Small Engine Component Technology (SECT) — 19910014892 (1986)

Part--Power Performance Comparison for the

LIST OF FIGURES (Contd) Title Figure 8S Part--Power Performance Comparison for the Commuter Application Projected Reductions in Commuter Parameters, Recuperated Engine Mission Results Commuter DOC Results for Recuperated Engines Technical Approach for Estimating Technology Benefits 243 Projected DOC Benefits (Percent) for Isolated Technologies 90 246 Commuter Mission Analysis Results Reference Missile Configuration with Rocket 249 | Propulsion System

Reference Missile Configuration with Turbojet i

Propulsion System Reference Turbojet Engine Configuration and Data Summary Missile Thrust and Drag 95 Several Missile Scenarios for the Future 262 96 263 Representative Reference Mission Advanced Compressor High-Temperature Material Requirements Axial Compressor Performance Projections 273 99 275 Cruise Missile Combustor Technology Projections i00 Cruise Missile HP Turbine Projections i01 Missile Cycle/Configuration Study Engine Options I02 Engines Sized by Minimum Acceleration Condition I xiii i

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i LIST OF FIGURES (Contd) Title Fiqure Page -' I,., Missile Cycle Study Results 282 1 !

Compressor Inlet and Exit Corrected Flow : t Variation 283 r; / , i Axial and Axial-Centrifugal Compressor Diameter Comparison 285 -, 286 "" Compressor Axial Blade Height Limits Missile Engine Weight, Size, and Cost Trends 287 i : Missile Engine Range and Cost Range Trends t_

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Selected Missile Engine ii0 Impact of Comprezsor Size Change 293 ..

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iii 294 "" Impact of Tu:bine Size and Loading Change °.

Performance Comparison of Reference and Year- 296 '.

2000 Turbojet Engines v Missile Range and Cost Per Range Comparison

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of Reference and Year-2000 Turbojet Engines 298 ,_ Comparison of Technology Benefits to Missile System

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3O2 Benefits of Additional Technologies t' Missile Range Comparison 303 -.

Materials Technology Schedule 305 " '

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Combustor Technology Schedule 312 Significant Increases in Cruise Missile Range !2 are Possible with High-Energy Fuels 318 "" t" Boron-Slurry Element Combustor Test Rig 320 GTEC Pure Airblast Nozzle :/ p * _° xiv

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LIST OF FIGURES (Contd) Title

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Volumetric LHV Advantage Results in Range Improvement Potential 123 325 Axial Compressor Technology Program Schedules 124 338 Turbine Technology Schedules 125 343 "Inserted" Blade Concept 126 344 Proposed "Spiral" Weave 127 345 System Technology Schedule 128 352 Proposed Fuel Pump/PMG Module 129 359 Cruise Missile Mission Analysis Results 130 Cross Section of Reference APU Without the Gearbox 131 374 Projected Centrifugal Compressor Efficiency Projected Centrifugal Compressor Efficiencies Projected Efficiency for High-Pressure Axial Turbine Stage Projected Efficiency for Low-Pressure Axial Turbine Stage Projected Radial Inflow Turbine Efficiency in Year 2000 Projected Annular Combustor Performance in Year 2000 APU Cycles Investigated in the SECT Study 138 384 APU Simple Cycles Using Axial Turbine Wheels 139 385 APU Cycle Performance Results 140 387 AGTI01 Regenerated Engine Cross Section ,_" r,P • h _ ._ . _ XV

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LIST OF FIGURES (Contd) Title

Figure page

Projected Performance for APU Based on Scaled AGTI01 388 142 APU Cost Comparison 389 143 APU Weight Comparison 391 .-2 144 APU DO(: Comparison 392 145 Reduction in DO(: for Selected APU Cycles 397 146 DOC Benefits from APU Technologies 399 147 APU Technology Benefit Comparison in DOC Per Hour 400 ) 148 Ceramics Technology Schedule 404 149 APU Radial Turbine Technology Schedule 410

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150 APU Compressor Technology Schedule 418 151 Systems Technology Scheme for APU 423

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152 GTEC Self-Acting Foil Journal Bearing Concept With Backing Spring APU Mission Analysis Results

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r, ° , ie o4 .4 xvi I i l LIST OF TABLES Table Title I Projected DOC Reductions for Rotorcraft, Commuter, and APU Engines II Projected Mission Benefits for Year-2000 Cruise Missile Engines III High-Payoff Technology Categories 1 i0 Rotorcraft Component Weight Savings i0 Crit.ical Sizing Condition 3 SECT Reference Rotorcraft Data 4 18 Materials for Rotorcraft Reference Engine Rotorcraft Reference Engine Design Point Data Rotorcraft Reference Engine Design Data 7 Modular Weight Breakdown 8 Rotorcraft Vehicle Partitions Basis for DOC Computations i0 Rotorcraft Trade Factors ii Year-2000 "Cold" Material Projections Projected Properties for Hot-End Materials Recuperator Technology Projections 14 186 Commuter Reference Engine Design Point Data Commuter Reference Engine Sea Level Data Modular Weight Breakdown 17 Commuter Vehicle Partitions Aircraft Parameters for the Economic Model 192 Commuter Owner/Operator Costs xvii

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LIST OF TABLES (Contd) Title Page Table Commuter Trade Factors 202 ,,° Performance Refinements from Task II 236 22 Turbojet-Powered Missile Weights 250 Turbojet-Powered Missile Aerodynamics 252 24 Reference Turbojet Component Materials 254 25 Reference (Turbojet) Engine - Design Point f, Data 258 Reference (Turbojet) Engine - Max Power .

Cruise Data 259 I 27 Reference (Turbojet) Engine - SLS, Uninstalled Data 260 Economic Model for the Cruise Missile 266 v Missile Trade Factors (Constant Missile Volume) 268 3O Additional Material Technologies for Cruise

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Missile Engine 272 Task III Missile Engine Performance T Refinements 292 32 Evaluation of Slurry Fuels for LRCM HF-3 324 TI ri Operational APU Duty Cycle 363 34 Reference Engine Design-Point Performance - Sea-Level, Static, ISA, Max, Uninstalled 367 r,.-.

Reference Engine Off-Design Performance - Sea-Level, Static, ISA, Uninstalled 368 APU Trade Factors 372 SECT Simple-Cycle APU Design Point Performance - Sea-Level Static, ISA, Max, Uninstalled 393 J 38 SECT Regenerated APU Design Point Performance - Sea-Level, Static, ISA, Max, Uninstalled 395 xviii [.

and.conceptualize year-2000 engines.

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The year-2000 engines were evaluated for the aircraft o and missions defined for this study.

Engine performance and aircraft cost/missilerange data were compared for the year 2000 versus current technol- / ogy engines. High payoff technologies were identified and benefits were quantified. Based on this, technol- ogy plans were defined for each of the four engines of h this study.

L The results of this study show that an aggressive small f • engine component technology program of high-payoff technologies can produce significant benefits for year-2000 engines. These benefits, as expressed in reductions in fuel burn and aircraft direct operating costs (DOC), are summarized in Table I for the rotorcraft, commuter, and APU engines.

PROJECTED DOC REDUCTIONS FOR ROTORCRAFT, COMMUTER, TABLE I.

r_ AND APU ENGINES I .I DOC Reduction / Reduction Low Fuel In High Fuel Fuel Burn Price* Price** o.

craft, commuter aircraft, tactical cruise missiles, and auxiliary power units (APU).

The purpose of this effort is to provide technology plans

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for guiding future research and technology efforts. The program addresses the technology requirements as envisioned for small

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turbine engines of 186 to 746 kW (250 to 1000 shp) suitable for use in commercial or military rotorcraft, commuter aircraft, tac-

I tical cruise missile_, and auxiliary power units (APU). Parallel

studies weTe conducted for these applications to identify high-

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i f payoff technologies for year-2000 engines and to establish tech- nology plans for guiding future NASA-sponsored research and tech- nology efforts to establish technology readiness by year 2000.

The studies documented in this report are in the following sec- !

tions: / r

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2.0 Rotorcraft Engines 3.0 Commuter Turboshaft Engines !

4.0 Cruise Missile Turbojet Engines I 5.0 Auxiliary Power Units The SECT study consists of four major tasks, which are des- cribed in the following paragraphs: Task I - Selection of Evaluation Procedures and Assumptions Year-2000 aircraft and missions were defined, along with !

_° current technology engines (reference engines), for baseline data. Environmental constraints were projected to assess the potential impact on year-2000 engines, and economic models were defined to facilitate evaluation of aircraft operating costs.

/ Trade factors were computed for primary engine features to allow for quantifying beneficial engine changes that might result from technology advancements to the year 2000. These studies are reported in paragraphs 2.1, 3.1, 4.1, and 5.1 for the respective engines.

Task II - Enqine Confiquration and Cycle Evaluation Technology projections and engine cycle and configuration studies were made to conceptualize year-2000 engine performance levels, weights, envelopes, and costs. Results were compared to i the reference engines, and trade factors were applied to quantify potential gains. Promising engine candidates were selected for

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further evaluation. These studies are reported in paragraphs

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2.2, 3.2, 4.2, and 5.2 for the respective engines.

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Task Ill - System Performance Evaluation

I Operating costs of aircraft with selected year-2000 engines

were computed and compared to data for the reference engines.

2.1 Task I - Selection of Evaluation Procedures and Assumptions I The following paragraphs present the study results for the I reference rotorcraft, mission, engine, projected environmental constraints, economic model, and the trade factors.

I 2.1.1 Reference Rotorcraft I A reference rotorcraft was configured and sized for this study using NASA computer program HESCOMP. (Reference I) i

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;_ The configured aircraft, which is a derivative of the

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!_' I Sikorsky $76 Mark II rotorcraft, represents a year-2000 refezence !

aircraft. Major technology projections, including weight reduc- improvements in the main rotor, and improved aerodynamics, were / I tions in the airframe and associated subsystems, figure-of-merit . applied to the $76. Additionally, the reference engines as defined for this study in paragraph 2.1.3 were incorporated into the reference rotorcraft.

The technology factors applied to the weight predictions represent the component weight savings as shown in Table i.

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The combination of fixed engine power and rotorcraft system

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takeoff performance requirements act to determine the rotorcraft takeoff gross weight. The critical sizing condition is as shown in Table 2.

_BI_EDtNG PAGE BLANK NOT FILMED

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The combination of fixed engine power and rotorcraft system takeoff performance requirements act to determine the rotorcraft takeoff gross weight. The critical sizing condition is as shown in Table 2.

power level is somewhat higher. Low specific fuel con- sumption at these higher power levels is of major importance for the ferr_ mission.

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MAXIMUM PgWER AVAILAIILE

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MAXIMUM SPE_

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TYPICAL '-*-----'-- OPERATING RANGE

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AIRSPEED

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engine-demonstrated components. The materials for the major com- ponents, as selected for the reference engine, are as listed in i Table 4.

Performance - The reference engine performance is based on exist-

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ing engine components as adjusted to represent 1985 engine-demon- Q strated levels. These components have been scaled as necessary

G5-281.,33

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Figure 4. Rotorcraft Reference Engine Configuration.

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, I , 4 |i ....i _i • i r i TABLE 4. MATERIALS FOR ROTORCRAFT REFERENCE ENGINE ° _kj Aluminum Accessory Gearbox Cases : / !

Compressor

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_p Titanium 6-4 Stage 1 Impeller : i Diffuser INCO 7i8 6-2-4-2 Titanium Stage 2 Impeller Diffuser INCO 718 Combustor : HS 188 HP Turbine LJ Stator - MAR-M 247 DS - Cooled Stage 1 Blade - MAR-M 247 DS - Cooled Disk - Waspaloy B Stator - MAR-M 247 Equiaxed - Uncooled Stage 2 Blade - MAR-M 247 DS - Uncooled Disk - Waspaloy B

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LP Turbine Stator - MAR-M 247 Equiaxed Stage 1 Blade - MAR-M 247 Equiaxed Q Disk - Superwaspaloy Stator - MAR-M 247 Equiaxed Stage 2 Blade - MAR-M 247 Equiaxed Disk - Superwaspaloy _p Shafts INCO 718 b>' ]

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to meet the power output of 746 hW (i000 shp). The resulting engine and component performances are summarized in Tables 5 _,d 6.

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The reference engine achieves 746 kW (i000 shp) at the sea

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level, static, ISA condition at a turbine rotor inlet temperature of i149C (2100F). At this condition, the compressor inlet cor-

information, materials definition, and measured and calculated weights for the IPS, engine controls, and accessories. This same

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computer program and reference engine weights for IPS, controls, and accessories were also used for follow-on computations for year-2000 engine weight estimates to achieve comparable weigh

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data. Engine size estimates _re an overall length of 40.4 inche3 and a maximum diameter of 14.4 inches.

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I i TABLE 5. ROTORCRAFT REFERENCE ENGINE DESIGN POINT DATA i J .

Design Point Performance-Sea Level, Static, ISA, IRP, Unlnatalled Wlth No Production Margins Overall Engine Performance CompOnent Performance 746 (1000) zP__ss Engine Rating, kW (shp) 0.285 (0.468) BFC (kg/h_)/kW, (lb/hp)/hr o W,/_/5, kg/s (ib/sec) 0.62 (1.36) Bypass Flow Turbine Inlet Temperature o Extraction, kW (shp) 5.5 (7.4) 0 _P/P, t 3.6 o HP Turbine, C (F) 1149 (2100) O LP Turbine, C (F) 769 (1416) tip Compressor 13.37 Overall Cycle PR o PR 4.73 3.604 (7.941) Inlet w_/6, kg/s (lblsec) o _AD, % 81.8 Compresso_ Inlet W_/6, o _Poly, % 85.3 3.037 (6.696) kg/s (lb/sec) 2408 (23,000) NLp, rad/s [rpm) o PR 2.85 o _AD, % 81.6 4679 (44,690) NBp, rad/s (rpm) o nl_oly , % 84.0 42,798 (18,400) Fuel LHV, kJ/kg (Btu/Ib) Overall o PR 13.5 o _AD, % 78.7 i O _polv, % 84.7 O Ex_t=W_/6 , kg/s 0.32 (0.69) L_ (lb/sec) TOTAL CHARGEABLE COOUNll FLOW • U% Combustor o _, % 99.98 o _ P/F, t 4 HP Turbine o W_'/6 , kg/s (Ib/sec) 0.50 (i. I0) 0 _AD, % 87.0 o Cooling Flow, % 6.8 O Interturbine, % 1.4 NONCHAR6EABL.E CHARGEABLE _P/P) UNCOOLEO LP Turbine o WV_-/6 , kg/s (ib/sec) 1.95 (4.30) O _AD, % 88.5 O LPT-NOZ p/p, % 2.5 *Stator and blade cooling flows are shown. The second-stage stator and blades ace un- cooled. The balance of the cooling flows (4.9_) are used for disk and firtree cooling.

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TABLE 6. ROTORCRAFT REFERENCE ENGINE DESIGN DATA Cruise Performance-Sea Level, Static, ISA, 37% Power, Unlnstalled With No Production Marglna.

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Overall Engine Performance Component Performance • I 276 (370) Engine hating, kg (ehp) iP__ss 0.381 (0.627) o WV_'/8, kg/e (ib/sec) 0.529 (I.I_71 SFC, (kg/hr)/kW (Ib/hp)/hr Bypasa Flow o Extraction, kW (ahp) 4.0 (5.3) Turbine Inlet Temperature o _P/P, I 2.73 853 (1567) o HP Turbine, C (F) HP Compressor o LP Turbine, C (F) 553 (1028) 8.50 Overall Cycle PR o PR 3.66 Inlet W_/8, kg/s (lb/eec) 2.678 (5.904) o _AD, % 83.3 O _Poly, % 86.0 _ompre_eor Inlet WV/8/8, 2.173 (4.791) kg/s (ib/eec) st a_!_a___ O PR 2.34 2408 (23,000) NLp, rad/s (rpm] O _AD, I 82.7 4679 (38,501) NHp, rad/e (rpm) o _Poly, t 84.5 Overall Fuel LHV, kJ/kg (Btu/Ib) 42,798 (18,400) o FR 8.55 o _AD, % 80.7 o _p_Iv, % 85.4 o Zx[t'W_/6, kgls 0.33 (0.73) (lb/sec) Combustor o _, % 99.97 o aP/P, % 4.39 HP Turbine o Wv_-/6, kg/s (lb/sec) 0.4_7 (1.096) o WAD, % 86.1 o Cooling Flow, % 6.8 o Interturbine, % 1.3 (_P/P) LP Turbine o W_-/6, kg/s (ib/sec) 1.859 (4.098)

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o _AD, % 88.6 o LPT-NOZ _P/P % 1.13

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o 21 J Weight 31.5 L--.

33.6 Compressor Combustor 51.3 L! IPS HP Turbine 34.5 LP Turbine 72.7 Controls and Accessories 79.6 TOTAL 303.2

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Cost - The reference engine cost was estimated at $185,000 (1985 dollars). This cost is based on a mean sell price for turboshaft engines in this size class, as shown in Figure 5A. The sell price range is based on a GTEC market survey of presently avail- able turboshaft engines in the 447 to 895 kW (600 to 1200 shp) range.

2.1.4 Environmental Constraints Environmental constraints for the year 2000 were projected for rotorcraft engines based on a review of existing regulations

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and projections. Two major areas, noise and emissions, were addressed to establish year-2000 guidelines as based on the fol- lowing sources: Noise Federal Aviation Administration (FAA) International Civil Aviation Organization (ICAO) Committee on Aircraft Noise (CAN) Aerospace Industries Association (AIA) General Aircraft Manufacturer's Association (G.%/4A)

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@ Figure 5A. Rotorcraft Engine Cost.

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Emissions U.S. Air Force AFR 19-1 (1979) International Civil Aviation Organization (ICAO) (1982) Environmental Protection Agency (EPA) (1984) Governmental noise limits and standards for helicopters remain in the infancy stage. The FAA proposed a set of noise limits for helicopters several years ago, but the proposed regu- lation was withdrawn because of significant inadequacies. The FAA currently is working on the subject and should develop a new helicopter noise proposal in the next few years. Significant FAA funds have been spent on the development of a helicpoter noise data base and standard measurement procedures. ICAO has pub- lished a noise standard for helicopters, which includes limits for takeoff, approach, and overflight conditions. In May, 1983, the ICAO CAN/7 delegates voted to recommend a 3 dB increase in the noise limits; however, the member states have not yet approved the change.

Research in helicopter noise, which will continue for the next 20 years, will concentrate on the reduction of rotor noise.

The technology funding for the reduction of helicopter gas tur- bine engine noise will be somewhat limited unless the associated environmental impact becomes a greater concern to the various governments. Overall, no great increases in the stringency of the existing standards are expected. The exceptions would be rules imposed by local governments without federal intervention.

The 3 dB relaxation in stringency proposed at CAN/7 should disappear by the year 2000, and the resulting noise limits should be similar to those currently adopted. Therefore, the Garrett SECT study has used the existing ICAO Annex 16 Chapter 8 _oise standards for helicopters.

Emission levels likely to be required by the year 2000 for a

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746 kW (1000 shp) size class turboshaft are as follows:

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Smoke levels - below the level of visibility, smoke number (SN) = 50 o Gaseous emissions - unregulated for civil engines

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Costs for the reference engine were based on estimated market

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sell prices for 1985, as described in paragraph 2.1.3. Costs for advanced technology, year-2000 engines were based on cost factors as applied to the reference engine cost for parts/components

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*Low fuel price: $0.264/liter ($1/gal)

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High fuel price: $0.528/liter ($2/gal)

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ROTORCRAFT VEHICLE PARTITIONS _! TABLE 8. ROTOF Costs Airframe minus nacelle Variable _-: I Airframe min Variable

tNacelle

Fixed Fixed Variable

J Fuel Tankage

Variable Fixed Variable Variable L" i Engine Variable I'_2 , Engine Cost BASIS FOR DOC COMPUTATIONS - TABLE 9.

Variable Fixed Costs Costs i" _ Assumptions Load Interest I Number of Engine Aircraft Rate I, Engines I (Excludlng 3000 Fuel* Spares) Imputed Interest Rate Airframe Spares 4 perce,,t Maintenance Depreciation I.

Potential Schedule Aircraft 1500 Engine Insurance Maintenance 2500 hr Schedule i Annual Use Crew Expenses Service Life Tax Rate 7 years Takeoff Crew Wages Gross Weight . .

(Structural 4338 kg Hanger Rent Limit) (9564 ib) Miscellaneous Payload 1667 kg (3676 ib) : I *Based on $0.264 and $0.528/liter ($i and $2/gal)

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/, m replaced by new technologies. These costs adjustments included a projected 15 percent cost reduction by year 2000 for current technology parts.

The DOC model used in this study resides within the HESCOMP program. Within this code, the following constant values were used:

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800 -

28.2°/,

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600 -

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400 -

|NGINE

200-

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MAINTENANCE 8.5%

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$1/GAL $2/GAL

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Figure 5B. Rotorcraft DOC Breakdown - Reference Engine.

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• • 0 made on the basis of payoff in aircraft direct operating cost (DOC). The DOC improvements were estimated through the trade factors (established in Task I), which relate changes to DOC in terms of changes in engine performance (SFC), weight, diameter, length, and cost (changes are relative to the 1985 reference engine). Size, weight, and cost were quantified for each engine of interest.

2.2.1 Technology Projections The initial task in configuring potential rotorcraft engines for the year 2000 was to establish the expected level of technol- ogy in that time frame. Inherent in these projections is the assumption that the technologies will be available by the year 2000. Technologies have been identified in three major areas: materials, aerodynamics/thermodynamics, and mechanical improve- ments. These technologies impact the cycle study in terms of efficiency levels, turbine inlet temperature limits, and cooling flow requirements, as well as turbine stage count and hub speed limits.

2.2.1.1 Materials Both hot- and cold-end material technologies have been iden- I.

tified for future engines. For the cold end, four key mate- rials have been considered for the rotorcraft application, as shown on Table Ii. These materials primarily allow reduction in engine weight and cost.

For the compressor, two materials offer the potential of reduced cost: high-tempe.-ature powder metallurgy aluminum alloys and cast titanium alloys. Powder metallurgy aluminum can potentially reduce cost and weight relative to materials "pre- sently used. As shown in Figure 6, material temperatures between 3O

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TITANIUM

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ALLOYS

YEAR-2000

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OGY LEVEl

ADVANCED

ALUMINUM

1985 TECHNOLOGY

ALLOYS

LEVEL

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(400

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CONVENTIONAL

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ALUMINUM

ALLOYS

-18

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4 8 12 16

COMPRESSOR PRESSURERATIO

• OBJECTIVE

• EXTEND USEOF

ALUMINUM ALLOYS

• PAYOFFS

• REDUCED WEIGHT

• REDUCEDINERTIA

• LOWERCOST

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Figure 6. Aluminum Metal Temperature Limit.

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454 and 482C (850 and 900F) will be possible by the year 2000, allowing application of aluminum for higher compressor pressure engines. Present aluminum alloys have been successfully tested to over 316C (600F).

Cast titanium alloys offer the possibility of 20 percent cost reductions through the elimination of expensive machining operations. Other materials of interest are polymeric composites for gearboxes and metal matrix composities for shafts. Polymeric composites are predicted to reduce weight by as much as 30 per- cent relative to present aluminum gearboxes. Metal matrix compo- sites for shafting are expected to be a required technology for year-2000 turboshaft engines in this small size class. The 50 percent reduction in weight, combined with favorable high-temper- ature strength, achieve the critical speed margins needed for the high spool speeds and small bore sizes of future engines.

In addition, four key hot-end materials have been identi- fied, as shown in Table 12. These include two metallics: super single crystal for turbine blades and vanes, and Ni3AI for tur- bine disks. Projections for two nonmetallics (ceramics coated and carbon-carbon) were also made for turbines, combustors, and transition liners.

Super single crystal increases the temperature capability of present single-crystal materials by approximately 56C (100F).

Super single crystal allows higher stress and loading levels and, for a cooled turbine, could reduce cooling flow requirements at a given temperature.

Nickel aluminide (Ni3AI) offers the potential of an improved strength-to-weight ratio by reducing weight approximately 15 per- cent relative to present astroloy materials. This leads to higher turbine hub speed capabilities, resulting in reduced tur- o bine stage count and/or aerodynamic loading.

