.

j MARKET ENGINE VEHICLE VEHICLE NEEDS G CHARACTER I STI CS CHARACTERI STI CS SIZE COST r a PERF --• Airframe Costs Mission $JE,y' GM • SFC Initial + P/M Operating L!D Limits MSTR • Noise GM COO '' • Emissions MJS ^Lra Engine Costs $JkW Shaft .^ Initial • Material Production Rate Power • Fabrication Operating Utilization TE-3347 Figure I. Interrelationship of vehicle system parameters and cost of ownership.

III. GENERAL APPROACH The Gate study was divided into the following tasks: • !--Market Analysis • II---Broad Scope Trade-Off Studies • III--Evaluation of a Common Core Concept • IV--Technology Program Plan MARKET ANALYSIS The market analysis task was structured for maximum aircraft company involve- ment so that the forecasts would reflect the industry needs as accurately as possible. Representative levels of advanced engine performance and cost anti- cipated were provided early in the market research phase to seed projections and establish GATE market impact. In this phase of the program, DDA identi- fied the turbine engine domain and forecast the engine market for turboshaft, turboprop, and turbofan engines by power class and production units to a 1988 scenario. Typical applications and corresponding missions were identified.

BROAD SCOPE TRADE-OFF STUDIES The trade--off studies were begun early in the program to postulate advanced engine characteristics for the market survey. In this task, technology ele- ments and propulsion requirements including noise and emission standards were identified. A parametric study engine matrix was defined and evaluated on a total vehicle cost of ownership basis. These efforts were supported by para-

metric maintenance studies to establish maintenance plans and criteria for the

general aviation market. optimum engines were selected f.or each mission and the potential benefits and were compared with current engines.

penalties

EVALUATION OF A COMMON CORE CONCEPT The potential for engine commonality was investigated to obtain the broadest range of market usage with a single basic engine.

TECHNOLOGY PROGRAM PLAN Conceptional engine desi gns were implied to form a basis of creating a tech- nology program plan to best serve the advanced engine needs of the general aviation market. The detailed plan included engine component and core pro- p rams .

b IV. DISCUSSION OF RESULTS 4a. Market Analysis APPROACH In developing the approach to the market analysis, it was necessary to {ormu-- late a plan consistent with the needs of the overall GATE study. Considera- tion was also given to the availability of data, and emphasis was given to maximum correlation with various members of the general aviation industry. A flow chart of the plan is shown in Figure 2.

As a prerequisite to projecting the market for general aviation engines, it was necessary to forecast the market for the various sizes and types of air- craft. Contacts were made with the three largest manufacturers of fixed--wing general aviation aircraft in order to determine the categories used by the industry in forecasting markets. DDA conducts ongoing studies of the rotary- wing market so these forecast categories were well established. The forecast categories relevant to GATE and the current production models in that category are shown in Tables I, II, and III.

Upon identification of the relevant forecasting categories, historical ship- ment data back to 1470 was accumulated for each category. At this time it was determined that data on shipments on U.S. sales by foreign general aviation

N;L PV*

ILIEI4'TI F5' ^ ;}^'.^ St AL u M AIRCRAFT CATEGORIES ACCUMCLA:E 'NCINE I IISTORIC.IL FORECAST BY IDENTIFY SHIPMENT IDENTIFY POWER CLASS LATE COST DATA E NGI N E D E W/GATE PERFORIIAXCE C;IANS GOALS ANALYSE FI`AL CATA & A1;CRAFT Mi.11KFT IDE\TIFF FORECAST TRENDS POTENTIAL I.[TH CATE -.PPLICATIOT- ?R£L01INARY REVIEW MARKET I;ITH FORECAST AIRFRAME t410 GATE MFCS.

. INDEPENDENT r:'A >[RCRAFT MARRET 7C7 CAST FORECASTS Ire,} t::ITE ENC I %E nREC.AT RY Pu{r ^^r ,:LA>s E TE-3348 Figure 2.

GAPE Task I methodology.

TABLE I. - FORECAST GATE CATEGORIES-FIXED WFING-PROPELLER DRIVEN Category Current aircraft I Trainer 2-4 place Under 112 kW Bellanca Citabria Series (150 hp) Cessna 150/152 Grumman American T-Cat/Lynx Piper Cherokee 140/cruiser Beech Sport 150 Beech Model 77 II Light single 4 place 112-150 kW Cessna Skyhawk/Skyhawl, HP engine fixed gear (150-200 hp) Cessna Cardinal Piper Cherokee 150/Warrior-180/Archer Beech Sundowner 180 Grumman Cheetah/Tiger III Light single 4 place 135-157 kW Beech Sierra 200 engine retractable gear (180-210 hp) Cessna Cardinal RG Mooney VA20 C,F,J Piper Ai-row -Turbo Arrow Rockwell 112B--112 TCA c IV High performance 4-7 place 172-231 kW Cessna Skylane (230-310 hp) single engine fixed gear Cessna 206--Turbo 206 Cessna 207--Turbo 207 Piper Cherokee Pathfinder-235 Piper Cherokee Six 260-300 V High performance 4-7 place 186-324 kW Bellanca 17-30; 3IA, 31ATC single engine retractable gear (250-435 hp) Beech Bonanza F33, V35 1 A36 Cessna Centurion 210, Turbo Centurion Piper Cherokee Lance Rockwell 114 Bellanca Rocket VI Utility 2-6 place 112-231 kW Cess a 180, 185 Skywagon single (150-310 hp) Mault Lunar Rocket MS-210,235 engine Piper Super Cub Bellanca Scout Helio Courier H-295, HT-295 VTI Single engine I place Under 336 kW Cessna AG Wagon AG fixed gear (Under 450 hp) Cessna AG Truck Cessna AG Carryall Piper Pawnee 235-260 Piper Pawnee Brave 285-300 VIII Single engine 1 place Over 336 kW Emair Mai 600 hp Haib 900 hp AG fixed gear (Over 450 hp) Grumman AG Cat A,B, 450 hp, 525 hp, 500 hp Weatherly 2018 Rockwell Thrust Commander 600-800 hp Ayres Turbo Thrust (PT-6) Marsh Turbo Thrust TPE 33I Frakes Turbo Cat t f I TABLE I. -- (CONT) Current aircraft 239-447 kW Beech Model 76 (320-500 hp) Beech Baron B55, B58, B58TC Cessna 337 5kymaster Cessna 310, Turbo 310 Piper Seneca II Piper Aztec F, Turbo Aztec F Ted Smith Aero Star 600, 601B 6/10 432-559 kW Cessna 402B 404 Titan passenger ( 580-750 hp) Piper Navajo C., CR., Chieftain Rockwell Shrike Cessna Pressurized Skymaster 5-10 place 336-634 kW under 4082 kg (450-850 hp) Cessna 340, 414, 421 (9000 Ibm) Beech Duke, Pressurized Baron Ted Smith Aerostar Piper Pressurized Navajo KII Twin engine 597-1268 kW Beech King Air C90, E90, A100, press (800-•1700 hp) B100 Over 4082.3 kg Beech Super King Air 200 p&GS 15 lb.)

( 9000 Piper Cheyenne

1^^L ppOR QUO-11 Rockwell Turbo Commander 690B

0^I

Cessna Conquest Swearinger Merlin IIIA, Merlin IVA Mitsubishi MU-2N MV-2P TABLE II. - GATE FORECAST CATEGORIES--FIXED WING - THRUST DRIVEN Current aircraft Category XIII Light turbo-jet/ 5-8 place Under 22,241N No aircraft currently turbofan (5000 lbf) thrust available Under 5216 kg (11,500 Ibm) GTOM XIV Med light 8-10 place 17,793-26,689N Learjet-24 Series (4000-6000 lbf) Cessna Citation turbojet/turbofan thrust 5216--6804 kg kg(11,500-15,000 Ibm) GTOM TABLE III. -- FORECAST GATE ROTARY WING CATEGORIES Category Current aircraft Under 1270 kg I Ultralight (recip) Enstrom F-28, 280 (2800 Ibm) GTOM Hughes 300 Hiller 12E Light 1270-2041 kg Bell 206 II (2800-4500 lbm)GTOM Hughes 500 III Light utility 2041-3629 kg Bell 222 (4500-8000 Ibm) GTOM Utility 3629-5216 kg Sikorsky S-76 IV (8000-11,500 Ibm) GTOM Bell 204, 212 205, manufacturers was generally unavailable. Since the worldwide general aviation market is dominated by U.S. manufacturers, the ground rule was established with NASA concurrence that only production by U.S, domestic manufacturers be included in the forecast. The forecast includes installed engines only and does not include spares.

Once the historical shipments by category were accumulated, it was possible to study growth trends of each category. Several approaches to the trend analy- sis were taken including use of a least squares method. This method has limi- tations in as much as there were dramatic changes in the 1988 projection de- pending on the starting year. Introduction of a new model in a particular category can cause dramatic changes in the total shipments for the entire category resulting in erratic behavior in the trend analysis. A more realis- tic approach was to study each category individually as it is impacted by varying market conditions.

During this phase contact was also made with the Aviation Forecast Branch, General Aviation Division of the FAA. M is branch has made projections of the general aviation fleet through 1988 as shown in Table IV. Using this data and subtracting historical attrition and exports resulted in a forecast of the new aircraft production required to support the forecast rates of fleetgrowth.

Subsequent contacts within the general aviation industry indicated that the forecasts obtained in this manner are far more conservative than those gener- ally used within the industry.

TABLE IV. - ESTIMATED ACTIVE GENERAL AVIATION AIRCRAFT BY TYPE OF AIRCRAFT (In Thousands) Fixed Wing Balloons.

As of Piston dirigibles, January 1 Total Single engine Muitiengine Turboprop Turbojet Rotorcraft gliders 1972 131.1 109.1 15.5 1.3 1.2 2.5 1.7 1973 145.0 120.4 17.3 1.4 1.2 2.8 1.9 1974 153.5 126.1 18.7 1.9 1.4 3.1 2.3 1975 161.5 131.9 19.8 2.1 1.6 3.6 2.5 1976 168.5 137.5 20.3 2.5 1.8 3.8 2.5 1977* 181.6 1'+7.7 22.2 2.9 2.0 4.1 2.7 1978* 190.5 154.3 23.6 3.2 2.2 4.3 2.9 1919* 196.9 178.9 24.7 3.4 2.4 4.5 3.0 1980* 203.7 163.8 35.9 2.6 3.7 4.6 3.l 4.1 1991* 213.3 170.8 27.5 2.3 4.8 3.3 1982* 226.0 180.2 29.5 s.6 3.1 5.1 3.5 1983* 233.5 195.3 30.9 5.0 3.4 5.3 3.6 1984* 237.3 187.5 31.8 5.3 3.6 5.4 3.7 1985* 243.3 191.5 33.0 5.7 3.8 5.5 3.8 1986* 196.4 34.4 6.2 4.1 250.7 5.7 3.9 1987* 258.7 201.6 36.0 6.7 4.5 5.8 4.1 1988* 267.0 207.1 37.6 7.3 4.8 6.0 4.2 kForecast.

;Iota --An active aircraft must '-lave a current registrati.)n and have been flown during the previous aI=n:ar Jrar.

It should he noted that historical data are estimates.

Further analysis was made using forecasts developed by ether companies inter- ested in general aviation. From these various sources, it was possible to prepare a preliminary forecast of aircraft production through 1988. Concur- rent with this effort analysis of typical mission capabilities of the various categories of aircraft was conducted.

Comparisions were made between some of the existing piston and turboprop en- gines to determine the relative merits of each type. Table V shows a compari- son of three similarly sized engines, a turboprop, a naturally aspirated pis- ton, and a turbocharged piston. It can be seen from this comparison that the turbine engine has considerable advantage in mass, frontal area., and low- turbo- altitude cruise characteristics over either of the piston types. ne prop has a cost and fuel consumption disadvantage over either of the two reciprocating types and an altitude cruise performance disadvantage over the turbocharged piston engine.

TABLE V. - ENGINE DOMAIN Lyc Lyc Allison Manufacturer Model I0-720-AIA TIGO-541-DIA 250-H17B 317 kW (425 hp) 298 kW (400 hp) Power max SL 298 kW (400 hp) Max cruise 5L 224 kW (300 hp) 237 kW (318 hp) 273 kW (366 hp) 237 kW (318 hp) 168 kW (225 hp) Max cruise (20,000 ft) -- Weight 257 kg (567 318 kg (701 Ibm) 88 kg (195 Ibm) Ibm) Shaft sEc (cruise) 764 g/W • s (0.45 764 g/W • s (0.45 1103 g/W • s (0.65 lb/hp-h) lb/hp-h) lb/hp-h) Frontal Area 0.465 m2 (5 (1.8 ft2) ft 2 ) 0.557 m2 (6 ft 2 ) 0.167 m2 T.B.O. 1200 h 300:1 h 1800 h List price $17,500 $29,700 $61,900 Overhaul cost $ 8 ,750 $11,250 $22,500 Curtis-Wright Corporation was contacted regarding rotary-combustion (RC) en- gines. We were advised that work is continuing on an aircraft type RC engine, but no production commitment has been made. Further contacts with airframe manufacturers indicate that they have no current plans to use RC engines in future aircraft. The performance characteristics of RC engines are more simi- lar to piston-type engines than to turbines. Therefore, for the remainder of the market study, it was assumed that rotary-combustion engine penetration, should it occur, would be included in the reciprocating engine forecasts.

Preliminary market forecasts were reviewed during visits to airframe manufac- turers. Helpful information regarding marIcets, forecasts, missions, and po- tential applications for GATE engines was received.

The 1988 market was estimated based on historical data broken down into market segments and judgmentally projected based on the impact of engine quality on aircraft economics. Market forecasts were made on the basis of GATE engines cost with having either a 20% better sfc with no change in cost or a 20% lower no change in sfc as compared to current technology gas turbine engines. The market forecasts were made initially to these ground rules and were repeated during Task III, after definition of the selected engine concepts, with sub- stantially the same results. The selected engine concepts did achieve a 20% PAGE C)RtGtNAL ^jp^^"I'Qf Q OF pOOR reduction in sfc, but with a cost penalty. The cost penalty, however, was judged to be offset by major reductions in mass, volume, and maintenance costs.

improvements in engine sfc or cost were assumed to have a substantially equal effect on aircraft acquisition and ownership costs. Improvement in engine sfc impacts aircraft costs through reduction in gross mass to accomplish a given mission, thus reducing power requirement and engine cost as well as reducing airframe structural mass and cost. Engine cost reduction has no effect on the aircraft or its performance and reduces aircraft cost simply by the reduction in engine cost. GATE trade studies generally show that a change in sfc has greater influence than a change in engine cost on total cost of ownership and that sfc and engine cost have an approximately equal effect on total aircraft acquisition cost using a current technology gas turbine engine as a base. In some applications, engine mass is as much of a driver on ownership costs as engine cost. (Reference sensitivity data in Figures 61, 62, and 63.)

As a result of the airframer reviews, it was not possible to identify a viable market for under 6672N (1500 lbf) thrust turbofan engines. It was felt by the airframe manufacturers that passenger capacity for turbofan aircraft powered by under 6672N (1500 lbf) thrust engines would be insufficient to justify their relatively high initial and operating costs, especially when compared with a more fuel efficient turboprop. In addition, there would be major dif- ficulties in certifying a turbofan-powered aircraft with only one pilot. The avionics requirements for aircraft operating in the realm for sufficient tur- bofan operating economy prohibit a low initial price aircraft.

MARKET PROJECTIONS Following the reviews with the airframe manufacturers, the forecasts for air- craft production through 1988 were revised to reflect inputs from marketing departments of these companies. The resultant forecast for the non-GATE-im- pacted, single-engine production is shown in Table VI. It is noteworthy that the retractable gear categories show the greatest growth. The superior fuel economy of these types when compared to fixed--gear models with similar power and accommodations is a significant factor in this choice.

The utility category shows relatively little growth due to increased foreign competition in the international markets.

The light agricultural category shows a small percentage growth from 1975 as a result of unusually strong sales in that year. When considered from 1976 pro- duction, a 3% g rowth through 1988 is projected.

TABLE VI. - SINGLE-ENGINE FIXED WING AIRCF.AFT U.S. PRODUCTION QUANTITIES (1975 to 1988) Aver ape annu l Class category 75 76 77 78 79 BO dl 82 83 84 85 B6 57 88 grouth(X) I Trainer 3040 2035 2115 2279 2411 2551 2699 2855 3021 3196 3381 3577 3794, 4001 5.3 it Light !fixed gear 3132 4066 3990 3979 4071 4163 4251 4359 4459 4562 4667 .174 4884 4.996 3.6 It( Light retr Rear 1170 1142 1';9 :301 1388 1451 1580 1586 1799 1320 2049 2196 2332 2498 5.0 IV High parf fixed 1832 1590 1853 1903 1955 2011 2067 2125 2184 2245 2708 2312 29177 2439 2.4 gear High per: retr 1079 1719 1874 1988 2066 2184 2293 2407 2521 2653 2786 2925 1017 322 8.8 2aar ut 7ctliry -25 757 905 820 836 937 335 843 851 360 969 918 687 895 116 9II AG--•ender 335 kW 315 569 45 775 Boo 775 775 Boo 810 310 820 320 830 830 1.01 (450 hp) 7111 AG--336 kW 329 392 395 370 390 400 4+1 410 472 444 456 468 48) 894 3.2 1 450 hpY S up i The multiengine non-GATE-impacted production forecast is shown in Table VII.

All multiengina categories show good growth through 1988, with the unpressur- I ized-cabin class being the weakest with some of the traditional customers for this class of aircraft moving up into pressurized aircraft.

The rotary-wing non-GATE-impacted production forecast through 1988 is shoran in Table VIII. The limited capabilities of the ultralight category of helicop- ters is expected to inhibit growth of these types. The light category is ex- j pected to continue on its long-term growth pattern. New models such as the 5^ Bell 222 and the Sikorsky S-76 will spur considerable growth in the light.

utility and utility categories. l TABLE VII. - TWIN--ENGINE FIXED WING AIRCRAFT U.S.

TABLE VIII. - ROTARY 14ING AIRCRAFT U.S. PRODUCTION QUANTITIES (1975 to 1988) Average annual catezory '75 '76 '77 '18 '79 '80 '81 '82 '63 '84 '85 '86 '87 '88 arovth (t) Ultralight 199 258 220 324 324 244 275 260 250 240 240 240 240 240 1.9 Light 447 402 457 491 545 579 590 602 624 674 665 657 660 3.0 Light utility -- -- -- -- 58 39 80 l00 Its 120 120 180 120 200 19.9 utility 166 98 90 100 158 161 176 135 150 150 150 210 260 290 4.4 GATE POTENTIAL CATEGORIES It was determined through the course of the studies, that to make full use of the advantages afforded by turbine engines such as lower specific mass and drag, it was necessary to design the aircraft from the ground up to use tur- bine power. An example of this is shown in Figure 3. The Dornier Skyservant and the GAF Nomad N24 were designed to nearly identical missions with similar levels of airframe technology and power plant size. The Skyservant vas de- signed to use a reciprocating engine, whereas the Nomad was designed for a turbine engine. It can be seen that the disadvantages of turbine engines such as higher cost and sfc can be offset if the aircraft is designed to make full use of the engines advantages.

It was determined through the studies of engine domains and airframes contacts that within the GATE cost and performance capabilities, that the areas with potential impact for GATE engines are those applications in which performance is of primary cons deration and initial cost of equipment is a secondary con- sideration.

Single Engine In the single-engine categories, it was determined that as a result of cost considerations the upper categories V, VI, VII, and VIII, all of which require E OnIGjNA^ PAG ^^"^°^ 13 '^ P Q _ 7- high performance, offered greatest potential for GATE engines. Figure 4 shows the high-performance, single-engine, retractable gear category as it might be impacted by the GATE program. The turbine penetration of the category would be at the expense of both piston-powered singles and piston-powered twins.

j The impact on category VF utility single engine is shown in Figure S. These are generally working class aircraft, frequently requiring 5TOL capabilities.

Turbine engines are beneficial to this class of aircraft as a result of their cold weather starting capabilities, high power-to-mass ratio, and ability to run an a wider range of fuels than reciprocating engines. A GATE program could help domestic airframe manufacturers compete more effectively with the growing number of foreign manufacturers in the worldwide market.

Agricultural aircraft offer another potential application for turbine en- gines. Reliability, performance, noise, and adaptability to a wide range of fuels make turbine engines attractive to agricultural operators. In the lighter category of aircraft, turbine engines can offer the productivity of much larger aircraft while maintaining the handling charactAristics of the smaller aircraft (Figure 6). In the larger AG category, the lack of avail- ability of large reciprocating engines and demands for increased productivity will be major factors in causing turbinization of the fleet. A GAVE program could accelerate this process (Figure 7).

DESIGNED FOR TURBINE POWER DESIGNED FOR REC:P POWER r

4J;F;M-r

Type Domier Skyservont GAF Nomad N24 Power Eye IGSO-540 283 kW (380 hp) Allison 250-8178 298 kW (400 hp) + 5% - 4% GTOM 4014 kg (8850 Ibm) 4173 kg (9200 Ibm) Passengers 2- 12 2 + 16 +3316 Cabin Volume 8.2 m3 (289 ft3) 12.5 m3 (440 ft3 ) +53% 76 m/s (170 mph) +131/0 Cruise Speed 85.8 m/s (192 mph) T. 0. 15 m (50 ft) 445 m (1460 ft) 293 m (960 ft) -34% -51 % Productivity 3283 ( 2040) seat km/h (mi/h) 4944 (3072) seat kmA (mi/h) 20.4 (48) seat (mi/gal) 17% Fuel 17.4 (41) seat k n/I (mi/gal) km/I 0,50 0.65 +30% Engine sfc i TE-3751 Figure 3. Aircraft comparison.

P`'Q^JP

,1 0-I Turbine Penetration ^^ J 4-7 Place 186-231 kW (250-310 hp) Mission: ' Distance 1667 km (900 NM) plus reserve load 5 passengers plus 36 kg (80 Ibm) W 422 kg (930 Ibm) Speed 103 m/s (200 kn) at 6096 m (20 000 ft) Runway 762 m (2500 ft) Current Price Range: $60,000-$125,000 7976 W/o `v/ GATE] GATE TE-3752 Figure 4. GATE category V--high-performance, single--engine, retractable gear aircraft.

®— Turbine Penetration 2-6 Place 112-231 kW (150--310 hp) Mission: a u Distance (185 km (100 NM) plus reserve Load 2 passengers and 363 kg (800 Ibm) = 517 kg (1 i40 Ibm) o Speed 67 m/s (130 kn) at 1219 m (4000 ft) CD Runway 305 m (1000 ft) E Special l=our missions without refueling D Z Current Price Range: $25,000- 580,000 W/ 1976 W/o GATE GATE TE-3849 Figure 5. GATE category VI--utility, single-engine aircraft.

The four categories of single-engine aircraft with potential for turbinization were forecast to total approximately 5400 aircraft by 1985. A GATE program nas the potential to expand this market to over 6000 aircraft with approxi- mately 1400 of these being turbine powered (Figure 8).

Twin Engine Engines used in the light twin, category IX, were considered too small to be economically turbine powered.

Twin-engine, unpressurized cabin class aircraft offer potential for some tur- bine penetration. These aircraft are frequently used by feeder airlines on the relatively short-haul missions. Noise and reliability are two major fac- tors favoring turbine engines for this type missions. Improved turbine en- gines can increase the turbine penetration in category X as shown in Figure 9.

The smaller pressurized twins, category XI also offer some potential for GATE impact. These aircraft are currently all reciprocating engine powered. The advantages of turbine engine, will be used in some of these aircraft by 1948 with increased penetration resulting from improved turbine engines. It is also expected that a GATE engine could give these aircraft sufficiently im- proved capabilities to expand the total market for the category (Figure 10).

0-Turbine Penetration 1 Place 149-•336 kW (200-450 hp) Mission: Distance 16.1 km (10 mi) shuttle Load 1 pilot plus 660 kg (1500 lbm) = 757 kg (1670 !bm) Speed variable's sea level Runway 457 m 0500 ft) Current Price Range: 540,000-575,000 w/o ti/ GATE GATE T E-3753 Figure 6. GATE category VII---single-engine aircraft aerial application.

A"M— Turbine Penetration ADA 1 Place 336-746 kW (450-1000 hp) Mission: r; 32 km (20 mil shuttle Distance Z lbm) = 1211 kg (2670 Ibm) 1 pilot plus 1134 kg (2500 Load 2^ sea level Speed variable - 457 m (1500 Ft) Runway Current Price Range: 565,000-5180,000 1976 w/o w/ GATE GATE TE-3754 Figure 7. GATE category ViIl--single-engine aircraft aerial application.

Aerial Appl i cation over 336 kW (450 hp) Aericl Apo I i coti on under 336 kW (450 ho) Utility V MM Q 40W O v C Z Hiah Pericrmance Retractable Gear w/o 1982 th GATE GA i E 1988 TE-3755 Figure 8. Single-engine, fixed-wing aircraft production forecast summary.

®- Turbine Penetration 603 603 "'t - to 0 ^Wa 6-10 Passenger 433-559 kW (580-750 hp) y u i Q Mission: ti Distance 1852 km (1000 NM) plus reserve Load 10 passengers plus 91 kg (200 Ibm) = 862 kg (i900 Ibm) E Speed Z 98 m/s (190 kn) at 2438 m (8000 ft) Runway 762 m (2500 ft) Current Price Range: $200, 000 .5350, fl00 GATE I GATE TE-3756 Figure 9. GATE category X--twin-engine, cabin class, unpressurized aircraft.

Category XII twin-engine aircraft over 4082 kg (9000 lbm) gross take--off mass (GTOM) produced by domestic manufacturers are all currently turbine powered.

Improved engines can allow this type of aircraft to compete more effectively in the international markets and with turbofan-powered aircraft resulting in an expansion of the total market (Figure 11).

Domestic production of twin-engine aircraft is forecast to be approximately 4500 by 1988. A GATE program could expand the total production by approxi- mately 100 aircraft with an increase in the turbine-powered versions from about 850 to 1300.

A twin-engine, fixed-wiLng production forecast summary is shown in Figure 12.

The reduction in the light twin total was due to loss of sales to high-per- formance. singles.

Rotate Wing Since all domestically produced helicopters over 1270 kg (2800 Ibm) GTOM are currently turbine powered; it was determined that the major area remaining for GATE impact was in the ultralight category. These aircraft are all currently piston-powered and their limited capabilities have restricted growth in the market. It was felt that a 20% less expensive engine resulting from a GATE program could result in increased turbine penetration and expansion of the total production for this category helicopter (Figure 13). The current avail- ability of several new engines for the larger helicopter categories as well as the existence of new engine technology programs discouraged the prediction of major GATE impact in this class. A rotary-wing production forecast summary is shown in Figure 14 as allocated based on projections of helicopter production; by class of helicopter.

Turbine Penetration a 5-10 Place 336-671 kW (450-900 hp)

Mission: 2

1667 km (900 NM) plus reserve Q Distance 7 passengers plus 69 kg (140 Ibm) = 603 ky (1330 Ibm) Load o c 123 m/s (240 kn) 6096 m (20 000 it) Speed 914 m (3000 ft) Runwcy E Z Current Price Range: $140,000-.`425,000 1976 w/o w/ GATE GATE T E-3757 Figure 10. GATE category XI--twin-engine, pressurized aircraft under 9000 Lb.

®— Turbine Penetration 6-17 Place 820-1417 kW (1100-1900 hp) Mission: a Distance 1852 km (1000 NM) plus reserve `u Load 8 passengers plus 73 kg (i60 Ibm) = 689 kg (1520 Ibm) a Speed 139 m/s (270 kn) ^d_) 6706 m (22 000 ft) P,unway 914 m (3000 ft) c Current Pr i ce Range: 5500 000-S1 400 000 w/o w/ GATE GATE TE-3758 i Figure 11. GATE category XII--twin-engine aircraft over 4077 kg.

4082 k9 (9000 1 bm irized r kg (9000 lbm) 3000 u Q I O essurized E2000 Z Twin 1976 1982 •x/o With GATE GATE TE-3759 Figure 12. Twin-en g ine, fixed-wing aircraft production forecast summary.

ENGINE PRODUCTION FORECAST From the aircraft production forecasts, it was possible to forecast the 1988 number of domestic airframe, manufacturer-installed turbine engines by power class. The market for reciprocating engines in 1988 is forecast to exceed 25,000 engines, all of which would be less than 299 kW (400 hp). The market for turbine engines is shown in Figure 15. Experience on the Detroi t _ Diesel Allison Model 250 engine would indicate an international market equal in size to the domestic market in 1988 and additional sales equal to 100 of the in- s t:alled engines for spare engines. The total world market would, therefore, oe approximately 120% more than the quantities shown in Figure 15. It can be seen that the greatest unit impact from a GATE program would be in the under 447 kW (600 hp) classes of engines. From a dollar value of market viewpoint it can be seen from Table IX that the smaller engines still maintain a slight edge in potential for GATE impact.

MARKET ANALYSIS CONCLUSIONS • 1988 turbine engine market is roughly equal in terms of dollar volume in all shaft power Classes; however, the best opportunity for GATE exists under 447 kW (600 hp) because no advanced technology engines are planned in this class.

9 GATE engine should be less than 447 kW (600 hp).

• Good market exists providing aircraft are designed to capitalize on the advantages of an advanced turbine engine.

®—Turbine Penetration I 2-3 Place 134-186 kW (180-250 hp) Mission: Distance 463 km (250 NM) Load 3 passengers = ?31 kg (510 lbm) Speed 51 m/s (100 kn) Current Price Range: $50,000-5100,000 GATE GATE TE-3760 Figure 13. GATE rotary wing aircraft category I--ultralight (recip).

PI — Recip 2000 ^ Uti Iity u Q Light W Utility c z 1000 Light ;00 Ul fro] ight GATE GATE TE-3761 Figure 14. notary wing aircraft production forecast summary.

v Q v c rn ro v c a `a 1000 >a E a z 599-746 299-447 448-597 298 (400) & Under (601-800) (801-1000) !401-600) Power Class—Ic y^J (hp) TE-3762 Figure 15. Turbine engine market forecast--1988.

TABLE IX. - 1988 ?1ARKET FORECAST-- GENERAL AVIATION TURBINE ENGINES gate With zate Without Est value* ,I VALUE Eat value* Engines IS pillions) ($ rsillions) engines (S villions) +24 5 650 29 Turboprop LINDER 29RkW(400 hp) 12C 61 a36 452 25 1102 299-447kW(401-600 hp) 1400 108 •29 1020 79 448-597kW(601-800 hp) 536 54 846 84 *30 398-746kw(SO1-1000 hp) .7 0 200 7 UNDER 299kW(400 hp) 0 Tur3ashaft 660 30 660 30 '-99-447kW(401-600 hp) 49 0 780 49 790 448- 597kW(60L-800 hp) 180 L5 0 598-746kW(801-1000 hp) 180 15 66724 (1500 1Ff) thrust Turbofan--no significant market under • !978: Turboprop $1481 W($110/hp) 7u 1 cehaft 5121/kW($90/hp) IN MARKET IDENTIFIED SURVEY - MISSIONS { TABLE X.

Pr-ce Shaft Power, Range, Payload, Speed. Runawav, Class, Fixea Wine kW(bp) km(Ha) knLlbm) mis(kn) a(ft) 4 thousand Fj heavy twtn* 5971800) 2222(1200) 844(1860) 154(300) 7b2(2500) luou Light twin* 336(450) 1667(900) 080(1500) 134(200) 7b2u5UU) 473 Cabin class twin 33b(450) 1016(2240) (Unpress)* 1852(1000) 113(220) 702(25UU) 4UU < AG ( 450)* 33b(45U) 3 h 1089(24U0) 51(LU0) 3U5(U000) 110 0G ( a 450) 597(800) 3 h 1542(3400) 51(iuu) 305(1000) 16U E Utility single 261(350) 741(400) 726(1600) 82(16U) 305(100) li0 } 11.P. single (RG) 248(400) 1852(1000) 578(127U) olU(2000) 118(230) 16U ti Helicopter Ultra)ight 2b1(350) 363(800) 373(180) 57(130) 120 Twin 201(350) 833(450) 544(1200) 641125) 5uU e *:fissions exercised in Task it trade studies.

• No market for small turbofan of less than 6672 N (1500 lbf) thrust in 1988 time frame.

• Penetration of fixed-wing aircraft market required for attractive engine quantities.

• Representative Missions recommended for investigation of gas turbine engine requirements for each market category are shown in Table X.

O RIGINAL PAGE IS OF POOR QUALITY 4b. Broad Scope Trade—Off Studies TECHNOLOGY FORECAST FOR MMET ANALYSIS t A study was made to establish representative trends for general aviation gas turbine engine cost, specific fuel consumption (sfc), and weight. These trends were used as a guide in estimating market potential of the small gas turbine engine, and also served to provide engine size —scaling relationships t I for the airplane/engine sizing computer program used in the trade—off studies. In the analysis of cost trends, relationships were also developed to judge the impact of production volume on unit costs. Turboshaft, turboprop, and turbofan engines were examined.

Effect of Engine Size on Weight, Performance. ? and Cost As the scale size of an engine of a given configuration is reduced in the pow - er range below 746 kW (1000 hp), there is a general trend for specific fuel consumption, specific mass (mass/power) and specific cost (cost/power) to in- crease.

One of the causes for these increases is the deterioration of component effic- iencies as component sizes are reduced.

A study was wade to determine representative trends of specific fuel consump - tion, specific mass, and specific cost versus shaft power for use as a guide in the Task I market forecast studies and in the Task II mission trade—off analyses. Scale effects on specific fuel consumption were run on two cycles

(1) an 8:1 compressor pressure ratio (Rd, 1311 K (1900 0 F) rotor inlet

temperature (RIT) current technology cycle and (2) a 14:1 R c , 1478 K (2200 o F) RIT advanced technology cycle.

Figure 16 shows the difference in compressor and gas turbine efficiency as - sumed for the two cycles using the 1.36 kg/s (3.0 lb m/sec) air flow 8:1 cur- rent technology engine as a base. Also shown are the ffects of scale on sfc. The sfc scale effects were similar on a percentage basis, thus a typical line is shown. Scale effects on specific mass and cost were run on configura- tions representative of current and advanced technology cycles. The results are shown in Figure 17 and 18 for specific mass and specific cost, respective- ly. The effect of these trends on turboprop and turboshaft characteristics for current and advanced technology is shown in Figure 20 throug!.a 25.

The shaft power range from 112 to 746 kW (150 to 1000 hp) was analyzed. The sfc versus power trend resulted from computerized engine performance cycle runs at a series of airflow levels compatible with the selected shaft power range at sea level static standard (slss) inlet conditions while holding R and RIT constant. The scale effect was accounted for by applying estimated scale effects to the values of compressor efficiency, gas generator turbine efficiency, power turbine efficiency, and percent turbine cooling air used in the cycle as engine airflow rate was raised.

The specific mass versus power trend was obtained by making computer runs at a series of scale sizes compatible with the airflow rate range used in the per- formance study already described. A DOA gas turbine masses estimating program was used. It calculates component masses and total engine mass using suitable input data for a given set of configuration features and cycle parameters with the scale size keyed to the airflow rate. The weights obtained were then re- lated to power by using the power versus airflow rate relationship established in the cycle performance scaling runs described above.

1 *s .— u 0 ,ter-- — ,^.....^ -'-^ 8: T Rc engine --T4:1 Rc engine Typical ^+nn+'Y'•w•'M^ 1' sfc at sfc" = 373 kW (500 hp) " 1.10 u T.00 0.90 1 1 3 4 0 2 5 7 8 kW/100 0 1 2 3 5 4 6 7 8 9 !0 hp/100 Shaft Power TE-3763A

Figure

16. T'urboshaft

at,:d turboprop engines sfc sensitivity to scale size

relative trend at slss T.O.

1.8 1.7 1.6 Q^^f^s^l 3 ; a 0 1.5 a: q 1.4 F a, v4^ 1.3 1.2 CL 1.0 0.9 0.8 0 1 2 3 4 5 6 7 8 ¢ 4V ! I ^ r r ^ r ^, I I I 0 1 3 4 2 5 6 7 9 8 10 ^P 100 Shaft Power TE-3764 Figure 17. Turboshaft and turboprop engines specific mass 1.7 i 1.6 * = Value at 373 kW 1500 hp) 1.5 U ^ 1.4 a o a.

o a 1.3 t o a e : 1.2 c a a 1.1 V 1.0 a 0.9 0.8 ' r ! ! r 0.7 ^!

0 1 2 3 4 5 6 8 kW/100 r ^ r r t r r „^ ^ r r 4 5 6 a 9 10 0 1 2 3 7 hpe 100 KA r Shaft Power rE-3765 Figure 18. Turboshaft and turboprop engines specific cost sensitivity to scale size--relative trend at slss T.O.

The specific cost versus power trend was obtained by using a DDA cast-estimat- ing computer program that is an auxilary to the previously memtioned masses program. The program calculates components and total engine manufacturing costs using as inputs the component masses plus component costs for the unity size engine. The component masses from the mass--estimating program are fed into the cost program. The unity size component prices were estimated by us- ing the DDA refined 'Materials Index Factor (MIF) method. The scaled compo- R nents costs were estimated by assuming that component cost varies with a se- lected power of the component mass. The costs were adjusted to a common pro- duction rate, total quantity, and specific calendar year economy.

These scaling trend curves were used in all of the turboshaft and turboprop mission trade-off studies. The desired scaled value of sfe, mass, or cost was obtained by multiplying the unity engine size values by the ratio of the trend curve ordinate values at the scaled shaft power divided by the ordinate value at the respective engine unity size shaft power.

Other important items which impact engine cost include: • Ecomonic year of production • Total quantity of engines produced • Production rate t> In the GATE cost studies constant economics was used to avoid the uncertain- ties of predictin future escalation rates; 1978 was used as the base year.

g As the total production, quantity of a given engine configuration increases the cumulative average cost normally decreases exclusive of other items. This is generally referred to as the learning curve effect.

As the production rate (engines per month) increases, the manufacturing costs normally decreases.

Effect of Engine Size and Production Quantity on Cost Figure 19 illustrates a typical combined effect of total production quantity and engine power size on specific cost. The effect of production rate is also included since the overall time is held constant. These trends assume con- tinous flow of production units over a 5-yr period. Tooling costs are assumed to increase proportionately to production rate. The figure shows effect of order-of-magnitude changes in production quantity. Typical aircraft engine production rates for small turbine engines are most nearly represented by the 1000 engine per year or 5000 engine quantity line.

Turboprop Engine Trends for Analysis Published competitive turboprop engine data in the power range up to 745 kW (1000 hp) was reviewed along with DDA production engine data to estimate cur- rent levels for specific fuel comsumption, mass, and cost.

112 kW (150 hp) 2.4 2.2 2.0 N o 1.8 4 C.

o^`a 373 kW 1.6 (500 hp) * = Value at 5000 engines and 373 kW (500 hp) a•o uiv 597 kw 1.4 (800 hp) 746 kW a {1000 hp) 1.2 500 Engines v U 1.0 ^V V G 5,01G17 Engines 0.8 50,000 Engines 0.6 500,000 Engine 0.4 TE-3766 Figure 19. Turboshaft and turboprop engines production volume power size impact on cost-relative trends.

f Figures 20, 21, and 22 are composite plats of this data in the form of rela- tive specific fuel consumption versus shaft power, specific mass versus shaft power, and specific cost versus shaft power, respectively. In general, these plots show that sfc, specific mass, and specific cost increase as engine power rating is decreased. Superimposed an these plots are the scale effect trend lines described earlier. The 8:1 R e , 1311 K (1900 0 F) RIT trend line is representative of a current technology engine (CTE) and the 14:1 Re, 1478 K (2200 0 F) RIT trend line is typical of an advanced, air--cooled engine. The advanced technology engine (ATE) trend line is lower than the current techno- logy engine trend line for sfc, specific mass, but hi g her for specific cost.

In interpreting the plotted competitive engine data, it is cautioned that the data may not be to a common definition base. For instance, it is not Icnown whether the published competitive engine shaft power and specific fuel con- sumption data are guaranteed values or the normally more favorable average values or projected values. The engine price terms are also difficult to de- termine. Sources do not always qualify costs on the basis of list or OEM, effective year and engine quantity. Where data detail permitted, costs were adjusted to a 1978 base year.

Turboshaft Engine Trends A review and analysis of current competitive and DDA turboshaft engines was made similar to the turboprop en gine review described above. The results of this review are shown in Figures 23, 24, and 25, showing sfc, specific mass, and cost trends, respectively. Kass data was adjusted to provide a standard drive shaft rotational speed as noted, and costs were adjusted to a 1978 base year.

1.6 = Value of 373 kW (500 hp) for CTE 1.4 CTE 1311 K 0900T) e1T 1.2 8;1 R, O u } O 00 0 O O G 1.0 (900 z O 0 0 00 o.s ATE 1478 K (2200 °F) RIT 140 Ra 0.6 0 200 400 00 800 kW I ^ I I I i 200 400 600 800 hp TE-3767 Equivalent Shaft Powe.

Figure 20. Turboprop sfc trends.

2.0F = Value at 373 kW (500 hp) far CTE 1.8 Q. CTE 131' K (1900 , `a F) RIT all Ra 1.6 a ATE 1,4 1478 K ( 2200 °P) RIT 4.1 RC a m 0 1.2 0 .

O 00 000 0 1.0 pp c 0 0 O 0.8 0 0.6 k w 0 200 hp Equivalent Shaft Power TE-3768 Figure 21. Tu rboprop mass trends.

p^^AL A

P. 44

. Or

is

1.8

p

U

^ Y'Y ATE 1478 K (22GWF) RIT } 1.6 141l Rr a R = Value at 373 kW (500 hp) 1.4 for CTE n 1.2 0 c a O O O o" 1.0 a O O 0 0,8 O 0 CTE U 131T K (1900°F) RIT O B:1 R C 8 g 0 8 V C y ti 0.6 kW 0 204 hp TE-3769 Equivalent Shaft Pawer Figure 22. Turboprop cost trends.

I.6 O 1,5 u 1.4

u

., - • = Value at 373 kW (500 hp) for CTE

c 1.3 O E O E 1.2 V O ` o e 0 I.1 O CZ O O 0 J 1.0 O O CTE ° $ O 1311 K (19ffF) RIT 0.9 8:1 R a O O O 0.8 A TE 1478 K (2200°F) RIT 14:1 R, k W/100 IL^^ 2 3 4 5 6 0 9 to F 1 hp/100 Shaft Power TE-3770 Figure 23. Turboshaft sfc trends.

Mass Adjusted to 6000 rpm Output 2.0 t $ 1,8 e p = Value at 373 kW (500 hp) for CTE 1.6 u a o.

CTE 6 1.4 311 K (1900°F) RIT f 8:1 Rz F e 1.2 0 O o O 0 O 1.0 ATE ^a 1478 K (2200aF) RIT 14:1 Rc O 0,8 2 3 4 f 6 7 8 9 to WIN I i ^ I I 1 ^ ^ ^ I I I I 2 3 6 7 8 9 hp/loo 1 4 5 10 11 12 1 3 Shaft Power TE-3771 Figure 24. Turboshaft mass trends.

2.2 _ CTE O R 2.9 s Y 1.8 1.6 1.4 u ATE 1478 K {2200 0 x ► RIT 1.2 r g 14:1 Ra 0 0

s

1.0 0 ° o CTE 1311 K (1900 0 F1 RIT a.la 00 RC 0.6 6 7 8 9 kW/100 I I !

I I S I I I I I I tip/100 1 6 8 2 3 4 5 7 9 10 11 Shaft Power TE-3772 Figure 25. Turboshaft cost trends.

Note that again significant improvements in sfc are shown for the advanced cycle, but that this gain is accompanied by an increase in en g ine cost. These trends were examined again in greater detail in the Broad Scope Trade-Off Studies.

COST PERFORMANCE TRADE STUDIES Approach Cost/performance trade studies were conducted for six defined missions repre- senting important market segments for the small gas turbine engine. For each of these missions, a representative basepoint aircraft configuration was es- tablished representing current aircraft design practice. Gas turbine engines were evaluated in a computerized system that generated comparisons in aircraft design, gross mass and economics. The payoff parameters considered were mini- mum acquisition cost, minimum direct operating cost, and minimum cash flow requirement as determined for the complete engine/airframe combination. Air- craft gross mass was also an important parameter particularly from the fuel conservation standpoint, but also because for comparisons conducted among gas turbine engines, the gross mass was the major driver on costs.

The evaluation process involved complete definition of the current technology basepoiat gas turbine engine as well as the matrix and candidate engines in terms of design and off-design performance, mass, dimensions, initial cost and maintenance cost. Engine-scaling procedures were applied as described in the section on Engine Size Effects.

The goal of the study was to improve as much as possible an a modern gas tur- representing perform- bine engine. For this purpose, a scalable study engine ance and cost characteristics of an advanced DDA Model 250 was selected to represent current production technology. The hardware configuration of this engine entered'production in 1978 at 485 kW (650 hp), Long production experi- ence with the basic engine frame has resulted in highly competitive perform- ance and price. This engine, designated CTE (for current technology engine) or CTE* (when the price is increased to reflect a hypothetical case of no pre- vious production experience on the model) forms the standard for measurement of cost/performance-related improvements.

Two engine matrices were generated using single and dual centrifugal stage compressors selected for performance and cost advantages in small engine ap- plication. These matrices provided a basis for selection of engine technology features. Pressure ratio, turbine temperature and turbine configuration were included in the evaluations. Candidate engines were defined and evalutated at

a nominal, 373 kW (500 hp) base, and in turn were compared to the GTE and CTE*

in both fixed-wing and rotary-wing applications.

Sensitivity studies were conducted to evaluate component and cycle parameters in terms of their impact on the general aviation aircraft design gross mass and economics of performing the specified desig p, mission. Selected technology elements were qualitatively or quantitatively -valuated.

Engine noise and emission characteristics were studied to determine impact of the new technology, and regulatory constraints were reviewed to forecast the probable limits for 1985--1990.

Mission Re q uirements

Typical missions were defined for all general aviation categories investigated in the Task I market study. Representative cases were selected for trade studies from market categories VII through XI for the fixed-wing aircraft and Category II for the helicopter as shown in Table X.

Fixed-wing missions used in the trade studies are shown in Figure 26 in terms of typical cruise altitude, air speed, payload, and range. Airframe specific cost values used in assessing the airframe cost in the economic analysis are

shown. Also, the major engine sizing conditions considered in the analysis

are specified. Figure 27 presents similar information for the rotary-wing aircraft. Cabin size of the reference aircraft was scaled appropriately to meet the payload requirements of the selected missions.

Airframe Characteristics Airframe characteristics based on aerodynamic and mass data Obtai,zed from the manufacturers were applied in the study. Airframe cost data was obtained from published information from such sources as GAMA (General Aviation Manufac- turers Association).

The basic analysis technique employed a completely described reference air- craft and simply resized it to exactly meet the mission requirement for each engine examined. Thus the primary aircraft characteristics such as design wing loading, aspect ratio, base drag, structural mass fraction, and airframe specific cost were unchanged except for alterations resulting from scale ef- fects and changes related to propulsion mass fuel economy and geometry associ- ated with each engine considered. Aerodynamics were synthesized to represent the geometric configurations as detailed in Table XI.

Cruise CIim 6 T.O. Power 56.6 m/s ~^ Reserves at T. 0- (110 kn) EAS 3048 m 2 min at (10000 ft)/VOPT T T . O. power for endurance ^+------ Range Unpressurized Light Heavy Ligh• Twin Agricu!tu-al Mission Twin Twin yes yes yes yes T.O. Allowance yes yes none Climb Allowance yes ( 9144/154 (30K/300) SV51 (5!./100) Cruise Alt/TAS, m/(m/s) (ft / kn) 3664/113 12K/220) 731 /134 ( 24K/260) 45 45 0 Reserve Time, min 45 1667 (900) 2222 ( 1200) 556 ( 300) 3 hr Range, km (NM) 1852 (1000) 1015 (2240) 680 (1500) 843.7 (1860) 1099 (2400) Payload, kg (16) (includes crew) 187.39 ( 85.00) 44,09 ( 20.00) Specific Cast, S/kg Empty Mass ( S/lbm EM) 110-89 (50.30) 124.56 ( 56.50) SL, 51.4 m/ s (100 kn) Critical Engine Sizing Cand 4.57 m/s (900 ft/min) rate of climb 0 Cruise & OEl Climb TE-3784A at 1524m (5000 ft) - wing missions.

Figure 26. Fixed- Reserves at SL Cruise SL/VOPT for endurance Range Light Light Mission Single Twin ORIGINAL PAGE IS 56.6 (110) 64.3 (125) Cruise Vel, m/s (kn) TAS QUAL17- + 0; POOR Reserve Time, min 30 45

Range, km (NM)

333 (180) 833.4 (450)

Payload, k9 (lb)

363 (800) 544 (1200)

Specific Cost, $/kg Empty Mass ($/!bm Empty Mass) 265 (120) 132 (60) Critical Engine Sizing Conditions

• SL m/s (kn) TAS

61.7 (120) 77.2 (150)

Vmax, M

305 m130 m/s TAS (1000 ft/57 kn) OEl 305 K (90°F) day NA yes

TE--3785 Rotary—wing missions.

Figure 27.

LE XI. -- DESCRIPTION OF REFERENCE A Fixed wing Rotary wing Unpreas- Aircraft type Heavy urized Light Light Lighc Light twin twin twin AG single twin Confieuration 9-11 6-8 7-9 1 3-4 5-6 Places (includes crew) Wing: p 1.58 (33) 1.58 (33) 1.58 (33) 1.01 (21) Loading, k a (Ib/ft2) 9 6 Aspect ratio 9 9 3 3 3 0 Sweepback, degrees 0.14 Thickness/chord, avg 0.14 0.14 0.14 0.45 0.45 0.45 0.63 Taper ratio Horizontal tail: 0.28 Area ratio, tail to wing 0.28 0.28 0.17 V ertical tail: 0.20 0.20 0.20 0.06 Area ratio, tail to wing Main rotor: 2 2 No. of blades 2.68 Disc loading, kPa (lb/ft2) 2.68 (5.6) (5.6) 9.07 0.07 Solidity 210 (690) 210 (690) Tip speed, m/s (ft/sec) Tail rotor: 2 2 No. of blades Solidity 0.16 0.16 187.5 (615) 187.5 Tip speed, m/s (ft/sec) (615) m3 *Hopper capacity, (ft 3 ) = 1.50 (53) (400 U.S. gal).

i In this type of analysis, the mission is fixed, as are the basic airplane characteristics, and changes in aircraft gross mass, wing area, and engine power size basically reflect differences in the quality of the propulsion sys- tem. Fundarlentally, gross mass changes calculated for a change in engine characteristics are a result of a change in propulsive mass fraction, which is the percentage of total gross mass required for the installed engine, fuel, and fuel system. Complex interactions are treated interatively to arrive at a solution, for example engine mass is affected by engine specific power (engine mass per unit of power) and the power requirement at the sizing condition. It in turn is dependent on gross mass and airframe drag which are influenced by fuel mass aad engine geometry, which depends on engine power size. Fuel mass is affected primarily by engine-specific fuel consumption and power require- ments at cruise.

Economic Model The economic model for general aviation aircraft developed for the GATE study is described in detail in the appendix. Cost data calculated for the fixed-- and rotar y-wing aircraft were: • Total acquisition cost, $ • Direct operating cost, $/flight hour • Total Cost of ownership, $ • Cash flow requirement, $ ORIGI%AL PAGE IS ®V POOR QUALITY Total acquisitor cost (TAC) is the total selling price of the engine/airframe combination. The engine is priced at list. Direct operating cost (DOC) is i., the hourly cost of operating the airplane and includes the cost of fuel and i oil, maintenance, depreciation and insurance based on an annual use. The to- tat cost of ownership (TCO) is the sum of the airplane cost (TAC) and the op- erating expenses over a prescribed period of time. The cash flow requirement (CFR) is the sum of the yearly net cash outflow for the specified ownership cycle. Cash outflow items include the initial payment, annual payments, and certain variable and fixed operating costs.

Cash inflow items include invest- ment tax credit and tax savings resulting from allowable deductions for depre- ciation and operating expenses.

Details of economic evaluation methods are shown in the appendix. Included is a summary of the cost standards used in the analysis. A partial summary is shown in Table XII. Fuel cost values used as a base in GATE are shown in Ta- ble XIII.

TABLE XII. -- ECONOMICS OF AIRCRAFT OWNERSHIP COST METHODOLOGY BASED ON; 190 ATA Standard (doc) Rockwell cash flow analysis (CFR) AEAA Paper 67-828 (doc) Aircraft operation cost summaries TDR AX-0000-590 (doc) Economic standards: (MY 1978) Fixed wing Helicopter Fuel cost, $/ 1($ /gal) 0.22(0.83) and .33(1.24) 0.22(0.83) and 0.33(A.2_4) Oil cost, $/1 ($/gal) 2.51(9.50) 2.5119.50) Depreciation period, yr 8 8 Labor rate (including burden), $/h 20 Annual use, h/yr 600,9(10 36o,600 Annual insurance rate, X 1 5 Annual rate of depreciation, % 25 25 Annual interest rate, X 10 io Down payment rate, % 10 10 Resale value, % 40 40 Rate of tax saving, % 52 52 Residual value, % 20 20 Hangar rental, $/yr 3540 -- Aircraft re istration $ + $/k ($/lb) 25 + 0 077(0 035 1 -- g ^ g Turbine engine OEM price: quant ; . ty, total units 5000 5000 Rate, units/mo 80 Units, for cum avg price 5000 5000 TABLE XIII. - COSTING Low Medium High Fuel cost, $/1 0/920* 0.21(0.79) 0.22(0.83) 0.24{0.90) 80 Octane 0.22(0.810 0.22(0.85) 0.25(0.95) 100 Octane 0.18(0.68) 0.20(0.77) 0.22(0.83) Jet fuel 1978 ( 1 March) Economic base: 7/11/77.

*Reference: Business Aviation t r i Engine Cost DDA used the material index factor (MIF) method for cost estimation of the engine concepts in this study. One of the chief advantages of the MIF costing method is its ability to measure the impact of advanced technology on the cost of a whole family of engines before they are detailed. A cost model is pre- pared from an engineering layout of a basic point design engine using the DDA MIF costing method. This cost model is factored from latest production and advanced development experience. Selected information from the MIF Cost Model is then programmed into the Design Math Model together with changes for the concept under study. From the computer programmed for the Revised Math Model, it is possible to get relative section weights, section costs, and engine costs resulting from iterations of various design parameters direct without benefit of further engineering drafting.

The materials index factor method of estimating engine costs is a series of calculations that has been derived from DDA refinements of original work by the late R. J. Maurer of the Naval Air Development Center (NADC).

After studying many different parameters used as cost indicators of aircraft engines, Mr. Maurer found that the best indication of engine cost was the ma- terials index factor. The MIF may be defined as the sum of relative indices OE cost ("weighting factors") times their respective required raw material weights. The material cost indices, as shown in Table XIV, are derived by multiplying the relative material costs by their relative factors of machina- bility. For engine applications, DDA considers that many of these indices are outdated and otherwise inaccurate. For these reasons, (1) the indices are used selectively--i.e., DDA uses special new indices for exotic blade and vane materials; (2) the indices are continually in process of modernization, using data similar to that gathered under DDA's recent contract N62269-76-M-6616 with NADC; and (3) errors are minimized in DDA costing by equating Maurer fac- tor data to actual costs determined for production engines.

TABLE XIV. - MATERL`1, INDICES FOR MIF TECHNIQUE Hajor case, disk, soccer, shaft Turbo blades 6 vanes Material. classification Ti A B C D Conv A 8 Conv Relative material cost 7.0 3-4 4-5 5-7 7-10 3.0* 1.5 2.0 1.0 Relative machining casts 1.5 1.9 3.1 4.0 3.5 1.0 1.5 2.0 1.0 Relative weighting factor 10.5 6.7 14.0 24.0 29.8 3.0* 4.0 2.5 1.0 v — .^ x c a u a O T ry Typical materials as ,d w n o o- H e - 0 0 ^. r - O N y} .+ .-. ^. !^V v N O N C C J I+ CI O 4 1 O y p.- O V 0. .O^ ^^ u uR *substitute value by DDA.

The Maurer factor is based on the total raw weight or material requirements of the engine. No recognition was made, in the original concept, as to whether or not a product has an efficient utilization of material or an inefficient one. The total Maurer factor can be changed only by changing design configu- ration so as to affect material mix and weight, or by changing the processing r efficiency to affect material utilization. For example, processing efficiency x may be improved by use of "near net" raw material shapes.

The DDA MIF method of costing a production engine compensates for materials utilization. First, the finished weights (FW) of all the parts are computed.

Material utilization factors (MUF) are estimated/computed on the basis of ac- tual part shape and form. The material indices (MI), which include the fabri- cating difficulty, are assigned from a modified Maurer factor table. These three factors (FW, 10F, MI) are multiplied to form MIF factors, which quanti- tatively represents both material and labor for each part. The MIF factors are divided by a value representing the average material utilization for small turbine engines built at DDA. It has been determined that this material util- ization value--e.g., the raw material weight divided by the finished material weight---for small DDA-built turbine engines is 3.36. The aforementioned quo-- tient is then multiplied by a K factor to give the material cost of the vari- ous engine. sections. 'These costs are then summed to give the total manufac- turing cost of the engine except for assembly, test, and proprietary acces- sories. The cost of assembly, test, and controls/accessories are estimated by ;3 DDA experts (by examining similar production costs) and added to the other engine costs factored by MIF from the FW of the materials: This then is the recurring manufacturing cost of the engine under study.