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:2 ,m The nonmetallics can greatly increase strength at tempera- ture, as shown in Figure 7, and also increase turbine material temperature capabilities. Both materials also reduce component weights significantly, relative to present metallics. The key improvement is the increased temperature capability that will

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allow higher turbine inlet temperatures without cooled turbine blading. For ceramics, the maximum material temperature limit is projected to be 1538C (2800F). Assuming a favorable combustor

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pattern factor (PF _0.12, as estimated for year 2000), engine cycle temperatures up to 1427C (2600F) are projected.

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871 927 982 1038 1093 1149 1204

11600) (1700)(1800)i1900} (2000) (2100)(2200)

G5-281-ll

TEMPERATURE,C (FI

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i i Figure 7. Comparison of Turbine Material Strengths.

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,,._ ROTORCRAFT/COMMUTER COMPRESSOR PERFORMANCE 0.90 1 I I

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0.86 --_,_ i_. _ , 2. i PTS

J _ 0.84 -ROTORCRAFT REFERENCE I _ 0.82 ....... ENGINE .......

0.80 0.09 0.140.180.23 0.32 0.45 0.68 0.91 !.36 Q5_ 12.01 i3.01 I1.01 (0.21 [0.3l10.4)iO.5110.71 II i EXIT CORRECTED FLOW, KG/SEC (LB/SEC) • REASONABLE STAGE LOADING i • AXIAL FIRST STAGE PIP = 1.65 • EQUAL WORK SPLIT FOR 2 CENTRIFUGAL • 3 PERCENTCLEARAKCE/BLADETiP WIDTH • 0.3 EXIT MACH NUMBER • OEVELOPMENT OF 3-0•VISCOUS ANALYTICAL COOESANO NONINTRUSIVE MEASUREMENT TECHNIQUESTO SUPPORT ANALYTICAL OEVELOPMENT (60*/4 • • REDUCEDIFFUSER VANE INLET LOSSESWITH 3-0 DIFFUSERS 110%) LOSSES 110%] OTHER IMPROVEMENTS !. I' • REDUCEIMPELLER SHOCK AND SECONOARY •OEVELOP HIGHER LOADED VANE DIFFUSERS (LOWERWEIGHT) Figure 8. Compressor Performance Projections.

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L Relative to the reference rotorcraft compressor, a 2.9 point improvement in polytropic efficiency is foreseen (approximately four points improvement in adiabatic efficiency) for a two-stage centrifugal compressor. This improvement in efficiency is pri- marily attributed to the anticipated development of 3-D viscous analytical codes. Additional payoffs are seen from improved clearance control, reduced diffuser vane inlet losses, and reduced impeller shock and secondary losses.

Improvements in combustor performance are expected in sever- al areas. As shown in Figure 9, combustor pattern factors will be reduced to the 0.10 to 0.13 range. Additionally, diffuser technology will be improved to maintain present pressure drop levels at increased inlet Mach numbers. Year-2000 combustors I will also have higher heat release rates, reduced size, and improved durability.

The HP turbine performance projections are based on flow size and stage mean work coefficient for a two-stage axial con- figuration, as shown in Figure I0. Inclusion of the work coeffi- .

cient, which is a function of mean blade speed, brings rotational speed and dimensional aspects into the cycle analysis. For the HP turbine, a two-point improvement is predicted relative to the "i uncooled reference turbine. The efficiency improvements are from projected reductions in vane/blade interaction losses, and rotor °.

tip losses.

The LP turbine performance projections are similarly pre- ° • sented in Figure Ii. A 2.9-point efficiency improvement is pre- dicted for LP turbines relative to the reference configuration.

The projected improvements are due primarily to minimizing vane/

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blade interaction losses, and rotor tip clearance losses.

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I ROTORCRAFT/COMMUTER [H/CI

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_ ENGINE REFERENCE YEAR 2000 ENGINES ROTORCRAFT/ PARAMETER _ R/C COMMUTER COMPRESSOR " 1"/" 4 "" EXIT MN u. o u. u.J COMBUSTOR -o o Z$P/P 4.0/o 4.0% PATTERNFACTOR 0.15-0.20 0.10-0.13 I COMBUSTOR rj 0.998 0.998 * CYCLEPARAMETERS (ZSP/P. r/) REMAIN THE SAME • IMPROVEODIFFUSERTECHNOLOGY FOR INCREASED COMPRESSOR EXIT MACH NUMBER • ADDITIONALTECHHOLOGY IMPROVEMENTS • REDUCED PATTERNFACTOR • HIGHERHEAT RELEASERATE • REDUCED SIZE • INCREASED OURABILITY 65.281-36 .... I_ - r t !

HP TURBINE I r._ll

, 2000

0.9; *' LIMIT- z _ 3.63(8.0) 1.8114.01

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/ t 2 PTS 0.45(I.0)

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COOLING )- °__ Wv_'/6 = 0.23 KG/SEC J REFERENCEI IENGINE 0.86 1.25 1.50 1.75 2.00 (0.5 LB/SEC) STAGE MEAN WORK COEFFICIENT. gJZ_H/UM2 --a

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• NO FAB CONSTRAINTS

• UNSHROUOED BLADE

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• 2 PERCENT TIP CLEARANCE

• UNCOOLED

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• TWO-STAGEAXIAL

EFFICIENCYIMPROVEMENTPROJECTIONS

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• REDUCEDROTORTIP LOSSES(55%)

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MINIMIZE VANE

BLADEINTERACTIONLOSSES(45%)

OTHER IMPROVEMENTS

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DESIGNTECHNIQUESFOR IMPROVED

PERFORMANCE/LIFE/COST TRADES

65-281-2 Figure i0. HP Turbine Performance Projections.

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LP TURBINE o.g4

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r, 2000

>.: 0.92 z ¢...p m

;'_CN 3.63{e.01

LL 0.90

,I

i J...

-- w,,,,,,-_"-,,11.6114.01 LU _ ! _ !_LIMIT '"lO.gll2.0l

¢,,l 0.88

i

,,x,., ROTORCRAFT w,./o/6 = 0.45 KG/S ,qc REFERENCEENGINE II.O LO/SEC]

l [ [ z

0.86

I

1.5 2.0 2.5 3.0 _.5

STAGE MEAN WORK COEFFICIENT.

gJ_H/UM 2

I

• NO FAB CONSTRAINTS

I

• SHROUDEDBLADE

I

• 1 PERCENTTIP CLEARANCE

• UNCOOLEO

l

• TWO-STAGEAXIAL

I

EFFICIENCYIMPROVEMENTPROJECTIONS

• MINIMIZE VANE/BLADE INTERACTIONLOSSES

I

i

OTHER IMPROVEMENTS

• DESIGN TECHNIQUESFOR IMPROVED

I

PERFORMANCE/LIFE/COST TRADES

G5-291-9

I

Figure Ii. LP Turbine Performance Projections.

I

, '> :I

• ." .:_,_____.:- ....... : .... 'i{

2.2.1.3 Recuperator Improvements in recuperator technology will come in two

i!

areas, materials and processes. Of particular interest in this study was the counterflow platefin recuperator design. A typical cross section is depicted in Figure 12.

/ ftj Material improvements are expected in both metallics and

tl

nonmetallics. As shown in Table 13, a nitride-dispersion- strengthened 300 stainless steel has the potential for increasing _J recuperator operating temperatures from a present limit of 816C (1500F) to between 982C (1800F) and 1093C (2000F).

It Ceramics could further increase the temperature limit above 1437C (2600F). Furthermore, the 3 to 1 density advantage of _J ceramics is predicted to reduce overall recuperator weight by ., approximately 50 percent. At present, only experimental heat exchanger modules have been evaluated with ceramics, but by the year 2000, ceramic heat exchangers are predicted to be operation- al. One key to making ceramics operational is improving the manufacturing process, as schematically shown in Figure 13.

2.2.1.4 Mechanical Technology

l

After material and aerodynamic capabilities had been estab- ° • lished, mechanical limits were set. As shown in Figure 14, GTEC uses established empirical correlations to set the mechanical inputs for the cycle evaluation. Considered in the cycle study .,,.

were turbine hub speed limits, turbine blade AN 2 limits, as well as TRIT constraints and cooling flow requirements. These were • ° set based on materials and mechanical technologies as projected for the year 2000.

I G ° • 42 ° ,.

L, f I

II

p L !

TABLE 13. RECUPERATOR TECHNOLOGY PROJECTIONS.

Metals 1985 Year 2000 Chromium-molybdenum steel Nitride-dispersion-strengthened 300 stainless steel I Fin Density: Fin Density: 37 fins per 2.54 cm 37 fins per 2.54 cm (i in.) (I in.)

Maximum Temperature is 816C Maximum temperature is 982C ii (150OF) for short times (I800F) to I093C (2000F) Ceramics Year 2000 1985 Experimental plate-fin HXs Operational plate-fin units Plain Fins: Offset Fins: 30 fins per 2.54cm 35 fins per 2.54cm (i in.)

(i in.)

0.04cm (0.015 in.) 0.03cm (0.010 in.)

thickness thickness

L

Maximum temperature is 1427C (2600F) to 1538C (2800F) .J • " 8 • ° 44 o .

C

I,} .,-4 I,) _J o E t_

')

_A_$3U MPTIO N$ • MATERIALS • LIFE REQUIREMENTS

ri

• CYCLE/CONFIGURATION I _d • AEROLIMITS MECHANICAL INPUT TO CYCLE* GTEC \ _._ EMPIRICAL \ • TURBINEHUB SPEEDLIMITS 366-488 M/SEC (1200-1600 FT/SEC) • AN2 LIMITS 33.3-77.4 (5.0 x I010 TO 12 x 1010) • TRIT LIMITS** • COOLING FLOW G5-281.39

[i

REQUIREMENTS T: !

T *LIMITS DEPENOENT ON MATERIALS, CONFIGURATION, AND APPLICATION L..

*'1427C 12600F) FOR UNCOOLEDCERAMIC TURBINE BLADES 2204C (40(X]F) FOR CARBON-CARBON TURBINE BLADES i

i

Figure 14. GTEC Empirical Study Approach.

# t

h

2.2.1.5 Cost r Weight r and Size Estimates To complete the input required for estimating aircraft DOC (with trade factors) cost, weight, and size estimates were re- quired for each engine of interest.

The cost estimates for _he candidate engines of this study were based on the reference engine cost and on a buildup of com- ponent costs as estimated by GTEC manufacturing. Each engine cost was therefore estimated separately, based on technologies employed, number and type of components, TRIT, and flow size.

The costs are projected to year 2000 but are expressed in year- 1935 dollars. The cost adjustments include a projected 15 per-

{

cent cost reduction for parts/components manufactured in the year 2000.

I

Engine size and weight were estimated with the WATE-S

I

!

1'

•t i i / SPECIFIC POWER Figure 16A.

Idealized Cycle Study Results with Constant Efficiencies.

T1

° , '&.

\ .;".; _-'--

", SINGLE.STAGE !

HPTURBINE

\7-

TWO-STAGE HPTURBINE

. _-"_.- I_ "-_

SINGLE-STAGE LP

"-.._ --_ \

UNSHROUDEO SINGLE-STAGE LP TUR61NE TURBINE COOLED METAL UNCOOLEO SHROUOEO METAL r SPECIFIC POWER GS-28T.41 ?

Figure 16B.

Cycle Study Results With Mechanical Limits.

5O ? ,

l i

IL

_jlI_,__:_A<_,_p_,_ ._ ., _.,-, :..

_" "| i

!

I

!

LP TURBINE I COMPRESSORS/ HP TURBINE ADVANCED METALLICS (cooled/uncooledl (cpr: 16-261 [cooled) CERAMICSAND ADVANCED METALLtCS (uncooled) I TRIT : 1204.164gc [2200-3000F] 2-STAGECENTRIFUGAL I Gs.2,,._cpr: 16-26l

I

i

I

Figure 17. Simple-Cycle Study Configurations and Options.

I

r 1 SPECIFICPOWERiNCREASE, PERCENT 0.290 0 10 20 30 40 50 50 70 80 I J 10.46J '%I 15._C " ¢; 1" I REF SFC = 0.285 [0.468) 28 _12900F) 0,274 2M. I M 24 w (.", REF SPECIFIC POWEI_= 246 (15t) 10.451 I ,_.,, -5 0.268 i , 10.441 ° , Z

0.261 I

(_ 10.431 1427C a: 12600F) _ a.

0.255 (0.42) -- 2800FI --- ° I " 0.249 IM-IM i' LL (0.411 © SELECTED CYCLES 0.243 24 _...

(0,40) -15

TRIT -- ,31,c

0.237

l I

(0.39) 460.3 263.0 295.9 328.8 361.7 394.6 427.4 °.

230.2 i160) 1180) (200) 1220) (240) 1260) 12801 (140) SPECIFIC PGWER, KW/(KG/S), [HP/ILB/SEC]] CONFIGURATION • 746KW (1000 SHP] • COMPRESSOR: TWO-STAGE CENTRIFUGAL I HP TUR61NESTAGE COUNT • HP TURBINE: MINIMUM STAGE COUNT [LP TURBINE STAGECOUNT • LP TURBINE: MINIMUM STAGECOUNT T • ADVANCEDMETALLIC TURBINES IM--lm [M : COOLEO. m = UNCOOLED) I _LP TURBINE MATERIAL/TECHNOLOGY HP TURBINEMATERIAL/TECHNOLOGY 65-281-100 Figure 18. Rotorcraft Simple-Cycle Performance Results - Metallic Turbines.

l '11 case. Several candidate engines, as indicated, were selected for further evaluation in terms of weight, size, cost, and DOC esti- mates.

Of the technologies studied, the use of ceramics for blades and vanes has the largest performance impact. Replacement of the advanced metallics allows turbine cooling to De eliminated, resulting in an SFC improvement of over 20 percent relative to the reference engine, as shown in Figure 19. The cycles shown are all uncooled, however, based on the material temperature limit of ceramics, some cooling would be required above inlet temperatures of 1427C (2600F). The cycles selected for further evaluation have therefore been limited to 1427C (2600F), as indi- cated.

Size, weight, and cost were estimated for each of the engines from the selected advanced metallic and ceramic turbine cycles. The resulting values are presented in terms of deltas relative to the reference engine, as shown on Figure 20. As indicated, the ceramic turbines achieve superior results relative to advanced metallics in terms of all four parameters, weight, length, diameter, and cost. One of the more promising cycles uses ceramics at a TRIT of 1427C (2600F) and a 22:1 pressure ratio. With this configuration, weight is reduced by 50.3 kg (iii ib) (-36.6 percent) relative to the reference engine. Diam- eter is reduced by 13.2 cm (5.2 in.) (-36.1 percent), length by 44.2 cm (17.4 in.) (-43.1 percent), and cost by $34,200 (-18.4 percent).

The payoffs for improvements in size, weight, and cost were evaluated in terms of their respective impacts on mission per-- formance by DOC. The DOCs were estimated by trade factors shown in Task I, derived from the reference engine as "flown" on the year-2000 reference mission and rotorcraft. Both high and low fuel prices, as defined in paragraph 2.1.5, were used.

SPECIFICPOWER INCREASE. PERCENT 4O 5O ilO 70 iiO 90 IO0 110 120 0.249 10.41) REFSFC = 0285 (0.468) REFSPECIFIC POWER = 248 [1511 / 0.243 / /

I0.401

-20 O225 10.37) 0.219 (0.36i 0 SELECTED CYCLES 0.213

i__A

10.35J 361.7 394.6 427.4 460.3 493.2 526.1 559.0 (220J 1340i 1240) 1260) (280] 13001 (320J SPECIFIC POWER,KW/IKG/S]. [HP/ILB/SECI] HP TURBINE: CERAMICS(C) MINIMUM STAGE COUNT LP TURBINE: CERAMICS IC) AND ADVANCED METALLICS [M. UNCOOLEDi 746 KW 11000 SHP) COMPRESSOR:TWO-STAGE CENTRIFUGAL G5-261-101 Figure 19. Rototcraft Simple-Cycle Performance Results - Ceramic Turbines.

(

OFp_R Qu_

,I

I

I[f • |lilSK

J

I

Improvements With Advanced Metallics and Ceramics.

I

ti

Figure 21 presents the resulting DOC values for the selected cycles in terms of deltas relative to the reference engine.

Again, the ceramic engines are superior to those with advanced metallic turbines. The ceramic engine, at a TRIT of 1427C (2600F) and a PR of 24:1, has the highest reduction in DOC.

Although increasing the pressure ratio from 22:1 to 24:1 has

If

additional DOC payoffs, this cycle requires an additional HP turbine stage which would increase maintenance costs (not

I1

accounted for during Task II). Based on these DOC results, the ceramic engine with a TRIT of 1427C (2600F) and a PR of 22:1 was selected for further evaluation.

I: I t Other cycle/configuration trades considered were i) replac- .-4 ing the two-stage centrifugal compressor with an axial-centrifu- gal, and 2) replacing single cooled HP turbines with two lightly

[i

loaded stages.

i I .o The axial-centrifugal compressor showed a slight performance T l advantage, 1.5 percent in SFC (Figure 22), compared to the two- stage centrifugal configuration (Figure 18). However, in terms of DOC, little or no improvement is projected beyond the engines with two-stage centrifugals, as shown in Figure 23. The SFC '4 advantage of the axial centrifugal is offset by the greater size and cost of this configuration.

The two-stage HP turbine also chows no performance benefit.

The additional cooling flow required for the second stage results in a reduction in overall engine performance, as shown on Figure 24 (compare to Figure 18). Engine weight, size, cost, and DOC are as presented in Figure 25, clearly showing the advantage of the single stage HP turbine, particularly in terms of DOC.

<

r'' 10%

'6o 1

143 141 138 138 139 0- 14f 10% I,,,- ..-;- ,_ 121

_ 8o-

63 64

N 50-

40- 20- CPR 16 18 20 20 22 22 24 26 22 24 26 TRIT 1204C12200FI 1316C 12400FI 1316C 12400F) 1427C {2600FJ _ .E_" MATERIAL METALLIC CERAMIC _ .., GS.281-80 DOC Comparison of Advanced Metallic and Figure 21.

Ceramic Engines.

, ,' ,7:1 "- "* +-+_ + ,i'1, :+ +, *'+' > +"

I!

SPECIFIC POWER INCREASE, PERCENT 0,280

0 I0 2O 3O 4O 50 60 7O 8O

I I l ,,,I, I iI I J I, 10.401 I , • Irx18C 12800F1 I 2M- IM o. 1427c I2NOF) 0.274 I 201 . t538c__ .

i0.45i o. 12o4c 12200F1 5 0.268 _12900F1 _. t316C t2400F)-- _ 10.441 0..

..se-

I 1427c I

e,l.- = 0.261 I _ (L_OOFI I .

ao [

_ 10.431

I / 24 1427C o.

". 0.255

i l

-,-" 10421 '26 16 "

o..<.j/ I -,%-.I,_o ,o

Cl0

_ 0.249 L+

I +-

_ 10.41)

L

0.243 10.401

i l

! 1

0.237 24 __:_12g_, SFC

!.5%

10.391 230 263 296 329 301 395 427 460

iJ

11401 11601 1180) 1200) 1220) t2401 1280) 12801 SPECIFIC POWER. KW/IKG,:S). [HP/ILB/SEC)) &, 0 SELECTED CYCLES T o 746 KW (lOOOSHPI T • COMPRESSOR: AXIAL + CENTRIFUGAL o HPTURBINE: ADVANCED METALLICS MINIMUMSTAGE COUNT ° o LPTURBINE: ADVANCED METALUCS -+ 65-281-85

i

22, Simple Cycle Performance With an Figure Axial-Centrifugal Compressor and i [..

Metallic Turbine.

I ._+

i

I 1,10 • ;.'1 e SPECIFICPOWERINCREASE.PERCENT 0.280 t 0 10 20 30 40 5O / (0.46) / 26"q 15,_8C] ' IM 0.274 12800F) 12M.

(0.45) I=., i,.,,- m.- 0.268 "5 z (0.44) 2M. Im t e',e.- -,e'- LL= e,n ,.,J w z 0.262 fj 2600F1 _ 1316C r 1-,,-, ,!

(0.43) ¢,.1 ,..,., t,L'_ 12400F ,,-,j e,,e,,, LLJ

12o4c

e'z" 0.256 t , !0 _ . .

(0.42) ¢/'J

IZ,OOW, _,;_le _r •

¢.3 I,.k.

0.249 i.

10.41) t_ 0.243 (0.40) 197.3 230.2 263.0 295.9 328.8 361.7 394.6 (120) 1140) (160] (180] 1200] (220) 1240) SPECIFIC POWER, KW/IKG/S), [HP/ILB/SEC)] o 746 KW (1000 SHP) 0 = 1204C (2200F) o COMPRESSOR: TWO-STAGE CENTRIFUGAL _= 1316C 12400F) • HP TURBINE: ADVANCED METALLICS TWO STAGE [3 = 1427C (2600F] © LP TURBINE: ADVANCEOMETALLICS O= 1538C (2800F) G5-281-I02 Figure 24. Simple-Cycle Performance with a Two-Stage (Cooled) Metallic HP Turbine.

6O I I I c¢1 Q,I °,,,4 t,_ .._ I o t..I .,,-I ,.-.4 O_ .4.,I o I .,.-4 t_ .,--4

:-"<: ""_: _"' ;"_ ":''"_i i: !_ ,_ ' "; ' ': ,_"_"7 '. "

• 0/_

LI

The impact of size, or scaling, was also examined. A 373 kW k (500 shp) engine was investigated with the baseline configuration to determine the performance impact. Comparing the downscaled engines (Figure 26) to the baseline (Figure 18), SFCs are in- creased by approximately 5 percent at the lower flow size because of a reduction in component efflciencies. For example, at a

H

pressure ratio of 20:1 and a TRIT of 2400F, SFC is 0.417 at the lower flow size, compared to 0.398 at the higher flow size.

U

2.2.2.2 Recuperated Cycles Study

J i

In addition to the conventional simple-cycle engines, a num-

[}

ber of recuperated cycles were evaluated for the rotorcraft ap- plication as well. The cycles were selected based on results from the commuter cycle studies discussed in paragraph 3.2. The resulting performance, presented in Figure 27, shows a signifi- cant reduction in SFC (up to 27 percent relative to the refer- ence), with an effectiveness of 0.7. Higher levels of effective- ness would further reduce SFC. The cycles shown use a variable LP turbine to take advantage of improved part-power performance (particularly critical for the rotorcraft mission) by maintaining

IT

a high temperature delta across the recuperator. Weight, size, cost, and DOC improvements are presented in Figure 28.

!i

2.2.3 Cycle/Engine Selection Based on the DOC results two cycles, a conventional simple .

cycle and a cycle using recuperation, were selected for further . ° evaluation in Task III.

The simple-cycle engine is configured with a two-stage cen- trifugal compressor with a pressure ratio of 22. The compressor is driven by a single-stage uncooled ceramic axial HP turbine.

The LP (power) turbine consists of two uncooled stages, and is L, i

h

I.-- Z I,,J,J ¢.1 e,e., O.

i

l_ CONFIGURATION (HEAT RECOVERY CYCLE) %1 ' • 746 KW )1000 SHP]

L J

COMPRESSOR: ONE-STAGE CENTRIFUGAL :j HP TURBINE: CERAMIC ONE STAGE LP TURBINE: CERAMIC VARIABLE GEOMETRY THREE-STAGE SPECIFIC POWER INCREASE. PERCENT 0.23] 30 40 50 60 70 [0.38) , L . I | I 1i J.

REF SFC = 0.285 [0.468) REFSPECIFIC POWER = 248 (151] !

,20 0.225 [0.37) _ = 0.7

i

_,p/P : 80/0 • ° I,-,,.

-j,, o.219 (o.36)

i

rz,- ° q (.,,.

o.213 _T 4 = 1427C ..

(o.35l r...1 I • ° 0.206 Io.34]

T "_r m-we

I to

• SELECTEDCYCLES

0.201 (0.33J J 296 329 362 395 427 [I 801 1200l 1220] (2401 [2601 SPECIFIC POWER. KW/IKG/SI. [HP/(LB/SEC)] G5-281-96 Performance of Recuperated Cycles for Figure 27.

Rotorcraft Application.

k I .... ' .JL .L_ l= cr_ e- I r,_ ¢.l

° i

q-i ql; > 4

° i

° t

I -.-4 q_ o _h oo ¢N HP TRIT was set at ceramic as necessary to remain uncooled.

1427C (2600F).