The K factor, as explained in a paper given at NADC, Philadelphia, 20 November 1975, by L. L. Robinson, B. A. Zolezzi, and D. K. Hanink, is a combination of four factors which must be weighed in order to reach an accurate dollar cost.

These four factors are (1) X factor related to factory efficiency, which also takes into consideration normal material utilization; (2) Y factor considers rate of production of a study engine in relation to the facilities normal rate of production; (3) Z factor relates a given production quantity of production which includes anticipated learning, and (4) T factor in which economic esca- lation is related to time. Since DDA does not have dedicated engine facili- ties, a parametrically derived K for a specific engine model is affected by abnormally low production rates on other engines being produced concurrently in the facility.

The determination of accurate costs depends on determination of an accurate value of K. The DDA method used for the determination of an accurate value of K was to examine the cost history of the T63 engine, compute and total the various materials index factors for the engine, multiply the sum of the MIF factors by K/3.36, and equate this total to a real average cumulative cost of the engine (MIF x K/3.36 = actual manufacturing cost). From this equation the K value can be determined for a given quantity of engines, rate of production, and economic period. The K value can be corrected for changes in these three variables. For example, a few years ago a K factor was established for the average cumulative cost per engine of 2000 T63/250 engines at a rate of 70 per month. K was found to be 20.8.

ORIGINAL PAGE IS OF POOR (QUALITY Since that time engine costs for labor and material, as experienced at DDA, have escalated 40% (T factor). A new value for K can therefore be calculated as 1.4 x 20.8 = 29.12. However, this K factor should also be corrected for production rate. The plant efficiency (X factor) is considered to be un- changed. It is estimated that procurement cost will increase by 6% due to decreasing the rate of production from 70 to 21 per month (Y factor). The 6% value is derived from analysis of current production costs and is predicated on accounting procedures that relieve a particular model of fixed plant over- head costs in nearly direct proportion to its reduction in rate per month--

i.e., DDA fixed costs are

not segregated and assigned to a given model "in perpetuity" as would be the case with a dedicated facility. The previous value of K can be corrected as K = 29.12 x 1.06 = 30.86.

This value of K must now be corrected for a total production quantity such as 2500 engines (Z factor). The average cumulative cost of the 2500 engines will

be

less than that of 2000 engines. The reduction in cost can be accomplished by application of the values contained in the 90% learning tables. Values for 2000 (0.3149) and 2500 (0.3044) are applied to K by the ratio (0.3044/0.3149). This results in a K of 29.82, which DDA arbitrarily elected to round off to a slightly more conservative value of 30. This is, in fact, the value of K used for costing of a recently proposed DDA production engine.

To accurately determine the learning rate experienced at DDA, a cost history

of the T63 engine was compiled. The engine was manufactured during the period

from 1966 to 1970. Table XV is a summary of the cost history for that en- gine. Engine cost was refined for economics and for lot size. The tabulation is graphically illustrated in Figure 28. Note that the first lot of five en- gines did not follow the log-log linear relationship as well as the remaining lots. Therefore, this point was excluded from the learning curve calcula- tion. The cumulative average costs of the first 13 engines were assigned a cost of unity, and all other cumulative average costs were shocm as ratios of unity.

TABLE XV. - T63 ENGINE COST HISTORY Lot Quantity Cumulative Cumulative number in lot total average ratio 1 5 5 1.0199 2 8 13 1.0000 3 13 26 0.8917 4 63 89 0.7392 5 70 159 0.7050 6 653 812 0.5037 7 1303 2115 0.4264 8 3858 0.3894 The learning curve was determined by two methods: 1. Graphical illustration on log-log paper (Figure 28) and a comparison with an MI slope analyzer.

177 F

(Derived from production record of costs for 250 HP helicopter engine) MIN 1 • i.:..

q w - - .:

i :: fit. ii:.: iia ii -

Att I 1rr { I^-( T ( IF•f .^ 1:1v 11•^ I 1 iI t 1 t^ (^ I ' t • { i - • i" r i i ri h' I i^^ ii 11- I. I, I ^T ,-, , ,!

1.( . -- - - -

0 % _

dt w I^ _ _a.: go - :i.: • ,Y' ^ - t 11V _t.

3F'ri . - t" '3 ` ii F _ - 1 ' ri:: - - r - •r-

Q

Ti-TT t-1 ( f. i^: i^ . j ... f^ i r^- }{ !^ I ^ i `^? i ^I {} ti ' I i II - I i 0.1 1 10 10.000 1,000 100 Quantity of engines 7 F-5053 _N Figure 28. Log—log linear learning curve.

y

N r/1 y:::.:.. ....._ ^.^,,.i>_,..»...-,.y...M..,.,....,-».. .-. ..c-.,-.a^.,v..:o.....w.....r•.,..=,..-r+e^rw7e^::=.t^+i.^:«:v. - -:-«.^..a- ^.,,..___,_,^^.,,..i i k ^ 2. Mathematical determination using the formulae.

1 cost_ Q2 N Log %L = N Log 2 Q2 cost - ^Q1^ where: Ql is the lower cumulative production quantity Q2 is the higher cumulative production quantity Ql cost is the cumulative average cost for production quantity i Q1 Q2 cost is the cumulative average cost for the production quantity Q2 N is the slope of the learning curve L is learning A single calculation follows: 1 cost _ (Q2 N Lag % L = i Log 2 Q2 cost - t Ql \ N= Log es%L Log 2 1.000 _ 3858 N 0.3894 13 2.5683 = (296.77)N _ Log 2.5683 N Log 296.77 Log % L _ 0.4096 Loo 2 2.4785 Log % L = 0.0499, as the slope is negative Log % L = —1.00 + 0.0499 = 9.9501 — 10 L = 89.1% Although the history of the T63 engine plots to a learning curve of 89%, re- cent DDA engine model values have been subjected to 90% learning curve fac- tors. This is judgmental and intends to reflect sut;n improvements as DTUPC programs which permit nets parts to be manufactured initially at somewhat near- er optimum efficiency than in the 1960's. This is expected to result in lower initial costs and, consequently, a "flatter" progression toward the cumulative average cost of 2500 engines.

,F Calculation of GATE Engine Costs An engine parts list derived from a given GATE point design drawing was used as a unity cost model. Materials were identified by the Design and Materials Engineering Groups, and finished weights were provided by the Weights Group.

All engine estimated costs were factored to a common basis of 5000 units at 80 per month and given in 1978 dollars. Historical cost data and recent cost study results for a wide range of DDA small gas turbine engine configurations were reviewed as described in the beginning of this section on Engine Cost Estimation. Factors derived from this review were applied using Materials Index Factor (MIF) methodology to obtain realistic acquisition cost estimates for GATE engines. A study was made of the different materials required to- gether with `ue material use and processing differences for each GATE config- uration. These physical differences formed the basis for factors used in dif- ferential costing; by the MIS' technique and were plugged into the design com- puter program that iterated design chaages and their resultant cost changes.

Cost data was generated by MIF methodology as described for those engine con- figurations that stressed design simplicity and minimum parts count. This cost data was then compared with similar cost data generated for more sophis- ticated engines that stressed performance. Cooled and uncooled turbine engine configurations were examined from a cost context, and these results applied in the cost/performance trade atudies.

By comparing candidate engine 26 to engine 29 shown in Tables XVI (SI units) and XVII (customary units), it can be seen that by air cooling the turbine and designing to a higher RIT, 2200 OF vs 1950 0F, the weight decreased 10% and the cost decreased 6% while the sfc remained about the same. The lower weight, lower cost, and slightly better sfc gives the 2200 OF air-cooled tur- bine engine an advantage when considering total cost of ownership (TCO). Sim- ilarly, it can be seen by comparing candidate engine 30 to engine 26, also shown in Tables XVI and XVII, that by designing with two smaller compressor wheels with a compressor pressure ratio of 14 and two smaller turbine wheels in engine 30, in place of one larger compressor wheel with a compressor pres- sure ratio of 10 and one larger turbine wheel (2200 OF air cooled), that the two-stage compressor and turbine engine is only 3% heavier, 2% more costly, but has a 5% improvement in sfc. Because of its better sfc, candidate engine 30 has a lower cost of ownership (TCO) than engine 26 despite a small increase in acquisition cost and weight.

TABLE XVI. CANDIDATE ENGINE DATA - units)--UNITY SIZE (SI Engine 26 27 28 29 30 Technology ATE ATE ATE ATE ATE Type of compressor 1-C 1-C 1-C 1-C 2-C Type of GP turbine I-A 2-A 1-R I-A 2--A Air-cooled GP turbine YES YES YES NO YES TABLE XVI. - (CONT) Design point performance, slss T.O.

Turboshaft and turboprop w/o gear box loss Re 10 l0 10 10 14 RIT, K 1478 1478 1478 1339 1478 Shaft power, kW 372.8 372.8 372.8 372.8 372.8 Airflow, kg/s 1.340 1.349 1.268 1.625 1.400 sf c , g/W • s 86.41 86.39 82.03 87.22 82.34 Turboshaft engine data (incl red gears, 6000 rpm output) Mass, kg 65.5 66.6 73.1 65.7 67.2 Length, m 0.511 0.531 0.503 0.544 0.643 Diameter, m 0.444 0.447 0.437 0.478 0.452 OEM Price, {$) * 62,200 63,422 58,830 66,170 63,562 Premature R.R./1000 hr 0.53 0.53 0.53 0.50 0.53 Maintenance cost, $/fl hr 30 hr/mo util 18.47 18.77 17.42 18.88 18.81 50 hr/mo util 18.05 18.34 17,02 17.83 18.38 Turboprop engine data (inel prop gear box, 2000 rpm output) Mass, kg_ 80.5 79.1 79.3 88.3 81.2 Length, m 0.907 0.927 0.899 0.940 1.039 Diameter, m 0.447 0.437 0.444 0.478 0.452 OEM Price, $ x 74,764 76,233 70,714 791537 76,401 Premature R.R./1000 hr 0.45 0.45 0.45 0.41 0.43 Maintenance cost, $/fl hr 50 hr/mo util 18.24 18.60 19.57 18.64 17.26 75 hr/mo util 17.59 17.93 16.64 18.90 17.97 *Cumulative average price 5000 engs, 80/mo, 1978 base year.

C - centrifugal A -- axial R - radial TABLE XVII. - CANDIDATE ENGINE DATA (customary units)--UNITY SIZE Engine 26 27 28 29 30 Technology ATE ATE ATE ATE ATE Type of compressor I-C I-C 1-C 2-C I-C Type of GP turbine I-A 2-A I-R I-A 2-A Air-cooled GP turbine YES YES YES YES YES Design point performance slss T.O.

Turboshaft and turboprop w/o gearbox. loss Re 10 10 10 10 14 RIT, O F 2200 2200 2200 1950 2200 shp 500 500 500 500 500 'Airflow, lbm/= 2.954 2.974 2.795 3.583 3.086 lbm/hp • hr ^sfc, 0.5114 0.5113 0.4855 0.5162 0.4873 TABLE XVIT. - (COMT) Turboshaft engine data (incl red gears, 6000 rpm output) deight, lb 144.4 146.4 144.8 161.2 148.2 Length, in.

20.1 20.9 19.8 21.4 25.3 Diameter, in.

17.5 17.6 17.2 18.8 17.8 OEM price, $ 62200 63422 58830 66170 63562 Premature R.R./1000 hr 0.53 0.53 0.53 0.50 0.53 Maintenance cost, $/fl hr 30 hr/mo util 18.47 18.77 17.42 18.88 18.81 50 hr/mo util 18.05 18.34 17.02 17.83 18.38 Turboprop engine data (incl prop gearbox 2000 rpm output Weight lb 174.4 177.5 174.9 194.7 179.0 Length, in.

35.7 36.5 35.4 37.0 40.9 Diameter, in.

17.5 17.6 17.2 18.8 17.8 ($) * OEM price, 74764 76233 70714 79537 76401 Premature R.R./1000 hr 0.45 0.45 0.45 0.41 0.45 Maintenance cost, $/fl hr 50 hr/mo util 18.24 18.60 17.26 19.57 18.64 75 hr/mo util 17.93 17.59 16.64 18.90 17.97 *Cumulative average price PAGE IS 5000 engs 80/mo, 1978 base yeVRIGINA1.

C -- centrifugal; A -- axial; R - radial QUALITY OF POOR In addition to the many design parametric iterations compared to a unity sized engine for effect on acquisition cost and TCO, additional technology elements were considered and their cost etiectiveness evaluated. Advanced state-of- the--art technology elements evaluated for their effect on cost when physically incorporated into the selected candidate study engines were dual property tur- bines, ceramic turbine stators, composite material gearbox case, Lamilloy* combustors, and axial-centrifugal compressors. The effect on acquisition costs was evaluated by MIF methodology from the differential bills of materi- als and processing complexity of the considered technology change. More com- plete technical description of these elements and tables showing their effect on cost are given in this report:.

Conclusions on Engine Costing In general the DDA MIF methodology has allowed costs to be plotted as a para- meter of design. This can be observed in the many charts showing costs and how they are affected by different specific design changes. The overall cost analysis also indicates that production experience is as much or more of a cost driver than material content resulting from engine technology. DDA stud-

ies show that the acquisition

cost cf an advanced technology engine (ATE) is only slightly more than that of a current technology engine (CTE) when simi- larly compared. The real life situation of comparing an ATE just going into production to a CTE of the same horsepower evolved from a long production run shows an appreciable increase in the acquisition cost of the newer engine.

This concept is shown

graphically in

Figure 29. In this illustration relative turbine engine costs have been adjusted to constant economics for an actual production run of over 12,000 engines. From this experience the cost of a *Lamilloy is a registered trademark of the General Motors Corporation.

2c s 650 hp ATE L " 650 hp 650 h p a 13% (without transferred learning) Actual Projected

u

A (with transferred U !learning) c fl — ——___,_ f

^_----______^

-^ -___ - C

317 hp I 420 hp - I 400 hp

1975 1978 T

c o a / a V lu !I 12 13 14 15 16 17 Thousands of Engines (small turboshaft engines) T E-3792 A Figure 29. Comparative average manufacturing costs---CTE and ATE type turbine engines.

hypothetical current technology growth engine (CTE) of greater power, 485 kW (650 hp), has been projected assuming no previous production experience (or transferred learning) applied to this design. The cost of a more advanced technology engine (ATE) of the same power, 485 kW (650 hp), but of different design has also been projected. The two differently designed 485-kW (650-hp) engines (ATE) and (CTE) show different average cost 's for 5000 engines. The cost is approximately 13% greater for the advanced technology engine (Figure 29 point A vs point B). If the 485-kW (650-hp) ATE cost at 5000 units is compared to the current technolog) 485-kW (650-hp) CTE average cost after an- other 5000 small turbines have been produced concurrently, the 485-kW (650-hp) ATE average cost of 5000 units (point A) is 29% greater than the 485--kW (650- hp) growth version CTE average (point C). Despite the 29% lower acquisition cost for the CTE (point C) than for the ATE (point A), the total cost of ownership (TCO) is significantly less for the ATE engines as is shown in Table XVIII.

Engine Maintenance The GATE maintenance cost analysis was divided between turboprop and turbo- shaft powerplants. This division resulted from different typical general aviation usages for twin-engine, fixed-wing and single-engine, rotary-wing applications. Tables XIX and XX show the principal difference in use (hours per year).

y TABLE XVIII. — "OPTIMUM" ENGINES Heavy twin Unpr twin He! twin NOMINAL SP, kW (hp) 820 (1100)* 403 (540) 298 (400) (* Flat Rated) Cyc le 14:1 10:1 14:1 1478 0K (2200 0F) 14780K (22000F) 14780K (22000F) Configuration 2 STG CENTR 2 STG CENTR 1 STG CENTR 2A GPT 2A GPT 1 RADIAL AIRCOOLED 2 SPOOL 2 SPOOL 2 SPOOL Lamiltoy combustor Tech elements Ceramic stators Composite GB Dual property GP turbine production base) Benefits compared to a current technology engine (with sfe improvement, % 20 20 Specific mass improvement, 2 23 23 24 CM, % -21 --11 -12 +1 +5 TAC, % -7 -3 TCO, Z -20 -11 Fuel reduction, % 32 23 24 TABLE XIX. - TURBOSHAFT MAINTENANCE STANDARDS Basic Study - Turboshaft Aircraft Single--engine helicopter Use 360-600 h/yr ORIGINAL PAGE IS OF POOR QUALMf Max operating time 5000 h TABLE XX. - TURBOPROP MAINTENANCE STANDARDS Basic Study - Turboprop Aircraft 'Fain engine Use 600-900 h/yr Max operating time 5000 h It is obvious that the annual utilization for a turboprop vs a turboshaft ap- plication varies from the same (600 h/yr) to almost three to one (900 h/yr/360 vary somewhat as an indirect function of h/yr). Since operating costs may annual utilization, direct comparisons between turboprop and turboshaft en- gines should be done where annual utilization is equal or nearly so.

A maximum operating time (MOT) of 5000 engine operating hours was selected.

This value was used instead of on-condition for several reasons. Two princi- pal ones were the desire to inspect the engines after 5 to 7 yr and the belief that 5000-hr design technology would be pretty well within the state of the art of small engine design by the time that GATE concepts entered production.

The high time operator would achieve the selected 5000-hr MOT within 7 yr. A low time operator or an aircraft in storage might run the risk of dry seals leading to severe oil leaks or engine damage. A few years ago, a major airline experienced a strike which meant that their turboprop-powered aircraft sat about 3 weeks. Upon reentry into service, a series of premature engine removals resulted which were thought to be caused by a "dry bearing" condition at start-up.

DDA has seen evidence of some engine distress through lack of use. Engines that have not been run often can be subject to damage immediately after starting while oil that has drained away from bearings and seals is pumped back into place. Since low annual time operators could fall into this category of inordinate wearout characteristics, a 5 to 7 yr time-related inspection could enhance the safety of operation.

Operators should plan for the benefits of turbine-powered aircraft. TBO val- ues can greatly exceed usual piston engine Levels. Inherent premature removal rates may vary significantly for different missions. Of course, operators are more concerned about operational removal rates since these directly impact aircraft availability and operational cost. Operational rates are the sum of engine-inherent (quality, design life, installation, etc) and engine-noninher- ent (foreign object damage such as rocks or birds, pilot error, improper main- tenance, dirty fuel, conspec oil, secondary damage resulting from aircraft problems, etc). Operational removal rates exclude removals for time (TBO) and convenience. Typical reasons for a convenience removal include using an en- gine from an aircraft on ground (AOG) aircraft to replace a re.noved engine in another aircraft thereby allowing the second aircraft to become operational.

Sensitivity studies can be made to assist potential turbine-powered aircraft users in the selection of engine-aircraft combinations for their intended use. These studies can help illustrate the risk of operating cost variance from the anticipated norm. If a sensitive cost driver, such as a high ratio of take-off power occurrences versus operating time, can be isolated, the user might be able to select equipment/maintenance plans to guard against high costs.

A typical sensitivity chart (Figure 30) shows the relationship of unscheduled engine removals as a part of total engine removals plotted as a function of TBO and premature removal rate for a given engine flying hour per month use.

Close review of Figure 30 shows increasing unscheduled removals after passing 3000-hr TBO and moving toward 7500-hr TBO. The rate of change is pronounced at 0.4/1000 premature removals and becomes greater as premature removal rate increases. Since unscheduled removals plague commercial operators from an aircraft availability and, therefore, revenue generation and reputation for reliability, these factors could play an important role in the decision con- cerning aircraft acquisition and operation.

The results shown in Figure 30 were determined using a DDA operating and sup- port cost computer simulation model (OS 590). This model was furnished to the USAF Jas CDRL Item No. A005 of Contract F33657-77-C-0425, Reduced Cost Turbine Engine Concepts. Various combinations of TBO and premature removal rate were passed through the simulation. Typical results are plotted on Figure 30. The data plotted as intuitively expected until the area for dominance between

u

D

Use = 50 fl hr/,no PRR = Premature Removals/1000 hr TBO = Time $etween Overhaul (hr) T8O = 10,-)00 TBO = 7500

v

TBO = 5000 ORIGINAL PAGE IS v C ^— TBO = 3000 OF POOR QUALRY s ^TBO = 2000 a PRR = 1.0 E o: PRR =0.8 -n J ^i PRR = 0.6 b r O U / PRR = 0.4 S

0 j

PRR = 0.2 0 10 15 20 25 30 Total Removals (hundreds) TE-3793 Fi,ire 30. T30 and PRR relationships--100 A/C, 15-yr forecast.

scheduled removals versus unscheduled removals was reached. The area of in- stability showed that spending money to increase TBO values substantially be- yond 5000 h had little payoff for operational availability.

The maintenance studies resulted in a recommendation set depicted in Table XXI. Again, turboshaft and turboprop distinctions are made. In addition, a furthex distinction results from the recognition of two major user class-s - corporate and non corporate. The turbos}:art engine usually lives in a more unfriendly environment than the typical turboprop. Rotary-wing vibration fixed-wing levels. Demanding mission profiles occur levels generally exceed more often in a rotary-wing appplication. Therefore, a maintenance driver, Unscheduled Removal hate, is substantially higher for turboshaft engines than a comparable technology turboprop engine.

The random removal rate leads to a lower noncorporate TBO for the turboshaft.

Corporate TBO was recommended at 3000 h. The tendency is toward super safe- tv. This attitude is most understandable when the consideration is made for risk of loss of senior people. This attitude is the result of the importance placed on continuity of operations involving key management personnel avail- abi?lLy. A similar consideration toward extra safety is shown by municipal and state aircraft operators who are entrusted with the safety of elected and other key government officials.

Finally, the calendar time/TBO relationship is influenced b y other co-a- siderations. Tax write-offs, resale value, use, and many u"her factors influence the values recommended. The financial opportunities vary for corporate and -ioncorporate operators. The recommended overhaul times were selected to give a nominal best business fit for each user and powerplant combination.

ar i i TABLE XXI. - SCHEDULED OVERHAUL RECOMti iIENDATIONS Recommendations Turboshaft Turboprop Scheduled overhaul Corporate 3000 h 3000 h 5 yr 5 yr Noncorporate 4500 h 6000 h 5 yr 5-7 yr Unscheduled removal rate Random removals/1000 h 0.3-0.5 0.2- 0.4 (INHERENT, MATURE) • Mature engine commonly defined as 10 6 operating hours.

• Inherent engine removals exclude: hod, error, etc.

The engine removal plan follows from the TBO and type of application assump- tions. The premature removal rate (PRR) for each basic engine configuration was estimated. A typical set of PRR estimates is depicted by Table XXII. It is readily seen that turbine temperature is a ?

E RR driver. A more careful re- view shows that the PRR threshold is higher for a turboshaft engine than for a turboprop. This threshold difference may be generally attributed to the more unfriendly environment common to rotary-wing applicatious.

TABLE XXII. - PROJECTED PREMATURE REMOVAL RATES Turboshaft Turboprop ENG PRR RIT* ENG PRR RITE' ENG PRR RIT* ENG PRR RIT* 1 0.53 (1) 16 (1) 0.39 (1) 0.52 1 16 0.38 (1) 2 0.56 (2) 17 0.60 (3) 2 0.41 (2) 17 0.44 (3) 3 0.61 (3) 18 0.53 (1) 3 0.45 (3) 18 0.39 (1) 4 0.55 (1) 19 0.61 (3) 4 0.40 (1) 19 0.45 (3) 5 0.58 (2) 20 0.55 (1) 5 0.43 (2) 20 0.41 (t) 0.63 (3) 20A 0.50 (1) 6 0.46 (3) 20A 0.41 (1) 7 0.59 (1) 0.63 (3) 21 7 0.42 (1) 21 0.47 (3) 8 0.62 (2) 22 0.55 (1) 8 0.45 (2) 22 0.41 (1) 9 0.68 (3) 23 0.63 (3) 0.49 (3) 9 23 0.47 (3) 10 0.62 (1) 24 0.53 (1) (1) 0.30 10 0.45 24 (1) 11 0.66 (2) 25 0.61 (3) 11 0.48 (2) 25 0.45 (3) 12 0.72 (3) 26 0.53 (4) 12 0.53 (3) 26 0.45 (4) 0.68 13 (1) 27 0.53 (4) 13 0.49 (1) 27 0.45 (4) 114 0.72 (2) 28 0.53 (4) 14 0.53 (2) 28 0.45 (4) 15 0.80 (3) 29 0.50 (1) (3) 15 0.59 29 0.41 (1) 30 0.53 (4) * RIT - 0) 1339 K(1950 0 F); (2) 1450 x(2150 0 (3) 1561 x(2350 (4) 1478 F); 0 F); K(22000F).

I The maintenance action distribution is a relative cost driver. A common main- tenance plan is assumed for both turboprop and turboshaft engines. The dis- tribution shown by Table XXIII provides for overhaul for time expired (TB0) removals and a'variable maintenance level distribution for PRR. The distribu- tion is really a function of failure cause and associated repair level. Val- ues used are nominal experience for small 298 kW (400 hp) class turboshaft and turboprop engines.

TABLE XXIII. -- MAINTENANCE PLAN • Engine Removal Plan TBO = 5000 h PRR = function of use and temperature Y Distribution of Maintenance Repair Level Time expired Depot/distributor Premature removal, Depot/distributor overhaul 40 Depot/distributor major repair FBO major repair 10 FBO minor repair FBO = Fixed Base Operator The functional maintenance was valued with respect to dollar cost. Each cost element was related to OEM price as shown in Table XXIV except for nominal FBO repair costs and engine removal and installation. The values shown are typical of total cost. No attempt was made to apportion labor and material costs for each repair level. It was further assumed that capital investment is recovered through burden or overhead charges contained in the overhaul /re-- pair cost structure.

The OEM price was related to factory cost for the en g ine. A 0.52 factor (Ta- ble XXV) was used to provide dollars for all elements above cost to manu- facture. Although values for each individual element may change from engine manufacturer to manufacturer, this factor was considered representative to support this program trade study.

TABLE XXIV. - BASELINE MAINTENANCE INPUT Distributor 0/H COST = (.65) (OEM price) Distributor Major repair cost = (0.33) (O/H cost) FBO Major repair cost = $500 FBO Minor repair cost = $100 Engine installation = $50 + 4 m-h at $20.00ih Engine removal = 3 m-h at $20.00/h L FBO corresponds to Fixed Base Operator TABLE XXV. - PRICE STRUCTURE List Price = OEM price x 1.5 Factory Cost X 1.52 = OEM price 0.52 factor includes: • General and administrative a Profit * Product liability 0 Warranty Current Technology (Baseline) Engine An engine configuration was selected to represent current technology in the small gas turbine field. It was used as a baseline engine for comparative purposes with the advanced technology matrix and candidate study engines.

Pertinent performance, size, mass and cost data for the CTE are shown in Ta- ble XXVI and XXVII at a shaft power size of 373 kW (500 shp).

TABLE XXVI. - CURRENT TECHNOLOGY ENGINE DATA (SI units) 1-stage centrifugal compressor 2-stage axial gas generator turbine 2-stage axial power turbine (unity size) Engine identification CTE CTE^ Performance, slss T.O.

Turboshaft and turboprop w/o prop gearbox loss Rc 8.5 8.5 RIT, K 1316 Shaft power, kW 373 373 sfc, /Ig/Ws 103 103 Turboshaft engine data (6000 rpm output) Mass, kg 86.6 86.6 Length, m 0.91 0.91 Diameter, m 0.56 0.56 OEM price, $ 49,399 56,424 Maintenance cost, $/fl hr 30 h/mo util 23.37 26.67 50 h/mo util 22.45 25.63 Turboprop engine data (incl prop gearbox. 2000 rpm output) Mass, kg 105 105 Length, m 1.11 1.11 Diameter, m 0.54 0.54 OEM price, $ 64,709 73,910 Maintenance cost, $/fl hr 50 h/mo util 25.16 28.73 75 h/mo util 25.20 28.77 CTE--Current Technology Engine CTE^---Current Technology Engine with price adjusted to no prior manufacturing experience basis.

Study Engines A matrix of 22 turboshaft engines and their turboprop deriuities were selected for the initial phase of the Task II Broad Scope Trade-Off Studies. The ma- trix consisted of 12 two-stage centrifugal compressor engines (Nos. 4 through 15), and 10 single-stage compressor engines (Nos. 16 through 25). The unity size of these engines was approximately 615 kW (825 hp). Unity size perform- ance, mass, price, dimensions, and maintenance cost were estimated for each of these engines for use as input data into the mission trade-off studies.

^^©^ 51

TABLE XXVII. - CURRENT TECHNOLOGY ENGINE DATA (Customary units) 1-stage centrifugal compressor 2-stage axial gas generator turbine 2-stage axial power turbine (unity size) Engine identification CTE CTE* Performance, slss T.O.