Based on DOC considerations in Figure 28, the heat recovery cycle was selected with an uncooled ceramic turbine at 1427C (2600F) TRIT. The single-stage turbine drives a single-stage centrifugal compressor with a pressure ratio of i0. A multistage variable uncooled power turbine is incorporated for better part- power fuel consumption, which will be ceramic as necessary. The fixed-boundary recuperator has an effectiveness of 0.8 with an 8 percent pressure drop (based on the results of the commuter study, section 3.2.3).

2.3 Task III- System Performance Evaluation In Task III, the two engines selected from Task II were evaluated in terms of their impact on overall air:raft system performance. A detailed mission analysis was conducted (using the HESCOMP model) for both aircraft/engine systems requiring an extensive matrix of off-design performance. Finally, the selec- ted engines were evaluated in detail with the HESCOMP economic model to determine aircraft direct operating costs (DOC).

Prior to the system performance evaluations, both selected configurations were further refined by the typical GTEC prelimi- na£y design process.

2.3.1 Engine/Cycle Refinements Several minor engine refinements were made to the year-2000 engines selected in Task II. These refinements were made follow- ing a more detailed design analysis of each engine component.

The identified reflnements and the resulting performance effects are summarized in Figure 29. For the simple-cycle, size, weight, 'I !

i z t ¸ _z i -.-- -,

<,l

Ill

! i0--

-i II. i,_l ii1 t,,.,I m

°t !=

I,,.i

<'._i_\ _

o t

_::_ =_

" i

_ ,,-i I_ l I!.

" t

_. ____ ,,m I,.,.,- _ • ° _- . _ _- m ,-4 o 0

,.-, 1

e,, - _

-

e:i ._ II . ,,,..._ r..

_ _ _ _-- _ _ • li..

i,--: c_ z ,,..,i

= i

67 J

,i ,#, and cost remained relatively unchanged. SFC and specific power,

I

however, improved by 3.1 percent and 2.9 percent respectively, primarily due to an increase in LP turbine efficiency. LP tur- '\, bine efficiency was improved nearly 5 points by the addition of a

L

second stage. This stage addition was necessary because the tur- / bine corrected work (AH/_) exceeded the limit as projected for

iJ

the year 2000.

/ r,

SSr

6_ For the recuperated cycle, engine performance changed only slightly, but weight, size, and cost estimates were revised. The

!i

changes in size, weight, and cost are primarily due to a reas- sessment of the recuperator design and cost estimates, resulting

H

in a size and weight reduction of 6 to i0 percent and a cost increase of 9 percent. These refinements were analyzed and found to have no effect on the optimum cycle selection.

2.3.2 Mission/Economic Analysis The mission and economic analyses fcr the rotorcraft appli- cation were conducted with the reference mission and aircraft using the HESCOMP mission/economic model, all as defined in Task r q I.

/ I To support the mission analysis, a matrix of off-design per- formance ccnditions and power settings were generated. A compar- ison of sea-level, static load lines (Figure 30) indicates that the recuperated engine has superior part-power SFC to both the advanced simple cycle and the reference engine. A Mach number

i

lapse rate comparison as presented in Figuze 30, shows little difference between the three engines.

The resulting mission performance is summarized in Figure 31. As shown, the key change is the reduction in fuel burn, 21.9 [ 0.467 SLS, ISA.]

IO,eOI _. 0305 10.801 REFERENCE t i _ 0.243 %_ YEAR.2000 • _.--,-,, SIMPLE _ 10.401 _._ YEAR-2000 RECUPERATED _ [0.201

i

I

0 ' _96' ' s_7' o95

I01 14001 18001 [I 2001 POWER. KW ISHP]

I

I YEAR-2000

RECUPERATEO 110801.

• E E E,,cE v,LOES []

\\\

6.9 2.6 2.6 1; L j' k\_! :_1 TOGW ROTOR FUEL WETTED BLOCK !i 4338.2 KG* DIAMETER BURN AREA TIME :i 19564 LBI 107.7 CM" (2.31 HR) o -10-- 537.5 KG* M2 (1185 LB) 68.1 (42.4 IN) (733FT2) * G5-261.78 Figure 31. Projected Reductions in Rotorcraft Parameters.

7O ..... i ...............................

-'-.-'_%h._mL_--_ ,'T_-7_ T ,:% _, . _ +1,,, _I percent for the simple _ cycle, and 41.6 percent for the recuper- i.b ;J ated cycle relative to the reference engine. The fuel burn reduction for the simple cycle is approximately the same as the percent reduction in design point SFC (21.2 percent). For the recuperated cycle, however, the fuel burn reduction is greater than the percent SFC reduction at the design point (31.8 per- cent). This is due to the superior part-power performance of the recuperator engine. Takeoff gross weight has also been reduced by 6.9 and 4.5 percent for the simple and recuperated cycles, respectively compared to the reference engine.

Figure 32 shows the recuperated cycle to be superior in DOC to the simple cycle at both fuel prices. Despite the negative impact of weight, size, and cost incurred with the recuperated engine, the SFC benefit is large enough to offset these aspects.

Relative to the reference engine, as shown in Figure 33, the advanced engines reduce DOC by 7.0 and 7.4 percent, for simple and recuperated cycles respectively, at the low fuel price. At the high fuel price the year-2000 simple cycle can reduce DOC by 8.7 percent. However, the recuperated engine, which has a 41.6 percent reduction in fuel burned, has a DOC reduction of 11.4 percent.

2.4 Task IV - Small Engine Component Technology Plan Task IV identifies and quantifies high payoff technologies for the rotorcraft engines and presents technology plans that are based on the benefits as projected for the high payoff technolo- gies.

2.4.1 Technology Identification/Benefits Tasks II and III performance and DOC results are based on a number of technology projections. These technologies have shown

!I

PR = 10,TRIT = 1427C (2600F] 160' PR = 22. TRIT = 1427C(2800FI kd c = 0.8, _P/P = 8 PERCENT =-. 14_.

,_ 120- _ _'}J $0.264/LITER

ioo-

I-1 $0.529/LITER "=-'= 80" ($2/GALi L (St/GALl 60- 40- "i t "/.,4/'./1 _ _ -,,-_ // 20" O :!

SIMPLECYCLE ENGINE "'-' RECUPERATED ENGINE Figure 32. Projected Reductions in Rotorcraft DOC for Simple Cycle Versus Recuperated Engines.

(

%, ( 1600- I

T

-8.7% -I 1.4%

t

,l

"-1400 a.

1424 'I i .... "1382 -7.0% -7.4%

i

;1282 iz'Te

1200-

I

Z ¢ %. 0¢%.

flOG

I

So.2_,/,,T_R Is,/_.,_J

$0.528/LITER ($2/GAL)

I

i

Figure 33. Helicopter DOC Results _or Simple Cycle and Recuperated Engines.

I

| | n I ii

i

benefit in terms of SFC, weight, size, or cost. Several of these technologies, such as metal matrix shafts, are difficult to quan- tify in terms of DOC, but their use is considered beneficial or necessary to meet engine design goals. The identified technolo-

[

gies are:

i

diameter, length, and cost data. Finally, trade factors were applied to the new engine parameters, which resulted in new DOCs.

Comparing the resultant DOC value to the DOC for the baseline r year-2000 engine with all technologies shows the improvement derived from that technology. The selected technologies are not

!I

independent from one another and are therefore not additive.

L 7!

Of the technologies quantitatively investigated, hot-end materials were found to have the greatest DOC impact, as shown in Figure 35. For example, ceramics for application in turbine air- foils contribute approximately half of the overall DOC improve- ment projected for the simple cycle engine. Ni3AI for turbine disks was found to have the second greatest DOC benefit, follow- ing ceramics. Removal of Ni3AI turbine disks and reducing hub o.

speeds to values consistent with today's disk materials resulted -i in reduced efficiencies and increased turbine stage count for the l" simple cycle.

°.

.I Compressor and turbine efficiency improvements also show significant DOC benefits. Combustor pressure drop reduction technology and improved compressor materials, however, resulted i in only small improvements in DOC.

° In addition to the technologies examined for the simple cycle, the impact of ceramic recuperator technology was quanti- fied for the recuperated engine. As shown in Figure 36, the recuperated cycle has a DOC advantage over the simple cycle.

Ceramics, both in the engine hot section and in the recuperator itself, are vital for the recuperated cycle. The use of a ce- ramic recuperator results in nearly a 3 percent decrease in DOC.

i _ • l SIMPLE-CYCLE] ENGINE J TOTALBENEFIT,$O.528/LITER($2/GALI TOTAL eENEFIT $0.264/LITEfl ($1/,_.GALI $0.254/UTER I$1/GAL) $0.sze/LITER ($2/GAL] I"_--_ "41 "-a I _'_ J ¢J_ ¢,_ _1_ r-- 1,-- , _ ll= g[_ IL;g Ni3AI CAST ALUM COMPONENT Ti PERFORMANCE Figure 3E.

Projected DOC Benefits (Percent) for Isolated Technologies.

................. "• -++ ;- _ ._,,4= -,

J

i

$0.264/LITER ($1/GALJ F-,, Z $O.528/LITER 1$2/GAL) ., F-,,- L Z

f!

_a ALL . .

ALL CERAMIC TECHNOLOGIES TECHNOLOGIES RECUPERATOR G5-l'81.28 -° Figure 36.

° ° Comparison of Projected DOC Benefits (Percent) for Recuperated Versus Simple-Cycle Engines.

°_ _° 2.4.2 Technology Plan GTEC's recommended plan fo[ Small Engine Component Technolo- gies for year-2000 rotorcraft engines is presented in this sec- tion. This plan addresses a broad spectrum of technologies in keeping with the technology benefits as presented in paragraph 2.4.1. The plan is presented in sections for the following tech- nologies: 2.4.2.1 Ceramics 2.4.2.2 Recupecators 2.4.2.3 Metallics for turbines 2.4.2.4 Turbine performance 2.4.2.5 Combustor performance 2.4.2.6 Compressor (centrifugal) performance 2.4.2.7 Materials for "cold" parts 2.4.2.8 System technologies These plans address the high payoff technologies needed to obtain "technology readiness" by the year 2000. Some technol- ogies may require verification and engine demonstration testing prior to commitment to an eDgine full-scale development. This activity is not included in the technology plans; it may be con- ducted during the latter program years (ii through 14).

2.4.2.1 Ceramics

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Ceramics merit an all-out effort that should include er,gine

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A. Design Study to Conceptualize SECT Engines with Ceramics The current application of ceramics to hot engine parts is limited primarily to experimental parts of low mass such as vane '1 " segments, rotor blades, and thin-wall structures. A notable extension of this experimental technology can be found in the I AGTI01 automotive power plant as pioneered by the GTEC/Ford team for DOE/NASA. The AGTI01 design, which is fully committed to ceramics (Figure 38), incorporates a small radial turbine wheel I that is integrally cast of silicon nitride ceramic material.

I This program will explore the design opportunities of ceram- ics as conceptualized for year-2000 rotorcraft engines. The ini- on design application I tial effort will be based the of ceramics as limited to parts of low mass, and will be based on projected I properties of fully dense monolithic ceramics.

I A second design study will conceptualize a year-2000 rotor- craft engine fully committed to ceramics (monolithics and COL=DO-- sites), including integrally cast axial turbine rotor(s). The I study will include consideration of solid-hub rotor(s) based on projected properties of composite ceramics. It will use an aft- I power drive arrangement to accommodate the solid turbine rotor(s). An update of this design is also proposed to establish w I rotor design and sizing for second-generation composite ceramic test parts (Reference Program D).

I Program Description I | This program consists of three conceptual design studies, as ° _ • shown in Figure 37, Technical Program A, and described below.

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• i E. COMPRESSOR A. COM8USTOR • SINGLE STAGE CENTRIFUGAL • PREMIX/PREVAPORIZING • 5:! PRESSURERATIO • CERAMIC MATERIAL • POWDERMETAL ALUMINUM B. REGENERATOR • VARIABLEINLET GUIDEVANES • CERAMIC MATERIAL • ROTARY-- EXTRUDEDMATRIX F. FOIL GAS BEARING • DEVELOPED SEALS AND BEARINGS G. CERAMICSTRUCTURE • TRANSITION LINERS. TURBINE C. TURBINE SHROUD,TURBINE STATOR • SINGLE STAGE• HIGH WORK RAOIAL • MATERIALSINCLUDE RBSN.

• CERAMICMATERIAL SINTEREOSiC, RSSIC • 1371C 12500F) MAXIMUM TURBINE INLET TEMPERATURE O. BALLREARING • RADIALAND THRUST LO_.S G5-281-35 AGTI01 Power Plant.

Figure 38.

J • 82 ..... V-- Monolithic._s - Analytical design will include an assessment i of fully-dense monolithic ceramic materials, manufacturing pro- ,I | cesses, and joining methods. Properties and geometry/mass limit- |_ ations for design will be projected for a year-2000 rotorcraft engine. Preliminary ,[zing_ of components will be made and the I SECT engine cycle (simple) will be reviewed and revised as neces- sary to track with the conceptual engine design.

_' I Conceptual board design will be conducted to configure the I engine based on technologies as projected for year 2000. Trade studies will be conducted as relevant to design options. A con- _ | ceptual engine design will be depicted in cross section. Ceram- | ic-to-metallic interface features will be conceptualized and depicted in supporting section views.

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i Composites - Analytical and conceptual engine designs will i_.. i be conducted in the same manner as for monolithics except that ceramic properties and geometry/mass limitations will be based on I projections for ceramic composites. Integral ceramic turbine rotors will be assumed for the HP turbine and for the first stage _' of the LP turbine. Components will be sized in accordance with I projected properties. Solid-hub rotor(s) will be considered I along with an aft-drive engine arrangement.

Analytical and conceptual engine design will be reviewed and program years on I updated in seven and eight, and will be based updated ceramic material properties. Integral rotors will be i I reconfigured and resized (to define test parts for Program D).

i Technical Approach The simple-cycle rotorcraft engine, evaluated in Task III, I will form the basis for this study program. The salient features of this engine are:

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,14 L_ o 746 kW (i000 shp), front drive

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o Two-stage centrifugal compressor o Reverse-flow annular combustor incorporating ceramic / !

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that offer the potential for higher toughness and consequent non- catastrophic failure modes. Progress of these composites result- ing from separately funded activities or from Programs D and E should be assessed and incorporated if/as feasible for the later

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years of this program.

Ceramic technologies from this program will benefit both _ rotorcraft and commuter aircraft applications and APUs.

.° o,i Program Description This program is scheduled as four major activities, from material/process efforts through engine environment tests.

Fully-dense monolithics are planned for the initial effort, which e.

is scheduled for seven years. A second iteration is shown, for i planning purposes, to establish technology readiness for the most suitable ceramic material (monolithics or composites) available in the 1992 time frame. Execution of the second iteration would me depend on program results through the first iteration, on the

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outlook for higher-payoff integrally cast ceramic rotors (i.e., Technology Programs D and E).

A pilot combustor and turbine stage will be designed to be representative of the SECT rotorcraft engine and to be compatible r with an existing GTEC test-bed engine. The pilot turbine stage would be designed to replace the first stage of the HP turbine.

° t .° Material vendors will be surveyed and the best available , high-strength, high-temperature ceramic materials will be pro- cured and tested to obtain the required design data. Ceramic _o vendors and fabrication processes (net or near net-shape) will be ..

chosen for each ceramic component. Fabrication evaluations will _4 be conducted to verify the fabricator's capability to produce _ high-quality components by the selected fabrication methods.

-i i Specimens and experimental hamdware will be tested.

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_° A separate study will be conducted to establish a compliant

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layer for the blade dovetail attachment to achieve long life and a high-temperature capability.

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.i ceramic vanes. The vane design will be based on the test engine

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requirements, and on prevailing design and fabrication processes Y / for monolithic or composite ceramics. Based on the detailed ..

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design, performance predictions will be made for comparison with test results.

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Fabrication - A set of ceramic variable-geometry turbine vanes (and spare parts) will be fabricated/procured.

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Test - A se£ of the ceramic vanes will be assembled, along with engine parts, into a test unit. This unit will be bench tested to verify proper function and to assess gas path leakages.

The unit will then be assembled into a test-bed engine and operated through its full range of operating temperatures, and _e with full gas path loads. Engine and turbine performance data will be obtained and compared with predicted values.

bo Technical Approach The selection of an existing GTEC test-bed engine will be

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I made to facilitate LP turbine operating conditions and test data b_ that are representative of recuperated rotorcraft engines as con- .° figured by the SECT study program for year 2000. Analytical .o design, including engine cycle evaluation, will be conducted to configur_ a set of ceramic variable-geometry stator vanes for the LP turbine section. This design activity will be based on mate- rials and manufacturing inputs for low-cost fabrication methods as envisioned for ceramics in the year 2000.

Testing of the vane set will include evaluation of leakage, performance, wear, and dynamic characteristics.

J °, P. Composite Ceramic Axial Turbine Rotor/Stator Ceramic materials have the potential for uncooled component use at high turbine inlet temperatures. Ceramics are lighter than comparable metal components and offer the potential for lower cost. The feasibility of a metal axial rotor with inserted

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ceramic blades has been demonstrated for short-life applications.

However, the full benefit of ceramics is better realized in an integral axial rotor. This eliminates the costly ceramic and metal machining required for the blade dovetail attachment, and

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eliminates blade stagger angle constraints on the aerodynamic design of the airfoil.

Since the stresses around a bore hole at the center of the rotor are expected to exceed the material capability, the pro- posed rotor would use a solid disk with a stub shaft or other attachment concept not requiring a bore hole. In the long term,

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ceramic composites are the prime material candidates for improved toughness of the rotor and reduced potential of catastrophic failure.

The feasibility of ceramic radial rotors having a large hub

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mass is discussed for APUs in paragraph 5.4.2.1 as a parallel or follow-on technology program.

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Program Description This program is scheduled as four major activities, from materials/process development through bench tests. One genera-

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tion of test parts is planned through bench/spin tests during the first six program years. Configuration of these parts will be

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established based on the engine conceptual design studies con- ducted in Program A. Second-, third-, and fourth-generation parts are envisioned for engine environment testing (Program E) in program years i0, 12, and 14, respectively.

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, Composite material systems will be surveyed and evaluated I for properties and suitability of processes. Design data will be r obtained for selected material systems. Specimen tests will be conducted along with attachment tests.

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/ The design of a pilot turbine rotor with a solid hub (no bore hole) and a matching stator will be conducted, based on .o results from Program A. They will be configured to be rep. esen-

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tative, in size and shape, of the SECT rotorcraft engine (as con- ceptualized for the year 2000), and will be designed for compati- bility with an existing GTEC test-bed engine. The aerodynamic and mechanical design of the pilot turbine parts will be based on attachment concepts, material systems, and fabrication approaches as selected during this task.

Fabrication and attachment development will be conducted for the pilot design/hardware. Rotor/shaft attachment schemes will be evaluated, and pilot parts will be bench- and spin-tested to demonstrate attachment concepts and other critical design features. The stators will undergo thermal shock tests.

Based on the results of the pilot design/test evaluation, a design update will be accomplished incorporating the latest aero-

dynamic technologies. The design will be completed and ceramic i7

parts will be fabricated for bench/spin tests and for follow-on i engine environment tests (Program E). Parts (metallic) will also be fabricated for cold aerodynamic rig tests, which will be con- ducted to fully map the turbine stage.

Based on the experimental results of this hardware, follow- on design/fabrication test cycles are projected for planning pur- poses to achieve technology readiness.

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Analytical and board design will be conducted to modify the

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selected test-bed engine (such as the Garrett FI09) to accommo- date the ceramic test hardware, as provided by Prcgram D. The

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gas generator section of the selected two-spool engine will be used for this testing. This will accomodate a solid-hub rotor

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(no bore) with a forward drive for the compressor section.

Engine refurbishment and special adaptive hardware will be fabri-

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cated.

Engine gas-generator testing of the experimental ceramic

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turbine will be conducted (program year 10) to the design temper- ature for the ceramic parts. Performance and mechanical data will be obtained and compared with design predictions. A tech-

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nical report will be published.

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Two follow-on tests are planned for years 12 and 14, based on successful progress for the composite ceramics Program D.

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E, Technical Approach Specifications for a turbine stage will be established that are representative of the SECT rotorcraft and commuter engine requirements as envisioned for the year 2000. The specification will also be compatible with the selection of a suitable engine test bed for verification tests.

Analytical design of the integral-ceramic, axial-turbine rotor will draw from the latest technologies available for aero- o_ dynamic and mechanical design, materials properties, and fabrica- tion/manufacturing processes. The aerodynamic design will take full advantage of high'blade stagger analyses as achievable with integral blade-disk rotors. The mechanical design will accept _° risks commensurate with engine benefits and will be based on the available material properties as established for the pilot and subsequent rotors.

_G 2.4.2.2 Recuperator This plan includes a comprehensive program to establish the B_ *i fabrication technologies necessary for a ceramic platefin recu- t/ perator. The technologies would be applicable to a recuperated

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engine as envisioned for the year-2000 SECT rotorcraft. This plan provides for an experimental recuperator program and an engine environment test program. Figure 39 presents the schedule for this plan, which is comprised of two discrete technology pro- grams (identified as F and G).

This plan does not include a technology initiative for regenerators. It is envisioned that this technology will be addressed separately by continuation of existing programs such as the NASA/DOE-sponsored AGTI01 vehicular engine program.

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Schedule.

Figure 39. Recuperator Technology

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i F. Ceramic Recuperator Fabrication Development k j_t Present metallic, flxed-boundary recuperator technology has i limited application for future airborne propulsion systems due to excessive weight and insufficient temperature capability. The T_ ] -, introduction of ceramic technology to recuperators reduces weight and increases the temperature operating range.

The SECT ceramic recuperator program objective is to evolve the appropriate fabrication technology to produce a finned-plate recuperator suitable for application to a specific small engine design. The specific goals of the study are:

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(a) Verify the fabrication technology necessary to produce full-size plate-fin recuperator modules and associated D_ ceramic manifolds, ducting, and ceramic/metallic inter- faces (b) Fabricate a full-size recuperator module (c) Test the full-size module under simulated engine condi- tions Program Description !

This program, which is scheduled for 60 months, includes experimentation with both a subscale and a full-scale recuperator module. The program is based on feasibility work already com- pleted on a core module. Figure 40 depicts these three modules of increasing complexity, while Figure 41 presents a detailed schedule for this technology program.

Task i t Module B Design and Fabrication - The fabrication technology will be demonstrated by constructing a partial-stack- / q.

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MODULE A RECUPERATOII MOOULE ICORE MOOULEi ISUBSCALEI G5-281.I0 Figure 40. Test Modules of Increasing Complexlty.

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height recuperator (Module B). The major effort of this task is to fabricate a working subscale recuperator module.

Desiqn Iterations - The initial reference design estab- lished at the outset of the program will be based upon the results of the initial fabricate and test cycle for Module A. Three design iterations are expected to achieve program goals.

o Thermal and Stress Analysis - This task will concen- trate on structural analysis to keep the stresses in the ceramic components within the engineering material strength for the various environmental loadings to which the recuperator will be subjected. Because the engine may undergo significant transient operation, investigation of transient thermal stresses in the recuperator may be vital to ensure the structural

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integrity of the ceramic recuperator.

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The recuperator operating conditions require that close attention be given to design for pressure containment. Mani-

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folds, port connections, and manifold end closures will be ana- lyzed. The analysis will include consideration of the loads and stresses in the core. Particular attention will be given to the bonding and sealing of ceramic ducts to the core. An objective of the design is to minimize stress concentrations in critical

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areas. A major portion of this task will be to determine the containment requirements/problems.

o Compression Moldinq - The purpose of the compression

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molding task is to fabricate the finned plates required for the construction of the Module B cores. Forming

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ceramic finned-plate heat exhangers is an iterative process involving increasingly complex shapes. The

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I/7_ i'_• i'_ expertise gained in molding the Module A and sample recuperator plates during previous programs will be used to form the Module B recuperator.

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Tooling for Module B will be procured as early as pos- sible in the program in order to initltate compression molding studies. The fabrication effort has been structured to allow three iterations, with approxi- mately four modules produced in each iteration. Form- ing parameters will be varied as necessary to produce acceptable individual finned plates. These plates will be used to assemble modules for use in performance, f I pressure, and thermal shock tests.

-s o Duct and Attachment Development - Ceramic ducts provide the transition between the ceramic core and the metal engine ducts. Preliminary duct concepts are shown in L Figure 42. These shapes are rather complicated and will probably require a strong development effort.

Several sets of ducts will be fabricated and used for L4 Module B testing. It is anticipated that the ducts will be fabricated from the same material as the recup- erator core.