Turboshaft and Turboprop w/o prop gearbox loss 8.5 8.5 R RIT o f 1910 1910 shp 500 sfc, Ibm/shp • h 0.61 0.61 Turboshaft engine data (6000 rpm output) Weight, Ibm 191 191 Length, in. 35.8 35.8 21.9 Diameter, in. 21.9 OEM price, $ 49,399 56,424 Maintenance cost, $/fl hr 23.37 30 h/mo util 26.67 50 h/mo util 22.45 25.63 Turboprop engine data (incl prop gearbox 2000 rpm output) Weight, Ibm 232 Length, in. 43.7 43.7 Diameter, in. 21.2 21.2 OEM price $ 64,709 73,910 Maintenance cost, $/fl hr 50 h/mo util 25.16 28.73 75 h/mo util 25.20 28.77 CTE--Current Technology Engine CTE*--Current Technology Engine with price adjusted to no prior manu€acturing

experience basis

The performance was obtained by running computerized cycle calculations at the slss T.O. design point and at a series of flight Mach number, altitude, and power setting conditions applicable to the particular mission understudy. For turboprop input data, shaft power was adjusted for prop gearbox loss and con- verted to thrust by assuming the following propeller efficiencies: Flight Condition prop Efficiency Climb-one engine operative 0.78 Climb--all engines operative 0.83 Cruise--all engines operative 0.915 The effect of power and bleed air extraction was included in the cycle calculations. The requirements utilized in this study are listed below and apply to the unity engines.

Fixed-wing Rotary-wing applications applications j Power extraction/engine, kW (hp) 2.94 (4) 2.94 (4) Bleed air extraction, kg/sec (lbm/sec) 0.045 (0.1) 0 (0) The unity size mas-^s were estimated with a computerized components mass anal- ysis program that takes into account a complex set of mass-sensitive parame- ters. The unity size prices were estimated by first estimating the factory costs using a cost-estimating computer program with input that includes compo- nent mass - and material index factors (MIF ` s). The factory costs were convert- ed to OEM selling prices by applying factors to account for such items as:

• General and administrative expenses

• Product liability • Warrantee ORIGINAL PAGE iS • Development OF poOR QUALM

• profit

The estimated engine dimensions were a by-product of the mass analysis. The engine maintenance costs were estimated with an operational and support com- puter program.

Dual- and Single-Stage Compressor Engine Matrix i Table XXV-111, gives an overview of the configuration, pressure ratio, and tur- bine rotor inlet temperature variations in the 22 matrix engines, 4 through 25.

TABLE XXVIII. - OVERVIEW OF MATRIX ENGINES (TURBOSHAFT AND TURBOPROP) UNITY SIZE : APPROX 615 kW (825 hp) Ident. Comp Turb No. of Stages and Types No.

Rc RIT, K ( OF) Comp Power Turb G.G. Turb x 4 10 1339 (1950) 2-C 2-A 2-A 5 10 1450 (2150) 2-C 2-A 2--A 6 2-C 2-A 2-A 10 1561 (2350) 7 12 1339 (1950) 2-C 2-A 2-A 8 12 1450 (2150) 2-C 2-A 2-A 9 12 1561 (2350) 2-C 2-A 2-A 14 1339 (1950) 2-C 2--A 2-A Y TASTE XXVIII . -- (CONT ) !dent.

Comp Turb No. of Stages and Types

^.i

No.

PIT, K ( OF) Camp G.C. Turb bower Turb RC i; 11 (2150) 2-A 2-A 14 1450 2-C 14 1561 (2350) 2-C 2-A 2-A 13 16 1339 (1950) 2-C 2-A 2--A 14 1450 (2150) 2-C 2-A 2-A 15 16 2-C 2-A 2--A 1561 (2350) 16 10 1339 (.950) 1-C I-RI 2-A 17 10 1-RI 2-A 1561 (2350) I-C i 18 10 1339 (1950) I-C 2-A 2-A f - 19 10 C 2-A 2-A 1561 ( 2350) I 20 10 1339 (1950) I-C I-A 2-A 10 1561 (2350) 1-C I-A 2-A 22 8.5 1339 I-C 1--A 2--A (1950) 8.5 1561 (2350) 1-C I-A 2-A 24 1-C 1-A 5.5 1339 (1950) 2-A 25 5.5 1561 (2350) I-C I-A 2-A A--Axial flow C-Centrifugal RI-Radial In,-low Performance and Cost Comparison of Dual and Single Stake Figures 31, 32 and 33 show the relative sfc, specific mass and specific cost of the 12 two-stage compressor engines. Figures 34, 35, and 36 show similar data for the 10 single-stage compressor engines.

sIss T.O.; Shaft Power Approx 615 kW (825 hp) Two-Stage Centrifugal Compressors Two-Stage Axial Turbines SIT 'Matrix Eng. No. 9 = 100°!0 1561 K (2350°F) 1450 K (2150°F) 1339 K (1950°F) fl 1339 K (1950°F) W.

1450 K (2150°F) 1561 K (2350°F) 9$ 6 8 10 12 14 16 Compressor Pressure Ratio TE-3786 Figure 31. GATE parametric engine study--two-stage centrifugal compressor (engine sfc).

slss T. 0. Shaft Power Approx 615 kW (825 hp) RIT i30 J:r 1339 K 0950'F) Two-Stage Centrifugal Compressors e^ Two-Stage Axial Turbines i m d Ly N jd 1450 K b (2150°F) u 'a m o_ V1 Q1 G_ W 1561 K (2350°F) 4 6 8 10 12 14 16 Compressor Pressure Ratio TE-3787 Figure 32. GATE parametric engine study--two — stage centrifugal compressor (engine mass).

e f sIss T. 0. Shaft Power Approx 615 kW (825 hp) c :'" - Two—Stage Centrifugal Compressors OF ^ a ^^"^ i25 Fs f^ Two—Stage Axial Turbines ' y !

x•>!

.

RIT 120 1339 K (1950°F) r; o n.

115— U o LI 1450 K (2150°F) a v u - 105 ^U U Q V'f 1561 K (2350°P) a

G_ _

G w 8 10 12 14 16 Compressor Pressure Ratio TE-3788 Figure 33. GATE parametric engine study--two-stage centrifugal compressor (engine cost).

slss T.O. Shaft Power Approx 615 kW (825 hp) One-Stage Centrifugal Compressor ,o Turbine Stages RIT u w 1339 K (19507) 8 ^^^^2 Axial 101] 1339 K (1950°F) 1 Axial 1561 K (2350°F) 2 Axial 1561 K (2350°F) ^^'" • i Axial I Radial 1339 K (1950°F) • l Radial 1561 K (23507) 4 6 8 10 14 12 16 Compressor Pressure Ratio T E-3789 Figure 34. GATE parametric engine study---one — stage centrifugal.

compressor (engine sfc).

Shaft Power Approx 615 kW (825 hp) Orie-Stage Centrifugal Compressor Turbine RIT Stages Two-Axial One•-Radial 1339 K (1950 °F) too ^^ ^^ One-Axial q a N 95 ii u

q

v

m One-Radial Two-Axial ^11561 K (2350°F) ^ ^^--^} One-Axial ea

I 1 i

4 6 8 10 12 i6 Compressor Pressure Ratio TE-3790 Figure 35. GATE parametric engine study---one--stage centrifugal compressor (engine mass).

Shaft Power Apprax 615 kW (825 hp) One-Stage Centrifugal Compressor Turbine 105 Stages RIT *0 o^ .0 Two-Axial ^^^ - 'Q One Axial 1339 K (19500F) U 95 OOne-Radiai U 0.,` u 1 r^ "" ORIGINAL PAGE IS N ^^^ ''••••••••• •Q Two-Axial OF POOR QUALITY 1561 K (2350°F) s5 One-Radial s0 6 d 10 12 14 Compressor Pressure Ratio TE-3791 Figure 36. GATE parametric engine study--one-stage centrifugal compressor (engine cost).

In general these curves show the following: • sfc decreases as compressor pressure ratio increases except for low tur- bine temperatures at high pressure ratios.

• Engine-specific mass decreases as turbine inlet temperature increases, • For the single-stage compressor engines, specific mass decreases with compressor pressure ratio up to a range of 8.5--9.5 Rc then increases for higher pressure ratios.

• For the two-stage compressor engines, the specific mass increases with compressor pressure ratio over the 10 to 16 Rc range studied except for the high turbine inlet temperature (1561 K (2350 0F)) engine that bot- toms out at about 12 Rc.

• The specific cost trends follow the specific mass trends already discuss- ed.

• For the single--stage compressor engines, the sfc and specific mass de- crease progressively as the gas generator turbine configuration is changed from two-stage axial to one-stage axial to one-stage radial.

Mission Evaluation of Dual Stage The dual-stage centrifugal, compressor engines (12 engine matrix) were evaluat- ed in each of the four faxed-win g and two rotary-wing vehicle/mission applica- tions.

The results for the unpressurized twin are shown in Figures 37 and 38. Figure 37 shows normalized values of aircraft design gross mass (gm), engine size specified in terms of rated shaft power at sea level static—standard day con- ditions (SP), and total aircraft (airframe plus engine) acquisition cost (TAC). Figure 38 completes the presentation of the unpressurized twin econo- mic results in terms of direct operating cost (DOC), total cost of ownership (TCO), and cash flow requirement (CFR). Both figures plot these normalized results as a function of the design cycle parameters compressor pressure ratio (Rc) and turbine rotor inlet temperature (RIT). The cycle trend shown in these figures indicate a preference for high turbine inlet temperature and a relatively flat trend in Rc for GW, SP, and TAC. The DOC, TCO, and CFR re- sults indicate a preference for Rc around twelve. It is noted that the GW, SP and TAC generally follow previously presented engine specific mass, speci- fic fuel consumption, and specific cost trends. However, the DOC, TCO, and CFR are showing the influence of "scatter" in the engine maintenance cost trends U .e., crossover in the 1450 and 1.561 K (2150 and 2350 0F) RIT results at loos ' Rc values).

Figure 39 shows the results for the pressurized light twin application. Be - cause of the similarity in trends and a desire to reduce the amount of data presented, only GW, TAC, and TCO results will be shown for the remaining fixed— and rotary—wing applications. The cycle trends in Figure 39 again indicate a preference for high RIT and a Rc around 12. Figure 40 shows the results for the pressurized heavy twin and Figure 41 the results for the light agricultural application. Generally speaking, the trends for these two appli- cations are similar to the unpressurized twin and pressurized light twin.

Gross Moss Shoff Power Total Aircraft Cost 1.15 1.15 1.15 % l.la l.la 1.10 V o:

0 0

V 1 .05 1.05 1.05 a 1.00 i.OG 1.00 14 10 12 14 10 12 14 16 10 12 16 16 Rc Rc Rc b RIT = 1339 K (1950°F) RIT = 1450 K (2150°F) A RIT = 1561 K (2350°F) TE-3794 Figure 37. Dual--stage centrifugal compressor en g ine matrix--gross mass, power, and total cost of ownership and cash flow requirement treads (unpressurized twin).

l.la i.10 -: Lto O a 1.05 I.05 1.05

U

1.00 1.00 I.aO 12 14 16 12 16 14 16 10 14 10 12 r Ra Re Re O RIT = 1339 K (19507) ORIGINAL PAGE 1.9 0 RIT = 1450 K (2150°F) OF POOR QUALITY 0 RIT = 1561 K (2350°F) TE-3795 Figure 38. Dual-stage centrifugal compressor engine matrix--direct operating cost, total cost of ownership, and cash flow requirement trends (unnressurized twin).

Gross Mass Total Aircraft Cost Total Cost of Ownership 1.10 1.10 1.r0 1.a5 v 1.05p 1.05 O Q U (^^ l`^ Q ^ 1.00 ^ ..^-^----^ 1.00 U 1.00 0. 95 0.95 0.95 14 16 10 12 14 16 10 12 14 Rc Re Re O RIT = 1339 K (i 950°F) q RIT = 1450 K (2150°F) d RIT = 1561 K (2350°F) TE-3796 Figure 39. Dual-stage centrifugal, compressor engine matrix----gross mass, total aircraft cost, and total cost of ownership trends (light twin).

Gross Mass Total :aircraft Cost Total Cost of Ownership 1.15 1.15 1.15 1.10u 1.10 0 1.10

a u

O 1-1.05 1.05 "^' 1.05 1.00 1.00 1.00 12 14 16 10 T2 10 12 14 16 10 14 16 R c RC Rc O RIT = 1339 K (1950°F) C3 RIT = 1450 K (2150°F) A RIT = 1561 K (2350°F) TE-3797 Figure 40. Dual-stage centrifugal compressor engine matrix--gross mass, total aircraft cost, and total cost of ownership trends (heavy twit._).

The results for the two rotary-wing applications are presented in the same format as the fixed-wing trends. Figure 42 shows the results for the light single-engine helicopter, and Figure 43 the light twin-engine helicopter.

Both applications indicate a preference for high RIT and, except for the TCO, exhibit a relatively flat trend for the range of R c values evaluated. With respect to the TCO, the influence of engine maintenance cost trends is again producing crossovers in the 1450 and 1561 K (2150 and 2350 0F) RIT results.

A summary of the cycle selections from the preceding trend curves for each vehicle/mission application are shown in Table XXIX. The selection criteria used was either minimum GPI and TAC or minimum TCO. All applictions examined prefer 1561 K (2350 0 F) RIT regardless of the selection criteria used. The fixed wing applications indicate a preference for R c ranging from 12 to 14, whereas the rotary-wing applications prefer 10 to 14 R c . The overall opti- mum selection from the dual-stage centrifugal engine evaluations was determin- ed to be the high pressure ratio (Rc = 14), high turbine inlet temperature cycle because it tended to provide the lowest gross weight and total aircraft cost at little compromise in total cost of ownership.

Gross Mass Total Aircraft Cost Total Cost of Ownership 1.15 1.15 1.15 * 1.10 d U 1.10 p 1.10

U

U

1.05 O 1.0.5 ti 1.05 1.00 1.00 zr----ter=^ 1.00 12 14 16 12 14 16 10 12 14 16 Rc Rc Rc FZ O O;F 0 RIT = 1339 K (1950°F) P®Ql? QUALITY 17 RIT = 1450 K (2150 °F) a RIT = 1561 K (2350°F) TE-3798 Figure 41. Dual — stage centrifugal compressor engine matrix--gross mass, total aircraft cost, and total cost of ownership trends (light agricultural).

Gross Mass Total Aircraft Cost Total Cost of Ownership • 1.10 1.10 1.10 O 1.05- 1.05 1 1.05 ^a V V t7 ° U 1.00 ¢ 1.00 _--+"^^ 1'..00- 0.95 0.951 0.95 \, 10 14 16 12 14 16 10 12 16 10 12 Rc Rc Rc O RIT = 1339 K (1950°F) a RIT = 1450 K (2150°F) a RIT = 1561 K (2350°F) TE-3799 Figure 42. Dual—stage centrifugal compressor engine matrix--gross mass, total aircraft cost, and total cost of ownership trends (helicopter—light single).

Gross Mass Total Aircraft Cost Total Cost of Ownership 1.10 1.10 1.10 4c * O U 1.05 1.05 01.05 t¢3 u [9 p,^ ' 1 -13 F"'

1.00 1.00 -Y^^- ^°

1.00 12 14 16 10 14 12 16 10 12 R c RC RC O RIT = 1339 K (1950°!=) 17 RIT = 1450 K (21507) A RIT = 1561 K (23507) TE-3800 Figure 43. Dual-stage compressor engine matrix--gross mass, total aircraft cost., and total cost of ownership (helicopter- light: twin).

TABLE XXIX. - SUMMARY OF CYCLE SELECTIONS Lowest GM and TAC Lowest TCO selection selection Rc RIT, K ( OF) Rc RIT, _K (°F) Fixed wing Unpressurized twin 12-14 1561(2350) 12 1561(2350) Light twin 12-14 1561(2350) 12 1561(2350) Heavy twin 12-14 1561(2350) 12 1561(2350) Light agricultural 1561(2350) 12-14 12 1561(2350) Rotary wing Light single 16 14 1561(2350) 10 1561(2350) Light twin 12-14 1561(2350) 10 1561(2350) Table XKX shows a GM comparison for a CTE versus ATE. The current technology engine has a single-stage centrifugal compressor (Rc = 8.5) with a relative- ly low RIT, whereas the advanced technology engine incorporates a dual-stage centrifugal compressor (Rc= 14) with high RIT. The GM comparison in Table XKX lists the magnitude of the mass increase obtained if a CTE is used instead of an ATE in each of the fixed- and rotary-wing applications. It is noted that the heavy twin produces the largest mass advantage for the ATE, i.e., the most stringent requirement in cruise altitude, velocity, range, and payload A.

indicates the Largest payoff for the ATE. Table XXX also lists the ATE sea I level.-rated power sizes for each of the applications. Using data provided in # Table Yom£, three of the six vehicle/mission applications were selected for the t remaining engine studies. The selected vehicle/missions and reasons for their selection follow: e Unpressurized twin---This application shows the largest fixed--wing aircraft 1 GM advantage for the ATE (compared to the CTE) in the 450 kW (600 hp)/and under non-flat-rated engine size.

Pressurized twin (heavy)---This application i i , dicates the largest GM advan- tage for the turboprop ATE.

• Light twin helicopter---This application shows the largest GM advantage for the turboshaft ATE.

The selected missions are denoted in Table XXX by an asterisk.

TABLE XXX. - CURRENT TECHNOLOGY ENGINES VERSUS ADVANCED TECHNOLOGY ENGINES GROSS MASS COMPARISON GM, % ATE (,ATE base) sea level—rated W hp) (flat rated) Fixed wing *^^ +14 365(490) Unpressurized twin +18 410 (550) Light twin ^^ ^^P^.^+^ +31 753 (1010) Heavy twin* Q ^^

P^

+7 507(680) Light agricultural Of Rotary wing +8 276(370) Light single 283(380) +14 Light twin* *Mission selected for further studies.

Mission Evaluation of Single Stage The single-stage centrifugal engines (10 engine matrix) were evaluated in the selected vehicle/miss1_on application, i.e.. unpressurized twin. heavy twin, and light twin helicopter. The results for the unpressurized twin are shown in Figure 44 as plots of normalized GM and TCO. The GM and TCO results are plotted as a function of: • Rc for lines of constant RIT.

RIT for lines representing a high pressure turbine (HPT) configuration for a constant pressure ratio of 10.

The cycle trends shown in Figure 44 indicate a preference for the 1561 K (2350 0F) RIT and an Rc of 10.

The radial inflow is indicated to be the preferred EPT configuration. The preference indicated for the radial inflow HPT configuration is a result of lower sfcs, acquisition costs, and mainte- nance costs relative to the one- and two-stage axial HPT. Comparison of the one- and two--stage axial to the radial inflow configurations shows the axial configurations to have 2 to 4% higher sfcs, 5 to 7% higher engine acquisition III costs, and 8 to 12% higher engine maintenance costs. The larger advantage for the radial inflow HFT is at 1950 OF because sfc and cost improvements are greater than the improvements indicated at 23500P.

Figure 45 shows the results for the heavy twin and Figure 46 the results for the light twin helicopter. In general, the trends for these two applications are similar to the unpressurized twin. It is noted that the heavy twin pro- vides the largest change in the normalized GM and TGO results (i.e., the sen- sitivity to changes in design cycle parameters is double that indicated for the light twin helicopter). Cycle selections from the preceding trend curves for each vehicle/mission are listed in Table XXXI under the selection criteria used. The overall optimum selection for the single-stage centrifugal compres- sor engine configuration was determined to be the 10 Rc , high RIT cycle with a radial inflow RPT.

Q Radial Inflow GGT .a Two—Stage Axial GGT O One—Stage Axial GGT Total Cost of Ownership Grow Mass 1.10 1.10 1.05 1 .05 RIT = 1339 K (1950°F) RIT = 1339 K (1950°F) 4.

4C O U p 1.00 1.00 U f-- A 0.95 0.95 RIT = 1561 K (2350°F) RIT = 1561 K (2350°F) 0.90 0.90-p

6.0 8.0 10.0 1340 1560 K 8.0 10.0 134 `r 0 i 0 K

4.0 4.0 6.0 1950 2350 AF R 2350 OF RIT at 10.0 RC RIT at 10.0 RC TE-3801 Figure 44. Single-stage centrifugal compressor engine matrix -gross mass and total cost of ownership trends (unpressurized twin).

Gross -Mass Total Cast of Ownership 1,10 1.10 RIT - 1339 K (1950°F) RIT = 1339 K (1950°F) 1.05 1.05 A 1.00 1.00

o

t0 u 0.95 0.9S 0,90 0.90 RIT = 1561 K (2350%F) RIT = 1561 K {2350T) r 4.0 6.0 8.0 10.0 1340 1560 K 4,0 6.0 8.0 10.0 1344 1550 K I t Ra 2350 O F Ra 1950 2354 of RIT at 10.0 R= 10.0 Ra RIT at 0 Radial Inflow GGT a Two - 5tage Axial GGT TC-3802 0 One-Stage Axial GGT Figure 45. Single-stage centrifugal compressor engine matrix--grass mass and total cost of ownership trends (heavy twin).

13 Radial Inflow A Two-st age Axial ORIGINAL PAGE IS 0 Ono-Stage Axial OF POOR QUALITY Grass Moss Total Cost of Ownership 1.10 ^ 1.10 1.05 1.05 RIT = 1339 K (1950°F) « _ ` RIT = 1339 K (i950°F) O u ^Y 17, 1.00 1.00 c9 0.95 0,95 RIT = 1561 K (2350'F) 50°F) ^r RIT - 1361 K (2350°F) 0.90 0.90 4.0 6,0 8.0 10.0 ' 1340 1560 K 4.0 6.0 8.0 10.0 1340 1560 K _ Ra R c - i II 2350 Of 1950 2350 of RIT at 10.0 Re RIT at 10.0 Ra TE-3803 Figure 46. Single-stage centrifugal compressor engine matrix--gross mass and total cost of ownership trends (helicopter-light twin).

f i TABLE XXXI.

- CYCLE SELECTION ONE-STAGE CENTRIFUGAL COMPRESSOR MATRIX GM & TAC Lowest TCO Lowest selection selection GGT GGT RC Fixed wing Rc RIT, K( O F) Configuration RIT,K(oF) Configuration 1lnpressurized twin 10 1561(2350) R OR I-A (1) 10 1561(2350) R 10 1561(2350) R OR 1-A 10 1561(2350) R Heavy twin Rotary wing Light twin 10 1561(2350) R OR I-A 10 1561(2350) R generator turbine configuration (1) Gas R = Radial inflow GGT I--A - One-stage axial GGT Candidate Engines Five candidate engines (Nos. 20 through 30, Table XVI--SI units and Table XVII —customary units) were selected for the final phase of the mission trade-off studies. 373 kW (500 hp) shaft power was selected as the design size for these engines. The shaft power size was chosen in recognition of the market potential near this power level and in order to focus on a high per- formance engine design in a significantly smaller power class than is current- ly available or under development.

Performance, mass, dimensions, price and maintenance costs were estimated for the use as input in the mission trade off studies. Estimation procedures were similar to those described earlier for the matrix engine estimates.

Tables XVI and XVII list the configuration type, estimated design point per-

formance, mass, dimensions, price, reliability, and maintenance costs for both

the turboshaft and turboprop versions of these five candidate engines.

It is noted that the TCO and DOC values quoted in the candidate engine compar- isons are for a fuel cost of $0.22/1 ($ 0.83/gal) and utilization rate of 600 hr/yr for the fixed wing applications and 360 hr/yr for the rotary wing appli- cation. The period of ownership is 8 years for both fixed and rotary wing (reference Table XII).

Engine Component Comparisons These five candidate engines provide a basis for the following engine compo- nents comparisons: 1. Axial versus radial inflow HPT configuration selection. Engine Nos, for the single stage 26 versus 27 Centrifugal Compressor engine versus 28 (Design RIT = 1478 K (22000F)) 2. Cooled versus uncooled Engine Nos.

one-stage axial HPT 26 versus 29 Selection for the singl,a-stage compressor 3. Single-stage versus dual-stage centrifugal compressors for the Engine Nos.

dual-stge axial LPT and HPT engines 27 versus 30 Comparison 1 Axial versus Radial Inflow HPT !

- The results of a single stage axial, A two--stage axial, and a single-stage radial inflow HPT engine evaluation in the unpressurized twin are shown in Table XXXII ( i.e. candidate engine 26 versus 27 versus 28). The results are shown as percent changes in GM, TAC, and TCO with the one - stage axial HPT con- figuration ( 26) as the reference. This table indicates the two- - stage axial HPT configuration ( 27) to be slightly heavier in GM (0.2 to 1.0%) and from 1 to 2% higher in TAC and TCO. The one - stage radial inflow HPT ( 28) results indicate a 0.5 to 2% reduction in GM and a 2 to 4% reduction in TAC and TCO when compared with the one - stage axial HPT engine ( 26). Therefore, the on- stage radial inflow HPT configuration ( 28) is preferred over the one or two- stage axial turbines ( 26 and 27) in this comparison.

TABLE XXXII. - AXIAL VERSUS RADIAL INFLOW HPT CONFIGURATION STUDY RESULTS (percent change from one--stage axial HPT engine 26) Mission Heavy Unpressurized Light twin twin twin helicopter Engine ID 27 28 28 27 27 28 HPT configuration 2-A 1-R 2 A 1 - 2-A R I-R GM, % +1.0 -0.5 *0.4 -2.2 +0.2 -2.2 TAC, % +2.0 --1.6 +0.9 +1.5 -4.1 -3.3 TCO, I -2.2 +1.8 +1.3 -4.6 +0.8 --4.0 s Comparison 2 Cooled versus Uncooled HPT An evaluation of a cooled versus uncooled HPT engine configuration (i.e., can- didate engine 26 versus 29). Both engines have a single - stage centrifugal compressor with a design R of 10, a one - stage axial HPT, and a two-stage c axial low - pressure turbine ( LPT). Engine 26 has a design RIT of 1478 K (2200 0F) ( cooled) versus engine 29, which has a design RIT of 1339 K (1950 oF) ( uncooled ) The results of this comparison are shown in Table ' XXXIII as percent changes in GM, TAC, and TCO, with engine 26 as the reference ( and indicate a preference for the cooled HPT turbine 26). Although the en- gine with an Uncooled HPT (29) is competitive i n the heavy twin, it is shown to be from 2 to 3% higher in GM, 5 % higher in TAC, and 4 to 5% higher in TCO for the unpressurized twin and light twin helicopter, respectively.

Engine 29 becomes competitive in the heavy twin because the constant bleed air extraction, 0.045 kg / s (0.1 lbm/sec), produces a relative performance improve- ment for engine 29 when going from the unpressurized twin sizing / cruise alti- € N— _ . is OF POUR QJAUTY tude and Mach 3657.6 m/0.35 Mn (12,000 ft/0.35 Mn) to the heavy twin con- dition 9144 m/0.50 Mn (30,000 ft/0.50 M n). The unity size maximum cruise power TSFC for 29 at 3657.6 m/0.35 Mn (12,000 ft/0.35 Mn) is approximately 0.5% higher than 26, however, at 9144 m/0.50 Mn (30,000 ft/0.5 Mn) the TSFC for engine 29 is approximately 2.5% lower than engine 26. Further dis- cussion of the design trades involved in choosing a cooled or uncooled turbine is provided in tl-e Technology Element Section.

TABLE XXXIII - COOLED VERSUS UNCOOLED TURBINE STUDY RESULTS (Percent change from cooled HPT engine configuration 26) Heavy Unpressurized Light twin Mission twin twin, helicopter Engine ID 29 29 -0.1 +2.3 +2.8 GM, % TAC, % 0 +5.0 +5.3 +4.8 TCO, % --0.2 +4.6 y Comparison 3 Single--Stage versus Dual-Stage Centrifugal Compressor An evaluation of a single--stage versus a dual-stage compressor engine can be made by comparison of configurations 27 and 30, respectively. The single- of 14. Both stage has a design R. of 10 whereas the dual stage has an R c engines have two--stage LPT and HPT and design RIT of 1478 K (2200 0F). The results of this comparison are shown in Table XXXIV as percent changes in GM, TAC, and TCO with engine 30 as the reference.

TABLE =IV. - SINGLE--STAGE VERSUS DUAL-STAGE CENTRIFUGAL COMPRESSOR STUDY RESULTS (Percent change from dual--stage compressor engine 30) Heavy Unpressurized Light train Mission twin twin helicopter Engine ID 27 27 27 Compr.

Configuration one-stage one--stage one-stage +18.9 +7.6 +2.7 GM, % TAC, % +22.7 +11.0 +1.5 +12.6 +2.5 TCO, % +24.1 The results in Table XX€IV indicate a preference for the dual--stage centrifu- gal compressor (30) for all aircraft configurations. Even though GM, TAC, and TCO were only slightly greater (from 1.5 to 2.7%) for the light twin helicop- ter, they progressed to 7.6-12.6% for the unpressurized twin and to 18.9-24.1% for the heavy twin.