J Because a significant effort will be required to develop a suitable method for forming the ducts, par- J ticularly if an unusual geometry is required, this task has been scheduled over a 22-month period. The forming

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method used to make the ducts (e.g., either slip cast- ing or injection modeling) will dictate the type of

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bond used in attaching the duct to the recuperator core. This high-temperature, high-pressure seal is a !L key development area for ceramic recuperators, o L_ W ...........

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AIR OUT AIR IN GAS OUT F GAS IN _I-14 !

o Figure 42. Conceptual Ceramic Duct Design for Recuperator Core.

.J 6, t L _ ___h"_--T_. _ .'_W_.* _%'%'I" _ _,_ " _-_ '_ ': i i iii ii j ii , t Assembly - After all recuperato_ components are fabri- cated, they must be assembled to form modules for use in testing. Fixtures are needed to align and hold the •j various components during bonding. Good alignment is important so that dimensional tolerances are main- tained. Techniques from the previous subtask that !

/ effect duct-to-core bonding will be used in assembling the modules.

o Processinq - The assembled modules must be run through o.

binder extraction and nitrlding cycles to convert the green module to a ceramic recuperator. Further pro- cessing cycle effort may be necessary to provide an acceptable strength recuperator for high-temperature, i high-pressure service. For example, if the binder is extracted too rapidly, bloating may occur, seriously degrading the material strength.

Task 2 t Module B Testinq - The Module B testing program is designed to perform independent pressure, thermal shock, and per- formance tests on the partial-stack-height recuperators.

The modules fabricated in Task 1 will be subjected to the described tests to verify the recuperator's capability to func- tion at the design pressure and temperature, as well as verify ° .

the heat transfer and pressure drop predictions. Thermal shock resistance also will be tested.

! o Fixture Design and Fabrication - A test fixture will be designed to pressurize Module B at ambient tempera- tures. Pressure differentials will be evaluated across the air- and gas-side layers, up to 1.25 times the design pressure differential. The test fixture will be I T" :I I00

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specially designed and fabricated so that it can be used to subject Module B to low-temperature (less than 427C [800F]) performance tests.

o Performance Test - The specially designed pressure/flow

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test rig will be used to measure the module's heat transfer rates and pressure drops. These performance tests, conducted below 427C (800F), will be used to make accurate predictions for heat exchanger perfor- mance at high temperatures. This is a standard proce- dure for evaluating high-temperature heat exchangers.

0 Pressure Test - The pressure test is a nonflow test to check the pressure containment capability of the recup- erator module. Pressurization to 1.25 times the design pressure differential is expected. If a module fails the test, fracture analysis wil be conducted to iden- tify the cause of failure (e.g., inadequate bonding, inadequate dimensional tolerances, or fabrication flaws). Efforts will then concentrate on the identi- fied problem area to develop a solution.

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Thermal Shock - When in service, the recuperator will

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be exposed to both a wide range of temperatures (ambi- ent to 1260C [2300F]) and rapid temperature changes.

To determine suitability for use under design condi-

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unit to the engine. An 18-month time frame has been allotted for this effort. An adequate ceramic-to-metal i joint is required for the high-temperature, high-pres- sure conditions. In addition to a mechanical seal, such as that shown in Figure 43, a permanent joint may L be feasible. One such joint has been demonstrated by sputter-coating silicon nitride with titanium and braz- *e ing it to a compatible metal component. Several alter- native solutions will be sought while researching this

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problem. 6-4 Design - Module C will replicate the plan form of Module B unless test results indicate the need for a .

design modification. The high-stress regions are expected to be localized at the air outlet manifold/gas ° inlet face region, based on previous analyses. To b_ reduce the maximum stress level, changes in the recu- perator plate configuration may be considered. The changes will be aimed at more uniformly heating the gas inlet face/air outlet manifold area. One possibility is to form channels or fins in the manifold to allow f either the incoming gas or the exiting "air to flow

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Compression Moldinq - Tooling for Module C will be pro- cured as the Module B fabrication effort nears comple- tion. Module C is similar in shape to Module B, but is larger in stack height. Therefore, only fine tuning of the forming and processing parameters is anticipated.

Three modules will be constructed for use in simulated engine tests.

Form Attachment Components - To complete the modules, i

ceramic ducts and other attachment components must be fabricated. The components required to effect a joint between ceramic and metallic ducts will be fabricated.

o Assembly - The full-size core will be bonded and the ducts and other attachment components adjoined. The size of Module C dictates the use of large fixturing

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:l tools to achieve the proper alignment. This assembly essentially replicates that of Module B, but on a larger scale.

be developed to maintain the desired temperatures, pressures, and flow rates for both air and gas streams.

The rig will be highly instrumented to record tempera- tures, pressures, and flow rates on both the inlet and

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outlet sides of the recuperator. This test rig will be used to qualify recuperator modules for installation in i, vehicles.

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O Construct Rig - After the test rig is designed, the

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unit will be constructed. The design task is scheduled

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,| I Dear the beginning of the program so that sufficient time is available to order and receive components for r" the test rig. Upon completion of the rig, a check run i will be made to ensure that the unit can provide a suf- ficient quantity of air and gas at the temperatures and pressures of interest. Also, the rig will be cali- brated before testing of any modules.

o Duct Interface Testinq - Potential ceramic/metallic duct joint designs resulting from Task 3 will be tem- perature- and pressure-tested. Laboratory tests are necessary to ensure that the components can survive the pressure and thermal stresses at elevated temperatures.

In addition, joints will be tested for sealing integri- ty under these conditions.

• ° o Pressure Test - MGdule C will be pressure-tested at ambient temperature to verify the pressure containment capability of the prototype.

o Thermal Cyclin_q - Modules will be exposed to cyclic variations in temperatures of 21 to 1260C (70 to 2300F) to ensure that the recuperator module can withstand the _hermal stresses resulting from these temperature extremes.

o Simulated Enqine Test - The simulated engine conditions test is the most significant test to be conducted dur- ing the program because it will verify the capability of the ceramic finned-plate recuperator to perform in a typical engine environment. Steady-state and transient test conditions are envisioned. The steady-state test T ,L./ i II] condition, which will be achieved by bringing the recu- perator up to temperature slowly, will verify the pre- dicted recuperator performance (heat transfer and pres- sure drop].

Transient tests simulate engine startup conditions.

The transient test will generate operational-level stresses because of the temperature gradients imposed.

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427C (800F) in order to generate heat transfer and pressure drop I: data. Since performance values measured at low temperatures can J i be accurately scaled to the higher operating temperatures, actual thermal performance of the module will be contrasted with pre- dicted values to substantiate a full-size reference design recu- perator.

) Upon completion of performance testing, modules will be ._° pressurized using large pressure differentials to determine the T pressure containment capability of the design. Modules that sur- vive the pressure tests will be exposed to increasingly severe thermal shocks to obtain thermal shock resistance of the unit.

After each thermal cycle, modules will be pressurized to deter- mine whether any damage was incurred as a result of the thermal cycling.

The data collected from these tests will be used to generate the Module C design. Three design/fabricate/test cycles are anticipated to evolve a satisfactory design for this recuperator.

The full-size recuperator (Module C) will be constructed and tested under simulated operating conditions.

The reference recuperator design will be modified, if neces- sary, based upon the results of Module B tests. Also, the asso- ciated parts required to seal the ceramic recuperator to metallic engine components will be designed. After the tooling needed to form these parts is obtained, full-scale prototype modules will be fabricated.

These modules will be subjected to independent pressure and thermal cycling tests and will then be exposed to simulated engine conditions. These tests require the design and construc- tion of a rig that produces the desired conditions. This design/ construction effort is expected to last a significant period of time because of the sophistication required. The intent is to construct the test rig in such a way that it can be used for prooftesting of each production module before installation in a vehicle. Because of the high reliability desired, it is antici- pated that each production recuperator module will be prooftested at stress levels in excess of design conditions to qualify it for installation in a vehicle.

Along with the two recuperator modules, the other components required to produce a functioning heat exchanger will be devel- L i - --_'C-IL-_ _,,_ , ...._,.- _,.i= ¸ , ,,_ _ . J- _____= . ' _M., , _':<k.. : -: , ,",,,,_ , ' , " ",'",. - "' ,_._,[e ." ""'_"'_' '- - " . i! ."_:-:,: ' < "/.,_ '. -:.: :_ .

_,_! oped. These include the ceramic ducts that connect the recupera- _'",'i tot core with the engine metal duct work, and the attendant seals and mechanical Joints required at this interface.

A ceramic plate-fin recuperator shows the potential of solv- ing the weight and temperature limitations associated with pre-

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sent metallic designs. The ceramic design could be the key to making the benefits of heat recovery (imposed fuel economy) avai- t, lable to airborne applications.

G. Engine Environment Test of Ceramic Recuperator The objective of this program is to conduct engine environ- ?

ment tests of a ceramic plate-fin recuperator. This experimental program would be based on successful completion of the ceramic

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plate-fin recuperator Program F.

Proqram Description This program consists of three tasks and provides for recup-

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erator design, fabrication, and engine environment testing as scheduled in Figure 39.

The recuperator will be designed and sized for the power class of the year-2000 rotorcraft and commuter engines. Compati- ble engine parts will be selected from existing GTEC engines and incorporated into a recuperated engine system.

° The ceramic recuperator and engine parts will be fabricated and component tests will be conducted. The components will be assembled into a recuperated engine test unit and instrumented

for performance and mechanical data. The unit will be submitted L

to mechanical shakedown, performance and controls, and acceler- ated mission tests. Teardown inspection will be conducted at intervals and results will be documented in detailed reports.

ii0 i | 2.4.2.3 Metallics For Turbines This plan comprises four interdependent technology programs identified as Programs H through K. As scheduled in Figure 44, the programs involve material and process efforts for turbine rotors with the objective of increasing operational speeds and temperature levels.

_. Ni3AI Alloy Composite Disk Current nickel-base alloy turbine disk materials such as Rene 95 and Gatorized TM IN100 are limited in specific strength by their relatively high densities and the inability of conventional gamma prime precipitation strengthening mechanisms to achieve very high strength without serious degradation of fracture prop- erties. In order to overcome these strength deficiencies, the classical approach to turbine alloy development has been the addition of larger quantities of gamma prime forming elements, which also tends to reduce incipient melting temperatures. This approach results in a significant decrease in the temperature window between the grain recrystallization/growth temperature and the incipient melting temperature. This further restricts pro- cessing to controlled grain sizes for amelioratiotl of unaccept- ably rapid crack growth and uncontrolled fracture. Reliance on gamma prime strengthening also inherently limits maximum service temperature of conventional alloys.

Lightweight, inherently stable turbine disk materials with acceptable strength and fracture properties can be developed through the revolutionary approach of metal matrix composites.

Intermetallic (Ni3AI) compounds display the low-density, oxida- tion resistance, and thermal stability necessary for turbine disk application; however, precipitate strengthening mechanisms are iii ........ I I ill

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NIsAI ALLOY COMPOSITE DISK • NEW RAPIO-SOLIOIFICATION.PROCESSlN6 iRSP) OF NiCKEL-ALLOY MATRIX WITH SILICON-CARBIOE FIBERS PROMISES HIGHER ROTATIONAL SPEEDS • MATERIALS COATINGS/PROCESS EXPERIMENTS • SPECIMEN TEST, DESIGN DATA T • EXPERIMENTAL PARTS DES/FAD/BENCH TEST LJ • ENGINE ENVIRONMENT TEST I SiC FIBER-REINFORCED SINOLE CRYSTAL TURBINE BLADES L.

• LOW-OENSITY HI6H-STRENGTH BLAOES PROMISE HIGHER BLADE SPEEOS FOR LPT • MATERiALS/COATINGS/PROCESS EXPERIMENTS • SPECIMEN TEST DESIGN DATA • EXPERIMENTAL PARTS DES/FAD/BENCH TEST • ENGINE ENVIRONMENT TEST

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J OIFFUSION.BONOEO LP TURBINE ROTOR IOISK/BLAOES) • 80NOING.PROCESS OEVELOPMENT BASEO ON CURRENT MATERIALS • 80NOtNG.PROCESS EXPERIMENT3 • SPECIMEN TEST. DESIGN DATA • EXPERIMENTAL PARTS OES/FAB/BENCH TEST • ENGINE ENVIRONMENT TEST K FOLLOW-ON OIFFUSION.BONOEOROTOR _tisAI ALLOY' OiSK + SiC FIBER-REINFORCEO SINGLE.CRYSTAL 8LAOES • APPLY 80NOING.PROCESS TO IIEW MATERIALS FOR OUAL-ALLOY ROTOR TO ELIMINATE SPEEO ANO AIRFOIL CONSTRAINTS OF INSERTEO BLADES ?

• BONDING-PROCESSEXPERIMENTS • SPECIMEN TEST OESIGN OATA • EXPERIMENTAL PARTS OES/FAO/9ENCH TEST • ENGINE ENVIRONMENT TEST SECT TECHNOLOGY VERIFICATION 65.Zlll-OO i i Material Technology Schedule for Figure 44.

Turbine Metallics.

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not available to provide sufficiently high tensile capability in

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this class of materials. Addition of lightweight, high-strength fibers such as silicon carbide are therefore necessary to meet

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this requirement.

parts for components in program years six and seven. Engine environment testing will be conducted in program year seven.

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seven.

J. Diffusion-Bonded LP Turbine Rotor (Disk/Blades) The diffusion-bonding technology required for manufacturing

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uncooled turbine rotors for the LP turbine section will employ the most advanced blade and disk alloys available in designs of

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maximum mechanical efficiency. Current small engine production turbine rotors use either inserted blades with dissimilar

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materials for blades and disks, or integrally cast rotors of a single alloy.

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In the inserted blade designs, turbine rotor speed and effi- ciency are limited by the stress concentrations in the disk rim

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as imposed by the mechanical attachments of the inserted blades.

The need for mechanical attachments also imposes limitations on

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turbine blade hub solidity due to the limited space available for the attachments on the rim of a small-diameter turbine disk.

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In integral cast designs, the mechanical attachment problems I

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are eliminated, but maximum rotor speed is limited by the lower strength of a cast alloy hub versus a high-strength powder metal- lurgy alloy hub that is better for hub strength. The cast alloy hubs are also inferior to powder or forged alloys in low-cycle- fatigue life.

Advanced technology diffusion-bonded turbine rotors can com-

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bine the best single-crystal cast alloy with the highest strength/weight turbine disk alloy available in the time period of 1995 to 2000. The single-crystal alloys may have a 56C (100F) improvement in stress-rupture strength over the best of today's

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alloys, GTEC's SC-180. The turbine disk alloy will be a deri- vative of the Ni3AI intermetallic compound. It is anticipated i that this alloy will have strength characteristics equal to today's strongest alloys, with a 10 percent reduction in density.

i

When the technology is available the cooled diffusion-bonded bladed disk will be a candidate for use in the LP turbine sec- tions of rotorcraft and commuter engines.

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Program Description q This metal-joining program consists of four major activi- ties, ranging from materials experiments through engine environ- ° ment testing, as shown in Figure 44. It is a seven-year technol- . .

ogy program to establish diffusion-bonded integral turbine wheel technology for the best blade and disk materials available today.

This process should also be largely transferable to new metallics T as envisioned for the year-2000 time frame.

Various diffusion-bonding concepts will be considered for final selection. The screening process will include specimen i ¸ a i16 i I

!

testing of the different candidate alloys, bonding processes, and bonding geometries. A two-year effort is expected to select and

i

verify the bonding method. Following selection of the bonding geometry and process, an effort of approximately two years will be necessary to develop the bonding process for the production

; I

environment. Efforts will be geared toward reducing production costs through minimizing scrap and process time. Nondestructive

il testing procedure_ will also be established at this time. Speci-

men tests will be conducted and design data will be obtained in parallel with these activities.

I

Following initiation of the process development, an initial

I I

design will be started. Selected for experimentation will be a L turbine wheel from an existing GTEC gas turbine engine, such as the FI09 turbofan. The detail design will examine the modifica- tions necessary to incorporate an integral diffusion-bonded wheel !, ': I in this existing configuration. This design will be fabricated r- and bench-tested.

A second design and fabrication phase will incorporate both design changes driven by iteration 1 and the benefits of an improved bonding process as the process efforts conclude.

The fabricated hardware from design iteration 1 will be evaluated by component and rig testing. Component testing will consist of a cold whirlpit overspeed test of several turbine wheels, either to the point of failure or to the point of verify- ing adequate burst margin. Hot testing will be conducted in an existing or a modified rig. Both steady-state and transient con- ditions will be simulated to determine actual stress and tempera- ture levels.

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t Hardware from the second design iteration will undergo full

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engine testing in the selected engine. The engine will be cycled to expose the turbine to typical transient operation conditions.

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Technical Discussion / /' The program will use a GTEC test-bed engine for the engine !

blade design to the stator exit flow field (vane/blade interac- tion) are required to achieve the year-2000 performance levels predicted for the rotorcraft, commuter, and turbojet engines.

Year-2000 efficiencies reflect a 1 percent increase in efficiency due to optimizing vane/blade interaction.

The scope of the program will involve analytical and exper- imental work. Analytical work will consist of developing a design system tool for predicting interblade 3-D flow field to i: _(' optimize the stator/rotor matching. The experimental work will include turbine component testing to understand rotor influence on the stator exit flow field. The test vehicle will be the GTEC/Army single-stage high-work turbine, which would be directly

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applicable to the rotorcraft HP turbine. The results of this technology program will extend to the axial turbine configura- tions of the rotorcraft, commuter, and turbojet engines.

Program Description

L

This program consists of three major tasks, from analytical modeling through rig testing of a turbine, as shown in Figure 45.

The program is scheduled for eight years.

L

The program contains both the basic research and the analy- tical approaches, with a combination of the two in the sixth year. For the first two years, the research approach consists of rig design, fabrication, and instrumentation. The GTEC/Army high-work, single-stage turbine will be researched. Instrumenta- tion used will be a nonintrusive type of measurement device such as a laser two-focus (L2F) system. The third and fourth years will include the rig testing program and analysis.

if /'r* = ¸" :.

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, b J The analytical approach will begin during the first year and q k will extend to the third year. The main thrust of this approach will be to develop a 3-D code that will accurately predict the f / stator exit 3-D flow field. This activity could be either a Navier-Stokes solution or a Boyle-code-type analysis. In the fifth year, the results of the rig testing will be incorporated I a into the code. Results of the rig tests will include J the rotor influence on the stator exit flow field.

.6 The :emaining three years (sixth through eighth) will involve use of the code to design a new turbine to the same velo- city diagrams as the GTEC/Army high-work single-stage turbine, and testing the design. In the final year, a final report will be written on the results of the program.

Technical Discussion Included in the year-2000 performance values is a one-point improvement in efficiency due to optimum matching of the stator exit and rotor inlet flow fields. This goal will require an extensive study of the effect of the rotor on the stator exit i flow field. To date, a limited amount of work has been accom- plished using a large-scale, low-speed test rig. The traditional probes that measure angle, pressure, and temperature cannot be used with the rotor in place. Nonintrusive instrumentation, such as an 52F system, will be required to map the interblade row flow field for the engine scale rigs.

' After the interblade row flow field is understood, the results will be incorporated analytically. An accurate predic- tion of the 3-D stator exit flow field incorporating the stator viscous effects, interblade row endwall viscous effects, and the rotor influence of the flow field is needed to design the rotor.

and commuter engines, high rotor shroud divergence, and high stator turning will be required.

The turbine design, as conceptualized, assumes ceramic inserted blades or a composite ceramic integral rotor. Either

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should eliminate or minimize the broach-angle constraints required for metallic inserted blades. This will permit

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increased blade stagger angles and higher reaction, which allows the higher stage performance required for the high-work, single-

I

stage HP turbine of the simple cycle rotorcraft engine and the first-stage blades of the LP turbines of both the rotorcraft and

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commuter engines.

The high-work, single-stage HP turbine required by the I_ [ rotorcraft engine incorporates a vane with high turning to achieve the performance needed for the year-2000 engine perfor- mance. To meet this performance goal, the limitations on the amount of turning that can efficiently be accomplished must be understood. The influence of the aspect ratio and the trailing edge blockage on the ability of the stator to meet the high training req,lirements must .iso be known.

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_'%_I The- prog ram will consist of explor ing sta tOE per for _ance and

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!_i_ ] turning for stators designed for a range of exit flow angles from 77 to 81 degrees. A matrix of seven configurations will be rig tested, with variances in turning, trailing edge blockage, and aspect ratio from a baseline design. The results of this study i

will be incorporated into the design system (Program L) to aid [i

J / configuration trade-off studies for the rotorcraft turbines.

Proqram Description This program consists of two tasks, from airfoil design through rig tests, as shown in Figure 45. The program is sche- I.

duled for three years.

!: The program consists of testing seven configurations of a high turning vane. The first year will consist of designing the seven configurations, which will include a baseline vane and two variations each of turning, trailing edge blockage, and aspect ratio. The exit angles will range from 77 to 81 degrees. Thus, the matrix to be tested will include three variations each of exit angles, trailing edge blockage, and aspect ratio.

Testing of the seven configurations will be conducted in the

J

third year. A report on the results of the rig testing will be issued at the end of the third year.

Technical Discussion High-work, single-stage HP turbines required by rotorcraft engines use stators with high turning angles to achieve the work output. Industry has shown that turning of 79 to 80 degrees can be satisfied with nominal trailing edge blockage levels and high corrected flow stators. For the rotorcraft HP turbine to be suc- cessful, the turning technology must be verified in a lower

|

flow regime. With the lower corrected flows (smaller machines), the aspect ratio decreases and the trailing edge blockage increases due to size limitations. The increased blockage and decreased aspect ratio influence on stator performance at high turning must be evaluated and understood.

The program consists of designing seven stator configura- tions. This includes three stators at different exit angles at some nominal trailing-edge blockage. The stators will be typical of the size, aspect ratio, and solidity required for the rotor- craft engine. Using one of these stators as a baseline, two additional configurations each will be designed around the base- line stator with different trailing edge blockage levels and aspect ratios. The stators will be designed using the latest technology in 3-D design, which will include tangential lean and endwall contouring.

Following hardware procurement, each annular stator cascade will be tested. Testing will include a stator exit survey at various pressure ratios. Stator performance and turning will be evaluated for all seven stators. Particular attention will be given to the radial distributions as well as to the global results. Meeting turning and performance requirements globally would not necessarily achieve the optimum stage performance if ..

the radial distributions are skewed.

N. Turbine Tip Clearance Sensitivity Reduction

' I

To allow close-coupling the HP and LP turbines on the rotor- craft and commuter engines, high shroud divergence will be required for the LP turbine rotors. To achieve successful designs of this type, mechanical technology advances are required to reduce axial excursions during operating point changes.

Reductions in axial excursions will allow closer running clear- , i |

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ances and will thus reduce performance penalties due to tip clearances. Furthermore, aerodynamic advances in the reduction of performance sensitivity to tip clearance are required to achieve the year-2000 performance levels predicted for the rotor- h i craft, commuter, and cruise missile engines. Year-2000 efficien- / cies reflect a 25 percent decrease in tip clearance sensitivity.

/ / • The main thrust of this program is the continuation of GTEC stu- dies already completed on casing treatment and rotor tip treat-

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ment.

The GTEC/USAF Case Treatment Study has shown that the smooth shroud has performance advantages over recessed shrouds. A goal of the proposed program is to extend this study using a high-work turbine as a testing vehicle. The high-work turbine will be used

I

as a model for extending this technology into the rotorcraft and commuter engines, which will require high-work HP turbines.

r ° The technical approach will extend the study on nonconven- tional tip geometry, such as winglets, into the high-work turbine technology. The GTEC/USAF LART program has shown significant performance advantages using winglets. A further aspect of the I "-4 program will be to study the effects of aerodynamic loading on tip leakage. One approach will tailor the vector diagram so that

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the tip section is unloaded; another approach will vary the chordwise position of maximum loading.

T Although the test rig will be a high-work, single-stage tur- | bine, the results of the technology program wou_3 extend over all the axial turbine configurations of the rotorcraft, commuter, and turbojet engines.

f t• Program Description This program, which consists of three tasks, provides for screening tests and a second design iteration, as shown in Figure 45. It is scheduled to maximize the advantages of companion Programs L and M.

/' / I | The program considers two areas for reducing tip clearance sensitivity, casing treatment, and rotor tip geometry. The rotor tip geometry will include nonconventional geometries such as winglets and experiments in blade loading.

The first year of the program will consist of defining the shroud and tip geometry to be tested, and designing the test con- figurations. The latter portion of the first year and the early portion of the second year will involve fabrication of the con- figurations followed by rig testing and analysis.

.k The final two years will follow the same procedure, super- imposing the results of the two study areas into a conceptual design. A final report on the success of this effort will be submitted early in the sixth year.

Technical Discussion Preliminary studies have demonstrated several approaches to reducing the stage performance sensitivity to tip clearance. Two approaches relate to geometry changes on either the shroud casing J or the rotor blade tip. Past studies at GTEC have shown perfor- mance advantages for smooth shrouds and winglets on the tip sec- tions. Continuation of the GTEC/USAF Case Treatment Study and the GTEC nonconventional tip geometry studies will be the main thrust of this program.

p r, ... _. ............ .......

_""._ Another approach will include a tip section aerodynamic

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i " loading study. In this approach, the vector diagram will be _:,i I tailored such that the tip section would be unloaded aerodynamic- | ,. ally, thus reducing the pressure g_adient across the airfoil. _.

The position of the maximum aerodynamic loading influence on tip / T performance will also be studied• !| / In summary, reducing the performance sensitivity to tip 1 clearance by 25 percent will require studies on shroud and tip geometry configurations, as well as on aerodynamic design philo- i sophies.

i

2.4.2.5 Combustor Performance This section presents the technology plan for improving annular combustor performance for year-2000 rotorcraft engines.

The plan consists of two discrete technology programs, identified as O and P in Figure 46.

The first program, which addresses pattern-factor for high heat release rate (HRR) combustors, will advance technologies

iJ

applicable to ceramic combustors. It uses metallic combustors for rig testing. The second program will build on this advanced technology base to obtain experimental data on a ceramic combus- . , tot, which includes the combustor liner and hot transition liners.

O. Low Pattern Factor Annular Combustors Significant performance requirement differences exist between the various SECT applications for annular combustors.

The rotorcraft and commuter applications require a low pattern factor and a durable, efficient, stable combustion system with altitude relight capability. The cruise missile engine applica- ( tion requires a low pattern factor, volume-limited, short-life t I1| PLAN YEAR LOWPATTERN-FACTOR ANNULAR COMBUSTORS • DESIGN PRIORITY ONPATTERN-FACTOR ANOPRESSURE LOSS. TESTMETALLIC COMBUSTOR AT LOW 6 TO ACHIEVE SECT HRR ANOMINIMIZEWALLCOOLING EFFECTS t • COMB ANDRIGOES/FAO • COMB MOOSANO HOT RIGTESTS P CERAMIC ANNULAR COMGUSTOR • OESIGN TO SECTCYCLE WITH PRIORITY ON PATTERN FACTOR • COMB ANDRI6 DES/FAD • COMB MOOSANDHOTRIGTESTS(2OOOF) SECTTECHNOLOGY VERIFICATION i F.

L

l g- Figure 46. Combustor Technology Schedule.

_ ] 129 i i + i ii i I_.

combustor to satisfy requirements for a low-cost, low-volume,

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low-weight engine. All SECT applications require nonmetallic materials to optimize combustion system perfor- D,.-I ,:: t: mance for the high cycle temperatures of this study. Until non- "s" metallic combustor materials are available, metallic combustors "/ can be rig tested at increased cycle temperature levels to eval-

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/ / uate concepts.

As combustor temperatures rise, increased design priority on low pattern factor combustion systems will prevail to limit tur-

I

bine inlet temperature hot spots. The additionll goals of mini- mum combustor pressure drop and minimum combustor size make pat-

tern factor reduction a difficult task. The use of nonmetallic L

materials will allow budgeting additional dilution air for pat-

tern factor control by eliminating wall cooling requirements. I

Design concepts to minimize pattern factor can be evaluated with metallic combustors by rig testing at low pressures and using i advanced wall cooling to minimize cooling air requirements.

Testing will be directly applicable to rotorcraft, commuter, and [.

APU reverse-flow combustion systems, and the analytical models will allow the extension of the results to through-flow combus- tors (cruise missile engines).

Program Description

!_

This program consists of two primary tasks, from combustor design through hot rig tests, as scheduled in Figure 46.

The schedule is based on the use of an existing high-temper- ature combustion rig with combustors adapted to existing plenums.

Existing analytical design tools will be used to define the ini- tial combustor/liner geometries. Following initial testing, an f iterative test-design modification procedure will be used on ana- lytically predicted improvements, and each test sequence will _p." .

f

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the degree of improvement obtained and will provide " :" data on where further improvement is required.

_, ! The main program activities are su_arized in the following: "_ o Define a fuel injection system to provide uniform, L ., well-atomlzed fuel distribution with coking resistance I suitab!e for the cycle temperatures of this study . (i_l_i_ 1427C [2600F] TRIT) _ o Define the combustion system to provide a well mixed, "" tern dilution zone to provide minimum pattern factor I I uniform, primary zone as well as an optimum mixing pat- _ o Rig test evaluate each configuration at selected test conditions to determine pattern factor, liner thermal _ _ gradients, and atomizer fuel passage wall temperature levels _ o Modify the configuration based on 3-D analysis and i I repeat tests, iterating until all design goals are /'_ " _c_ie_ed n n_final redesign and test cycle will be used i I o t e o - etallic design.

' Technical Approach Selection of design goals will include pattern-factor reduc- I tions of 35 percent (to 0.12 levels)for the rotorcraft simple cycle. Completely uniform wall temperatures are not achievable.

I However, hot-spot thermal gradients will be minimized via proper fuel injection and combustor aerodynamic design. The design goal I level for wall thermal gradients will be as required for pro- • jected ceramic material properties.

: _ 131

I

i '

p -- u Atomizer coking resistance is important for future rotor- I: craft and commuter applications with high combustor inlet temper- ature (particularly with recuperated cycles). The design goal will be to maintain fuel-passage wall temperatures at acceptable i, levels, currently estimated to be below 204C (400F). Effective I thermal insulation will be required to minimize stem heat trans- fer and adequate tip cooling, with airblast atomization.

!: These programs will provide much needed information on design goal feasibility for pattern-factor control, fuel atomizer coking resistance, and liner thermal gradient control. Analyti- cal design methods to optimize the aerodynamic design of nonme- tallic combustors will also be initiated. Pattern factor will be a design priority for the ceramic combustor, with a goal of 0.12.

I

i

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_k - r PLAN YEAR L I ANALYTICAL MOOELING AND COOE VERIFICATION FOR i ° AXIAL/CENTRIFUGAL COMPRESSORS • COMPUTER COOE FOR CALCULATION OF 3"0 TRANSONIC ANO VISCOUC FLOWTHROUGHBLADE/VANE ROWS. TO BE VERIFIED BY NONINTRU$1VE FLOW FIELO MEASUREMENTS !

/ r

_/////////_, i

• ANALYTICAL MODELING/COOE VERIFICATION • AXIAL AND CENTRIF AERO P, I6 DATA • COOE CALIBRATION

Id

• VERIF COMPR: DES/FA8 o • • VERIF COMPR: AERO RIO DATA ° PERFORMANCEOPTIMiZATiON OF IMPELLER/OIFFUSER/ X-GVER FOR A FIRST-STAGE CENTRIF • MEASURE EFFECTS OF SPEED. BLaDE/SPLITTER NUMBER. AND DIFFUSION RATIO IN TEST MATRIX TO OPTIMIZE DESIGN/TRADES • IMPELLERS 131: DES/FAB/AESO RIG TEST • IMPELLERS ANO DIFFUSERS: DES/FAG/AERO RiG TES'I • X-OVER: DES/FA8 • STAGE AERO RIG TEST PERFORMANCEOPTIMIZATION OF SECONO-STAGE CENTRIF BASEOON FIRST-STAGE EXiT CONOITIONS • MEASURE EFFECTS OF MODIFIED INDUCER BLAOE ANGLES SET TO ACCEPT INLET SWIRL AND THROUGH- FLOW DISTRIBUTIONS • IMPELLERS [31: DES/FAR/AERO RIG TEST • IMPELLERS AND DIFFUSERS: DES/FAG • STAGE AERO RIG TEST CENTRIFUGAL COMPRESSOR CASING TREATMENTS • MEASURE EFFECTS OF GROOVES.SLOTS ANO HOLES AS CONFIGURED FOR REDUCEDTIP LOSSES AND AOOEDSURGE MARGIN • STAGE DES/FAB/AERO RIG TEST IMPROVED DIFFUSION SYSTEMS • MEASURE DIFFUSER CONFIGURATIONS SELECTED FOR LOWER INLET LOSSES, IMPROVED SURGE MARGIN, AND INCREASED DIFFUSION RATES, ANO EVALUATE SHORTENED90 OEGREE BENDS r • DIFFUSER DES/FAB _t I I • FLOW RIG TESTS • STAGE TESTS !_1 SECT TECHNOLOGY VERIFICATION

_

d., .1D "I Technology Schedule.

Figure 47. Compressor _'_ 134

i1

Technology Programs R through U could be run independently to the SECT cycle. However, improved data validity can be expec- ted with the interrelationships shown on the plan scheule.

I

Analytical Modeling and Code Verification for Axial/Centrifugal Compressors To achieve the centrifugal compressor performance levels projected for the year-2000 time frame, compressor aerodynamic design technology must be improved in several areas. This is

I

required in advance of full-stage compressor development in order to thoroughly explore compressor dedign options and ensure that

I

proper design--parameter selections can be made.

The objective of this program is to develop 3-D viscous analytical codes as a design tool to significantly improve tur- bomachinery compressor performance. The obvious benefit of designing blade rows with improved predictions and reduced losses is the immediate application in all types of engine applications.

|

The program is structured toward prediction and verification of internal flow calculations that have been substantiated by inter- .

nal flow measurements so that the code can be used with confi- dence.

This ongoing program is the key to reducing compressor blade row losses, with reduction of as much as 60 percent anticipated - | for year-2000 applications.

,,.,,.: Program Description _I{'" i. TM This program consists of five major tasks, from analytical _J modeling through aerodynamic rig testing of a verification com- pressor, as shown in Figure 47. The analytical activity for this program will be the development of a computer code to calculate ,,..,__ • | • !

j t e ,, three-dimenslonal transonic and viscous flow through a turboma- chinery blade/vane row. Concurrently, a test technique for non- intrusive flow-field measurements will be refined to aid in the f' t I _

J

modeling development and checkout of the analytical prediction Z code.

i

Technical Approach 3-D Viscous Modelinq - The development of a 3-D viscous, transonic flow analysis code as an integral part of a design sys- tem requires an improved physical understanding of all relevant flow phenomena. The viscous flow calculation will model losses, work (rotating blade rows), a,d flow turning as well as quantita- t_ tively predict wake formation. The analysis will also include the effect of tip leakage flow through the clearance gap of a rotating blade row and a method for treating separated flow regions.

Three-dimensional viscous flow calculati._ns are currently under development by several researchers. The extensive amount of literature available will be reviewed to evaluate the best approach to creating this code.

T Development of the code will require calibration with test data, which is currently limited in open literature. Special emphasis will be placed on the development of a turbulent visco- sity model for calculating shear stresses near walls and a tip leakage model for inclusion of this effect on flow development and losses in unshrouded rotors.

The code will be calibrated using the data acquired from the nonintrusive flow-field measurements. When a suitable number of comparisons are made, the test compressors will be redesigned using the 3-D code. The revised (verification) compressors will J F ..

" ,i I .....

,I.," then be tested and compared to predictions, with the code models A • altered as required based On these comparisons.

%- Nonintrusive Flow-Field Measurements - Laser velocimeters provide a unique solution for measurements within the flow field of a rotating blade row. The most commonly used laser system is / the laser Doppler velocimeter (LDV), but this system does not lend itself well to blade channels with small blade heights. A second system, the laser two-focus (L2F), is less commonly used but has several advantages over the LDV for the measurements required within small turbomachinery blade channels.

The L2F system (Figure 48) measures the time of flight bet- ween the two parallel light beams. Particles are seeded into the flow and follow very closely the mean flow direction. Due to the highly concentrated parallel light beams (approximately i00 times

I

greater than in the LDV), the backscattered light from each particle gives a significantly higher signal-to-noise ratio.

ensure high performance potential.

'\, A diffusion system and crossover (deswirl) ducting will be designed based on the latest available technology. These areas

I I

/ will be tested using the best of the three impellers to establish

f.1

J i first-stage performance capability.

ii

Technical Approach

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A preliminary design study will establish the exact range of the parameters to stay within reasonable limits on stress levels and practical fabrication methods as projected to the year 2000.

The compressor will be sized and configured to allow use of

i!

one of several existing basic compressor rigs. The compressor 7_ configuration will be based on GTEC-proven design techniques based on a recent state-of-the-art compressor.

To ensure high performance potential, the detail blading design will employ the most advanced, full 3-D methods available.

U

Test instrumentation and the technique used will be the most advanced and definitive available.

S, Performance Optimization of Second-Stage Centrifugal Compressor Based on First-Stage Exit Conditions .

The objective of this program is to establish optimum !

second-stage design parameters in view of the flow conditions exiting the crossover duct of the first stage of a two-stage cen- trifugal compressor. This technology is required in advance of full-stage compressor development in order to thoroughly explore T compressor design options and ensure that proper design-parameter selections can be made.

will be written.

Technical Discussion

I

The typical flow conditions at the inlet to the second stage

+I

of two-stage centrifugal compressors have skewed profiles in both pressure ratio and angle. The present program is directed toward

-I

parametric investigation of design techniques to improve second- stage performance based on these skewed profiles. The impeller "L d J

[l

designs will be modified to incorporate typical measured inlet gradients in pressure and swirl. Test results will be used to modify downstream diffusion systems to match any new impeller

exit profiles as a result of the changes in impeller design. !i

Subsequent tests will be made to verify the improvement in stage performance for this technology item.

/ ! T. Centrifugal Compressor Casing Treatments .d The objective of this program is to reduce the tlp-clearance sensitivity and improve the surge margin of centrifugal compres- sors. A reasonable amount of casing treatment study has been done for large axial compressors. A much smaller technology base is available for small centrifugal compressor stages. The bene- fit of this program will be to ascertain the impact of passive shroud treatments on surge margin and tip clearance of small com- pressor stages. These efforts are needed in advance of full- .i stage compressor development in order to thoroughly explore com- pressor design options and ensure that proper design-parameter selections can be made. It is projected that a one-point improvement in stage efficiency can be achieved at constant su:ge margin.

Program Description This program is scheduled as one task, from stage design through aerodynamic rig, as shown in Figure 47. The program will °.

evaluate a series of centrifugal-compressor casing-treatment con- .o !

cepts including grooves, slots, and holes in the impeller shroud.

The position of these configurations will be varied along the shroud to determine their effect on tip clearance losses and .i surge margin. The proposed program will include design, fabrica- P, tion, and test of these concepts.

L _4 J Technical Approach The proposed program will examine the effect of the various shroud configurations on surge margin and tip clearance losses.

An existing rig that has a clearance control spindle capability will be used to evaluate the clearance effect. The shroud con- figurations will include some that are positioned in the knee and trailing-edge region for this clearance study.

Due to the strong interaction of the stages at part speed, the surge margin investigation will be done on a two-stage rig.

In the two-stage configuration, the first stage operates to the left of its normal surge characteristic at part speed. There- fore, the casing treatment configurations have their most influ- ence in this situation.

O. Improved Diffusion Systems An extensive program is proposed for reducing diffusion sys- tem losses as well as maintaining existing performance in smaller envelope constraints. The benefits derived from this program are a higher-efficiency centrifugal stage and/or lower weight and reduced cost stages.

Proqram Description This program, set up to establish technology for improved diffusion systems, consists of three major tasks, from diffuser design through stage tests, as shown in Figure 47. The program is directed at initial flow-rig testing to examine diffuser con- cepts followed by stage testing of the most promising configura- tions. Diffuser configurations will include those believed to lower inlet losses as well as improve surge margin, plus dif- fusers that increase the diffusion rate downstream of the throat

I

!

.i of the diffusers. Aggressive 90-degree bends will also be evalu- ated to lower the compressor envelope.

These configurations will be tested on a static-flow rig to be developed as part of the program. This flow rig will have the capability to simulate Mach and Reynold's numbers and to vary inlet total-pressure profiles and swirl distribution. Each of the 3-D diffuser designs will be fabricated and tested on both the static-flow rig and an existing in-house single-stage com- pressor rig. This dual evaluation will establish the validity of the static rig testing for subsequent diffuser effort. Dur- ing the test sequence the analysis methods will be reviewed for possible improvement, and the test configuration will be reana- lyzed as necessary.

_° At the completion of this analysis, the most promising con- cepts will be combined into a single-stage test configuration to demonstrate the benefit of this technology.

Technical Approach Separate static-flow rig and full-stage tests will be con- ducted to sort potential diffuser configurations into two areas.

The first application is to reduce losses in the inlet region of the diffuser. The second application is to maintain existing performance levels, but to reduce stage weight and cost by reduc- oo ing compressor envelope diameter.

The first method involves determining the effects of swept leading-edge and nonsymmetric endwall configurations on compres- sor performance. In addition, the effects of endwall circumfer- .o ential communication slots and diffuser bleed will be examined in terms of surge margin and rematch capability.

• SPECIMENTESTS. DESIGNDATA _ • EXPERIMENTAL PARTS DES/FAR/BENCHTEST _ _.

• ENGINEENVIRONMENT TEST "/_,_

i

SECT TECHNOLOGY VERIFICATION °_ 65-291,.50

I

Proqram Description

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This program consists of five major tasks, from casting ex- - _. [ perimentation through engine environment tests, as scheduled in Figure 49. The second-stage compressor impeller of an existing J[ 1-je %: J turbofan engine has been selected as the component to verify this technology. The impeller will be mechanically redesigned to o I" l',i f reduce its weight, using the higher strengths attainable in CST- processed Ti-6246. Ti-6246, having the highest strength capabil- ity among the alpha/beta family of titanium alloys, is the mate- i' g t rial selected for casting the compressor impeller.

Conventional investment casting techniques, as currently practiced by the industry, will be used in producing the net- shape-compressor impeller. This involves fabricating hard tool- .o ing to generate the impeller wax pattern, casting the impeller, and hot isostatic pressing (HIP) to fully densify the titanium casting. Casting iterations and tool rework will be required to establish dimensional reproducibility of the component as well as to optimize yield.

° Concurrent with the casting process optimization, heat treat process development will be conOucted. Its objective will be to achieve the best balance of mechanical properties in a cast _8 Ti-6246 titanium impeller. Results of these programs will lead to the generation of a design data base. Component evaluation, followed ultimately by engine testing, will provide validation of ° .

the technology.

. ° W. 454C (850F) Aluminum Alloy and Coating The advanced gas turbine engines at GTEC for propulsion sys- tems and auxiliary power units (APU) for both military and com- mercial aircraft, demand advanced alloy systems. Lightweight alloy systems possessing high-strength, high-temperature capabil- ities form the material basis for these modern gas turbLne engines. At GTEC, aluminum alloys are used for applications below approximately 182C (360F), with titanium alloys used for applications from 182 to 482C (360 to 900F).

#

!

The emergence of rapid solidification processing (RSP) tech-

l

nologies has made possible the development of advanced metal alloy systems with significantly improved basic material proper- ties. RSP powder metal (PM) aluminum alloys specifically made for elevated temperature service offer significant payoffs for gas turbine engines. These advanced aluminum alloys offer an enhancement of high-temperature strength via a large volume frac-

k tion of thermally stable, dispersed, intermetallic phases. These

aluminum alloys demonstrate the potential for replacing titanium alloys in 182 to 343C (360 to 650F) applications, with a 35 per- cent weight savings. This weight change results in a decrease in inertial momentum and a more responsive engine. In addition, aluminum alloys are less costly than titanium alloys.

Advanced aluminum alloys that show great potential for high- temperature service are those based on the binary AI-Fe system,

,[

with ternary additions (Cr, Nb, Hf, Mo, Ti, Si, W, Zr, V, and rare earth elements) used to increase the thermal stability of

I

the dispersed iron aluminide prec%pitates. Two PM aluminum alloy L systems that have demonstrated superior mechanical properties up i to 343C (650F) are the PM atomized AI-Fe-Mo (Pratt & Whitney) and Ai-Fe-Ce (Alcoa) alloy systems. Two other alloy systems that have demonstrated high-temperature capabilities are the AI-Fe-V

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and the Ai-Fe-Zr alloys produced by the planar flow casting (PFC) proprietary technique of Allied Chemical (based upon the melt-

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spinning principle). These metallic ribbon alloys have been shown to contain higher volume fractions of dispersed interme- tallic phases than PM atomized alloys, a difference attributed to orders of magnitude difference in the maximum cooling rate avail- able with the PFC technique (106 to 107 C/sec) as opposed to atomization techniques (104 to 105 C/sec).

A major drawback of these aluminum alloys is that they are more susceptible to particulate erosion damage than titanium o

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J alloys. Small gas turbine engines in helicopters, ground power units, and other low-altltude al)plications (particularly over dusty, unimproved land areas) routinely ingest considerable . I amounts of sand and dust. Engine components operating directly in this erosive airflow consequently suffer progressive degrada- "i tion in both physical and performance characteristics. There- fore, erosion protection for critical airfoil surfaces is needed to utilize the benefits of advanced aluminum alloys in these engine applications.

Program Description This propram consists of five major tasks, from materials experimentation through engine environment testing, as scheduled _a on Figure 49. Aluminum alloys produced by the Allied PFC tech- nique represent a great potential for the development of 454C (850F) engineering material• Metallic ribbon alloys (which are ° , later crushed to a powder form and consolidated into a billet) made by PFC have a larger volume fraction of dispersed phase than does PM-atomized aluminum alloy. This results in enhanced mate- rial strength properties. Research by Allied indicates that a 35 to 40 percent volume fraction of the dispersed intermetallic phases are required to meet the Air Force's 277 MPa (40 ksi) ultimate strength goal for a 454C (850F) aluminum alloy• Im- proved metallic ribbon alloy systems are the selected approach to achieve this goal.

The resulting operating speed limitations place restrictions on the design configuration and performance of multispool turbine engines. Operation below the first bending mode is achieved by restricting rotor speeds, minimizing shaft lengths, and using large shaft diameters (especially for the low-pressure rotor).

However, reduced LP rotor speeds reduce the maximum attainable _o turbine performance, while a large LP shaft diameter imposes larger bearing diameters and disk bores on the high-pressure (HP) rotor than are usually desirable. These size constraints impact the HP spool design by imposing higher bearing loads and greater rotor system size and weight for the required disk burst margins and low-cycle-fatigue lives.

Use of an advanced shaft material with a high ratio of elas- "i tic-modulus-to-density (E/p) will increase the shaft critical speeds and minimize the constraints as previously discussed.

This critical speed advantage can be employed in one or more of the following ways: The rotational speed of the shaft can be increased The shaft diameters can be decreased . , The unsupported length of the shafts can be increased .J t::

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ql , 'j .... j_; i.:.X:-, .: J Increased shaft speeds can facilitate improved performance via lower airfoil loadings or reduced stage counts (for a given cri- tical speed margin). Reduced shaft diameters or longer unsup- ported shaft lengths can provide reduced mechanical complexity of the shaft bearing package and reduced engine weight. Reduced diameter shafts will be required to acco_nodate small disk bores (for bore stresses) for the high rotational speeds projected for year-2000 engines.

Silicon carbide fiber reinforced titanium is an advanced metal matrix composite material offering high E/D values for shafting applications in conjunction with high torsional strength cipability and demonstrated fabrlcability. This material system exhibits E/0 values two to three times higher than those of con-

l

ventional shafting materials such as steel- at_ nickel-base alloys. In addition, titanium/silicon carbide turbine engine components have been fabricated at GTEC and elsewhere in the

I

industry using established diffusion-bonding techniques.

Prior metal matrix composite shaft efforts have stressed retrofitting advanced shaft designs into existing engines, which necessarily sacrifices overall system benefits. The proposed program emphasizes the design of the rotor/shaft system with the

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power turbine rotors of SECT rotorcraft and commuter engines as prime candidates.

Proqram Description L_ This program consists of four major tasks, from materials I.

experimentation through shaft rig tests, as scheduled on Figure % 49. The program will be conducted on shafting and rotating groups sized to be representative of the SECT rotorcraft engine.