Candidate Engine Comparison A summary comparison can be made for these five candidate engines to indicate the best selection for each aircraft application. Normalized GM and TCO re- sults for each of the four single--stage centrifugal compressor (Rc = 10) engines plus one dual-stage centrifugal compressor (Rc = 14) engine are shown in Figures 47 through 52. Figures 47 and 48 show the results for the unpressurized twin, Figures 49 and 50 show results for the heavy twin, and Figures 51 and 52 the results for the light twin helicopter application. Fig- ures 47 through 50 indicate the dual-stage centrifugal candidate engine 30 to be the optium selection and the single-stage centrifugal candidate engine 28 to be the "next best" in the fixed-wing applications. Figures 51 and 52 indi- cate candidate engine 28 to be the optium selection and engine 30 to be a close "next best" in the light twin helicopter.

Engine 28 performed better than 30 in the twin helicopter because no bleed air was provided by the engines, whereas, in the fixed wing applications, bleed air was extracted from the engines. As a result, installed performance for engine 30 was better than 28 in the fixed wing vehicles; however without bleed extraction, the performance of engine 28 was calculated to be approximately equal to that of engine 30.

1.20 Gross Mass (7720 f bm) GM * - 3502 kg ^^^^'^'^,q 1.10 G 7.00

1 R IA 2A +GGT Stages

1A 2A

1478 1478 + K 1478 1478 1339 0.90

RIT

(2200) (2200) (1950) (2200) ----'F

(2200)

2-Stage Compressor -------^-^-= ^

1-Stage Compressor Re =10 Rc-14 29 30 2b 27 28 Engine Cycle Number TE-3804 Figure 47. GATE candidate engines-gross mass trends (unpressurized twin).

Total Cast of Ownership 1.20 TCO* = $909,300 x 3.10 O

U

O

U

r—

1.00

GGT Stages 2A 1R 1A 2A IA .1478 1478 1478 1339 1478 K RIT 0.90 CF (2200) (1950) (2200) (2200) (2200) 2-Stage Compressor 1-Stage Compressor -- T^ ^^----- - RC = 14 RC = 10 26 28 29 30 Engine Cycle Number TE-3805 Figure 48. GA'Z'E candidate engine--total cost of ownership trends (unpressuriaed twin).

Gross Mass GM* = 4218 kg (9300 Ibm) 1.20 1.10 (7 1.00 ^-- GGT Stages 12A IA 2A 1 R lA 1478 1478 1339 1478 -^ K RIT 0.90 °R (2200) (1950) (2200) (2200) (2200) T-Stcge Compressor--- -----+I 2-Stage Compressor I- R = 10 R =•14 c c 28 30 26 27 29 Engine Cycle Number TE-3806 Figure 49. GATE candidate engines--gross mass trends (heavy twin).

Total Cost of 1, 3a Ownership .

TCO* $1,652,960 1.20 — 1.10 U 1.00— 0.90 =7 10 Rc Rc =

27 @ @ @

Engine Cycle Number TE-3807 Figure 50. GATE candidate engines - total cost of ownership trends (heavy twin).

Gross Mass (4310 16m) GM* = 1955 kg 4k ^E 1.00 2A IA ]A 1 2A - ---GGT Stages 1478 1478 0.90 1478 1339 1478 — K RIT - (2200) (2200) (2200) (1950) (2200) —'F 2-Stage Compressor -]-Stage Compressor Rc = r = 10 14 R 26 28 29 TE-3808 Engine Cycle Number Figure 51. GATE candidate engines-gross mass trends helicopter-light twin).

( Total Cost of Ownership 1,20 TCO* = 5613,590 1.10 4t O

U

O ^ i.00 G GT Stages 0.90 RIT OF ]-Stage Compressor 2-Stage Compressor Rc = 10 Rc = 14 26 27 28 29 30 Engine Cycle Number TE-3809 Figure 52. GATE candidate engines-total cost of ownership trends (helicopter-light twin).

Table XXXV shows the effect of bleed and poorer extraction on shaft specific fuel consumption (sfc) for engine cycles 28 and 30.

TABLE XXXV. - INSTALLATION EFFECT ON SHAFT SPECIFIC FUEL CONSUMPTION (Unity size engine) Cycle 28 30 Power Exc, kW ( hp) 0(0) 2.98(4) 2.98(4) 0(0) 2.98 ( 4) 2.98(4) Bleed--(lbm/sec) 0(0) 0(0) 0.045(0.1) (';d) 0(0) 0.045(0.1) sfc @ slss T.O.P. 486 0 . 0 . 492 0 . 529 0.488 0 . 491 0.509 Bleed, % of 0 0 3.8 0 0 3.3 WITOT At slss/takeoff power setting, with zero bleed, engines 28 and 30 sfc values are essentially the same. When 0.045 kg/s (0.1 lbm/sec) bleed is extracted, the sfc for engine 30 is shown, to be 4% lower than that of engine 28. (Note that bleed, when expressed as a percent of engine airflow, is 3.8% for 28 and 3.3% for 30.)

It is noted that bleed air for the 10:1 engine (28) was extracted at compres- sor discharge and interstage for the 14:1 engine (30). The performance of 28 could probably be improved by bleeding the diffuser at a lower pressure than compressor discharge.

ATE Versus CTE Study Results A comparison of the engine characteristics for the GATE-selected engines (28 and 30) versus two versions of a current technology engine (with and without production base) are shown in Table XXXVI and Table XXXVII. Table XXXVI shows the turboprop characteristics and Table XXXVII the turboshaft. Note that the asterisk indicates the no production base CTE. Figures 53 and 54 show the relative GM and TCO results in bar chart form for each engine and mission ap- plication examined. Relative DOG, TAG, and fuel are also shown in Figures 55, 56, and 57. Figure 53 indicates significant GM advantages for both engines 28 and 30 in all three mission applications. The reductions in gross mass are summarized in Table XXXVIII.

ORIGINAL PAGE IS OF POOR QUALITY TABLE XMI. -- ENGINE COMPARISONS (TURBOPROP) Identification 28 30 Technology ATE ATE CTE* CTE Type of compressor I-C 2-C I-C 10C Type of GP turbine 1-R 2-A 2-A 2-A Compressor • , ,essure ratio 10 _4 8.5 8.5 GP rotor inlet temp, K( OF) 1478(2200) 1478(2200) 1316(1910) 1316(1910) Air--cooled GP turbine 1-R 2-A none none Shaft power, slss, kW(hp) 373(500) 373(500) 373(500) 373(500) sfc, pg/W s (Ibm/hp-h) 82.0(0.485) 82.3(0.487) 103(0.61) 103(0,61) Mass, kg(.bm) 79.4(175) 81.2(179) 105(232) 105(232) OEM price, 1978 $ 70714 76401 73910 64709 Specific mass, g/W(ibm/hp) 0.21(0.35) 0.218(0.358) 0 282(0.463) 0.282(0.463) OEM specific price, $/kW($/hp) 190(142) 205(153) 198(148) 173(129) Maintenance cost (50 hr/mo util), $ fl./h 17.26 18.64 28.73 25.16 TBO, h 5000 5000 1500 1500 (TURBOSIFAFT) ENGINE COMPARISONS TABLE XXXVII. - Identification 30 Technology ATE ATE CTE CTE* Type of compressor 1-C 2--C 1-C 1-C Type of GP turbine I--R 2-A 2-A 2-A Compressor pressure ratio 10 14 8.5 8.5 GP rotor inlet temp, K( OF) 1478(2200) 1478(2200) 1316(1910) 1316(1910) Air-cooled GP turbine 1-R 2-A none none Shaft pawer clss, kW(hp) 373(500) 373(500) 373(500) 373(500) Airflow, kg/; (lbm.sec) 1.27(2.79) 1.40(3.09) -- -^ p g/Ws(,bm/hp • h) 82.0(0.485) 82.3(0.487) 103(0.61) 103(0.61) a ft , kg (11m) 65.8(145) 67.1(148) 86.6(191) 86.6(191) Mass, OEM price, 978 $ 58830 63562 56424 49399 Specific mans, g/W(Ibm/hp) 0.176(0.290) 0.180(0.296) 0.232(0.382) 0.232(0.382) OEM specific price, $/kW($/hp) 158(118) 170(127) 152(113) 133(99) h 17.42 18.81 23.37 Maintenance cost (30hr/mo util) $/f1 26.67 TBO, h 5000 5000 1500 1500 TABLE XXXVIII. - PERCENT GM REDUCTION (CTE as reference gross mass) Engine ID 28 30 Heavy twin 8 21 Unpressurized twin 7 11 Light twin helicopter 12 Figure 53 also indicates the preference for the deal-stage centrifugal engine (30) over the sin g le-stage centrifugal engine (28) in both fixed-wing mis- sions. The twin helicopter mission results show both 28 and 30 at the same GM.

Figure 54 indicates significantly lower TCO results for both engine 28 and 30 in all three missions applications. The TCO reductions are summarized in Table XXXIX.

TABLE XXXIX. - PERCENT TCO REDUCTION (CTE with production base as reference) Engine ID 30 Heavy twin 20 Unpressurized twin 11 Light twin helicopter 8 6 Figure 54 also indicates a greater TCO advantage for the dual-stage centrifu- gal engine configuration (30) in the fixed-wing missions. The twin helicopter results indicate the single-stage centrifugal engine (28) to have a TCO ap- proximately 2.5% lower than the dual-stage centrifugal engine.

! Heavy Twin Helicopter - Twin Unpressurized Twin I GM* = 4218 kg (9300 Ibm) GM* = 3502 kg (7720 Ibm) GM* = 1955 4 (4310 Ibm) 1.30 1.20 1.10 t, 1,00 1 1 1 11 1 ( 1 I 1 1 1 Y 1 1 1 1 I 1 I 0.90 CTE* GTE 28 30 GTE* CTE 28 30 CTE* CTE TE-3810 • -- No Produclian 8asc Figure 53. Gross mass comparison--best advanced technology engines versus current technology engines.

UnpromrJud Hallcoplat - Twin Twin TCO• = S6131590 0 TC0 . & S9091300 1.30 1.20 1.00 i.

0.90 2 0 8 ( D 28 CT E* CTE 28 30 CrE • CTE CTE • CTE

&

TE-3811 No Fmdwctlen Base Figure 54. Total cost of ownership comparison--best advanced technology engines versus current technology engines.

Unpressurized Twin Helicopter--Eight Twin DOC*= $267.30 DOC* = $202.90 1.2 IAt p IS 49Z D;?

QAt/7-v 1.1 1.0 . I . . . I . . I . . .

1 1 1 I 30 CTE* CTE CTE* CTE

30 28 30 28

CTE* CTE

(D a

'qh — No Production Base T E-5296 Figure 5j. Direct operating cost comparison--best advanced technology engines versus current technology engines.

L 1.3 Helicopter--Light Twin Unpressurized Twin = $421,810 TAC* = $390,860 TAC* 1.2 U Q ^ 1.1 U Q 1.0 0.9' 28 30 CTE* CTE 28 30 CTE* CTE 28 30 CTE* CTE * — No Production Base TE—.5297 Figure 56. Total aircraft cost comparison--best advanced technology engines versus current technology engines.

Helicopter—Eight Twin Heavy Twin Unpressurized Twin Fuel* = 1665 i6 Fuel* = 970 16 Fuel* n 2200 16 1.4 1.3 m e 1.2 1.0 34 CTE* C E 28 30 CTE* CTE CTE* CTE 28 (DS No production Base — T E-5298 Figure 57. Mission fuel comparison--best advanced technology engines versus current technology engines.

The DOG trends shown in Figure 55 essentially duplicate the trends shown by the TCO results in Figure 54. The TAC results shown in Figure 56 indicate the current technology engine with production base to be the least cost selection for the unpressurized twin and the Light helicopter applications. However, for the relatively more expensive heavy twill aircraft, the advanced technology dual centrifugal compressor engine is shown to be optimum. Figure 57 indi- cates the large reduction in mission fuel usage obtained with the advanced technology engines (28 and 30) when compared to the current technology engines in each of the three applications. This trend in higher fuel usage is a sig- nificant factor in the higher TCO and DCO levels calculated for the current technology engines.

Tables XL and XLI show a comparison* of a CTE (with and without production base) and a two-stag, ca?,..t fugal compressor, high pressure ratio, high tur- bine temperature ATE in the unpressurized twin vehicle mission. These tables are presented in order to illustrate the magnitude of the various cost compo- nents calculated in this study.

SUMMARY TABLE XL. - UNPRESSURIZED TWIN and fuel. = 0.22 $/l (0.83 $/gal) Use = 600 h/yr Representative ATE CTE CTE* ENGINE - ID 3368(7425) 3835(8455) 3835(8455) Gross mass, kg (Ibm) (base) (0) GM, X Required SL-rated SP, kW(hp)

• Initial cruise

400(535) 365(490) (3.658 km(12,000 ft)/113m/s(2ZOkn) /MCP 400(535) • OEI climb 328(440) 369(495) 360(495) (1.524 km(5OOOft)/56.6m/s(11Okn) 1TOP Cruise (3.658km (12,OOOft)/113m/s(22Okn)TAS) 177(46.7) 234(61.8) 234(61.8) • Fuel rate, 1/h (gal/h) • TSFC/power setting, 12.2(.43)/86 12.2(.43)/86 10.2(.36)/87 mg/N's (lbm/lbf • h)/% MCT Price, $ 160,880 160,870 180,870 Airframe

•

231,400 218,560 202,590 a Engines (2) 379,440 383,460 412,270 Total (-1.0) (Base) (+7 5) * TAC, % I83.99 215.72 229.64 DOC, $/fl h (-14.7) (Base) (+6.4) DOC, Z 822,190 1,038,500 TOC, $ 973,940 (+6.6) (-15.6) TCO, % (Base) *Denotes no production base.

Table XL shows the significantly reduced fuel rate for the representative ATE resulting from a combination of reduced aircraft size (12% reduction in GM) and the more fuel efficient engine (16% better TSFC's). Comparison of the Note: This comparison was completed using the mission/cost results obtained from uninstalled engine performance data, i.e., no customer bleed or power extraction penalties. It is also noted that the ATE engine in this study was an early study engine configuration with respect to performance and cost.

Performance was essentially the same. However, the cost numbers are large when compared to the final level of costs for the final candidate ATE engines.

TABLE XLI. - UNPRESSURIZED TWIN SUMMARY-- CONTINUED DOC BREMOWN--$/fl h Util = 600 hr/yr and fuel = $0.22/1 ($0.83/gal) ENGINE CTE CTE* ATE BREAKDOWN A/C Engine A/C ENGINE A/C ENGINE Fuel and oil - 52.34 -- 52.34 --- 39.58 Insurance 3.02 3.38 3.02 3.85 2.68 3.65 A/C Maintenance 18.16 - 18.16 - 17.47 - Engine Maintenance - 52.51 -- - 59.96 35.17 Depreciation 37.68 42.21 37.63 48.21 33.52 45.54 - Registration fee 0.52 -- 0.52 0.47 - Hangar rental 5.90 - 5.90 -- 5.91 Subtotals 65.28 150.44 65.28 164.36 60.05 123.94 A/C plus eng total 215.72 229.64 183.99 *Denotes no production base.

production to the no production base CTE shows a 7.5% increase in TAC and an approximate 6.5% increase in DOC and TCO. Comparison of the CTE (with produc- tion base) and the ATE shows a 1% reduction in TAC for the ATE. The DOC and TCO comparison shows an approximate 15% reduction for the ATE.

Table XLI shows the aircraft and engine DOC breakdown for the representative ATE and current technology engines. This table shows that 70% of the total DOC is engine oriented. Approximately one-third of the engine-oriented DOC is fuel and oil cost, one-third is maintenance, and one-third depreciation costs. The engine maintenance costs (MC) tabulated in Table XLI are mainte- nance costs for two engines sized to meet the unpressurized twin requirements.

Turbine Engine versus Piston Engine Trade-Offs The competitive positions of an advanced turbine engine (Candidate Engine 30) relative to a "typical" current and advanced (1988 technology) naturally aspi- rated piston (NAP) engine was studied in both the unpressurized twin and light twin helicopter applications.

Table XL1I shows the ATE and assumed NAP engine characteristics used in the study. These data indicate significant advantages for the piston engine in both sfc and price along with a large weight disadvantage with respect to the ATE. It is noted that TCO and CFR values could not be calculated in this com- parison due to the lack of piston engine maintenance cost. The piston engines were assumed to have a cooling air drag equivalent to 7% of the total aircraft drag and a prop efficiency of 85%. The lower prop efficiency was used because of the thicker blade section required to absorb piston engine firing order stresses.

Figure 58 shows the gross mass (GM) and total aircraft cost (TAC) results for the unpressurized twin application. Both current and advanced piston engines are shown to have a significantly higher GM than the ATE (30 and 7%, respec- tively). However, the piston engines indicate a 35% ro 45% lower TAC depend- ing on the technology level.

TABLE 1XLI1. -- ENGINE COMPARISONS TURBOPROP VERSUS PISTON 30 P1 P2 Identificaton NAP-C NAP-A Technology ATE Type of compressor 2-C 2-A Type of GP turbine Compressor pressure ratio GP rotor inlet temp, K( O F) 1478(2200) 2-A Air-cooled GP turbine Shaft power, slss, kW(hp) 373(500) 298(400) 298(400) sfc, ug/w-s (lbm/hp-hr) 82.3(0.487) 68(0.40) 59(0.35) Mass, kg(lbm) 81.2(179) OEM price, 1978 $ 76401 Specific mass, g/W(lbm/hp) 0.218(0.358) 0.821(1.35) 0.55(0.9) OEM specific price, $/kW($/hp) 205(153) 31(23) 31 (23) 18.64 ?

Maintenance cost (50 hr/mo util) $/El h ?

TBO, h 5000 ? ?

Figure 59 shows the GPI and TAC results for the light twin helicpoter. In this application the current piston engine has a GM 15% higher than the ATE, and the advanced technology piston engine has approximately the same GM as the ATE. The piston engines are shown to have a 30 to 35% lower TAC.

NAP-C: Naturally Aspirated Piston—Current Technology 0V NAP-A: Naturaliy .As p irated Piston-1988 1'echnaloav GM* = 3456 kg (7720 lbm) TAC* =.$421,800 1.3 1.0 1.2 ic U C7 0.8 U ^ 1.1 r Q - 0.6 1.0 0.9 0.4 ATE NAP-C NAP-A ATP NAP-C NAP-A TE-3824 Figure 58. Gross mass and total aircraft cost comparison----advanced gas turbine versus piston ( unpressurized twin).

NAP-C: Naturally Aspirated Piston--Current Technology NAP-A: Naturally Aspirated Piston-1988 Technology GM* = 1960 kg (43i0 !

TAC* = 390,900 1.2 1.0

U

0.8

U

i^ ti 1.0 0.6 0.9 0.4 ATE NAP-C NAP-A ATE NAP-C NAP-A TP-3825 Figure 59. Gross mass ars total cost of ownerslEip comparison---advanced gas turbine versus piston (helicopter---light twin).

Figure 50 shows the effect of reducing OEM price of the ATE on total aircraft cost (unpressurized twin only). This figure indicates that for the ATE to be economically competitive with the piston engine (PP.1, the OEM price of the ATE would have to be approximately 35 to 150% higher than the PE prices. As a result, DDA forsees the ATE initially replacing CTEs, not the piston engines.

It is noted that along with the previously indicated gross mass reductions, the turbine en g ine provide other advantages. The following is a list of the benefits of a turbine engine when compared with a piston engine.

* Lighter weight

0 Smoother operation

improved ride comfort

•

• allows use of more efficient prop • eases installation requirements Lower frontal. area 0 Reduced installation volume 0 Greater reliability 0 Longer TBO 0 Multiple fuel usage * Competitive installed fuel consumption 0.8 U U 0.6 0.4 4 5 b 1 2 3 7 ATE Price r OEM $ Nominal NAP Price r OEM S T E-5299 Figure 60. Total aircraft cost versus ATE to NAP engine price ratio (unpressurized twin).

Technology Elements The technology elements evaluated in this study effort include: s dual property turbine Ceramic turbine stator Advanced composite gearbox case

a

Lamilloy* cimbustor Uncooled RPT blades and vanes

0 Axial--centrifugal compressor

The purpose of each variation was to help determine the most cost-effective design, thereby substantiating engine configuration choices. A brief quanti- tative review of each technology element follows. The results of a quantita- tive evaluation using the unpressurized twin application of the dual property turbine, the ceramic turbine stator, and the composite gearbox case are also *Lamilloy is a registered trademark of the General Motors Corporation.

i { L4 included in the discussion. Utilization rate of 600 h/yr, a fuel cost of $0.22/1 ($0.83/gal) and a depreciation period of 8 yr were used for the com- parisons.

Dual Property Turbine Small gas turbine engines generally use cast turbine wheels with integral blades. The prime material property required in the blades is resistant to t,.e high gas temperatures. The wheel bore is at a much lower temperature but must resist high tangential stress hence requiring a high strength material.

A dual property turbine consists of a cast ring with integral blades of one material diffusion bonded to a hub of another material. The two materials can be selected to a more nearly optimize blade and wheel design than is possible by using a single material. The improved blade material can increase part service life reducing overhaul costs. The wheel can be somewhat lighter and shorter since a higher allowable bore stress is possible. The wheel may be made of powdered metal and then hot isostatic pressed (HIPed) to improve the powdered metal density and strength.

While the extra operation required to produce a dual property turbine wheel makes the wheel more expensive than a monolithically cast wheel (both with integral blades), the dual property wheel produces a more optimum design in terms of component weight and life. The dual property turbine has been incor- porated in engine 29 and quantitatively analyzed and compared to engine 29 CT. Both engines are of an uncooled design. Engine characteristics applied in the trade-off study are shown in Table XLIII. Results are reported in the following section which indicate a significant savings in total cost of owner- ship for the engine configured with the dual property turbine as shown in Table XLIV.

TABLE XLIII. - ENGINE COMPARISONS--TURBOPROP Dual Cast property Turbine rotor configuration, turbine turbine Engine ID 29 CT 29 Type of compressor 1-C 1-C Type of GP turbine (stage/configuration) 1-A 1-A Compressor pressure ratio 10 10 GP rotor inlet temp, K ( OF) 1339(1950) 1339(1950) (uncooled) (uncooled) Power Turbine (stage/configuration) 2-A 2-A SP, slss, kW (hp) 373(500) 373(500) sfc, μ g /W's (lbm/hp'h) 87.22(0.5162) 87.22(0.5162) Mass, kg (Ibm) 88.5(195) 88.5(195) OEM price, 1978 $ 72987 79537 Maintenanze cost (50 h /mo util) $ 44.10 19.57 TBO, h 1500 TABLE XLIV. - DUAL PROPERTY TURBINE EVALUATION (UNPRESSURIZED TWIN).

Engine ID 29CT 29 Vanes Cast X40 Cast M509 HPT blades M246 Cast Assy M247 HIPed into HPT wheels M246 Cast Assy PA IN792 assy LPT blades and IN713 Cast Assy IN792 Cast assy and HIPed wheels M4, kg (1bm) 3747 (8460) 3747 (8460) ,ACM, kg (lbm) (base) (0)

TAC, $ 459,760 484,750

ATAC, $ (base) (+24,990)

ACC), $ 1,331,340 1,056,940 A TCO, $ (base) (-274,450) Ceramic Turbine Stator DDA is currently engaged in developing a number of ceramic components for small gas turbine engines under a NASA contract (ref NASA CR--135230). Ceramic materials offer the advantage of having a low coefficient of thermal expan- sion, maintaining their strength at high (turbine) temperatures, and resisting oxidation and sulfidation. These properties make them suitable for considera- tion as turbine stator materials. For any given compressor pressure ratio there is a fairly wide range of turbine temperatures that will yield near op- timum specific fuel consumption. The highest turbine temperature consistent with good fuel economy will result in the Lmallest engine as specific power output increases with turbine temperature. Use of high gas temperatures cur- rently requires special costly stator metals and coatings to resist surface deterioration and cracking from differential. expansions. Air cooling is re- quired to help keep metal temperatures at safe limits. Ceramic material sta- tors would require no air cooling; hence, cycle performance would be improv- ed. The vane sections with their thin trailing edges are currently subject to cracking from differential thermal expansions caused by part geometry and by variation in gas temperatures through the stator nozzle passages. The low thermal expansion characteristics of ceramics would reduce this cracking pro- blem, which causes overhaul replacements and cost. The stator also usually forms a gas path seal with the turbine rotor drum and/or blade tips. The low- er thermal expansion characteristics of ceramics is expected to result in a more stable seal clearance throughout the range of operating conditions and environments permitting use of smaller clearances and hence again improving sfc. Engine 31 was configured to reflect mass and cost changes estimated for the use of a ceramic HPT stator (vanes and blade tip shroud) for comparison with engine 26 with its standard metal, air—cooled vanes and coated tip shroud. Engine characteristics developed are shown in Table XLV. A quantita- tive analysis and comparison was then completed for this item.

TABLE XLV. - ENGINE COMPARISONS--TURBOPROP Silicon Conventional X40 Carbide Turbine stator configuration i 26 31 Engine ID 1-C 1-C Type of compressor 1-A 1-A Type of GP turbine (stage/configuration) Compressor pressure ratio 10 10 1478(2200) 1478(2200) GP rotor inlet temp, K (OF) 2-A turbine (stage/configuration) 2-A Power 373(500) 373(500) SP, slss, itW (hp) 86.41(0.5114) 86.41(0.5114) sfc, fcg /W°s (lbm/hp'h) 78.9(174) Mass, kg (1bm) 78.9(174) 74764 74516 OEM price, 1978 $ Specific mass, kg/kT 1bm/hp) OEM specific price, 4/kW ($/hp) 18.24 18.17 Maintenance cost (50 h/mo util), $ 5000 5000 TBO, h Table XLVI indicates that both TAC and TCO slightly favor the engine incorpo- rating the ceramic turbine stator.

TABLE XLVI. - CERAMIC TURBINE STATOR EVALUATION (UNPRESSURIZED TWIN) Engine ID 26 31 GPI, kg ( lbm) 3751 (8270) 3751 ( 8270) :AGM, kg (base) (lbm) (0) TAC, $ 461,430 460,260 ATAC, $ (base) (--1170) TCO*, $ 1,010,550 1,008,300 A TCO, $ (base) ( - 2250) ^600 h/yr use and 0.22 $/l (0.83 $/-al) fuel cost.

Advanced Comp osite Gearbox Case The accessory drive gearbox cases for the basic study engines is a cast alumi- num material. Accessories such as starter, generator, alternator, oil pump, and fuel pump are mounted on the face of the case. The accessories are driven by the gears arranged inside the case in a gear train driven by the engines main rotor usually by means of a radial drive shaft. The main loads on the gear case are maneuver loads, which cause bending loads at the accessory/case t mounting flange. The accessories used with small engines are themselves small; hence, flange bending loads are usually small. The case casting toler- ance and minimum wall thickness are as low as possible but are often dispro- portional to that required. Also the low modulus of aluminium may permit ex- cessive deflection during periods of high operating loads causing gear and bearing misalignment and consequent wear or early failure. Advance composite materials offer varied combinations of material properties and fabrication techniques from which an improved case material may be selected to fit the requirements of any potential application. Ultimate strengths and modulus in the range of two to five times aluminum are available. Composites can be mold or layed-up and these and other techniques can be combined to build the de- sired geometry and properties in each section of a part. Composite density may be equal to aluminum but part mass will likely be much less for the same stiffness. Engine 32 was comfigured with estimated composite accessory gear- box case mass and cost for comparison with engine 26 as shown in Table XLVII.

A quantitative analysis and comparison were made with results as shown in Table XL,VIII, which shows GM, TAC, and TCO results slightly favoring the use of an advanced composite gearbox case.

TABLE XLVII. -- ENGINE COMPARISONS--TURBOPROP Cast Fiberglass Gearbox case material aluminum polyamide 26 32 Engine ID Type of compressor 1-C 1-C Type of GP turbine (stage/configuration) 1-A 1-A Compressor pressure ratio 10 GP rotor inlet temp, K ( O F) 1478(2200) 1478(2200) 2-A Power turbine (stage/configuration) 2--A

SP, slss, kW (hp) 373(500) 373(500)

sfc, A g /Ws (lbm/hp"h) 86.41(0.5114) 86.41(0.5114) 78.9(174) 78.9(174) Mass, kg (lbm) 74764 OEM price, 1978 $ 74579 Specific mass, kg/kW (lbm/hp) OEM specific price, $/kW ($/hp) 18.24 21.53 Maintenance cost (50 h/mo util), $ TBO, h 5000 5000 Lamilloy Combustor High performance engine designs impose combustor operating conditions of high pressure and of high. inlet and outlet temperatures. These conditions pose formidable liner wall cooling problems especially when combined with the high surface-to—volume characteristics of small reverse flow annular combus- ratio tors. Lamilloy construction is one of the most effective methods known for cooling at these severe design conditions. It uses 50% less cooling flow than convection/film--the next best system. DDA has conducted rig and engine tests and analysis an a variety of cooling methods including:

TABLE XLVIII. ^- ADVANCED COMPOSITE GEARBOX EVALUATION (UNPRESSURIZED TWIN) Engine ID _ 26 32 GM, kg (lbm) 3751 (8270) 3747 (8260)I AGM, kg (1bm) (base) -4.54 (-10) TAC, $ 461,430 460,120 A TAC, $ (-1310) (base) TC0*, $ 1,010,550 1,008,040 J TCO, $ (base) (-2500) *600 h/yr use and 0.22 $/1 (0.83 $/gal) fuel cost.