The program will use Avco SCS-6 silicon carbide fiber in a Ti- 6AI-4V matrix, with the fiber volume fraction and orientation i, J _qm

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dependent upon the outcome of a preliminary rotor design study.

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After selection of the basic shaft manufacturing approach, a con- cept validation study will be conducted during which subscale shafts will be manufactured using the fiber loading and diffu- sion-bondlng parameters required for full-scale shafts. These shafts will be subjected to torsion, bending, critical frequency,

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and nondestructive evaluations to validate the basic design and fabrication approaches selected for the full-scale shafts.

Mechanical property design data will then be applied to panels fabricated using the fiber loading and processing parameters

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selected for the full-scale shafts, with a final shaft design iteration conducted based on the design data. A pilot manufac- ture delonstration involving 30 full-scale shafts, five of which will be fully machined for shaft rig dynamic testing, will follow. The remaining 25 shafts will be subjected to extensive i bench and component tests to assess the consistency of the manufacturing process and methods for process scale-up.

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_8 Y. Polymeric Composite Gearbox The development of polymeric composite gearboxes will result in a weight reduction of about 25 percent, as compared to cast _8 aluminum gearboxes, and I0 percent compared to cast magnesium gearboxes. It will also eliminate the corrosion problems fre- quently encountered in magnesium gearboxes. In addition to the weight reduction, an initial cost reduction of about 25 percent is foreseen when the composite gearbox is compared with a cast aluminum or magnesium part. The composite gearboxes will be generically applicable to accessory gearboxes for rotorcraft and commuter engines, for APUs, and for the reduction gearboxes required in co,unuter turboshaft engines.

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PL_.NYEAR "

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112 3 4 5 9 7 e g _01I!121314 NONCONTACT FACESEALS !

• ACCOMMODATE HI6HER SHAFTSPEEDSWITH LOW AIR-OILLEAKAGE RATES FOR HIOH PRESSURE/ TEMPERATURE SECT ENGINES • ANALYTICALMOOELING • SEAL OES/FAB • SEALBENCH/GIGTESTS • ENGINEENVIRONMENT TEST' AA 9RUSH SEALS • COMPACTSEALSFOR LOW AIR.AIRLEAKAGERATES ] FOR HIGH PRESSURE/TEMPERATURE SECT ENGINES • ANALYTICAL MODELING • SEAL 9ES/FAB • SEAL BENCH/RiGTESTS • ENGINEENVIRONMENT TEST ADVANCED HIGH.TEMPERATURE LUBRICANTS AB

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_a • INCREASE MAXOIL TEMPERATURES FROM140-232C (300-450FJ TO 310427C (OOO-800FJ TO ACCOMMOOATE HIGHRPM/TEMPERATURE SECTENGINES

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• SPECIFYREQUIREMENTS/SURVEY PRODUCTS • VENDOR FORMULATION OF NEW LUBRICANTS • ENGINEENVIRONMENTTESTING i SECT TECHNOLOGY VERIFICATION q 68-211142 Figure 50. Engine Systems Technology Schedule.

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_._ SELF.ACTING

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GTEC ',. - imoj PROOUCTION EXPERIENCE :BE _WET-FACE _: 89 II001 r'ii/fl,,- //-/."//,I/i////_ SEAL • • n • i • u • • i 0 O 61.0 121,9 182,9 12OO1 14001 16OO1 SURFACESPEED. M/SEC {FT/SEC) G5-261-1

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Figure 51. GTEC Experience and Projections for Seal Capabilities.

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Labyrinth seals, although not imposing the surface speed limitations of conventional face seals, do have other design limitations. These seals result in greater leakage and, conse- o, quently, higher heat input and increased debris contamination of

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the bearing cavities.

Floating ring seals have operating limits similar to those of face seals. However, a major drawback is that leakage from this type of seal, as it wears, approaches that of a labyrinth seal. This is especially true at higher speed, pressure, temper- i ature, and excursion conditions.

The noncontact, self-acting gas film seal concept offers the : I best potential to meet the sealing requirements of advanced gas turbines. Figure 52 shows the details of a Rayleigh step-pad noncontact seal geometry. This is one example of many lift-pad geometries that can be used for noncontact seal operation.

The effect of self-acting lift pads is depicted in Figure 53, which shows parallel sealing faces operating without rubbing contact because of a balance between the opening and closing forces. If the seal tends to close, the lift pad hydrodynamic .i force increases to prevent a rubbing contact. Thus, a condition of no rubbing contact will prevail except at start-up and shut- ° • down.

Program Description This program consists of four major tasks, from analytical modeling through engine environment testing, as scheduled on Fig- ure 50. The noncontact face seals program will entail extensive analytical and experimental efforts. Various aspects of the seal operational modes have been analytically modeled, and it is a o ° topic of considerable current interest. As in the case of most o .} k¢ v., *t GAS FILM PL PL SEAL OAM PL -" LOW PRESSURE PH PH : HIGH PRESSURE PH RAYLEIGH STEP.PAO SHROUDS G5-29t,.3

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Figure 52. Rayleigh Pad Seal Operation.

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_d NO CHANGEIN GAS PRESSUREFOR MOVEMENT OF PARALLEL FACES POSITION OtSPLACEOPOSITION f ,60 SPRING SEAL ROTOR / GAP OPENING FORCE

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CARBON .J ROTOR HYOROSTATIC LIFT PAD CLOSING FORCE • I 1, L--SEAL[O GAS PRESSURE HYOROOYNAMICFORCEOF LIFT PAOS FOR INITIAL POSiTiON .o "_"NYOROOYNAMIC FORCE OF LIFT PAOS FOR OISPLACEO POSITION G_2814 Figure 53. Mechanical, Pneumatic, and Self-Acting Forces on a Rayleigh Pad Seal.

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analyses, however, only special cases have been considered, and many modes of operation observed in practice remain untreated by analysis. The noncontact face seals will remain principally an

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& experimental effort, with the analysis as an aid to understanding I the physical princiDles involved.

The method of fabrication of the intricate lift-pad geo- metries and the identification of suitable materials for these geometries are also significant problems to be resolved.

Extensive analysis will be performed on the lift-pad geom- etry and seal steady-state operating characteristics during the initial seal design phase. These analyses will include: i o Seal leak rates using compressible-flow analysis

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o Lifting pad hydrodynamic analysis o Pressure balance computations o Seal stability analysis

i

o Thermal and mechanical distortion analysis.

Technical Approach

i

This program will be initiated with the design and

I

fabrication of a seal test rig specifically designed to evaluate noncontact face seal operating parameters. To investigate the full operating range of noncontact face seals, the test rig must be capable of operation beyond 10,470 rad/s (100,000-rpm) speed, 206.9 N/cm 2 (300-psi) differential pressure, and 1200F sealed gas temperatures.

Current GTEC-sponsored R&D efforts will provide a technical basis for the initial lift pad geometry selection. These designs will then be tested to the full extent of the seal operating parameters. Performance deficiencies will be determined and the design will be iterated and retested.

Seal operational qualities to be determined in this advanced seal test rig Includez 0 Rubbing contact operation power loss o Rubbing contact seal distortion 4_ Qa o Carbon lift-off speed o Carbon touchdown speed o Noncontact power loss Noncontact seal distortion o °i o Seal dynamic modes o Seal leak rates o Seal wear rates o.

Appropriate fabrication techniques and proper material must o I be defined. The rotor surface must withstand contact seal oper- ation at low speeds and during many start/stop cycles. The

-I

requirement for a hard protective coating to satisfy these i requirements compounds the problem of machining intricate lift- o.

pad geometries into this surface. Some processes have already been attempted, with limited degrees of success.

i

An important consideration is the influence of tolerance on o_ lift-pad operation. Due to the small dimensions involved, large- percentage tolerances may be required to allow for the use of conventional machining processes. The final selection of lift- pad geometry may be determined by the particular design's opera- tional sensitivity to tolerance.

An important requirement for a stable, low-leakage face seal is that the mating surfaces remain plane during operation.

Extensive mechanical and thermal distortion analysis will be per- o° formed to ensure that this operating condition is satisfied.

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!

Seal performance testing to intermediate speed, differential

I

pressure, and gas temperature of 4188 rad/s (40,000 rpm), 103.4 N/cm 2 (150 psi), and 316C (600F), respectively, can be initiated

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on an existing GTEC seal test rig. A new high-speed test rig specifically designed to run noncontact seals to operating limits

'I

of 10,470 rad/s (100,000 rpm), 206.9 N/cm 2 (300 psi), and 649C (1200F) is currently under construction and will be available for this program.

The ongoing GTEC R&D efforts on noncontact seal effort for gas turbine mainshaft sealing applications ensure a solid technical base from which the designs for the extended operating conditions can be achieved. The availabiity of a suitable test rig that requires only minor modifications to accomodate a wide range of noncontact seal designs and operating experience on an R&D Raylelgh pad noncontact seal enhances the probability of a successful design with program cost effectiveness.

AA. Brush Seals The perfo:mance and reliability of high-speed gas turbine engines depend on the seals. Labyrinth-type seals have been used as the primary air-to-air seal in gas turbine enoines. However, these seals cennot meet the increased demands of performance and the required low leakage rates of the small gas turbine engines as conceptualized for the SECT Program.

Future gas turbine engines will require seals to perform at higher temperatures, pressures, and peripheral speeds. To meet the increased performance demands, lower leakage rates are also required. Even with the technology advancements in labyrinth seal land materials and coatings, the advanced operating require- ments limit growth potential. Labyrinth seals, unaffected by increased temperature.and periphera_ _peed, do have other design B* limitations. These seals, which operate with a running clear- ance, result in high leakage rates and potentially decrease the i engine performance. Radial excursions of the shaft typically lead to labyrinth seal run-in and higher leakage rates than is acceptable for future gas turbine engines. In addition, advanced gas turbines are required to be lightweight and compact.

Increasing the number of seal knives on a labyrinth seal ° .

incroases sea! performance, but it also increases seal axial length. To meet the requirements of advanced gas turbines, high density air-to-air restrictors for rotating shafts are required.

The brush seal concept offers the potential to meet the increased sealing requirements of gas turbine engines. Figure 54 shows a typical brush seal tested at GTEC. The lay of the bristles and the bristle stiffnesses permit a low starting torque . • requirement and reduced power consumption. The brush seal offers the reduced leakage rates that are required for advanced gas tur- bine engines. Besides offering excellent leakage characteristics -0 for air-to-air type applications, brush seals might also be con- sidered for air-to-oil bearing sump sealing if the initial debris generated from the bristles can be diverted around the bearings.

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Proqram Description This program consists of four major tasks, from analytical

!I

°., modeling through engine environment testing, as scheduled on Fig- ure 50. Development of brush seals will entail extensive ana- lytical modeling and testing. Some aspects of the seal operating characteristics have already b-en tested. From this testing, certain geometries have shown to be superior. Continued develop- ment of this seal geometry will be conducted, primarily with rig testing. Analytical tools will be used as an aid to understand- ing the sealing physics. The analytical tools will be used to evaluate seal leakage rates, power consumption, initial starting torque, and life predictions.

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LOW" PRESSURE PRESSURE "I HIGH" SlOE SIDE IPHI IPLI BHISTLES GS_ZBI-15

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Figure 54. Typical Brush Seal.

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• i ,o ,o A sizable effort will be devoted to the fabrication of the

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brush seal, Fabrication techniques are a large portion of the seal cost. Methods will be investigated to reduce fabrication _ ;b _ cost without impeding seal performance.

&-4 Technical ADDroach Current GTEC efforts have defined a base brush seal geom- !

_d etry. This design will be tested to define seal operating param- eters. Areas of improvement will be identified, modif{cations made, and the new design tested. Seal operating parameters to be determined by rig testing i_Lclude: o Air leakage rates relative to: Rotor speed Shaft eccentricity Bristle density Bristle-to-rotor interference Operating time O Power consumption Bristle wear rates O O Rotor coating wear rates • • Extended life characteristics O The seal will be tested over the entire operating conditions of temperature, pressure differential, and peripheral speed.

After extensive rig testing has verified the seal design, the seal will be evaluated in an appropriate engine test bed.

Attention will be given to the material selections for the brush seal bristles and the zotating journal. The journal sur- face must be capable of withstanding the bristle contact for the _L° I i_" _ associated speeds and temperatures. Wear of the the jour al sur- ,,, face will decrease sealing performance. The bristles must with- i | stand high temperatures and be compatible with the journal coat- I ing. Some materials have been identified, and initial GTEC test ! results are promising.

Advanced technology engines will operate _ith higher inter- nal pressures, temperatures, and shaft speeds (approximately 65 percent higher) than current designs. At the same time. the size more efficient. " These factors will place a higher demand on the _i I of these engines will become smaller as thc components are made _ I oil used to lubricate and cool the mechanical components in the factors: increased heat generation and sizing of lubrication sys- engine. This increased requirement on the oil comes from two • tam components.

I The higher temperatures, pressures, and speeds will increase the lubrication system heat load due to increased bearing, gear, I and seal losses. The conduction losses into the bearing cavities (in the turbine end) will be increased because of hotter secon- i dary cooling air, and because there will be less room for insula- tion on the cavity walls (due to reduced engine size). Moreover, churning losses will be increased because of smaller bearing I cavity size.

i I Because the heat load will increase, the lubrication system components (especially the oil cooler) must grow in size and

I weight in order to keep current oils at an acceptable temperature

i level. Currently, this temperature requirement is 135 to 149C I i I (275 to 300F). At some point, the size and weight of the lubri- cation system components become large enough to be considered i : unacceptable. 169 ..... "_'_'_-_ • •'%% L _ '_ ...... -h| ......

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There are some oils with temperature c/pabilities of 204 to 232C (400 to 450F), but they either degra/d'equickly or are very expensive and therefore not suited for general gas turbinp, use.

The solution to this problem will be advanced gas turbine lubricants that can operate at much higher temperatures oE 316 to 427C (600 to 800F). This will permit the lubrication components, -} and other engine components, to be smaller.

Q Program Description i d This program consists of three major tasks, from specifica- tion requirements through engine environment testing, as sched- uled in Figure 50. Creating a high-temperature lubricant for use in gas turbines is a project that will requi,e coordination with oil companies like Exxon, Mobil, and others that are heavily involved in the research and development of lubricants.

i Ql The program will be conducted in five primary activities: _t o Establish the detailed requirements for an advanced o ° p high-temperature lubricant. This will include foaming, viscosity, storage, compatibility with elastormers, • i temperature capability, and cost.

I

The rotorcraft mission analysis revealed significant reduc-

I

tions in fuel burn, of 21.9 and 41.6 percent, for the simple and recuperated cycles, respectively (Figure 55). The selection of

I

either a recuperated or a simple cycle is strongly dependent on fuel price. The recuperated and simple cycles have similar DOCs

I

at $1/gallon: approximately 7 percent lower than the reference engine. At the higher fuel price, the advantage clearly shifts toward recuperation (11.4 percent reduction versus 8.7 percent

:I

for the simple cycle).

These fuel burn and DOC reductions are strongly dependent on several key technologies, shown in Figure 55. These include advanced materials, component performance, and system tech- nologies.

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RECUPERATED

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SIMPLECYCLE 5_ 41.§

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$1/GAL KEY TECHHOtOGIES . ° • ADVANCEOMATERIALS (CERAMICS,NI3AI.

CERAMIC RECUPERATGRI • IMPROVEDCOMPONENT PERFORMANCE {TURBINE. COMBUSTOR.

COMPRESSOR AEROI • SYSTEM TECHNOLOGIES {METAL MATRIX SHAFTS.

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Rotorcraft Mission Analysis Results.

Figure 55.

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Fairchild Metro III, for which considerable data exists, as the ii..

current technology baseline.

via wing area/aircraft drag reductions as follcws:

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PRECEDL_(/ PAGE BLANK NOT FILMED

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o_ "i OIMENSIONS o 0 • 46-CM Ila-INCHi AISLE WlOTH ° • • 183.CM (72.1NCHJAISLE HEIGHT • 46-CM [IB-INCNI SEAT WlOTH I Q b • 38-CM 115-1NCHI FLOOR.TO.SEAT CUSHIONBOTTOM • 2.5 M3/M OVERHEAO STORAGE • 30-CM 112-1NCHIWlNOOWHEIGHT ..'

• 8.35-CM {2.5-1NCHICABIN WALL THICKNESS • 8.U-CM 12.75-1NCHI FLOORTHICKNESS . .

UNDERAISLE • 152-CM (fIG.INCH)EXTERNALWlOTH • 198.CM 178.1NCHJ EXTERNAL HEIGHT • SEATS MOVE 5-8 CM [2-3 INCHES)INTO AI=LE DURI;",G FLIGHT FOR GREATER COMFORT. SEAT RETURNEDTO POSITION o.o AGAINSTWALL FOR EASIERBOAROING • 2.5_ M2 127.60 FT21 FRONTALAREA "i °o ° • °.

65-281-64 o , for Reference Commuter.

56. Standup Cabin Figure ° .

. .

. ° O Reduction of aircraft weights through the use of com-

4 posite materials

Strut-mounting the engines/nacelles to eliminate the J interference drag at the wing-fuselage-nacelle channel and to avoid the ineffective area of the wing covered by the nacelle m ..

Allowing for improved surface finish due to use of com- posites With the baseline configuration established, the aerody- namics for the reference engine were then estimated. The meth- odology used for estimating aircraft drag was calibrated by gen- erating _:he known Metro Ill polar. The same methodology was then used to calculate the drag polar for the year-2000 technology commuter aircraft.

l

Figure 57 shows side and _ont views of the SECT reference aircraft and p_esents a data summary.

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3.1.2 Reference Mission For Commuter

J

A turboprop commuter mission typically consists of several stages with varying lengths. The operator/owner of a 19-passen-

i

ger commuter would typically have a different route than does the operator/owner of a larger or a smaller airliner. Therefore, a statistical examination was made of route analysis requests made

I

to a 19-passenger commuter airframe manufacturer. Results are shown in Figures 58 and 59. Current mission stage lengths range

I

primarily between 139 km (75 nm) and 278 km (150 nm). The mean route Segment is 272 km (147 nm); the most often recurring rcute

i

segment is 139 km {75 nm). Figure 58 shows that more than 50 percent of the takeoffs and landings occur below 305m (!,000 ft) altitude, and approximately 70 percent of them occur below 610m (2,000 ft).

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150.4 FT) ttt_ T_ 11.50 M _ !

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IS;.4 FTI I

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• 19 PASSENGERS • TWIN ENGINE 748 KW (1000 SHPI EACH • CRUISE L/O = 7 TQ 8

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• EMPTY WEIGHT - 3088 KS (8808 LBSI • OPERATING WEIGHT EMPTY (OWE] = 3,343 KG (7370 LBS) • PAYLOAO= 1637 KG (3610 LBSI • FUEL = 953 KG (2100 LBS) • TOGW = 5933 KG (13.080 LBS] • ° FEATURES • BONDED ALUMINUM LITHIUM

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ALLOYS -- FUSELAGE • COMPOSITES -- EMPENNAGE.

NACELLES. MAIN WING _20t_

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)" SECT Reference Com_nuter Aircraft.

Figure 57.

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;A ).78

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_t 114211 PERCENT OF ALL AIRFIELOS IiI_211.84 o

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Figure 59. Survey Results of Present Airfield Altitudes.

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Based on the statistical survey, a year-2000 reference com- muter mission was established as follo_,Jz 4J :. o Route Seqments A four-segment mission was selected , with the following stage lengthsl , Stage 1 2 3 4 Length, km 139 278 185 I]9 (nm) (75) (150) _,;,_' (75} 0 Takeoff Altitude Since mo_t takeoffs occur at low

J

-4 altitude, sea level was used for takeoff and landing.

T O Ambient Conditions ISA conditions were selected for the flight envelope.

Cruise Al_itude This was computed for the reference case and will be engine-dependent for the year-2000 engine commuter systems.

An analysis o_ optimum altitudes based on reference aircraft (paragraph 3.1.1) and the reference engine

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(paragraph 3.1.3) shows a dependence on block-time parameters. Since block time is controlled mainly by

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the time to climb to a cruise altitude and maximum #be cruise speed, the optimum altitude is selected based on the trade between maximum cruise speed versus altitude and time to climb. The maximum cruise speed varies d t, only slightly with altitude _or this case. Thu:3, the gain in maximum cruise speed possible at higher alti- tudes does not o_set the time to climb to higher alti- tudes. The optimum altitudes for minimum operating cost _or the stage lengths studied, therefore, was found to be at the altitude where VMo (maximum operating L ....

speed) intersects maximum cruise speed (approxi- mately 2438m [8000 ft]).

o Cruis_ and Climb Speeds Because time is normally the most expens_v- parameter in a route analysis, the cruise and climb speeds are set by engine max cruise and max climb power settings.

o Ground Handllnq A constant fiv_ minutes at Idle for taxi on the ground was used between each route segment.

o Fuel Ta,_kac_ The fuel tankage on the conjurer aircraft design was sized at a maximum to allow the largest num- ber of route segments before refueling is required.

o Fuel Reserves The flight fuel reserves were selected to be 45 minutes at maximum cruise speed at the cruise altitude for the last 139 km (75 nm) route segment.

Since the shortest route segment is flown at the lowest altitude, this segment will result in the maximum reserve fuel.

o Payload The missions were flown with a full 19-pas- senger load. A passenger weight of 86.2 kg (190 ib_ (passenger plus baggage) was used.

o Approach One minute at flight idle was used for the approach leg.

o Refueling Time Penalty If the weight and/or SFC of an engine did not allow the aircraft to fly all route seg- ments without refueling, a 15-minute refueling time

t

penalty was scored against the engine/aircraft block time.

_ "_ t_gure 60 depicts the reference mission for the commuter and F summarizes primary data. _ r t 3.1.3 Reference Enqine rot Commuter !

The reference engine for the turboprop commuter application uses the same engine core technology as derived for the rotor- craft reference engine (paragraph 2.1.3); as such, it is based on component performance that represents 1985 engine de_nonstrated ievels.

Confiquration - The reference engine has the same general con- figuration as the rotorcraft reference engine and is shown in Figure 61. Differences, as shown, consist of deletion of the IPS and the addition of an offset propeller gearbox and an annular segment inlet.

Materials - The materials used in the SECT reference commuter engine are identical to those listed for the reference rotorcraft engine as presented in Table 4. Additionally, the commuter engine requires a propeller gearbox with an aluminum housing.

Performance - The commuter reference engine performance is based on the _,me component performance levels as the rotorcraft refer- ence engine. The engine is sized for 746 kW (1000 shp) at take- off and is matched for optimum performance at a cruise flight condition as summarized below: O Engine Sizing Point - ISA, SL, static condition at L takeoff (T/O) power setting (TRZT = 1149 C (2100F), kW = 746 (shp = i000) J + i + +,..1__ _ f ,L

• MAX CRUISEPOWER

• CRUISE ALTITUOE =

2438 M (8000 FT)

CITY A

• TOTALDISTANCE =

z'l_ CITY S 741 KM 1400 NM)

• FUELCAPACITY =

CITY C 953 KG 12100 LBS)

• BLOCKFUEL=

'1

CITY O

817 KG (1801 LBS)"

• BLOCK TIME =

CITY E

;i

2 HOURS. 37 MINUTES

i STAGE "INCLUDES45 MINUTE/RESERVE OISTANCES.

KILOMETERS ORIGIN INAUTICAL PASSENGERS CRUISE DE3TINATION CARRIEO SPEEO MILESI A-R 13O 175i 19 470 12541 O-C 10 278 115O1 47Z 12'581

[

'1 19

C-O 185 I1001 472 I_I l, O-E 19 139 1751 472 12'8,81 i15-_II.I04

i

I

Figure 60. Commuter Reference Mission.

,I

!

HK:O_.'sv_,aiL'_tlkr% '_1'Ik • ':.-*'_45 _Ii"D_,,- ' Q o" 0 e • - 1 In[ _. ° , oO P , P

i

L REVERSE.FLOW PROPELLER ANNULAR OFFSET 1 GEARBOX l ANNULAR SEGMENT--I TWO.STAGE AXIAL INLET J COMBUSTOR POWER TURBINE oil I

/

oil _4 TWO.STAOEAXIAL C_NTRIFUGAL NP TURRiNE HP _OMPRESSOR |MAXi • 1149CI2100F TWO.STAGE l PR • 13.5 _o _o 1lS-Zll I.,14 -.

.a iF u!

Figure 61.

Commuter Reference Engine Configuration.

ZY

t

o Engine Design Point - ISA, 5486m (18,000 ft), 569 km/hr (307 knots) TAS condition at _aximum cruise power set- tinq (TRIT = I093C [2000F]) In sizing the engine, an iteration between the off-design

I i

sizing point and the selected design point was necessary. A_ a result of the of_-design operation at the sizing point, the com- pressor inlet flow of the commuter reference engine is slightly greater than that of the rotorcraft reference engine 3.2 kg/s (7.1 Ib/sec) versus 3.0 kg/s (6.7 lo/sec). The resulting turbo- prop engine and component performance data for these cases are summarized in Tables 14 and 15.

I The commuter reference engine achieves 521 kW (699 shp) (2 percent propeller gearbox loss assumed) at the design point. For this case the TRIT has been defined as 56C (100F) below the 1149C (2100F) takeoff power setting. Design point SFC for the refer- ence engine is 0.261 (kg/hr)/kW, (0.430 [(Ib/hr)/hpJ).

Weight - The commuter reference engine weight has been estimated at 191 kg (422 ib), including accessories and propeller gearbox, as shown in Table 16. This weight was generated by the WATE-S program, which included FI09/TSEI09 measured and calculated weights for accessories and controls and typical measured propel- ler gearbox weights from current turboprop en0ines. Length and diameter were estimated at 42.1 and 15.4 inches, respectively.

Cost - As with the rotorcraft reference engine, cost was esti- mated based on a mean sell price as depicted on Figure 62A. The sell price range is based on a GTEC market survey of presently available turboprop engines in the 447 to 895 kW (600 to 1200 shp) range. The resulting commuter reference engine cost is D $195,000 (1985 dollars).

L

i m (:ONNUTIIII UIPIIIUINCII |NGZli8 DII||QN IIOZNT DATA '_ TABL| 14.

Design PoLnt Perforaanae - S,4EU (1O,_00 fL), Seg kL/hc |30? knots) TAJ, ||A, UnlnsteLLed With No Pcoduat|oe Nerglne Overell IngLne Perfornlnce COnl_nent Perforxnnce e9 BP ¢oa_rsssor |n_lne Xetlng, kW (mhp) 511 (19t) Let |rigs -t SiC, (kg/htl/kw, (lb/hr|/hp 0.2|1 (0.¢]0) o PX 4. */] o _AD I1.1 Turbine Inlet Temperature oe iS.]

o _Poly O HP Turbines C (r) 109] (2000} o LP Turbine, C iF) ?]T (1352) o PR 2.|5 Overall ¢y=le PR L).|$ o qPoly OS.J |nlet W_O/d, kg/i (lb/en=) ].]|0 (7.120) NLp, Cadre (rpe) 2401 (2],000) Overall NHp, rld/i (rixJ) 4542 (4],J80) o PR 13.5 m, O qAO_ _ 76.T rue1 LNV, kJ/kg (Btu/lb) 42,791 (18,400) o _Polyt_ 14.7 _e o |sLt W40/4, kg/s [1b/see) 0.34 (0.74 Coabuetor ,De o _, _ 99.94 O _P/Pw t 4 _p Turbine

q

o W_0"/d, kg/s ¢lb/sec) 0.S] |1.11} o _&D, t |T.O o Cooling Pier,* t 6.1 o Intecturblne, I 1.4 (AP/P) LP TurbLne o W_L, kg/e (1b/nee) 1.96 (4.)1) o L-_I'-MO| 4P/P, t 2.

Propel let Gearbox I o _AD, t 88.S o _, t 98.0 Starer end blade cooling flovs ere shove.

The second-siege s_l_or end blades e_e uncooled.

The balance of the ¢oolLng flovs (4.9t) ere used re[ dLek and |1rices ¢oolLng.

J, i i; ORIGINAL PAGE IS OF POO! QUN..rrY £t

(i

• vI I

!

I

?All[.It iS, COtg4UTl[It ur|uNC| ItNGZNll IIIU4 LI[V|L DATA

+ I ?46 (iO00)

1149 (2100) 703 (2441) 2.70 12.60 82.0 04._ ).109 (i.854) 24Ui (23,000) 4542 (43,333) 42,798 (18,400) kg/s 1.94 (4.28) 08.0 1.7_ )States end blade cool_ng flovl ire shown. The l; uncoiled.

+-+--++.-- ++ .....

.... .... ,+ ++_+C.+-+ +_ _.d_"]r,_+ilmie"_--. " ._ . "I "_ .... ko_ I'_,.": Table 16. Modular Weight Breakdown. ||

il

_: /'' Module Weight (lb) _ Propeller Gearbox 151.7 :, ' Compressor 36.5 !I ; Combus tot 51.6 ,. HP Turbine 36.0 _i _ LP Turbine 81.3 Accessories 64.6 !I Total L , , 3.1.4 Environmental Constraints "_ Environmental constraints for the ye_r-2000 were projected _.

for commuter turboprop engines based on the sources as given for the rotorcra_t engine study in paragraph 2.1.4.

In-depth studies performed in the early 1980s by the organi- zations listed in paragraph 2.1.4 concluded that any increased stringency In the existing rules and standards would result in _] negligible beneEits to the environment, especially In the com- muter and small business Jet alrcra_t classes. In addition, r| Eur_her signiElqant reduutlons in alrcraEt noise are not likely without substantial Increases in the Eunding oE the applicable ., acoustic technologies. ThereEore, It Is reasonable to assume _I that noise limits imposed in the year 2000 wlll be equal or T_ slmilar to the most stringent rules existing today. _i - - T_f llllCl IANGt TUItDOSHAFTENOINE PRICE RANGE COMMUTER REFERENCE ENGINE

'1

I

&, o_! ' 740 _ 095

leOOl 17001 IO001 llOOi (I0001 II1001 112001 • 106511 KW IIIRPI/EH61HE ISLE Ilk ITATICi im.211.74

.;]

Based on Figure 62A, ReEerence Engine Costs Mean Market Sell P=ice.

iJ

• purpose of the i'''l Garrett SECT study, the Stage 3 take- .,I _ _1 o_f, sideline, and approach noise limits of the FAA FAR, Part 36 .'| regulations for large commuter aircraft (those carrying more than _' ', 19 passenger_) were used. ICAO Annex 16, Chapter 3 t_oise stan- P dards are essentially identical to the FAA Stage 3 limits. Small commuter aircraft studie_ should use the FAA FAR, Part 36, Appen- 'i dix F regulations that set. noise limits for a 305m (i000 _t) ° . level flyover. The ICAO Annex 16 standards are identical. A _e _ new takeoff noise measurement procedure currently is being Oevel- oped by the FAA and ICAO to replace the existing level flyover procedure for small aircraft. IIowever, this procedure should no_ - • increase the stringency of _he level flyover, and its use in this ..

° .

study was not considered to be of any additional benefit on a comparison basis.

--I _e The EPA regulations do not limit gaseous emissions from tur- boprop e;,gines. Moreover, no smoke constraints exist for small turboprop engines less than 999 kW (1340 np).

..5

The U.S. Air Force emission goals for engines are described in AFR 19-1 and are applicable to fixed wing, manned aircraft .[ only. Smoke emission levels are to be below the visibility threshold, based on exhaust diameter, with a maximum limit of SN = 65. Carbon monoxide and hydrocarbon goals are aggressive I and require _evels that result in an idle bombustion efficiency of 99.5 percent for engines with an idle pressure ratio above 3:I, and a combustion efficiency of 99.0 percent for engines with a,_ idle pressure ratio below or equal to 3:1. NO x goal levels are to be less than 50 percent of the Air Force defined "uncon- trolled" level at takeoff and climbout modes. These stringent emission goals, however, will not impact flight safety or combat f effectiveness for Air Force engines.

/ m Small aircraft engines have been found to cause only a minor impact on overall pollution levels and hez:ce have no reqL_ired gaseous emission standards {ICAO, EPA} for commercial usage.

Moreover, the stringent emission goals for military engines may be exceeded for small engines to prevent any impact on combat effectiveness. The engine specification for omissions levels is typically selected by the engine manufacturer consistent with small engine, state-of-the-art _mbustion efficiency predictions.

Emission levels likely to be required by the year 2000 for a 746 kW (i000 shp) size class turboprop are as follows: O Smoke levels - below the level of visibility, smoke number (SN} = 50 o Gaseous emissicn_ - unregulated for civil engines 3.1.5 Economic Model fo. Commuters This paragraph presents the economic modeling as used for the commuter engines of this study.

F A GTEC DOC economic model was used that is similiar to the one in a 1980 NASA Report (reference 5). The cost estimating models are based on the vehicle partitions as shown in Table 17.

The basic aircraft parameters for the economic models are listed in Table 18.

In synthesizing the total weight of an aircraft, it is con-

!

venient to partition the weight into a number of major compo- nents. GTEC LCC and DOC models used for this study made use of a TABLE 17. COMMUTER VEHICLE PARTITIONS | -t Component Cost oe Airframe minus nacelle Fixed Nacelle Variable Fixed Propeller Fixed Payload "t Missions Fixed Variable Fuel tankage o, TOGW limit Variable Q Fuel load Variable Variable Engine [ • i Variable Engine cost _m TABLE 18. AIRCI_FT PARAMETERS FOR THE ECONOMIC MODEL i000 NO. Of Engines, excluding spares el 10t Spares !

G.O Potential aircraft 500 Annual use 2000 Service life TOGW (structural limit) kg (ib) 5,670 (12,700) oe Paylold [Passengers at 86.2 (190 ib) each] 1637 (3610) .e Range km (rim) (Reference Engine) 741 (400 &, t • p, e: t i.

_t consisting of four major elemental airframe fixed

ToGwmodel

weight, airframe variable weight, installed engine weight, fuel _nd tankage weight. These can be expreseed a_ fractions of TOCW.

3.!.5.1 Airframe Weight PartitLons The fixed weight for all the commuter airframes consists of the crew, suppor_ systems, instruments, avionics, fuselage (minus the engln_ nacelles/pods), wings, empennage, landing gear, a_d systems (controls, actuation, pumps, wiring, and piping). These parts are fixed weights due to the structure of the study. The crew, support systems, instruments, and avionics are fixed because they ar_ not affected by the engine. Since the upper limit of power has been fixed at 746 kW (1000 shp) per engine for the study, the ai:craft size is also fixed. As weight

! decreases, block time is reduced, which improves DOC. This, in

I turn, fixes the fuselage (minus the engine nacelles/pods which

are affected by engine volume and weight), the wing, the empen- nage, the landing gear, and the systems weight.

-I

Some other weights were fixed or variable due to the influ-

I

ence of regulation, or of deslgn/market approaches of typical manufactures. These are described, following, for each airframe.

I

The fuel tankage except the fuel bladder (including pumps, pipes, collector plenums) is fixed. This is assumed because many commuter airframe manufacturers design the airframe with the max- imum tankage that the wings can hold. This is done so that the resulting airframe can then be used for both the executive and the commuter market. For the executive market, a large fuel sup-

I

ply is desirable to ensure that the aircraft appeals to the largest market. The maximum fuel tankage also allows a commuter to make several stops before refueling, thus minimizing block time. The maximum tankage is normally limited by the wing size.

On-board tankage is normally sold as a special kit and will not be considered for this study. Where fuel could be reduced, the fuel load was reduced to Ely the missions.

ri

,,6 The payload for this study was limited to 19 passengers.

This was due to the assumption that SFAR 41 type regulations will be in effect in the year 2000 for this class of commuter air- craft; SFAR 41 defines a small a[Tcraft at 19 passengers, maxi- oQ mum. A typical industry weight of 86.2 kg, (190 Ib) for passen- ger and baggage was assumed.

The change in aircraft drag due to increase in engine volume is approximated by the follow_ng: &| AL DRAG = CD q Sw - K2 _-v-T-- q Sw J Where: ,j C D = Coefficient of drag q = Dynamic pressure Sw = Wing area aircraft nacelle K 2 = Constant, from other -: studies Frontal area of new engine and reference engine Length of new engine and reference engine _O °.

The turboprop commuter airframe used in this study is simi- lar to the Fairchild Metro III, but it has modifications from the NASA Small Transport Aircraft Technology study to upgrade it to a year-2000 airframe, as discussed in paragraph 3.1.1. The basic m, cost parameters are generally based upon the data from the turbo- r prop aircraft operating and maintenance parameters from NASA L!

Report CR-165176 (reference 5). The fuel burn and block time to fly the mission was obtained by computing these parameters on the e.I ,I GTEC mission computer programs, SUPERM. " oo _ _._ TheThe calculaticalculations performed herein assume that the conunuter " can take off at the maxim:_m all-up weight of the reference air- _I - craft plus or minus some additional weight for year-2600 technol- ,.._ ogy engines (which may have additional components such as regen- ,., • erators). This assumption was examined by reviewing the maximum ' } weight at takeoff of the _¢_.,ent Metro III. The Metro II% has a / Garrett TPE331-10 engiue that has 746 kW (i000 shp) thermodynamic Q _ power and is gearbox flat-rated at 701 kW 1940 shp). That places the Metro Ill in the same size class as the SECT commuter. The maximum all-up weight of the MetLo Ill is 6577 kg (14,500 Ib).

With 701 kW (940 shp), it can take off at ISA at 6577 kg (14,500 [ Ib) up to 2438m 18000 ft) altitude. This allows for 816 kg llS0u • Ib) of additional installed engine weight (IEW). Thus, the assumption of adequate takeoff weight capability of the reference aircraft appears reasonable.

!i

The engine and nacelle weights can vary wlth the engine.

b The IEW consists of the bare engine weight, the weight of the cowlings, engine mounts, inlet, starter/generator, batteries,

I

connections, pipes/hoses, lubricants, oil cooling system, and exhaust system. The ratio of the nacelle plus engine weight is related to the engine weight by a constant, K I. K 1 was obtained from a study of GTEC designs of turboprop engines and nacelles in this shp category.

IEW = K 1 x engine weight The fuel (plus reserves) weight was varied with the engine performance. Only enough fuel was put on board to fly the four mission segments plus reserves. Thus, reducing engine weight I, improved specific range and reduced fuel burn (holding the engine SFC constant). A good dispatcher would normally load only enough fuel to fly the mission, o The commuter TOGW was calculated in the following form: TOGW • AF_ + K1 x engine weight + fuel (burned + reserver,) where AFFW - Airframe fixed weight TOGW can be approximated by mantpulat£ng the Brequet range equa- tion for fuel burn fraction.

The annual cost of owning and operating the referenced com- muter aircraft was structured into fixed and variable costs, as shown in Table 19.

i r ....

Based on e_pty weight, the model prepared by J.R. Humphreys t'," (Reference 6) was used to determine business aircraft development cost. Its mathematical for_. isz D o_ ADC- 741,000 (_'_t 1.49 / / _4 Where: _o r i ADC - Air_rame development cost ACEW - Aircraft empty weighv o.

_e The basic airframe cost, except f_: nacelles, is fixed. The changed development cost for nacelles was not considered to have &J a significant impact on total development cost, so a constant was used.

ee I o, 3.1.5.4 Manufactur[nq Cost Model Like development cost, manufacturing cost can also be esti- mated as a function of aircraft weight and engine thrust. The ma airframe and engine manufacturing cost inputs to the equations e described below are based on acquisition cost (sell price). The w airframe manufacturing cost model selected, which is based on data from severa_ business aircraft manufacturers, considers only o_ fixed and variable airframe weight. Its mathematical form is: .e AMC = [BMW X BMC + CW X CC + EC] QA o.

Where: .a o • AMC = Airframe manufacturing cost A* .4 l BMW = Bonded metal components weight BMC = Bonded metal cost per pound !

G

l&

CW • Composite components weight CC • Composite components cost per pound electronics packages EC • Cost oE electronics packages • Number of aircraft QA i i

I

Engine costs were estimated for each engine separately.

(3)

EM - Routine maintenance + UnscheduCed power section o; maintenance + Unscheduled line replaceable unit o.

maintenance Maintenance burden, which was also estimated by the Aero- t., space Corporation model, is dependent on the magnitude of air- frame and engine maintenance costs:

"1

I s.e (4) MB = 0.152 (AM + EM) + 6.4445 For the year-2000 recuperated engine, a number of changes

Ti

were required to properly reflect the addiVional complexity of the year-2000 recuperated engine. Airframe maintenance was con- sidered to be the same as Equation 2.

Engine maintenance and maintenance burden, however, required oO further defintion. Engine maintenance costs can be categorized as labor and materials costs, and as scheduled or unscheduled.

| Based on this deEinition, the maintenance cost for the year-2000 recuperated engine can be modeled as shown in Equation 5.

]i

(S) EM = 1.15 (Unscheduled labor maintenance*) + 0.75 (Scheduled labo_ maintenance*) + (Scheduled and unscheduled material mdintenance*J The maintenance burden definition was also revised to reflect the additional support equipK_at and parts stock required for the recuperators and associated hardware.

1400-

37.2%

1200-"

I

22.9% .rTiii.::. !: 7 + , ,,+

=- I000-

i: I 11.3%

I-.. CMW clli. 13.9%

,,,i 800-

[I

I.l 22.4% 18.3%

l= 600-

:2._o: __ _2.1%

400-

17.0% 20.9%

i'

I

200"

7.0% 8.6%

i

7.1% , 8.7% i

$1/GAL $2/GAL

I

PiBure 629. Commuter IX)C Breakdown - ReEer.ence ZngLne.

I

ORIGINAL P AC-_ 19 OF PO0_ L'L'_.qLITY 3.2 Task II - Enqlne Configuration and Cycle Evaluation The cycle/conflguratlon studies for the commuter application parametrically considered a range of potential combinations in terms of turbine rotor inlet temperature (TRIT), cycle pressure ratio (CPR}, component types, materials and associated efflclen- cles, cooling flows, p_essure drops, and leakages 4s projected for year-2000 capabillti_q. From the range of engines consi- dered, a finai engine selection for Task Ill evaluation wa_ made on the basis of payoff in aircraft direct c, perating co_t (DOC}.

The DOC improvements were estimated th_oug_ the trade factors established in Task I, which relate changes in DOC in terms o_ c_ changes in engine performance (SFC), weight, diameter, length, and cost (changes are relative to the 1985 reference engine).

Size, weight, and cost were quantified _or each engine of inter- est.

].2.1 Technology Projections Q_ The initial task in configuring potential commuter engines for the year 2000 was to establish the expected level of technol- o, ogy in that time frame. Inherent in these technology projections is the assumption that they will be available by the year 2000.

Technologies have been identified in three major areas: materials, aerodynamics, and mechanical improvements. These technologies impact the cycle study in terms of efficiency Q.

levels, TRIT limits, and cooling flow requirements, as well as turbine stage count and hub speed limits.

The technology projections for the commuter application are t essentially identical to those discussed for the rotorcraft application and therefore will not be repeated here. Material projections are outlined in paragraph 2.2.1.1, followed by aero- dynamic technologies in paragraph 2.2.1.2. Recuperator and v "-" T" '' 1 i mechanical technologies, as well as cost, weight, and size esti- mates are discussed in paragraphs 2.2.1.3 through 2.2.1.5.

Technology improvements for regenerators were investigated for both the commuter and the rotorcraft applications, The pro- Jections for regenerators are based on GTEC experience with the NASA/POE-sponsored AGTI01 automotive engine, shown in Figure 63.

As shown, a 7 percent leakage rate is projected for advanced seals with a 1000-hour life. Temperature capabilities up to I093C (2000F) are pLoJected for the regenerator core.

3.2.2 Cycle/Engine Studies The parametric cycle study for the commuter application has been generated with essentially the same ground rules and methods as discussed for the rotorcraft application in paragraph 2.2.2.

For the commuter application, three cycle types have been evalu- ated: two heat recovery cycles (recuperated and regenerated) and a conventional simple cycle.

3.2.2.1 Recuperated Cycle Study The primary emphasis of the commuter study was placed on cycles with waste heat recovery, particularly recuperation. A wide range of configurations and cycle parameters were consid- ered, as outlined in Figure 64.

The initial configuration consists of a single-stage centri- fugal compressor, reverse-flow combustor, single-stage HP tur- bine, and two-stage LP turbine. Ceramics are used in the combus- tot and turbine vanes and stators. In addition to the initial configuration, a number of component and material options were considered fcr the compressor and turbines.

,.'_: _'T,,",_','T___"t _ _ ' 41° _ .... _i , ,

lJ

fi

; .%.,,E rs _4 ,,L ° .

Q,o i TECHNOLOGY PROJECTIONS: ol • LOW'LEAKAGE, LONG.LIFE SEALS 17% AT IOOO.HOUR LIFE) . ° • HIGHER-TEMPERATURE CAPABiLiTY CORE T_ IINLET TEMPERATURE - 1093CI2000FI I b Qe t,A T1 !

APPLICATION • AUTOMOTIVE I

T_

'.1 POWER- 74,6KWIlO0 SHPI

I

SFC• 0,183 i0301 oe G5.29t.128 .e i L.

Projections Based on AGTIOI Figure 63. Regenerator Technology Experience.

"|i t

@

b _ J

.!

) ,,,, INITIAL CONFIGURATION] _ __r.-'.'_-I_ ''--'::=:z_=- "_ ''_-'_ b f7 OPTIONS _ OPTIONS i ' OPI_IONS "OPTIONS J SINGLE J _._ SINGLE AXIAL _ SINGLE/MULTI. • FIXEO-BOUNOARY [ • CERAMICS J STAGE AXIAL RECUPERATOR O.6 TO 0.8

zsT,_c_ • ,LovA,,C. I" C--AM.csl _'_

METALLICS • AOVANCEO , _p/p = imi_ AX ,_ CARBON/CARBON METALLICS 6 TO 10°'o • FIXEO/VARIABLE GEOMETRY STATOfl F, f, = 4 TO 12 TRIT : 1204C 12200F) TO 1316C [3400F) 65.261.93 II r .°1 Figure 64. Commuter Recuperated-Cycle Study Configurations and Opt ions.

,11 kl _"

t

b

J

i 'o_ w ° • = - _ _ i i'Li: For the compressor, both two-stage centrifugal and axial- centrifugal configurations were compared with the slngl_ centri- fugal design. Pressure ratios were evaluated from 4:1 to 12_I.

I For the HP turbine, only axial configurations were consid- ered; however, both advanced metallics and carbon-carbon were examined and compared with ceramics. TRITs ranging from 1204 to 1872C (2200 to 3400F) were investigated.

.I Axial stages only were considered for the LP turbine.

Advanced metallics, in addition to one initial ceramic material, were evaiu=ted. Both _ixed-geomeLry and var[ble-geometry LP =l turbines have been examined.

°_ A range of effectiveness from 0.6 to 0.8 was evaluated for the fixed-boundary recuperator. Pressure drops from 6 to I0 per- cent were evaluated. A ceramic coun_erflow plate-fln recuperator was considered as the initial configuration.

3.2.2.1.a Performance The performance (SFC versus specific power) of the initial configuration over a range of CPRs (6 to 12), TRITs of 1316 to 1538C (2400 tc 2800F), and effectiveness (0.6 to 0.8) reveals that an SFC improvemen_ of 35 percent is possible relative to the 1985 reference engine. As shown in Figure 65, increases in spe- cific power of up to 85 percent are also possible. The optimum CPR, which depends on both TRIT and effectiveness, falls between 8:1 and I0:i.

[i

!, FiQure 65. Recuperated-Cycle Performance Results for Commuter Appllcatlon.

• j .__ , ;I i ,J, J combustor pattern factors. The 1538C (2800r) cycles would there-

[l

fore require some cooling, thus reducing their performance and _b desirability. In terms 6f performance, it is clear that the highest effectiveness and temperature (0.8 and 1427C [2600F]) result in the best SFC and specific power (at a CPR between 8:1 and I0.I).

One of the penalties incurred _ith a recuperated cycle is

!J

the pressure drop in the recuperator itsel_ and associated mani- folding and ducting. To determiae the impact on engine perfor- mance, a range of pressure drops _rom 6 to i0 percent was inves- tigated, as shown in FigUre 66. The pressure drop Variations studied pertain to the total incurred in both the hot and cold streams in the recuperator.

Varying recuperator pressure drop has little impact on the optimum _cycle pressure ratio, but results in approximately a 0.5 percent increase in SFC per one percent increase in pressure drop.

Therefore, the recuperated cycles inherently desire high effect- Iveness and low pressure drop in terms of performance, which unfortunately have an unfavorable impact on recuperator size and weight, as well as on cost.

3.2.2.1.b Size and Weight The impact of effectiveness and pressure drop (_P/P) on size and weight can be found in Figure 67. As shown, both weight and 7" size increase with effectiveness and with decreasing pressure drop. A reduced pressure drop has a large impact on recuperator length, which in turn, increases overall weight.