..^ nu ni nnlu FI IrL.rlalllll.Irr • Transpiration (Lamilloy) • Convection film--roughened walls • Impingement film • Convection film---smooth walls e Tangential film

• Convection

a Effusion All the basic engine configurations in this study contained a Lamilloy com- bustor which is considered necessary for combustor life.

Lamilloy is a DDA patented material made by diffusion bonding several layers of sheet metal that have been photo chemically etched with interconnected holes and grooves. The finished material has an accurately controlled poros- ity that can be varied as desired over the sheet area. Individual sheets are formable and can be welded during manufacture and/or repair. A relatively simply and smooth combustor shell results from ase of this material. Less pressure drop results when transpiration cooling is used hence, cycle perform- ance is improved. Porosity is varied over different areas of the combustor to suit the local heat load conditions thus minimizing combustor tharm nal distor- tions. Service life should therefore, be improved and overhaul cost reduced.

The low-cooi:.ng flow requirements of the Lamilloy combustor permit use of in- creased cooling flow for other purposes, One study indicated that effective use could be made of this air to improve turbine inlet termperature patterns, which would result in a 10% decrease in inherent turbine failures.

^-"'t

r

A combustor design using conventional materials consists of a series of over- lapping tubular deflector rings connecting two end cover bodies. These de- flector rings are assembled with crimped bands interposed at overlap area to form passages admitting cooling air into the combustor at axial intervals.

Cooling air is also admitted through punched holes in the end bodies, and de- flector plates are also commonly used around each fuel nozzle as it enters through the combustor wall. Thus a large number of individually pre-formed pieces must be welded together during fabrication of a conventional combus- tor. More area of material is required in the conventional design as a result of the crimped bands and overlaps. More labor time is required to form, lo- cate and weld these individual pieces.

Up to eight sheets of Lamilloy must presently be butt-welded together to form a combustor. Currently production of the lamilloy sheets involves a consider- able amount of hand labor and sheet size is limited.• Cost is, therefore, fairly high. Improved facilities are planned that should cut material cost by two--thirds. Present technology does not permit the use of an automatic weld- ing process on Lamilloy because of its inherent porosity. The use of manual welding also results in increased cost. Combustor cost estimates have been made for conventional versus Lamilloy designs which show current Lamilloy part prices up to 2.5 times that of the conventional part price. The new facili- ties noted above would reduce the Lamilloy part cost by 50%. Further cost reduction studies are being male aimed at eventual cost equality.

One specific design study done for a GATE-type engine with a foldback combus- tor yielded the following results: Combustor construction Conventional Lamilloy Material Hastelloy X Hastelloy X Weight, lbm 8 8 Price* $2218 $3648 Combustor life, h Unsatisfactory 5000 *Note: Price based on 5000 engines at 80 per monkh and expressed in 1978 dollars.

The two combustors were comparable in weight. While the Lamilloy combustor was projected to cost 64% more than the conventional design, it is interesting to note that a conventional combustor design with satisfactory life could not be achieved.

Three plys or layers of etched sheet stock are commonly used. There is a min- imum ply thickness required for handling during etching and for part stiffness and pressure--loading considerations. The Lamilloy sheet often is thicker and may be heavier than the conventional material sheet. Since less Lamilloy ma- terial is used because of its butt-weld construction, combustor part mass is usually 'aqual to or less than that of an equivalent conventional design. Con- siderable detail design time and effort is required to define a combustor suf- ficiently to ensure equivalent performance and hence to permit good mass com- parison. Such effort was considered outside the scope of this project. No quantitative mass analysis could, therefore, be accomplished for this item.

The payoff comes when Lamilloy is developed so that a Lamilloy combustor is at least no heavier and possibly lighter than a conventional combustor. Tempera- ture profiles - into the turbine will be improved and the life of turbine blades and vanes increased.

Uncooled HPT Blades avid Vanes Air-cooled blades and vanes are cast hollow so that cooling air can be passed through their lengths. Sheet metal baffles may be inserted into the hollow to help distribute the air to obtain uniform cooling of the blade or vane mater- ial. A complicated care is required for each blade and vane. Core location is critical to providing specified wall thickness. Casting scrap rate is higher compared to an uncooled part. Cost is higher due to the added core and the increased scrap rate. Air-cooled casting costs may be up to to 3 times that of uncooled parts. Blade masses are usually only slightly differen°..

Wheel design stresses would be similar.

A lower TIT cycle temperature is needed with the uncooled design 1339 K versus 1478 K (1950 0 versus 2200 0 F). The lower temperature attempts to provide same turbine bore temperature at maximum power for the uncooled turbine as was estimated for the air-cooled turbine to achieve equivalent design safety mar- gins. This lower cycle temperature changes the sizing and matching of the various engine components to obtain the desired power output. Greater cycle airflow is required in the uncooled design. Overall diameter and length are increased. Klass and cost of the compressor and combustor components increased si g nificantly. These cost increases were over twice that of the turbine blades and vanes. Even w*'.th the loner TIT turbine wheel temperature, esti- mates indicated that service life would be reduced to 1500 h. A quantitiative analysis and comparison was completed for this item. Engine configuration 29 was established as an uncooled version of engine 26. Since the engine cycle was so great a contributor to the overall changes, the results of the analysis were presented as Engine Component Comparison. 2 on page 65.

Axial-Centrifugal Compressor Consideration was given to evaluation of an axial-centrifugal compressor in the program. Preliminary review of this design indicated little likelihood that the concept would be competitive costwise with the dual centrifugal con-- figuration. The rationale leading to this conclusion may be outlined as fol- lows. Work (compression ratio) in the axial and centrifugal stages should be nearly equal for maximum surge margins. A practical pressure ratio of 1.4 is assumed for each axial stage. Three axial stages would be required (with a centrifugal stage) for an overall pressure ratio of 10. Fcur axial stages would be required for an overall pressure ratio of 14. These overall ratios span the range of interest. Each axial stage is usually cast separately, and its cost is less than that of a centrifugal compressor. Two axial stages would be likely to cost more than a centrifugal stage. Thus the dual centri- fugal is cost effective throughout the pressure ratio range. In addition, the dual centrifugal compressor is expected to operate without the need for bleeds or variable geometry. The axial-centrifugal very likely would need bleeds and/or variable _ ,,_^ometry even with equal Mork split. If a single axial stage were used, variable geometry would be required to provide surge mar g in (by matching the operating characteristics of the two components at off—design conditions). Bleeds and particularly variable geometry would add cost. T would also add complexity and unreliability to the control system.

Centrifugal compressors above a pressure ratio of about 8.5 usually require acceleration bleeds. This is in the range of the single —stage axial plus cen- trifugal compressor. In the dual centrifugal configuration, however, each stage would be below this value so no bleed should be necessary. A more de- tailed analysis of this configuration was considered outside the scope of this program.

Sensitivity Studies Three types of sensitivity plots were developed for the GATE design studies.

• Type I--GM, TAC, and TCO sensitivity to engine sfc, engine mass, and en- gine cost

a Type II--TCO sensitivity to use and fuel cost

0 Type !!I--TCO sensitivity to component efficencies, cooling air, And leak- age TXpe I Sensitivity data was generated to show the effect of changes in engine sfc, engine mass, and engine cost on the unpressurized twin, pressurized twin (heavy), and light twin helicopter design gross mass and economic parameters (TAC & TCO). Figures 61, 62, and 63 show sensitivites with the current tech- nology as a base. These figures indicate that sfc is the primary driver on TCO for this relatively low—cost engine. Figures 64, 65, and 66 show compar- able sensitivites using a two—stage centrifugal compressor, high IZ c , high RIT advanced turbine en gine as base. Note that cost is the primary driver an TCO for this relatively high cost representative ATE.

Type II Figures 67, 68, and 69 summarize fuel cost and use trades comparing a repre- sentative advanced engine to the current engine technology as affecting TCO.

The three figures indicate that as fuel cost and/or use increase the greater the TCO advantage of the representative ATE over the CTE. Also, the more stringent the mission requirements (higher cruise altitude, velocity, range, or payload) the greater the advantage of the ATE.

III Type Engine cycle studies were made on one of the matrix engine (Engine 21) to de- termine the sensitivity of engine horsepower and sfc to variations in various engine component performance items at the slss T.O. design point. The items varied were: d Compressor efficiency

e Gas generator turbine efficiency

GM* 3937 kg (8680 lbsrl) TAC* = 401,010

].05 1.05 U Q 1.00 U 1.00 0.95 0.95 ^.

1.00 1.10 0.90 0.90 1,00 1.10

P/P*

P/P*

1.05

TCO* = $1,018,&10

P/P* Legend O Reference Point

U 1.00

U sfc

U

F-

L" Engine Moss

0.95

A Engine Cost 0,90 1.10 1.00 TE-3812 P/P*

Figure 61. Engine sfc, mass, and cast sensitivity data--•CTE

(unpressurized twin).

OM* = 5343 kg (11,780 16m) 1.10 TAC* = 5931,480 1.10 r 1.05 4 1.05 l7 1.00 - - - 1.00 0.951 0.95 f 0,90 1,00 1.10 0.90 1.00 1.10 P/P' 1.10 P/P* TCO* = $2,056,260 1.05 P/P* Legend OQ Reference point

U

" 1.00

C sfc / CI Engine Mass / n Engine Cost 0,95 0.90 1.00 1.10 P/P* TE-3813 Figure 62. Engine sfc, mass, and cost sensitivity data--CTE (heavy twin).

PAGE 1S

o P OOR QUAI-fry

(^

1.10

TAC* - $375, 980

GM*- 2227 kg (4910 16m)

1.05

P

V

- 1.00

0.95 ' - 0.95 0.90 1.10 1.10 1.00 0.90 1.00

i; P*

P/P* TCO* _ $651,810 ..05 P/P* Legend O O Reference Point

V

O sfc

O 1.00

F i

13 Engine Mass

n Engine Cost

J^

-0.95 0.90 1.00 1.10 TE-3814

Figure 63. Engine sfc, mass, and cost sensitivity data--CTE

(helicopter-twin).

GM* = 3502 kg (7720 16m) TAC* - $421,810

1.05 1.05

V

Q

1.00 w _ _ 1.00

Q- ^f i 0.95

0.95 I-- ./

0.90 1.00 1.10 0190 1.00 1. TO P/P* P/P* 1.05 P/P* legend TCO*= $909,300 x O Reference Point

O

V

p sfc

V A/

/r

N

Engine Mass

0.95 t ( I (^S Engine Cost TE-3815 0.90 1.00 1.10 P/P*

Figure 64. Engine sfc, mass, and cost sensi :i.vi.ty data--ATE

(unpressurized twin),

1.10 1 1.10 GM* 4218 kg (9300 I6m) TAC * = $B66,670 u 1.05 ^ 1.OS

^ a

1.00 1.00 ----- - V iI r r 0.95 0.95 1.00 1.10 0.90 1.10 0.90 1.00 P/P* P/P* 1.10 TCO* = $1,652,960 1.05 O P/P* Legend

U

q O Reference Point V 1.00 _ sfe ^ p Engine Moss G Engine Cost 0.95 T E-3816 1.10 1.00 0.90 P/P* Figure 65. Engine sfc, mass, and cost sensitivity data--ATE (heavy twin).

1.10 ^- 1. t o ^- GM * = 1955 kg (4310 16m) TAC * = $390,860 * 1'05 1.05 u

^ a

0 err 1.00 (9 i.00 - - ---- i 0.95 0.95 ' 0.90 1.00 1.10 0.90 1.00 1.10 P/P* P/P* TCO* = $613,590 1.05 y O P/P* Legend u OO Reference Point U o sfc F- 1.00 _ _ q Engine Moss Engine Cost 1- r 0.95 - ---!-- 0.90 1.00 1.10 T E-3817 P/P* Figure 66. Engine sfc, mass, and cost sensitivity data--ATE (helicopter-twin).

hr/yr 1100 . CIE '.00 Pr/yr ..... .............

GTE ..........

r V - 900 . . . ...... ........ . . ....... ......... .. .....

ATE 0.25 0.30 0.35 5/1 0.50 0.60 0.70 0.80 o.90S/gal, 1.()0 1.10 1.20 1.30 TE-3818 Fuel Cast Figure 67. Total cost of ownership sensitivity to use and fuel cost (unpressurized twin).

hr/yr C) 2200- U OS hr/yr cl S I?jUU- Nit 0.30 0.35 0.20 0,15 0.25 S/P 0.50 0.70 0.80 0.90 1.00 1.10 1.20 1.30 0.60 $/Sol TE-3819 Fuel Cost Figure 68. Total cost of ownership sensitivity to use and fuel cost (heavy twin).

i ii r-^ $ 600 hr/yr f CTE r- ATE hr/yr CTE ATE 4.15 0.20 0.25 0.30 0.35 5/,t 0.50 0.70 0.80 0.60 0.40 1.00 1.10 1.20 1.30 /gal S TE-3820 Fuel Cost fuel cost Figure 69. Total cost of ownership sensitivity to use and (helicopter-twins).

• Power turbine efficiency • Turbine cooling airflow rate Overboard air leakage rate

•

Each item was varied separately while holding all other values at their base values.

This engine data was then used in mission trade studies to determine the sen- sitivity of vehicle TCO to engine component performance variations. Three of the six vehicle/mission combinations were studied. The results of these stud- ies are summarized in the following figures:

• Figure 70—unpressurized twin

• Figure 71--pressurized twin (heavy) • Figure 72--light twin helicopter Environmental Considerations A forecast of the regulatory environment was made for 1985 to 1990 for the general aviation gas turbine engine. ;`Noise and emission regulatory require- ments predicted for the small turboshaft and turboprop engine for this time frame appeared to be satisfied by the technology now in hand.

Total Cost of Ownership TCO* = $849,490 ?oka9e Base Cycle RIT = 1551 K (2350°F) 1.02 R c = 10:1 Overboard Leakage = 0.2% ' 1 0

O = 0.80

U - 9GGT = 0.844 t - i.00 p PT = 0.846 n s^ cs Turbine Cooling Air = 4.3% ^— dTurN 0.98 GGT Campr 0.

PT 0.94 -4 -2 0 +2 +4 Point Change from Base Value T E- x"21 Figure 70• Component efficiency, cooling air, and leakage sensitivity data (unpressurized twin).

Total Cost of Ownership TCO* = $1,298,680 1.06 1.04 Leakage Base Cycle 1.02 RIT = 1561 K (2350 °F) O Rc = 10:1 u Overboard Leakage = 0.2% O 1.00 _ U T1C = 0.80 ~ /^ "GGT = 0.844 17 PT = 0.846 0.98\ Turbine Cooling Air = 4.3% tf / Tur6 CA GGT Compr 0.96 PT 0.94 -4 -2 0 +2 +4 TE-3822 Point Change from Base Value F igure7I• Component efficiency, cooling air, and leakage sensitivity data (heavy twin).

Total Cost of Owbiership r!

TCO* = $605,660 G 4 "t ^i Base Cycle RIT = 156T K (2350°F) 1.04 Rc = 10:1 Overboard Leakage = 0.2% )/8 Leakage 7c=0.80 1.02 GGT r 0.844 nPT= 0.846 1.90 = Turbine Cooling Air 4.3% GGT 0.98 COMPR PT 0.96 +4 0 +2 -4 -2 Point Change from Base Value TP-3823 Figure 72. Component efficiency, cooling air, and leakage sensitivity data (helicopter-twin).

Noise Turboshaft En g ine Noise Small turboshaft engines radiate noise from the compressor inlet, engine case, reducton gears, and the engine exhaust. Noise radiating from the exhaust is generated by several sources jet, turbine, and a third component, generally called "core noise," thought to originate in the combustor. The first two sources are generally negligible for turboshaft engines since the exhaust ve- locity is very low (122 m/s (400 ft/s) or less) and turbine tones occur at very high frequencies (above 20 KRz).

Engine case and gear-radiated noise are also usually very low in terms of con- tribution to the engine noise signature, thus leaving compressor and core/com- bustion noise as the dominant sources. The noise emisson characteristics of the small turboshaft engine have changed substantially in the past 16 yr as a direct result of engine aerothermal cycle development and improved component efficiencies. The cycle trend toward increased pressure ratio and turbine temperature to obtain reduced fuel consumption has changed the small turbo- shaft noise signature from exhaust noise (core/combustion) dominated with little or no noise contribution from the compressor to compressor noise domi- nated as shown in Figure 73.

The net result of the engine development trend to date has made the current engines about 3 dBA more quiet than the initial engines even though takeoff power has more Lhan doubled for the DDA production: engines shown.

Initial p roduction Engi 10 d

^' J

>^

Current a Producti m o= Solid Symbol--Compressor Noise Open Symbol— Exhaust Noise 50 100 200 400 1,000 2,000 5,000 10,000 7/3 Octave Band Center Frequency--Hz T E-3826, Figure 73. Comparison of initial and current production engines at takeoff power.

GATE Engine Noise The engine descriptons selected from the GATE matrix for final consideration are shown in Table IL along with a current production engine which was selected as the base for noise comparisons.

TABLE IL.

- GATE TURBOSHAFT AND TURBOPROP ENGINES-UNITY SIZE 373 kW (500 hp) CANDIDATE ENGINES-DESIGN POINT DATA Current Engine no. 26 27 28 29 30 Production Technology ATE ATE ATE ATE ATE CTE Type of compr 1-C 1-C 1-C 1-C 2-C 1-C Type of GP turb 1-A 2-A 1-R 1-A 2-A 2--A Aircooled GP turb yes yes yes no yes no Performance, slss T.O.

Re 10 10 10 10 10 8.14 RIT, K ( oF) 1478 1478 1478 1478 1478 1322 (2200) (2200) (2200) (2200) (2200) (1920) Shaft power, kW (hp) 372.8 372.8 372.8 372.8 372.8 484.7 (500) (500) (650) (500) (50(1) (500) Airflow, kg/s (lb/s) 1.340 1.t19 1.268 1.625 1.400 2.495 (2.954) (3.583) (3.086) (5.5) (2.974) (2.795) jig/W • s sfc, 86.41 86.39 82.03 87.22 82.34 100.03 (lb/hp • h) (.5162) (.4873) (.5114) (.5113) (.4855) (•592) ar Current core/combustion noise prediction methods do not accurately predict the reduction in noise radiated from the engine exhaust shown in Figure 73 and, therefore, are , not considered reliable indicators of the effect of engine cy- cle changes on exhaust noise. However, predictions made using the methods described in references I and 2 agreed reasonably well with test measurements for the base engine (+3 dB) and when applied to the GATE engine cycles showed a noise reduction with increasing pressure ratio. Noise reducticns of 2 and 7 dB, respectively, were obtained at a pressure ratio of 14. Correlating the historical noise reduction trend with engine sfc and projecting to the GATE sfc range indicated a 5 to 7 dB reduction (Figure 74).

The engine used as a noise reference base has a single-stage centrifugal com- pressor which has a supersonic inducer tip relative Mach number above about 85% speed making Multiple-Pure-Tones NPT) the dominant inlet noise. Figure 75 shows the inducer tip relative Mach number as ,a function of coTrn !5iion ratio for advanced single- and two-stage centrifugal compressors. As car be seen, the change from the base engine to the GATE single stage R c 10 is small (0.02 Mach) and occurs in a Mach number range when MPT strength usually begins to roll-off so that only a small change is expected. Using a two-stage centrifugal compressor to obtain a pressure ratio 14 will reduce the inducer tip Mach number required to 1.25 with a corresponding reduction in MP T noise.

Figure 76- presents a comparison of a single-- and two-stage compressor which are similar to those used in the noise base engine and the GATE R c 14 engine. This figure shows that, for a constant size, the increase in Mach number for the single-stage compressor should result in about a I dB increase and the decrease in Mach no. for the two-stage should give a 2 dB decrease in inlet noise. Scaled to GATE engine size, the change relative to the base engine should be a I to 2 dB decrease for single-stage compressors, and a 4 to 5 dB decrease for the two-stage compressor.

1st Production Year m 1962 1971 ---' I o 10 dB O U7 1978 Change to Single-Stage GATE .P 6 a) N } j €tonge Centrifugal Compressor o r Base Engin and Low Emission Combustor 0.55 0.50 0.60 0.65 0.70 Fa9/W's

^..^. 1 l ..^__I..--

80 90 Too 110 120 1 t, brnAp .

sfc TE-3827 Figure 74. Turboshaft engine exhaust noise as a function of sfc.

I .4U ^_Single Stage 1.36 ^$ase Engine 1.32 a Z 1.28 a Dual Stage_, .a . 24 1.20 I.16

$ I0 12 14

R^ T E-3828 Figure 75. Inducer tip Mach number for single — and dual—stage centrifugal compressors.

C n o Q L m -a ° f

I

Two Stage l0 d$ >

O

J_ ` Single Stage

o m °- x 1.0 1.L 1.4 Relative Tip Mach No.

T E-3829 Figure A weighted sound power level for single — and two—stage 76.

centrifugal, compressors.

i Effect of Turbine Power on Propeller--Driven A/u and Helicopter Noise Turboprop, The use of turbine power reduces propeller--driven aircraft noise in two ways.

0 The direct contribution of the engine is reduced. The exact contribution of the current reeiprr , cating engines has not been defined but for aircraft of less than 907 kg (2000 lbm) gross mass the propeller and engine gene- rate about the same overall sound power under static conditions. In flight, propeller noise is diminished, expecially the higher blade pass harmonics, because of the improved air inflow conditions, while engine exhaust noise is not so affected and can be easily identified in the air- craft flyover signature (3) x as seen in Figure 77• • Turbine power permits the use of thinner, more efficient propeller de- signs, primarily because the propeller does not have to accomodate the firing order stresses experienced with reciprocating engines. Figure 78 shows that noise reductions of the order of 5 dB can be obtained through this change in propeller design. (3) The use of turbine power also offers a third, indirect means of raduca.ng the certification noise levels for propeller driven A/C. Certification noise lev- els for aircraft certified under section F of FAR Part 36 consist of two parts 6 s 7P NOISE SP E CTRUM SP 6 CYL ENGINE 2.8LAQE 2400 RPM 2P 6X 5 305 METER AOL 2X 5 qP OX 3P iP ^I BP 1QP 2P 0 I IX I I I w 40 IP + f j P - PROPELLER Z 35 HARMONIC

l^

X - EXHAUST r^ ^ 1 1 il^ 11 I ^f I ^ HARMONIC 1s ORIGI'MAL PASS lS .8 1 2 6 R 10 70 OF poop, QUALITY PREOUENCY -IQO HERTZ TE-3830 Figure 77. Darrow band frequency spectrum for the flyover noise of a typical single-en gine, general aviation aircraft.

* Numbers in parentheses refer to references listed at the end of this report.

PROPELLER DESIGN EFFECTS ON NOISE K a ro as 1 VAL PACE IS Z ^^ n po ol', QUALlry ^ 80 W . ^ b^T• \QS^\ Q S^\`tea 9^ S -^E a,^^ ^\^ 2. 1 LADES S^ I A— 3-BLADES

I

0.75 0.80 0.05 0.90 0.95 0,70 too HELICAL TIP MACH NUMBER TE-383 Figure 78. Influence of propeller section properties on the observed noise.

(1) the actual measured level and (2) an aircraft performance correction.

Taking advantage of reduced power plant mass (0,213 k/kW (0.35 Ibm/hp) for turbine versus 0.912 kg/kW (1.5 Ibm/hp) for reciprocating engines) to increase climb performance would produce an additional noise reduction increment.

Helicopter The bulk of the current helicopter fleet is turbine powered, with only some very early designs and very light craft using reciprocating engines. Turbine engines powering current helicopters in the 907-1546 kg (2000-4000 Ibm) class make a small contribution to total helicopter sound power. Engines of the GATE technology level will contribute somewhat more primarily because of the increase in compressor inlet noise. An assessment of the possible contribu- tion was made by modifying the base engine noise to simulate the GATE candi- date engine No. 29, and scaling to sizes appropriate for the Hughes 500D and Sikorsky S61 helicopters. The engine noise was added to helicopter noise as observed during hover, and the combined noise used to determine the EPNL level for a 152 m (500 ft) flyover at cruise speed. The net effect was no increase on flyover noise. Differences between engine and rotor noise directivity ac- count for the low sensitivity of flyover EPNL to compressor noise. Peak en- pine noise occurs well forward of the helicopter, whereas peak rotor noise occurs aft so the two sources do not combine on a peak basis. The choice of a single--stage compressor or inlet configuration (side inlet rather than front) could alter this result. Treated inlet ducts may be required in GATE--powered helicopters.

D Certification Noise Levels t _, 1 The purpose of noise certification is to insure that the best available noise i . I reduction technology that is technically feasible and economically reasonable is incorporated in aircraft of new design. As a result, reductions in the certification noise levels are paced by introduction of new aircraft that dem- onstrate the effectiveness of the noise-reduction technology incorporated in their design. In this respect, engines shown in the GATE study probably will not affect propeller-driven aircraft certification levels in the mid--1980's, but would provide the technology base for reduction at a later time. Figure 79 shows measured noise levels for certificated (5) propeller-driven air- craft. Certification noise levels for FAR-36 in 1980 are shown, along with a decrement based on the technology just now emerging in repsonse to the 1980 requirements and a additional decrement showing the estimated benefit of in- troducing GATE-level technology.

Figure 80 shows the helicoptor certification noise level proposed by the FAA (7), which is very probably close to the actual certification requirements through the mid-1980s. Any reduction in requirement is wholly dependent on rotor noise reduction. Reducing rotor--type Mach number has been shown to be a nonproductive means of reducing rotor noise (8), however, the data spread shown in Figure 80 indicates the possibility of noise reduction through rotor design.

Q V Certificated A/C " 90 Noncertificated A/C q v FAR-36 for Type El O Single Engine o q Certificate after 1 6 Twin Engine O q i January 1980 cu t p

z

y 80 © Est 1985 m

W s

•

a

Est 1990-5 CD ^ with GATE m p z m P , I 1 1 Sa ^.1 - L --- I- ^- 0 6000 kg 1000 2000 3000 4000 5000 ^^ I ,^ 1 I i I I 10,000 12,000 Ibm 0 2,000 •!,000 6,000 6,000 Gross Mass TE-3832 Figure 79. Certification noise levels for small propeller-driven aircraft.

Flyover at 152 m (500 ft) OCH-47C 102 p 0 10 212 O SA-330 O 109A

54B ^

m O Lynx D5A-321 ^OCH7 Z

0 SA-342 and 350

0.

W OSH-3 as5t0el" P 9 O47G O205L 0 500C J .o Z 300C l.; 10 20 30 40 48 2 3 103 lbm ? E•-3933 Gross Mass Figure 80. Possible noise limit for type certification of helicopters.

Emissions t Originally, the GATE study included means to incorporate the considerable im- pact the Federal regulations for control of aircraft exhaust emissions have on the design of general aviation engines. Because rule making was still in pro- cess, the exact amount of impact was not known but was expected to depend on the following factors:

s The severity of the control the Enviromental Protection Agency applied to

i sanall turbine engines in 1985 and beyond.

! The discriminatory nature of the regulations as they are applied to dif- ferent GAZE engine applications i.e., do they preclude use of the "optimum 1. core 2 • The level of emission control technology that has been developed and is applicable in the time frame of this study q e However, early in the program, the EPA advised DDA that they intended to de- regulate engines used for general aviation. Subsequently, they published pro- "` posed rule making, which confirmed this approach, by deleting the requirement to control gaseous pollutants. As a result of these actions, pollution con- trol requirements and their impact were dropped from the GATE study.

i

i i ( i EPA ,Regulations The control of emissions from aircraft engines is specified in the United States by the regulations of the Environmental Protection Agency (9). These regulations presently call for the control of exhaust emissions (hydrocarbons, carbonmonoxide, and oxides of nitrogen and smoke) and of fuel-venting emis- sions from all engines including those used in general aviation. Turboshaft F engines, which are not considered a significant source, are not regulated.

In response to these regulations, the GAVE study was designed to include those cost and mass penalties which would be required for engines to comply with the stringent emission control requirements for general aviation.

! Change in EPA Direction Early in the contract period, the EPA advised industry of their intention to deregulate general aviation engines along with other proposed changes in regu- lations. The EPA's thinking was confirmed through the distribution of Draft Changes to the 1973 rules. This intent (to deregulate general aviation) has now been substantiated by the publication of a rules change proposal (10).

The EPA rationale for this change in position follows: "Recent studies have concluded that the air quality impact at major air termi- nals is much more significant than that at the smaller, general aviation air- ports and further, that the major aircraft contributors at the major air ter- minals are the commercial aircraft and not general aviation aircraft."

For large general aviation airports where general aviation traffic is a sig- nificant fraction of the total, it was found that CO approaches the ambient air quality standard only at the runway where people are exposed for only brief periods and, therefore, cannot be construed to be a problem. HC and NO, emissions were found to be less that 100 ton/yr at any general aviation airport. 'Therefore, it is clear that emissions from general aviation airports do not provide strong justification for a program of federal standards appli- cable to general aviation aircraft (10).

Additional substantiation of EPA's action was given when they stated that the pollution control of general aviation engines was not as cost effective as that of other sources (i.e., automobiles), and that the, "resources available to the EPA for the pursuit of achievement of the national air quality stand- ards are substantially limited," and could best be expended controlling other sources.

e Impact of Deregulation The result of the EPA proposed deregulation of general aviation turbine en- gines was to divorce all consideraton of special pollution control require- ments from the GATE study. This was done because the remaining EPA require- ments for control smoke an fuel venting are state-of-the-art requirements that are already incorporated in the design and production of present DDA small turbines. This change in impact is shown in Figure &l.

Technology Pal lutants Engines Impact o ^ n GATE Controlled Level Controlled • Requires Additional Develop HC • Gas Generators Have Turbafan CO Research Increased Cost, Weight, Start NO Complexity of SMIL Turboprop • Conflicting Requirements Program for Turboprops and Turbofans (1) Production Fuel Venting .

End Turbofan Smoke • Emission Control Not Production of Fuel Venting a Factor Program Turboprop (2) (1) EPA Rules for Aircraft, 40CFR87, 17 July 1973 (2) EPA Proposed Rules Changes, 40CFR87, 24 March 1978 TE-3837 Figure 81. Impact of EPA emissions rule making on GATE study.

Su - mmary The engine configurations selected as optimum for the various GATE applica- tions studied are shown in Table XVIII. Table XVIII indicates the size (kW (hp)) cycle, configuration, and technology elements for the selected engines along with a listing of their benefits compared to currenr engine technology in small gas turbine engines.

In summary: • Most benefit in aircraft economics for a GATE engine was found in the 597-746 kW (800-1000 hp) class.

Current technology gas turbine engines can be offered at lower prices principally as a result of an existing production base.

i

? i The advanced turbine engine in the 224-447 kW (300--600 hp) class offers

the potential for lighter more efficient aircraft than can be achieved with current turbine engines and at moderate reduction in ownership costs, s The low cost of the p iston engine compared to turbine engines is the pri- mary factor in its favor.

s G 4c. Evaluation of a Common Core Concept ^, APPROACH i1 Having identified that the GATE engine should be an advanced technology, air-- cooled, high pressure--ratio engine in the 373 kW (500 hp) class, attention was directed to obtaining the broadest application of the gas generator core to i [ general aviation requirements. The obvious solution was to configure the en- gine so that turboprop, turboshaft, and turbofan variants could be obtained with the least modification and development pr:,blems.

ENGINE CONFIGURATION The engine configuration was reviewed in terms of maximum usage of the gas generator core defined in the trade studies. Turboprop and turboshaft ver- sions were desired to satisfy the fixed ,ring and helicopter markets. The pos- sibility of a turbofan variant was also considered. Although no viable market for a small turbofan engine was forecast for 1985--1990, the life of the engine frame would be expected to reach at least the year 2000. By that time a small fan engine application could appear, thus it was decided to configure an en- gine which would also be adaptable to a turbofan conversion. By core com- monality to all three variants, cost reduction through increased core produc- tion could be achieved.

The trade studies indicated a strong trend toward high pressure ratio and high turbine temperature. Since the studies were broad based, iterations through the total engine design process were not made; therefore, where differences between individual engine configurations were small, selections could not be made. A two-stage centrifugal compressor engine at a pressure ratio of about 14 showed up well with an aircooled two stage axial turbine and a two stage power turbine. A simpler engine using a single-staged centrifugal compressor at a pressure ratio of 10, and a cooled radial inflow turbine was also highly rated. Axial centrifugal compressor configurations were not examined in de- tail because of their inherent cost disadvantage at the higher pressure ratios preferred in the general aviation application.

Foldback combustors featuring transpiration cooling were applied to achieve a more compact engine and provide adequate liner cooling.

The choice of a sing?e versus a two-shaft engine included the considerations discussed below.

Performance Matching propeller and engine aerodynamic characteristics at off-design mission operating conditions are much improved with the two-shaft design which permits independent selection (scheduling) of the gas generator and the power turbine speeds. The taro-shaft system also permits improved overall engine design point performance since each turbine component can be designed for maximum efficiency at its primary operating speed. SFC and DOC are consequently significantly improved.

A two-shaft engine would be better for low flight speed applications re- sulting from greater turbine expansion ratio.

The gas generator of the two-shaft engine is not affected by large tran- sient load variations which can cause a 20% rotor speed droop.

M ass and Cost An output shaft failure mechanically unloads the driving turbine permitt- ing high instantaneous acceleration until datected and corrected by the control system. Such acceleration of a single-shaft rotor system is slow _ enough to be easily controlled since the compressor would still be absorb- ing power. The power turbine of a two-shaft rotor system must be better protected from such a possible overspeed condition by a more complex con- trol system, and/or controlled turbine blade shedding. Attention to these requirements during the engine design can result in achieving satisfactory safety levels with little increase in mass. Location of the combustor over the turbine, for instance, adds some containment and current FADEC control incorporate automatic power turbine governing plus a redundant protection system.

The - single-shaft system has fewer operating parameters; therefore, fewer condition monitoring-control system sensors are required. The control system will be somewhat simpler to design but should not vary significant- ly in price, mass, or reliability since in either design a full authority digital electronic control (FADEC) is assumed.

The single-shaft T56 uses the engine compressor to generate negative torque and negative propeller thrust during flight idle operation on land- ing approach. Propeller blade angle is still positive, and positive pro- peller thrust can be achieved quickly by increasing engine fuel flow and power. The use of negative propeller blade angle with positive torque can ` achieve the same results but is usually more difficult to control with the same sensitivity. This latter mode would be required with the two--shaft system configuration. However, small aircraft may not need propulsion system contribution to decelerate for landing relying instead solely upon wing and tail control surface (and landing gear) drag, which may be en- tirely sufficient.

A prior DDA design study resulted in a single-shaft version of the free power turbine XT701 engine. Comparative calculated masses were 500 and 535 kg, respectively, indicating only a 7% mass penalty against the free power turbine system.

The single-shaft system generally has fewer parts, which often translates into slightly less engine cost. Fewer rotor support bearings may be re- quired.

j The free power turbine system permits a smaller, lighter starter and starter drive train since only the gas generator rotor must be accelerated rather than the complete engine and propeller. Similarly, the propeller . brake, if used, would be smaller since less rotating inertia mass must be decelerated.

The propeller is not mechanically coupled to the high drive-powered com- pressor. In event of a turbine failure, the propeller is not immediately decelerated, and no large drag forces are imposed upon the airframe. In a twin engine application, such a drag would be asymmetric and would impose t'b high aircraft tail structure loads.

The possibility of such deceleration forces in single-shaft systems such as the ':R 56 has led to the incorporation of a safety coupling teat de-- P H couples the propeller from the engine when a preselected negative torque value is exceeded. Stronger, heavier aircraft tail structures would otherwise be required.

Counterrotation of the gas generator and power turbine rotors is generally preferred. Gyroscopic loads of the two rotors are thus opposed and tend to cancel each other. Therefore, lighter engine and A/C structure may be i possible.

Operational The two shaft engine can have reduced prop speed for low noise considera- tions, since the two turbine speeds can be selected independently.

The gas generator rotor can be operated on the ground by itself (without propeller rotation) for on-wing check-out of engine accessory performance thus facilitating maintenance and repair operations. Aircraft accessories are generally driven by the power turbine shaft system and would still require propeller rotation during installation check-out.

Commonality With respect to commonality considerations, the two-shaft e:gine permits better turbine performance for shaft engines and permits matching of tur- bine-to-fan speed for fan engines.

P014ER RANGE A nominal 373 kW (500 shp) engine designed to operate with an air--cooled tur- bine has considerable latitude in rated power and thus a broadened set of go-- tential applications with fairly small sacrifice in efficiency as shown in Figure 82. Examination of the 14:1 compressor pressure ratio engine configu- ration No. 30 indicates that the engine derated 20% by reducing turbine tem- perature 125 K (225 0F) would increase sfc less than 5%. Increasing turbine temperature 56 K (100oF) could provide approximately 5% more power with a 1% penalty in sfc while retaining the two-stage power turbine. These estimates include appropriate change in vane and blade cooling air and turbine efficien- cy and involve relatively inexpensive resetting of the turbine nozzle areas.

Using a three--stage power turbine would yield somewhat more power and better sfc as shown, but would be a more costly modification.

Engine Selection Engine 30 appeared best overall in all applications studied. Engine 28 was somewhat better in the helicopter. A comparison of the 10:1 R c engine (28) against the 14:1 Rc engine (30), which has been shown to be optimum in the fixed wing application, provides the following GM and TCO trade--off percent- ages.

i

Heavy Twin Engine 28 would produce a 17% heavier design vehicle than engine 30 and would have a 19% higher 'NCO.

Unpressurized Twin Engine 28 would produce a 5% heavier design vehicle than engine 30 with a 6% higher TCO.

Helicopter —Twin Engine 30 would produce approximately the sane design vehicle gross mass au engine 28, but would have a 2.5% higher TCO.

O Engine No. 30 373 kW (500 shp) 1.1 a o a G 4.1 a M Two Stage Power Turbine m ^` OThree Stage Power Turbine o k ,u w 1.4!

•O C rj a O ^u zuu INJ 104 50 0 K I 300 200 100 0 100 OF Change in Rotor Inlet Temperature TE-5046 Figure 82. Effort of rated rotor inlet temperature.

4d. Technology Program Plan DDA prepared a plan that delineated our view of the proper content of a gov- ernment sponsored program designed to develop and demonstrate advanced tech-- nologies for small general avaiation turbine engines. The program plan i n- °' cluded the scope, schedules, and cost projections. !

Part of the task was to provide DDA's viers of NASA's role in the GATE pro- z^ gram.

DDA recommended a role strongly oriented to sponsoring meaningful basic research at the engine component level, to develop the data base required to ease the i ntroduction of new technologies into engine development programs at In performing this function, NASA should also the engine company facilities.

sponsor studies such as GATE as well as preliminary design studies to serve as a catalyst to encourage the application of new technology. The strong part x NASA played in early aircraft gas turbine engine component applied research was cited for its excellence and continued usefulness in guiding design choi- ces in engine development efforts. Although core demonstrators sponsored by NASA may have merit in basic investigations of the interrelationships of com- ponents, most major efforts in work of this scope should be oriented toward a development o b r d Commercial e eZo m nt eff rt and a use sponsored, C mmerc^. l en an development and .

P P g P certification should be user sponsored and based on market drivers.

Specific recommendations related to NASA's role in GATE were to perform and sponsor component research applicable to a 373 kW (500 hp) class gas turbine en g ine to provide basic data to achieve Reduced manufacturing cost Improved component performance Engine/airframe integration Improved gearing Figure 83 shows the overall plan from the GATE study and recommended GATE fol- low on to a continuing program o f research and development on all gas turbine engine components and systems. Further appliciation studies are recommended, and the need for core demonstrator work is shown as R & D efforts on the small engine components mature. Approximately 25 detailed technology programs were defined as required in support of the DDA concept of NASA's role in GATE.

Titles of these programs and areas of highest benefit are shown in Table L.

_ _ CY 78 79 1 80 8T 82 S3 B4 35 86 87 NASA SPONSORED GATE Stud GATE Follow-On En enn /airframe i Seal s s _ inns Tu Combustors trloterials lnfe ra'ion C ore 1] manstratem - Ao ITcntiart Studies Im i Eng ine Develo men ► Pr ram MMONIM"It Demonstrator mw Certification _ F Figure 83. Task IV---technology plan.

CIP 1. 1 3,

PQOf?

i

N^ d..

zll ;^ f c ^b f TABLE L. -- GATE TECHNOLOGY PLAN Benefit Title Material and Processes Dual property titanium impeller Reduce cost Reduce cost Lamilloy combustor .fabrication tech Rejuvenation of turbine components Reduce cost of ownership Improve performance and life Abradable seals Improve quality Structural control for cast TI components Thermal barrier coatings Improve life Gearbox and Shafting Composite shafting Reduce cost and weight High performance PM gears Reduce cost Composite material gear housings Reduce cost and weight Propeller reduction gear general arrangement Reduce cost and weight Improve bearing reliability Failure tolerant ball and roller bearings Spiral bevel gear load capacity tests Reduce cost and weight Advance Structure Ceramic turbine vanes and tip shrouds Reduce cost Fluid Dynamics Improved sealing, small gas turbines Improve performance Improve performance Inlet configuration studies Exhaust configuration studies Improve performance Aerotherm High efficiency small air-cooled turbine Improve performance Small high pressure ratio compressor Improve performance No. I-stage dev of 2--stage centrifugal compressor Improve performance Reduce cost Low-cost prechamber combustor Controls Control system conceptual design Define requirements Reduce cost, improve reliability High reliability, low cost electronic control dev Fuel handling system development Reduce cost, weight, size I Application Advanced twin turboprop concept study Define payoff L5p V. SUMMARY OF RESULTS MARKET ANALYSTS _ General aviation aircraft sales were forecast to increase substantially over the next 10--yr period, realizing an average annual increase in US general avi- This represents an increase ation and helicopter airframe production of 4%.

of from approximately 16,000 units per year in 1976, to over 25,000 units in . 1988.

During this time, the piston and turbine engine are expected to share the pro- will remain firmly established in the smal- . pulsion market. The piston engine ler power sizes up to 224 to 298 kW (300 to 400 hp), while turbine power could Other shaft power concepts such as the rot- predominate at the higher powers.

ary combustion engine are not expected to enter into significant contention.

Turbofan engines have found good acceptance in general aviation; however, in the sizes of interest in the GATE study, i.e., less than 6672 N (1500 lbf) thrust, no substantial market was anticipated. Customer acceptance of propel- lers has not been a problem in the GATE class of aircraft, and although jet power is attractive in terms of speed, the potential of the turboprop seems Turboprop sDeeds are sufficiently high to insurmountable for the short term.

cause little difference in block-to-block times for the average trip. In ad- dition, the turboprop fuel economy advantage could be an increasingly import- ant deterrent to the penFwration of the turbofan engine into smaller general aviation aircraft.

Potential markets identified for a GATE engine were nearly equal in total dol- lar volume above and below 447 kW (600 hp). The market was heavily dependent on GATE engine acceptance for fixed-wing applications. Above 447 kW (600 hp), current development programs in the engine industry are expected to result in a strong base from which to launch commercial turboshaft, turboprop, and tur- bofan engine programs. Below 447 kW (600 hp), no new technology programs had been identified. As a result, it was recommended that the major part of the study effort be directed toward an engine in the 373 kW (500 hp) class to de- fine the most viable engine concept, and a supporting technology program for the concept.

BROAD SCOPE TRADE-OFF STUDIES Air vehicle classes and related zLss3,ons representing important market seg- ments were used in trade studies to identify the best general aviation turbine engine concepts. Pay off parameters considered were minimum acquisition cost, minimum direct operating cost and minimum cash flow as determined for the com- plete engine/airframe combination sized to meet design mission requirements.

' Aircraft gross mass was also an important parameter since it was a major driv- er on costs and engine power required.

The "bogey" in the study was a "current technology" gas turbine engine. DDA's latest Model 250 small gas turbine engine which entered production in 1978 was used to represent the current technology base. This engine is a highly compe- titive machine in terms of performance, mass, and cost and represented a formidable state of the art. For the purpose for this study, the engine was considered scalable to other sizes, and its cost was adjusted to the study standards for two cases assuming (1) a hypothetical case with no production `S base or inherited learning and (2) inherited learning, ire., the situation a which exists when a manufacturer introduces a new model in a long production Pun of similar . models. The cost difference resulting from "learning " was found to be a major factor in comparing the costs of new advanced engines with those already in production.

I Parametric engines were defined in terms of design and off-design performance, mass, geometry, and acquisition and maintenance costs. Related hardpoint en- gines and components were used to guide the parametric designs. Scaling pro- cedures were developed to appropriately modify engine characteristics as the engines were sized to match the varying airframe and mission demands. Engine parametric designs covered a range of pressure ratio from 5.5 to 16 at turbine rotor inlet temperature from 1339 to 1561 K (1950 to 2350

0 0. Two-stage

centrifugal compressors were examined with two-stage axial gas generator tur- bines and two-stage axial power turbines. one stage centrifugal compressor engines were studied with gas generator turbine variants including two--stage axial, one--stage axial, and radial inflow. Sensitivity studies were accom- plished at the component level to measure impact on the vehicle gross mass and economics.

The parametric studies showed the GATE engine sliould be high pressure ratio and air cooled. Significant economic benefits for the complete aircraft were found as a result of improved sfc and engine mans, even though engine cost was somewhat higher than a new turbine engine using .urrent technology. Table LT summarizes engine configurations and results for three applications repre- senting a heavy twin, an unpressurized cabin class twin and a light helicopter twin.

TABLE LI. - TASK II "OPTIM(lr1" ENGINES Qnpr Hel Heavy twin twin twin 298(400) 820(1100) 403(540) Nominal shaft power, kW(hp) (*flat rated) 14:1 14:1 10:1 Cycle 1478K(22000F) 1478K(22000F) 1478K(2200 0F) i-Stage centrifugal 2-Stage Centrifugal Configuration 2-Stage Centrifugal 2-A GPT i-radial air-cooled 2-A GPT 2 spool 2 spool 2 spool current technology engine (with production base) Benefits compared to -12 -21 -11 GM, % +5 +1 TAC, % --7 -8 -20 -11 TCO, % 32 23 24 Fuel reduction, % r L1 i r i The engine concepts selected for CATE featured high pressure ratio compres- 1 1O•I sor either 1 .1 pressure ratio an a two -stage centzfuga , or pressure 4.. t ratio in a one-stage centrifugal; the turbine temperature selected was 1478 K (2200 0 F) with an air--cooled two -stage axial, high-pressure turbine for the 14:1 engine, and an air-cooled, radial inflow turbine for the 10:1 engine.

Other engine features included Lamilloy combustor, ceramic turbine stators, composite gearbox and a dual property high-pressure turbine.

The GATE engines realized significant improvements in airframe size and eco- nomics compared to the current technology engine. Aircraft gross mass was re- duced from 11 to 21%, acquisition costs were lower by 7% to higher by 5% de- pending on the application, but total cost of ownership over the 8--yr period used in the study was lower in each case by 8 to 20%. These economic com- parisons apply to the case where the advanced engine competed with an existing . current technology engine (i.e., with price advantages consistent with inher- ited learning). In addition, mission fuel requirements were reduced by 23 to 32% which results from a 20% improvement in engine fuel efficiency and a 23 to 24% improvement in engine specific ,Hass. These engine improvements react strongly on the aircraft by reducing the gross mass required to perform the mission with consequent reduction in drag and engine power size.

The relative engine performance and costs for the best advanced engines com- pared to the current technology en g ines are shown in Table LII for turboprop and shaft engine configurations. Note that the advanced engine cost increases are from 3 to 13% when compared to a new engine using current technology.

Although cost is of paramount importance in the commercial world, the trade TABLE LII. - RELATIVE PERFORMANCE AND COST--ADVANCE}] VS CURRENT TECHNOLOGY GAS TURBINE ENGINES 3731M500 hp) Selected GATE Technology Technology Current Pressure ratio Baseline Baseline turbine temp, K(OF) engine engine 10 14 (inherited (no inherited 1478 22200) learning) 1478 X2200) learning) % Change % Change % Change Turboprop Reference 20% better none 20% better sfc 23% Lighter 24% lighter 100 none mass 4% less 3% more 12% less cost 100 40% Less 35% less 12% less maintenance cost Turboshaft 20% better none 20% better sfc 23% lighter 24% lighter 100 none mass 4% more 13% more 12% less cost 100 35% less 30% less 12% less cost 100 maintenance studies showed that the new engine advantages in fuel economy (sfc), mass, and maintenance cost more than offset the initial cost increase on a cost of own- ership basis.

Sensitivity studies comparing the turbine and piston engines showed that en- gine costs were the biggest factor favoring the piston engine. The advanced turbine engine appears competitive in terms of installed performance, and has a large advantage in engine mass, however, the cost differentials are extreme as shown in Figure 84.

EVALUATION OF A COMMON CORE CONCEPT An engine configuration adaptable to turboprop, turboshaft, and turbofan vari- ants with minimum redesign was chosen featuring a separate power turbine with provisions for forward centerline power output. This type of engine offers the greatest flexibility for turboprop and turboshaft: applications, and offers increased potential for furture turbofan engine deri:vitives. Although market projections for the late 1980s indicated a lack of demand for small turbofan engines, the possibility for eventual need for commercial or military applica- tion was not overlooked.

An analysis of the market potential for the advanced technology GATE engine indicated that the GATE program could have a considerable impact on the future of general aviation. In the 400-600 shp class, an additional 650 engines per year was forecast for 1988. The advanced technology small engines, in allow- ing considerable downsizing of the aircraft for various missions permits the building of highly fuel-efficient aircraft. While there may be some question about the industry's capabilities for using this technology in all product lines in 1988, the turbine-powered aircraft market of the 1990s can be completely dominated by the advanced technology engine-powered vehicles.

TECHNOLOGY PROGRAM PLAN The GATE study showed significant advantages for an air-cooled, high pressure ratio small gas turbine engine. The recommended size in the 373 kW (500 hp) class addresses to a new frontier for small engine technology as the next step following the current U.S. Army program for a technology demonstrator at 597 kW (800 hp). Figure 85 shows how specific fuel consumption trends in small gas turbine engines could be affected.

c$/hp MY 78 Economics $/k W OEM Price 250— ATE Turboprop ATE Turboshaft 150 — 200— 150 — Turboprop Engines 100— Turboshaft Engines 50— 50 — Rig __ ^ _ .._^_^;_ Piston Engines 01 1 1 1 1 1 1 1 oL 300 400 500 600 700 800 0 100 200

L

jono 0 200 400 600 800 hp Shaft Power TE-3851 F igure 84. Specific cost trends.

LL 0.70— 0.66 — iia 0.62 — 5fG 0^ 10() .56 0.5 4 0,50 so 0.46 3 4 6 2 10 kW/1000 3 4 6 8 10 hp/1000 Shaft Power TE-3332 l ag.

Small gas turbine technology Figure 85.

The role recommended for NASA was one of active accomplishment and support of meaningful basic component research in compressors, combustors, turbines, seals, controls, gearboxes, and shafting including on--going efforts in materi- a18 and engine/airframe integration. It was also recommended that NASA serve as the catalyst to encourage the introduction of useful new technology into ^o engine designs through application studies and core demonstrators as component R & D work matures. Engine development and certification programs, however, should be user sponsored based on market needs.

Twenty—four suggested program plans were described and provided to NASA in support of the recommended role.

APPENDIX GATE Cost Analysis Computer Program Notes This computer program was devised in customary units only; thus, only the customary units are used in the description of the program and the sample problem.

t INTRODUCTION The cost analysis program developed fo= the NASA General Aviation Turbine En- gine (GATE) study provides the following cost information for fixed as well as t rotary-wing aircraft: • Total aircraft acquisition cost, dollars • Direct operating cost, dollars/$fl h • Total cost of ownership, dollars • Cash flow analysis, dollars This appendix is a presentation of the methodology used in the DDA cost analy- sis program. It begins with a description of the input data required. Dis- cuss ions of the total aircraft cost, direct operating cost, total cost of ow- nership, and cash flow analysis follow. Finally, a sample problem illus- trating the cost presented.

analysis computer program input and output is INPUT DATA The input data required by the cost program comprise the mission anc air-- craft/engine sizing results obtained from the DDA Mission Analysis Computer Program. They also serve to establish a number of cost parameters.

Mission and Aircraft/Engine Sizing Results The following is a List of mission and aircraft/engine information communi- cated to the cost routine: Symbol TB Block time--total mission time, h i D Block distance, statute miles (SM) VB Block velocity, statute miles per hour (mph) FB Block fuel, lb TF Flight time--block time less ground maneuvering time including T.O., h GWE Aircraft design gross weight, lb IME Aircraft empty weight, lb lb WENCI Engine weight, J yi

Engine price (Original Equipment Manufacturer), dollars

c Engine maintenance cost, dollars/flight hour ($/fl hr) Cost Parameters The cost parameters listed in Table I,III were established to complete the cost analysis routine input data for fixed, as well as rotary- -wing aircraft. These values reflect 1978 base year economics. A second fuel cost and utilization (shown in parentheses) were used in the cost calculations to obtain sensi- tivity data. An inflation factor of 0.0% was observed throughout this cost analysis.

TABLE LIII.---COST PARAMETERS Cost parameter Symbol Fixed wing Helicopter Fuel cost, $/gal CFT(CFTI)

0.83(1.24) 0.83(1.24)

Oil cost, $/gal COT 9.50 9.50 Depreciation period, yr DR 8 8 Aircraft less engine spares, % SPA 0 0 Engine spares, % SPE 0 0 Labor rate (including burden), $/h U 20 20 Annual utilization, h/yr u(n) 600(900) 360(600) Annual insurance rate, % XIRA 1 5 Annual rate of depreciation, % ARD 25 25 10 10 Annual interest rate, % AIR Doran payment rate, % 10 DPP 10 Resale value, % RSV 40 40 Rate of tax saving, % RTS 52 52 Residual value, % RV 20 20 Hangar rental, $/yr HRY 3540 Aircraft registration, $ + $/lb IC+(CPP) 25 + (0.035) TOTAL AIRCRAFT COST The total aircraft acquisition cost was calculated by summing the engine and the aircraft-less-engine costs as follows: r TAG = (CE*ECOSTX) + (EWLE*ACAS) where TAC = total aircraft cost, $ i EWLE # = aircraft empty weight less basic engine weight, lbm ACAS = aircraft cost, $/lb CE = number of engines ECOSTX = list price of single engine, $ ECOSTX = 1.5*OEM where - OEM = Original Equipment Manufacturer's price, $ The Original Equipment Manufacturer's (OEM) price is calculated by multiplying the specific cost for the scaled engine (in OEM dollars per shp) by the scaled engine shp rating.

The specific cost for the scaled engine is established by multiplying the un- ity engine specific cost by the engine scaling effect.

The specific cost for the unity engine is supplied by DDA's Value Engineering Department as input data to the cost analysis program. The engine scaling effect is established by using Figure 86.

The aircraft costs listed in Table LIV were used in the GAPE study (11).

TABLE LIV. - AIRCRAFT SPECIFIC COST'S (Includes Avionics) Aircraft Specific cost, $/lb Unpressurized twin 50.30 Light twin 56.50 Heavy twin 85.00 Light agricultural 20.00 Light single helicopter 60.00 Light twin helicopter 120.00 The fixed-wing aircraft cost curve (Figure 87) supplies the specific costs for the light, heavy, and unpressurized fixed--wing aircraft. The light twin was assigned a specific cost value based on the Beech Baron 58P and Cessna 414 aircraft plots. The Piper Cheyenne provided the data for the heavy twin air- 1.7 — Cost Cost Scaling ratio Scaling ratio

x

(scaled shp) (Unity shp) (scaled shp) (unity shp) 1.6— 1.5 1.4 1.3 1.2 IX — 0.9 — 0.8— ($/Shp) 0.7 1 2 3 4 5 6 7 8 9 10 11 shp/TOG at sis T.O.

T E-3847

Figure 86. GATE turboshaft and turboprop engine sealing ratios for en g ine-

specific and maintenance costs.

too

1000 20110 3000 4000 5000 6000 7000 8000 noo

A/C Empty (less engine) Weight—lb TE-3848 Figure 87• Cost curve--fixed-wing aircraft.

!

craft specific cost. The cost relationship between the pressurized (Cessna 414) and unpressuriz 3 (Cessna 4O2B) twin aircraft applied to the light twin € specific cost results in the specific cost for the unpressurized twin.

The specific cost for the light agricultural aircraft was based on the Rock- well Thrush Commander aircraft (12). The Bell Helicopter acquisition costs i curve (Figure 88) provided the specific cost for the light twin. helicopter.

DDA assumptions formed the basis for the specific cost assigned to the light single helicopter.

I

DIRECT OPERATING COST i .

Direct operating cost (DOC) is the cost of using and maintaining an aircraft.

ATA DOC calculation methodology (13) was used for fixed and rotary wing air- craft unless otherwise indicated. The total DOC was determined by summing the following items:

• Fuel and oil cost

• Hull insurance • Aircraft less engine maintenance cost • Engine maintenance cost • Depreciation t ^j 90 y Q 4 6 7 8 1 2 3 5 9 10 Aircraft (less engines) Empty Weight (1000 16) with Basic Equipment TE-3850 Figure 88. Helicopter acquisition cost.

• Aircraft registration fee* • Hangar rental* *Not included in the rotary wing DOC calculation.

Fuel and Oil Cost hs 1.02*((FB*(CFT/6.5))+(CE-k.135*(COT/8.1)*TB))/D FOC -here FOC fuel and oil cost, $/SM FB block fuel, Ibm CPT cost of fuel, $/gal (refer to Table LIII) CE = number of engines installed COT cost of oil, $/gal (refer to Table LIII) block time, hr TB D block distance, SM The rate of consumption of oil was assumed to be 0.135 Ibm/hr/engine- The oil density was 8.1 lbm/gal; jet fuel OP-5) density was 6.5 Ibm/gal.

Hull Insurance HI = (XIRA*TAC)/(U*VB) where HI = hull insurance, $/SM XIFA = annual insurance rate, % ( refer to Table LIII) U = annual utilization, hr/yr (refer to Table LIII) VB = block velocity, mph Aircraft Less En g ine Maintenance Cost--Fixed and Rotar y Win g Aircraft Fixed Wing Aircraft less engine maintenance labor and material cost for the fixed-wing aircraft was calculated per Ref. 13.

AFMLF = aircraft less engine maintenance labor cost, $/SM ^m XKE'CA = labor man--fours / flight cycle

XKFCA = 0.05*(EWi,E/1000)+6-(630/((EWLE/1000)+120))

^y

where

EWLE = aircraft empty weight less basic engine weight, lbm

XKFRA = labor man--hours =

/ flight hour 0.59*XKFCA

. TF

= flight time, hr

RL = labor rate, $/ hr (refer to Table LIII) Material Cost AFMMF = ((CFHA*TF) + CFCA )/( VB*TB) where AFMMF = Aircraft less engine maintenance material cost, $/SM CFHA = material cost, $/ flt hr = 3 . 08*TAC/106 # CFCA = material cost, $/flight cycle = 6.24*TAC/106 Rotary Wing Aircraft less engine maintenance labor and material costs for the rotary wing aircraft were calculated using the following equations derived from Bell Heli- copter data (14).

Labor Cost AFML = (( ( 2.345*10-'8)*EWLE- ( 2.729 * 10^5))*EWLE * RL+3.506)/VB where i AFML = Aircraft less engine maintenariEe labor cost, $/SM Material. Cost AFMM = ((7.338*(2.718*^'((2.633*104)*E=dLE)))/VB*APCOST i.. where AFMM = aircraft less engine maintenance material cost, $/SM APCOST = aircraft less engine material cost adjustment factor r ., i where Helicopter Adjustment Factor Light single 0.5 Light twin 1.0 Engine Maintenance Cost Q R Unity engine maintenance labor and material costs were obtained from the DDA Operation and Support Cost Program (05590) (15). Scaled en gine maintenance costs were obtained by using engine scaling effect in a manner similar to en- gine acquisition cost adjustment (refer to "Total Aircraft Cost" heading in this appendix).

Depreciation Depreciation was figured as follows: DEPR = (1/VB)*((TAC+SPA*(TAC--(CE*ECOSTX))tSPE*CE*ECOSTX) /(bR*U)) where DEPR = depreciation over U hours, $/SM SPA = aircraft less en gine spares, % (refer to Table LIII) SPE = engine spares, % (refer to Table LIII) DR = depreciation period or years of ownership (refer to Table LIII} Aircraft Registration Fee and Hangar Rental Aircraft registration and hangar rental were not included in ATA's report (13). They were included in the DDA general aviation fixed--wing aircraft DOC but not in the rotary-wing DOC calculation.

Registration Aircraft registration fee varies with the level of use and was calculated as follows: R = IC+((CPP*GW)/(U*VB)) where R = registration fee, $/SM IC = initial charge, $(refer to Table LIII) CPP = charge per pound, $/lb (refer to Table LIII) GW = gross weight of aircraft, lbm t Hangar Rental Hangar rental.also varies with the level of use and was calculated as follows: HR = HEY/ ( U *VB ) where HR = hangar rental, $/SM HRY = annual hangar rental, $/yr (refer to Table LIII) t..

Crew costs were not included in this cost analysis for general aviation air- craft. A breakdown of DOG was included to identify the components that make l up the DOC and their relative significance with respect to total operating costs. The following equation was used to convert DOC units of dollars per statute mile to dollars per flight hour: DOCH = DOCM*VB where DOCH = direct operating cost, $/fl h DOOM = direct operating cost, $/SM TOTAL, COST OF 014NERSHIP Total cost of ownership (TCO) is an indication of the cost to purchase and operate the aircraft over a specific period of ownership. As defined by DDA, TCO includes the following cost items: Total aircraft cost O ® Fuel and oil Lost • Aircraft less engine maintenance cost 0 Engine maintenance cost TOD was calculated as follows: s.

TCO=TAC+((FOC+AFML+AFMM+ENML+ENMM)*VS*U*DR) where TOO = total cost of ownership, $ ENML = engine maintenance labor cost, $/SM ENMM = engine maintenance material cost, $/ SM CASH FLOW ANALYSIS The cash flow analysis consisted of two components---cash outflow and cash in— flow---both in terms of dollars per year over a specific period of operation.

r# ,..

The constants defined in the following paragraphs are assumed to be valid for rotary and fixed-wing aircraft unless otherwise stated (16, 17).

Cash Outflow The cash outflow is composed of cost elements which were "paid out." The cash outflow includes the down payment, an annum, payment on the aircraft Loan, and the cost of operating the aircraft.

Down Payment The down payment is required only in the first year of ownership and is calcu- lated as follows: DPY(l) = TAC*DPP where DPY(1) = down payment, $ for year 1 DPP = down payment rate, % (refer to Table LIII) Annual Payment The annual payment on the aircraft loan decreases as the time of ownership increases and is calculated as follows: AP(M) = 12*XMLP+XIN(M) where AP (M) = annual payment, $ for year M M = indication of year, (1 through DR years) XMLP = monthly level payments, $/mo XMLP = (TAC-(TAC*DPP))/(12*DR) XIN(M) = annum, interest, $ for year M (refer to 'Table LIII) XIN(M) = (B1-((M--I)*12*XIILP)-(XMLP*66)/12)*AIR where B1 = first unpaid Valance, $ B1 = TAC-(TAC*DPP) AIR = annual interest rate, % (refer to Table LIII) Note that the monthly level payments do not include interest. Annual interest is the sum of the monthly interest based on the monthly balance and one- twelfth of the annual interest rate.

jj i " Operating Costs The total annual operating cost was determined by summing the variable and fixed costs.

Variable Costs .. Variable costs are those items which are influenced by use and/or fuel. costs.

The variable costs include: Fuel and oil • *Aircraft less engine maintenance labor and material 9 Engine maintenance labor and material Fixed Costs

Fixed costs are those items which are unaf.ected by use and fuel costs. The

fixed costs include: Hangar rental • Insurance 0 Aircraft registration fee The total cash outflow is obtained by summing the previously described yearly cost elements over M years.

Cash Inflow The cash inflow was composed of items considered as "income" with respect to current corporation income tax procedures. These included an investment tax credit, a tax saving, and a cash sale.

Investment Tax Credit The investment tax credit was applicable only in the first year of ownership and was equal to the down payment.

Total Tax Saving The total tax saving was found by using the following equatLor.

TS(M) = (D(M)+XIN(M)+TAOC(M) +BA(M))*RTS where TS(M) = tax savings, $ for year M D(M) = depreciation, $ for year M 1.

;1 S v; 1 Depreciation --- Declining Balance Method for N Years N = number of years declining balance method is used P '^ Y e.r N = LOG(RV)/LOG(1s0 ARD) where RV = residual value, % (refer to Table LIII) ARD = annual rate of depreciation, % (refer to 'Fable LIII) D(M) _ (1.0-ARD)**(M-1)*TAC*ARD The declining balance method of depreciation used during the first N years of ownership applies a constant rate each year to the book value of the asset at the beginning of the year (18).

Depreciation --- Straight-Line Method for Remaining 'Years of Ownership (N1) N1 = -number of years straight-line method is used N1 = DR-N D{M} = YD where YD = yearly depreciation, $ = RD*TAC where RD = annual straight-line depreciation rate, _ ((1.0-ARD)* N-RV)/Nl The straight-line method of depreciation results in a constant depreciation charge each year and is used during the last years (N1) of ownership.

TAOC(M) = total annual operating cost, $ for year M BA(DR) = book adjustment, $ for year DR -(TAC*RV) where RTS = rate of tax savings, % (refer to Table LIII) The book adjustment was considered a negative inflow in the last year of ow- nership.

Cash Sale The cash sale was applicable only in the last year of ownership and was equal to the estimated resale value. The cash sale was figured as follows: CS(DR) = TAC*RSV where CS(DR) = cash sale, $ for year DR ASV = resale value, % (refer to Table LIII) The total cash inflow was obtained by summing the previous yearly values through DR years.

Net Cash Outflow

The yearly net cash outflow was equal to the yearly cash outflow less the

yearly cash inflow. Summing these yearly net cash outflow figures over the period of ownership resulted in a net cash outflow for the entire ownership cycle.

SAMPLE PROBLEM

A typical set of the input and output data follows.

Input Data IV complete The cost parameters (refer to Table LIII) and the data in Table the input to this sample problem.

TABLE LV. - SAMPLE PROBLEM INPUT Value used

Var iab le

Block time 4.62 h 1151.60 SM Block distance Block Velocity 249.23 mph 1420.82 lb Block fuel Flight time 4.59 h Aircraft design gross weight 7425.0 lb Aircraft less engine empty weight 3570.9 lb Engine weight 186.2 lb

Number of engines 2

Engine price (list) $109,278.93 Engine maintenance cost 30.73(29.19)$/fl h Ai rcraf t cos t 50.30 $/lb Output Data The input data are converted to the printed output, as shown in the following pages, when the computer cost analysis program is executed. Note that four separate printed outputs of the direct operating cost breakdown, the operating cost summary, and the cash flow analysis tables are delivered for each air- craft/engine design studied. These are organized as represented in Table LVI.

TABLE LVI. - COMPUTER OUTPUT ORDER aircraft Rotary-wing aircraft Fixed-Wing Output Fuel cost, Utilization Fuel cost, Utilization, number $/gal h/yr $/gal h/yr 1 0.83 600 360 0.83 2 0.83 600 900 0.83 3 1.24 600 1.24 360 1.24 900 1.24 600

u

GATE COST ANALYSIS OPERATING COST: DIRECT AIRCRAFT COST- 37-442.375 AIRFRAME COST-160384.500 ENGINE COST- 109278.937 0.0013 INSURANCE- 0.0254 INS FUEL COST- 3.1575 FUEL [DST- 0.2754 OIL COST- 4.046Y COST- 0.002 EMG LABOR CO57 a 0.0444 ENG LABOR COST= A/F LABOR COST- 0.7aT4 A / F MATERIAL 9 DEPRECIATION- 0.3172 DEPRECIATION- 0.2116 ENG MATERIAL COST- 0.1972 EMG NATERTAL C11STa 0.1074 0.0237 HANGAR RENTAL- 0.0156 REGISTRATION- 0.4019 REGISTRATION- 5.0013 HANGAR RENTAL- r ' U'b DIRECT OPERATING COSTS- 210.30 DIRECT OPERATING COSTS- 176+63 DIRECT UPERATI'1C COSTS- 229r69 ^ 1%.03 OIRECT OPERATING CYSTS- TOTAL, COST OF OMNERSNIP 943454.37 12107+1I.03 10;1559.37 1353441.00 TOTAL COST OF OWNERSHIP- rr ma-sasar:sarraasa area a.sa+arsyisasrsar*ass*raasasawaxsasssesaasr_s *** DIRECT OPERATING COST RESULTS aDOC AND SCR - r-- --+ TOTAL COST OF UMNERSMIP RESULTS ?

^^ rtr TOD **+ FUEL CAST UTILIZATICM DOC -ra rs as -ra - HRS/YR s s i/F MR r 3 •/GAL - $ - re- r- n tww -fesa++ra-- ter*rssass errse resaswsss n tax aas rew 600.0.7 210.30 Y484S9.37 0.83 --- -s- 176.63 LZIC791.20 0.33 900.00 a-- 17416x4.37 ^a°+ r.24 600.00 Z29.69 - 900.00 146.03 135044L.00 --• 1.24 rrs rsa e*asssrresr.aarasar^-s ► \awsxs-.s-eserseretsrsassesa-sa-a sax ssaa*sxa-.

.....•.•a\-rr P-a.-srM r^t^srr - CPERATIN4 C!751 { SPr ARLION rsaaDl R` . C1 rr •* n - hO0.00+ RS/YN -+• FUEL . 0.231/GALLON UTILIZATICH- rr • i / FLL HV SlSr1 A/C Eyr.1tiF5.- er A/C EsrGINrS • t-ra-a•rr n afs.•o•-. •rr\•.a•a •- ar.-rfs.• ar ..r .4..E-- *a W-- 0.16 FUEL G ^IL \\ sir 3.54.5 \r C.tl C.01 Z.6y INSURANCE rr 17.40 hl..b•r MNT LOR C MAT 0.07 0.25 es s- 1,2.42 r- L.13 C.in !

3EPRECIAT704 •- u.D: --- • ^+.^7 -^' REGISTRATION a- A.QL •• sr HANGAR RENTAL G.C2 -^- -,• rr 4-0.04 T"3.22•• xr TOTAL • 0.74 J.6J • •• 0.84 :13.^C #M A/C * EN" TOTAL •- 5551 - at ar•f \srf rrra-•f rr \Vl.s-aa a--aaa-1a--\ri-raa--P-f!!l sf v'• •aitxx• asa- 1 -.

•.aaar-wrrarr•ssw CIR ECT O P ERATING CC57 8-EAftr)UMN .... ....... r.a.afr.

rs •• .r FUEL : 0.63V GALL M UTILIZATION- 4L0.00r4s/ry wi n r 1/S" 1/FLT *. A/C ENGINE°_ • A/[ Ftif:1HF5+^ o.fs \. x.as-rrrs trees aas- • -5251-..

TQ,pPr* C.16 --- FUEL C OIL ^r t 0.01 1.75 ..i?++ •+ INSURANCE Z.. .l r.

P •- .NT L6R £ MAT U.L7 ^.73 + l7.4C SN.?A•• rr s- O.CQ 7:.'+9 3C.36•• AEPRECIATIOM 0.12 • .- - • ^.?7 REGISTRATTON 4-.I,- ---

•a HANGAR REN T 0.02 ^- ? . 97

AL rr C...2 cti .RO •- + IBC .7^^• TOTAL • 0.18 r\ .a as A/C • EMGS TOTAL 0.71 • 17e.r^ •• -- aaaw----55555--r-s•-r -ra •mr• warrrserrarrf *\\* r r. \rar\arrrrrrr..........

resrswww y [aara[*se*OIREC nn aaarwrsrr yrr T OPERATING COST BREAROCAN' arrser ar -- •+ FUEL rr n 1.243/GALLON UTILIZATION. 600.00NIFS/YR sy .r a* A/t 3 FLT rr A/C f/SENGiN£S yEN GINES*s r• sasessear rarrssaes r rrr[ssrss rwwrrrrrs [w ar FUEL L OIL ^-- 0.24 w ..^- SB.9T^i rr nn INSURANCE 0.01 Gal • 2.68 3.64:: r n a rr s* AMT LSR C MAT G.CT 0.23 17.30 61.46:* a * a r.

i OCAECIATION G.13 0.18 * 33.32 4L^.w3lf •+ REGISTRATION G.CO -^ + C.47 as y a: HANGAR RENTAL O.t#Z as --- n ..4C rw r n ^* TOTAL * 0.24 0.60 n 60.08 169.6;:: ae A/C • ENCS TOTAL 0.92 • +.

228.68 asrweseasraer*a es. ra essrrrrrasr ry.

► +st y wrssuaasawwrrrrasasrwswewraarrw n sa *60 0*4*aeasa yarOIRECT OPERATING COST 6R6AADUWN rraarrrsaesss•ssawr s n e• nn FUEL : 1.243/GALLCM UTTLIZATICNs 900.00MRS/YP !/FLT ew n • f/S>• n HP •* A/C ENG7N#S A/C • ENGINES *e ra swrsa q ae r**ruses [ rrr.rrrrr :[#Parr*.

ra r R• y r FUEL C OIL --- r .^ ?w.5lar 0.I6 • +* sr INSURANCE ...:I C.Gi 2.68 3.e4 nn Y^ rr MNT LBR Z MAT 0.C7 C.23 • 17.SO h1.46^^ ar : ar - +* OEPRECIATIGN 0.13 0.18 33.52 45.53er s * n se [r REGISTRATION 0.00 • 0.47 ^-- sr s [ n sr n y HANGAR RENTAL 0.C2 -- 5.40 •.

n a TOTAL n 0.26 0.6C n 42.06 150.22 nn nn A/C e • ENGS TOTAL 0.84 + 210a30 • n * e .w rewey assssaaarys*r*es.'wsawesrrrrrr * n rsaraaw[srrarasr*rrrrraaa srraras.

CASH FLUY ANALYSIS f C.a3 UIILlZXTln4 . b ohs/YR FUEL C n ST CC.30 4 6 YEAR 1 2 ti 7 r TOTAL CASH OUTFLI7A C. C. °. 0. C. C.

iMG14PAYMENT 374 44 . 0. 174x..

7488ta 7C.612 : 66?43. 6674. +7Ftb. S35?7. 44266. 449x9.

ANNUAL PAYMENT -795 ZU.

78747.

OPERATING CO5T 78747. 78747. 78747- 7x747. 78747. 7e7 .7. 76747, C2Y979.

14#359.

TOTAL CASH GUTFL A4 1x1572. 14%01. 14CE2:. 1?n•53. l:22tl4. 1213316. 123747. 1147.•43.

CASH INIFLCN 0., 0.

INVESTMENT TAX CREDIT 37%44. G. C. G. C. 37444.

TAX SAYINGS •314.0. 40014. 14.

t]t:PRECIATION 94661. 71145, ' 4718. .71d. 4718. :x.:3554..

13656. 1^ .6.0, INTEREST 32103. Z7g2•. 193tl7. 1.119. nhBl. 2312. l36JZ2.

78747, 78747. 707.7. 78747. 74147. TE7.7. 7RT47. 79747. es4974.

OPERATING CYST IN AD.XISTHENT G. 0. -. La s1. 0. . -7-a6H. - 15888.

177817. 1: !+ • F0. °0/.47.

TOTAL tXP"SE 205a01. 1"762. 1381!•4. 943t5. 9889. 945b6b.

157417. *?4a•. Np49h. 7184 0. twlm. -gL44. 46811.

TAX SAVINGS ATC.52 PCT 5143. 517746.

C. C.

CASH SALE C. G. J. 0. .. :51777. :517 Ti.

..a". a TOTAL CASH L4FLOw 14..°61. x7465. BO%gb. 7184C. 64-t?. 46824. 154 L9. 107467.

X6611. S6Ay4. 44044. eAQb7. 711+5. 4:756. d1Iti1. -33173. .J976.

NET CASH CUTFLCM OPERATINT. CRST Su"Affy ^R6/Y4 9UEL COST - f D.ai OTILIfA T lt ' 1 =t.00.ZD VARIABLE COSTS FlvtV C9`.TS S 9:AFAL s 3541.00 FUEL LOST 275^;R..32 -4NGAa OIL COST te9.97 AERFRA',E 'AINTE•1ANCE 37°4.42 LABOR IOC176 Iw ^t- c v*CE :63 MATF 0 TAL 475 all.

4 LINTENANCB W-1NE LAOC AFURAFT RE6l ST4Al T-T4 :tl4.a7 Q 72SC.UO ?94°5.99 MATSR IAL TOTAL VAR IABLI COSTS 71I27.I^ Tr TAL .11EP Ct:STS 7b2G 10 A WN T = 12. 1c I31AL .ItEL COSTS i ?A?7.?O A4TtC1PAT r 0 RIA L + LI:. + NUtl..t 6....

ynUq _ Pro vAAlA9LE CL'ST - t lla-4• I-t XEO T - ! PEg•^nI.4 CTTS 37.75 1DTAL OPERATING CCST s ^wk 171.2` A NNUAL TOTAL = p£RATI4C CCST = 1 19747.27 ^....+en «.rr'.-w.s ¢er-•+.wsn`M1•F-^"SCT_^F- -J-!'Ae=-.,. _ ^^JS^: =^_==- ^[.

CASH FLOW ANALYSIS FUEL COST = i 0.67 UTILIZATION n 900.J0 kMS/VR s_ I: up YEAR 1 T 3 4 5 6 7 0 107AL CASK OUTFLOW !

OOWNPAYPENT 37944. 0. C.

O. 0. 0. C.3. 379.4• ANNUAL PAYMENT 512;6.

5'3537. 474322. 49268 44999.

9. K 39.

1115 111539. 7I1 3 111534. LIl5a4.

111539. 11I53 11153!. 5923125 TOTAL CASH OUTFLOW 224304. 122161. 1771iltT. 1736I3. 4 16 345. 164076. 160507. 1$6535.

1w0b775. !

CASH, INFLOW INVESTPENT TAX CREDIT 37444.

0. 4. C. 0.

J. C. 0. 37944.

^. TAX SAVINGS5 OEPRE:CI ATION 94861. 71145. 53359. 400I9.

3LO14. .7111 4710. 4711. 30555+.

INTEREST 9$193.

27920. 236~6. 19387.

15110. 10676. 6581. 231 2. 06OZZ.

, OPERATING CCST 111539. 11153-0.

111539. 1115 , 44. 11153 4 .

:11939. 11I539. 121539. e92312.

BMK ADJUSTMN T 0. 8. Oa .. D.

0. D -75a5dw 7S8H6.

7 1 1TAL EXPENSE 238593. 21d609. 152564• 17094-0. 150672.

i2T107w 122538.. +2651. 12579911.

TAX SAVINGS ATO.52 PCT 124462. 10951T. 9&046.

02592. P1.69. 66196. 63576.

22194. 6!141h9.

CASH , SALE C. 0.

0. L. 0. 0.

0. 151777. 131177.

TOTAL CASH INFLOW 162012. 1O Q 517. 98048. 55992. 214E9.

660-6. 03570. 173971. b43Y&0.

YET CASH OUTFLOW 623.`1. 72634. 79834. 24722.

47275. 9&9&0. 96931. -17433. 564895.

.,• CPERATTNa COST SUmAKY FUEL COST A 0.53 UTILIZATION • 900.00 HRSIYR VARIABLE COSTS FTxkC COSTS FUEL COST S 35335.22 HANGAR R E NTAL A 3541.CC OIL CJ35T 264.96 AIRFRAME OXINTENANCE IvSU PAtvc E 3794,42 MATEF TAL 15638.64 ENGSHE MAINTENANCE 1 AIRCRAFT REGISTRATION 2x4.87 R nAiO TOTAL VARIABLE COSTS 103918.69 TOTAL FIYEO COSTS 7620.30 e TOTAL FIXED COSTS A 762C.3C ANTTC7PATF3 ANNUAL FLIGHT HDt1RS "L. 8.47 VARIABLE COST : 7 115.47 PER "OUR FI XEO COST x 3 A.4T PEA kOUA TOTAL OPERATING COST t 123.93 PER HOUR r TOTAL ANNUAL OPERATING CCST S 111538.04 •- ^', • 'T L 3$ CASH. FLTW A411Y515 f ^A QUALITY S L.24 UTILIIITIP4 =60C.U0 HRS/VR FUEL CCST t. 7 YEAR 1 2 ' 6 TOTAL CASH OUTFLOW!

37944. C. ". :. ^. D.

DONNPAYKeNT C. 0. 37944, 63537. - 7Lacl. 74612. ot.14i. t+2G7^. S74C6. .9266. 4+999. 7952x.

ANNUAL PQYM EkT a7z N4. P• . N. ids.

OPERATING COST 9C'tf a, 9D° 9fl3C • F. 4D34S. 9G3AS. 923115. 723379. - 24/2:4. 10:947. 567?9. I;2454, 1461-0[. '434?2. 1346~3.

TOTAL CASH 011FLOW 135364. 12+x7.2.

CASH 3394x. C. C. r. u. C. L. 0. 37944.

7NVESTME41 TAX CREDIT T AX SAVINGS A, 71145. ^;34n, c0 CI' t,. ].. +715. 4716. .715. .

D1:PAEL A7IDk 94861. 3[.3654.

270?%. 76 a, 14987. 1•I[P. 115.0. e•5&l. 1312.

INTEREST 32143. 238422.

40 ;A e . .,C?Pe..

P 9439 S. 907&•. 9tiA.. 0:46;. QN3 K e . 40_3!1. 723079.

O ERATING CfST 0. L. C. J. 7. 6. -758BR. -755 BA.

SPOK AOJUST--EHT C.

1674ru.

TOTAL EXPENSE 217439. 169471. t4i1791. 13,yL4. 10!9.3. 10168+. 21527. 11800.

7P4A9. i u " .

TAX SAVINGS ATO.52 PCT 113C6et. 96717. 97046. 77891. 52376. IL2% . 566156.

2. D. C. C. L. C. U. 151777. 151777. - CASH SALE 97044. 77541. 7.-C^- 5EG4 5, 51876. 162971 '1551179.

TOTAL CASH INPLC4 151.12. 96517.

NET CASH CLTTFL3M 52197. 6249C. [96AG. iso?b. 56777. -27547.

- 7 •Ea A7 t1+G L^5T 5ii^9ACr F UI-L COST e S 1.24 11TILIlATI-1 -6.L .8C -OVYR VARIABLE COSTS FIXED OUSTS FUEL r 4ST t 3SI46.2Q wa4rAR RE`ITAL 3541.00 OIL COST I69.47 AIAFRAME .aI4TENA#CE 371..42 LAECR I45U9ANCF 1CZ76.67 RATER TIL .25.76 ENGINE 4A1NTt:NaNCE A1wr AAF7 = :6..67 LAFOR e+ISl^e atl^N 7380.CC - +ATER IAL «9495.94 TOTAL VARIABLE COSTS FIXED LC5T 1 N62G.3G 6276..6: T^TAL en TOTAL FEED COSTS 1 762C.3C ANfICI b AT ANNItAL -CUWS ocu. = 17.7L ^. : y PER SOUR VARTAOLF CCST 117..4 ^.p ^UU R .IIFO CJST i I..7n oeo ..nt1e TC7AL G P I V ATING COST : A 1•C .n+- TOTAL ANNUAL .11-ERATI4r- + -0364.97 LI ST .

REFERENCES s i i 1. R. K. Matta, Combustor Noise Prediction, draft prepared for SAE A21 subcommittee for core noise, June 1976.

2. D. C. Mathews and N. F. Rekas Jr., Direct Combuston Generated Noise in in Turbopropulsion Systems-Prediction and Measurement, AIAA No.

76-579, July 1976.

3. R. K. Rathgeber and D. E. Sipes, The Influence of Design Parameters on Light Propeller Aircraft Noise, SAE 770444, March 1977.

4. Anon, Certificated Airplane Noise Levels, DOT Advisory Circular, AC No: '-- - 36 -1B, December 5, 1977.

5. M. Steele, General Aviation Activities in Noise and Emissions, AIAA 77-310, January 10, 1977.

6. Anon, Comments on helicopter Noise Standards, Working Parer of U.S.

Representative to ICAO/CAN Working Group B, June 1977.

7. M. A. Bowes, Helicopter Noise Reduction Design Trade-Off Study, 1977.

FAA-AEQ-77-4, January 8. D. C. Gasaway, Cockpit Noise Exposures Associated with the Operation of Fined and Rotary-Wing Aircraft, SAM--TR-70-21 (AD705964), April 1970.

9. "Control of Air Pollution From Aircraft and Aircraft Engines, Emmissions, Standards and Test Procedures for Aircraft," Environmental Pro- tection Agency, Title 40, Chapter 1, Part 87, Federal Register, Vol 38, No. 138 (July 17, 1973) p. 19089 10. "Proposed Rules—Control of Air Pollution from Aircraft and Aircraft Engines, Emissions Standards and Test Procedures for Aircraft," Federal Register, Vol. 43, No. 58 (March 24, 1978).

11. M. L. Cross, Deviation, of Cost Curve for Fixed Wing Aircraft-GATE DDA TDR No. AKOOOO-007, October 1977.

12. "General Aviation Aircraft Fleet," Flying Magazine, Ziff Davis Publication, 1976.

13. Standard Method of Estimatinz Com p arative Direct O perating Costs of Turbine Transport Airplanes, Air Transport Association of America, December 1976.

14. Bell Helico ter Textron Visit/GATE Stucj 6 J,^ ^ 1977, distributed by J. C. Gill, Detroit Diesel Allison Division, General Motors, August 3, 1977.

15. E. T. Vance, Logistic Su port Cost Model, (Program 05590), DDA Report No.

C1689, September 1977.

16. Managerial and Engineering Economy, Van Nostrand, Princeton, N.J., 1964.

17. R. A Murdick, 'lime Value of Money," Machine Design, inn 13, 1972.

A 18. G. A. Fleisher, Capitol Allocation Theory: The Study

of Investment

USC, 1969.