V ¸ s Several other recuperator configurations were also examined, including plate-Fin and tubular designs in metallic or ceramic, and with either counterflow and cross-counterflow flow paths, as ?

Id- :-!

l

-- " -'rg i ,ill,, i t' Y, A S/ECIFtC P0WIP,. PlPtClNI' 10 20 30 40 F,0 _ 70 80 90 100 o to - • ; I; fo..lel---_ ,o ,o _o . "' " ,. .-uH_ I P

' ! l- Nx 1

0.Z_7] lO I| A/yp TOTIL . APIPc0i. o 4. &P/PIll . ,. _, .

S_._tJtWO.$tA_! I Taft lUOC tta_l fiEF SFC 0.292 10.481) 0.195 c(n.xlatc I,C(IIAIdlCI, I I [ IIEF _ICIFI_ POWER• 241.2 1146.7) 10.32J 1 [ . I I.. 1 1 329 362 42_ 493 111101 f2201 12801 13001 SPECIFIC POWER. KW/IKO/$I, IHP/ILII/SICII I$_111,1Z4 I',• j

i

Figure 66. Impact of Recuperator Pressure Drop on Performance.

,,/'.

J m JJ _ L I .

It

t2

]]

/ :i e4 CERAMIC PLATE.FIN COUNTERFLLlW Q_ PLENUM HEIGHT - 7.62 CM 13 IN.)

TURDINE DUCT RAOiLiS = 15+_4 CM (5 IN.]

TRIT- 1427C i2E,L_FL PR - 8 +b ,.1,PIP -- OiAMETER /.-I_l.

RECUPERATORCORE 60.04 --- CORE PLUS OUCTINO/CASIHG 5261 - .+ /.,.6% _ 55.88" _'_--_ _ .,,._ _ _§°/° _e 8% .. s -_ s" -- / f"" 45.72'

U

.. -" _//," .Ir/o 1181- 12OO1- _.,-, _/ /_10%

"'r"///_ I 52zl

_0 1141- IlOOI- o _4

o'o o'.7 o'.o o'o o:1 o:0

EeeeCTIVENESS emCTmM.8

111411.11| " °, . .

m, F:

i

Impact of Recuperator Effectiveness and Pressure Figure 67. O4 Drop on Weight and Size.

? ]• Ik;+.

.--._ - I P,, J shown in figure 68. The optimized deoign is dependent upon weight and cost, and possibly, size constraints. With respect to weight, it Is clear that either the metallic tubular two-pass crossflow design or the ceramic plate-fin counterflow designs are the most desirable, depending upon selected effectiveness.

Specific power is considered an indicator of englne slze and weight° As such, several engines have been selected to present the trend of weight as a function o_ specific power. A constant TRIT line at 1427C (2600F} Lvarylng pressure ratio} as well as a constant pressure ratio llne at i0:I {varying TRIT) is shown in Figure 69. Overall engine weight, as well as recuperator and

I

engine core weight, decreases with specific power, as indicated.

Limiting the investigation to TRITs of 1427C {2600F) and lower

!

shows that weight is starting to minimize in the i0 to 12 pres- sur_ ratio range.

I

3.2.2. l.c Cos...__

I

It has been shown that weight decreases with specific power.

,e.e, + (,..1 TUgULAA --.J ,dE TWO PASSCROSSFLOW

_- 91

"" 12001-

CERAMICPLATE-FIN I'-- COUNTERFLOW

"" 45

0,..

= I100)- ¢.1

td, J ,e.e,

o:s 0]7 o'.a I

EFFECTIVENESS

4_ 65.281.117 _° o ,¢.0 ?.

F:' t

i

Weight Comparison oE Recuperator Types.

Figure 68.

2, i.l i,'/, '/q _t -" SINGLE/TWO.STAGE TRIT = 1316C 12400FI CERAIAI_ ICERAIaICl .0.7 AP/P,IP/, PR.,IO o,.

TRIT. 1538C 12800FI .o Pii.,10 ii TRIT . 1427C 12600Fi 15001 WEIGHT" _' TOTAL AEFEIIENCE ENGINE WEIGHT 1300l IIECUPERATOli A WEIGHT

O

ENGINECORE IlOOI A 0,-,, ,.,-,.- ..,.. _. 0,-- ---0--.0 WEIGHT

O

_INCLUO|$ |NGiN[ CORE, AECUP|RATOR, ACC|$$ORI|$. ANO GEAROOX i I I I | I I

,i 329 345 362 376 395 41l 427

°o [190) (200) 1210J [220J 12301 (240] (250J 126O1 SPECIFICPOWER,KW/(KG/S), IHP/fLO/SECI]

I

65-261.122

i'|

L

,I

Figure 69. Recuperated Engine Weight Trend As a Function of Specific Power.

i I

IC.

J I .REF • SI95K _. TraIT• 131E 1244X)FI e • 0,7. ,_ PIP. 8% COST

• I°°1 !'

"'5 L.J 185'1 A_ TI

c...,c

i, 175- ,1 ,,Q 170- "-.. ! O .0 -.15 165- ""Q_,. 10 12 ENOINE q, tO "0,'@ COST tO ,,a 180- 155- :| 1O' :I o_ 14" I0 "_ °I A "_.. I_0 RECUPERATOR I0 "'"o 10 COST 0 , i

' ' 3_ ' ' ; '

11901 12001 12101 1220l 12301 12401 12501 12OOI SPECIFIC POWER. KW/IKG/SI,[HP/ILB/SECI] mml.lm .° T o , _d Figure 70. Recuperated Engine Cost Trend As a Function of Specific Power.

!

• : 800

o.s o'.o o_7 ola

65-261-121

I

Fiqure 71.

Impact of Effectiveness and Pressure Drop On Recuperator Cost.

b /T.

I i Impact on mission performance (DO(:). The trade factors, gener- ated with the year-2000 aircraft and the 1985 reference engine, place the greatest _mphasls on SFC and cost.

L

Plotting DO(: per trip as a function of SFC and specific power (Figure 72) shows that improving these two parameters does

L

in fact _esult in DOC savings. As indicated, DOC optimizes at a pressure ratio of i0:i at the 1427C (2600F) TRIT.

It has been demonstrated that although high cecuperator I effectiveness and low pressure drop are highly desirable in terms °o of engine performance, a penalty is paid in terms of size, weight, and cost. The DOC results ultimately determine where these tradeof_s optimize in terms of recuperator effectiveness and pressure drop. As shown in Figure 73, the recuperator opti- mizes near 0.8 at the low fuel price, but at the high price an effectiveness above 0.8 is indicated. Pressure drop, on the other hand, has little impact on DO(: (for the range evaluated) at either fuel price. Clearly, the need for pressure drops below 7_ ,f !

8 percent is not indicated.

For the initial recuperated configuration, the DOC results . { point to the selection of a pressure ratio of I0:£ at a TRIT of e: 1427C [2600F), with a recuperator effectiveness and pressure drop of 0.8 and 8 percent, respectively. Further, trades involving compressor configuration and turbine materials were made as dis- cussed in the following paragraphs prior to a final recuperated engine selection.

3.2.2.1.e Configuration Trade Studies Both two-stage centrifugal and axial centrifugal compressors were considered in comparison to single centrifugals. As shown in Figure 74, there is some improvement in performance to be gained, particularly at higher pressure ratios.

f- ..¢ i Ii I ii ! I II r. ", ......

!

I f • LT. /,,_/P • 0% --I-" _ IfllILEITWO-ITNHE -- -I I0 _Uili;: pC|UHICI II -I lg- It " • _0 O, PI I| p .-" .. _'.laO- )_f

-4

I " _ 15,1114; l 12,Ill • I lIl t__?

• 1427C mill

i!i

m 1 0 TIIIT • 1536C t2llaFi . _Of , , ', IO.SSl IO.,S,q _ I lli_ClFIC POWER. KW/IKG/SI. IHP/ILiI/SECU SFC.Ill/Hlil/lW, IILI/HliI/HPI

I

Figure 72. Reduct:J.on In Commuter DO(: AS I runct:J, on of_ SPC I and Specific Pover.

I 219

I

'i

O_IG_NAL PAC._ IS

I OF poO._ (:,JALn'Y

- i j

t_

f r1.1

iJ

T1

100 _ / TRtT• 1427C12800F), PR 10

! r """" 0"

I_ "_ _'_"" TWO.STABE IO.Z84/LITEIi I/ULI I

I CERAmiC tCEn_M,Cl

"1

':t,_,--is_uLI J

-t

o.o o'.7 o_o i i'o

EFFECTIVENESS &PIP. PERCENT O&.BI41t

I

Figure 73. Reduction in Commuter DOC As a Function of Recuperator Effectlveness and Pressure Drop.

L

i °,20 V J °, i ,_l_l_\_ ,_ _,_ -,.

The penalty for these alternate compressor configurations is in added complexity, weight, size, and cost. The DO(: results, as chown in Figure 75, indicate that little or no benefit is derived ,j from the added compressor complexity, particularly in the i0:i PR range.

IJ

Other configuration trades considered were in the turbine ._aterials, where both advanced metallics and carbon-carbon were Q_ evaluated. Advanced metallic cycles/configurations are compared to the ceramic cycles/configurations in Figure 76.

Replacing the cooled metallic turbines with uncooled ceramic turbines has a substantial impact on the cycle. The use of ceramics can further improve SFC by 10 to _5 percent relative to the advanced metallic engines. Also, the metallic turbines show SFC increasing with TRIT due to the impact of increasing cooling Q_ flow.

Carbon-carbon versus ceramics is compared as shown in Figure 77. Carbon-carbon allows TRIT to be increased substantially

;I

without the penalty of cooling. The cycle, as shown, clearly benefits. Unfortunately, carbon-carbon is not expected to be _vailable in man-rated engines (turbine section) by the year 2000.

In terms of DO(:, the ceramic turbines are clearly more I1: desirable than advanced metallic designs, as shown in Figure 78.

I i E The use of ceramics can reduce DOC by $50 to $100 per trip (4 to 7 percent additional DO(: reduction) relative to advanced i.

metallics. Carbon-carbon does have a DOC advantage over ceramics, but as previously discussed, it is not projected to be .!

I

I L available for this application.

,1 J I I 1 z l 40 _29 345 362 376 395 411 12001 1210) 1220) (230) (2401 1250) p SPECIFIC POWER. KW/IKG/SI, IXP/ILB/SECII I |5-281.107 ,,£ %

I

DOC Comparison for Different Compressor Figure 75.

Configurations.

I

{

• .,4IV ............ - ..... .__ .J

lJ

T ¸ ASPECIFIC POWER, PERCENT

BASE 20 40 _ 80 100

I I I . A • I , ! ; 0.208 .I M 10,441 ; LF I'UROINE 18_ : LP, TURDINE II 0.255 METALLICS I 'n" it, ..J. 1. : I A0vA'(;{0 I "!'" '''lz_.'6 I .n i ,_'-'_ 10.42l ICOOLEOI 1427C p , _--- I I20OOF I ./ % 1318C _'_ SINGLE SINGLE/TWO.STAGE O,,, Z _J i2400FI I 8 STAGE 0.243 OC I,-,,- "" 10,401 Z o LLJ 2O 0,,.

_ O.Z31 T 10.381 A " 1316C 12400Fi -- -- U-- TRIT i -p 1427C 1538C I-1 • 1427C I2O00FI .1 120OOFI 12600F] 0,219 • • 1538C 12800FI m

-25

10.301 ,,-4 I0 0.200 IUNCOOLED)

CERAMICS j

i0.341 *J

3O

0,195 TWO.STAGE LPTURBINE LPTURSINE I SINGLE.STAGE (0,321 ..

A 290 329 230 203 301 395 427 469 493 1!801 1200i 11401 11601 12201 12401 12601 12601 13001 . ° I5.ZOI.H SPECIFIC POWER. KW/IKG/Si,[HP/ILO/SEC]] Figure 76. Performance Comparison of Ceramic Versus Advanced T Metalllc Turbines.

i

I f _d

I

G

of Ceramic Versus Performance Comparison Figure 77.

Carbon-Carbon Turbines.

it i i u

• kl 1

i.!

1 ,

t

q

,++l / r 1_72C l?OOC [2400FI, ] 111,40C 133001 r) ._J_l IOMII/LIT|II I t_+31C 13000FI ,,,,% ."_" --i il2/IIALI J 200- 121100fl _r,, _ 1427¢ /q

lm-,mt, I \

izeOOFl TIIIT • IdO-

.,,o ] k I_?uu I \

1318C

leo- t+4m,_,.,, / I cAneON/cAneOl+ i

I TURBmES I

i TunemES i

,+.,c, .

'i ,ha

,_o-1 _,o_ /Ira,,,,-]

L

/'3'0_".,_ _ PR - I0

IPt4OOFl"_ e- 0.7, ,x PiP - 0%

O0- l u?c _"_" i METALLIC I

12eOOFI mac i TURBINES ] l IO_/UTtn Izam'I _ SIN6LE/TWO-STASE

"-I i_eAL)l/ ii

l,.+ 4o- ,,_ j, I I I I I i I I 1 290 329 302 395 427 400 493 5211 5541 501 II001 12001 12201 12401 12001 12801 13001 13201 13401 13001 + SPECIFIC POWER. KW/IKS/SL IHP/ILS/SECI] I i.lll.I II I + + :} I • i Figure 78. IX)C Comparison of Ceramic Versus Advanced _" Metallic and Carbon-Carbon Turbines. ]

L

.,Q

• , , L._'r" *. ' ,'!

3.2.2.2 Reqenerated Cycle Stud_ The regenerated cycle evaluation is slmilar to the recuper- ated study, with the added complexity of leakage. Only one engine configuration was considered that consisted of a single- stage centrifugal compressor, axial ceramic turbines, and a ceramic rotary regenerator.

Figure 79 shows a plot of SFC versus specific power with a range of pressure ratios from 4:1 to 10:1, TR_'_ from 1316 to

I

1538C (2400 to 2800F), and regenerator effectiveness from 0.8 to 0.95. Regenerator pressure drop and leakage are held constant at

|| _ -- ....

J ++++_..-h ..... _,+ ......... _ _..++ +.+ +

++i+ ..... _.,_+__, _

I

SPECIFIC POWER, PERCENT t.

-2O BASE 20 40 60

0.243 - LEAKAGE ,, 7% 10401 !

• °"'i ++''

.2O

0.231 0.90 [ CERAMIC- \ ._ 10,36) f_ 4 [ SINGLE/TWO.STAGE "%.i _' _I ! CERAMIC m im 0,219

,_._ PL_ _+.J I l

I,-,, 10,36] -- -st- z l.,#,,J ,i+++ t_3 O,207 _ 0.875 N

,,! I ILPI-.."_N I

l.l,J [0.341 -- _

3O

O., m +...+ . + 5.1.,, 0.195

i I__S I!J II""

O'J <I 10,321

-35

tJ 0.95 0,163 10.301 -- TRIT =

i_,,oo+:]-/- __J- r ! I LJ

tJ

4O

TRIT = 1427C [2600F)I 0,i70 TRIT -- 1538C 12800F! '

I0.261 I

! ! I

230 263 296 329 362 395 1120) 1140J 1160l 11801 12001 12201 i240J

'i

+..+

SPECIFIC POWER. KW/IKG/S),[HP/ILD/SECI] REF SFC = 0.293 (0.4811 REFSPECIFIC POWER- 241 (146.71 I J q 65-201.120 e+ Regenerated Cycle Performance Results _or Figure 79.

., T Commuter Application.

i

_o !.

I ,I

A SPECIFIC POWER. PERCENT

-2G BASE +20 +40 +60 +70

1 I " I t I - -25

i PR=" I REF SFC : 0.293 (0.4811

SPECIFIC POWER - 241 (146.71

l (0.35] |\ . ]\'TRIT = 1316CI I l"

I o.,o, I I =

i 0.20, i "_, !o',, _''1 ,,,-

• 10.331 _,"_k[\ //l_ TRIT= 1427C ]

i 0.195 I _'_,J ,_7'/; 12600FII

(0.321 ' '_"'_'_% '_ J, I 10% i

r _

t i

!

0.180 I l "_,._",_,_ _ " \ 7%TRIT= 1538C35

" _ (0.311- _ : 0.075 ,.--"-_ -,_, -r,..\",._ \ _ , ,,,o,.,,,,.

I -__,- _-__,".

0.183 io I SINGLE/TWO-STAGECEn,AMIC ! l J

m

• (0.301 197 230 263 2nli 329 362 395 427

!120) 1140) 11601 11801 1200) 12201 12401 .1280)

I

SPECIFIC POWER, KW/IKG/S). [HP/(LB/SEC)] 65.201-120

I

Impact oE Regenerator Leakage on PerEormance.

FiQure 80.

I

e TRIT 1427C 12600FI.PR- 6] _,P/P_ 5%

]

LEAKAGE 1 I ! !

0.84) 0.85 0.90 0.95 EFFECTIVENESS LEAKAGE iG% 7% 4% LEAKAGE I0% 4% I "" 5 12) ! | ! !

0.80 0.85 0.90 0.95 EFFECTIVENESS 65-Z81-tIO Figure 81.

Impact of Regenerator EfEectlveness and Leakage on Size and Weight.

[,,.,. : f

_f

l

200" SELECTE, RECUPERATE_ ENGINE AT

,oo. , ,o.s28/uTe, (,2/a,tl

-- 160" REGENESATEO .,.-,..---. I=OS_e/L,TE,I

_II :'_ _ 140" / TRIT = 1427C (2600Fl. FR : 6 " ,_u" _"_)iZ LEAK= 7% _ _ CERAMIC 100"_ & TW _., O-STAGE ! $0 264/LITER CERAMIC 801 Ill/GALl

OOt I s"---" _ I$o'2e4/uTEn i

I

i" I REGENERATED I ($1/GALI /

l 40 o:, o:o - _o

I "°I

1801 [ $0.52B/LITER I

I

=_ ,oo! ,,, I 1=2/_'cl i

I _ I I,,_'/,': 50,ol

_| e

-. ,2o 1 L_-O.aTs !

IOO_

I [ lO.528/LITER

I _

801, l ($2/GAL)

, oo 1

LEAKAGE,PERCENT

• _ I

65-281-L_

I

Figure 82. Impact of Regenerator Effectiveness and Leakage on DOC.

.J ll III regenerated engines cannot match the improvement in DO(: achieved by the ceramic recuperated engine at either oF the Fuel prices selected For this atudy.

3.2.2.3 Simple-Cycle Study A conventional simple-cycle engine was also selected for the !

commuter application, based on the results From the rotorcraft slmple-cycle study as discussed in paragraph 2.2. Due to the similarities between the commuter and rotorcraft trade factors, .° similar DOC trends were expected For the two applications, there- by not affecting cycle selection. As shown in Figure 83, an engine with ceramic turbines at a pressure ratio of 22:1 and a TRIT oF 1427C (2600F) had been selected for the rotorcraft appli- ° .

cation. This cycle was rematched to include a proper accounting of the inlet, propeller gearbox, and associated losses, as shown.

3.2.3 Cycle/Enqtne Selection A i Three cycle types have been investigated for the commuter _° application: recuperated, regenerated, and simple. All three are compared in terms of DOC in Figure 84. As shown, the recu- perated cycle is the clear winner in terms of total DOC at the T' high fuel price of this study. At the low fuel price,.it is a _ toss up between recuperated and the conventional simple-cycle approach.

Despite its SFC advantage, the regenerated engine loses out at both fuel prices. As indicated, the unfavorable contributions of weight, size, and cost are enough to offset the SFC advantage.

Based on the DOC results at the high fuel price, only the recuperated engine, as described below, was considered for fur- L.

ther evaluation in Task III: T" w_ o Compressor - Single-stage centrifugal, PR = I0:i l--.

i m

J

'ql .... , ...... _,--,_.. ..... . ._'" .t -- ROTORCRAFT STUDY 0.249 (CERAMIC TURBINES) (0.41 1316C 0.24_ (2400F]

!i _o o o.o

0.237 _r"_,la Jl : IC + 2C .._ ilc +tm ,20,,,,,,,._ -'"1 1538C -i] 0.231 _.-_22 1427" _Z_"''"_,,_ I2800FI te._!:649 10.36) 24" t_ n,-.._ "-- ( 0.225[--'1" I _Z4--...4p t .,_|

__ 10.37112c,+,,,,- I -]_-_6_._ i 22

| _ [I I I,c-,,: I ,', [

_ °.219L ] [ _ I 204[

t '°36'[1 l/ i

0.213[ [ [ _ [ : [ I [0.35) 362 395 427 460 493 526 559 (220] (240) (260J 1280) (300) (320] (340) I SPECIFIC POWER,KW/(KG/S), [HP/(LB/SEC)] I REMATCHEO • COMMUTER ENGINE OAT__A I TRIT = 1427C (2600F] • PR = 22 I • SFC = 0.238 (0.392) • SPECIFIC POWER= 406 (247) e _ • WEIGHT= 131 KG(286 LB) _ • DIAMETER = 23.4 CM (9.2 IN.)

I sr,.zel In • COST= $154.I O0 Rotorcraft Study.

I

Figure 83. Simple Cycle for Commuter Selected Erom _ - --- ,J,

[i

[J

1]

UO.ZO4hSTEn

zoo-

m 180- PR = 10 _ PR " 6 ,526/LJTER ($2/6ALI .j TRIT - 1427C12OOOF) TRIT - 1427C(2OOOF') 16(]- ,..- ,,. c = 0.8 -- c " G.875 z_P/P : 8% Z_P/P : 5% _ _,_]- LEAK : 7% i_ i"R - 22 0-- Ilt, m _, ,= TRiT . 1427C (2600F] -- 120- s,,-,, t.J I"-"

]]

.. i00- fJ fJ _ _ fj f_ fJ me fj f_ r_ fj

"Z Z ""

f/ f_' fj _ r_ fJ fJ _ f_ fj 20- _ _ _z _ _/_ -_ z[/,I G RECUPERATgD REGENERATED_ SIMPLE gl.l I| Figure 84. DOC Comparison of Recuperated, Regenerated, and Simple Cycles.

]]

I

engine selected in Task If. These refinements were made follow- ing a more detailed design analysis of each engine component.

The identified refinements and the resulting performance effects

I

are summarized in Table 21.

I

For the recuperated cycle, engine performance has changed only slightly; however, weight, size, and cost estimates were

I

revised. These changes are due primarily to a reassessment of t_e recuperator design and cost estimates. This results in a

I

r ° t , "a I i

li

!

U ,,,18 _1_ am .

_:Ov 0 C 0 .., IU.,.* _1Vl ...* i 4j o r,t ,,.¢ _ j_.j 0 (j 0_., ._ _/ ¢1/ _ ut Ulm _1 ¢_ ,..1 _ 0 .

,'_.ulwr, J I I I 0 0 0 .,_ :; ,-,M ,,,. -._,.,_.o . _. _.

!i o i

>-

o_, .',:..,',, • _ c;o.-,_ .._.., Nr "_=_' o,. • • _, _ 0o ®,-:'_" .d"_' • ,-¢ I'_ U't ,_ aO • (_ r"l L'_I " "

g_

I 041 U _[J_Oo o_' ' "' _O,,o __, , ,, o:,,,, T"

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size and weight reduction of 2 to 6 percent, and a cost increase of 11 percent. These refinements were analyzed and found to have no effect on the optimum cycle selection.

Task IV identifies and quantifies high payoff technologies for the commuter engines. It also presents technology plans that are based on the projected benefits.

g p_ I

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I

o_ e.

'l 0M i0.9i| "_ \ SLS, ISA I /

]l

MAXCRUISE POWER I 1OK _e

o.z4 _SA l

i°'!

tw ;l ,..j _'H1 %_ 1_.401 • %_ ._. --REFERENCE 119e51 I

n

0.10 %_ _.-- _ ----RECUPERATEO t2Ol)O_ I 10.31"

i

_" "" ""_"'_ 5K I_lO 0,17 _" 10K n

_o _7 _7 7to

IO.ti0 t_ 10.281 278 463 043 14001 [8001 I0001IlO00l i20oi f1501 f250i f350t el KW ISHPI FU6HTSPEED. TAS.KM/HR 110101'31 _O HI,lIT :1 iO Figure 85. Part-Power Performance Comparison for the Commuter Application.

. _ ,, + e" !; i

|

I

i

I _+0

O_ 0,.

"REFERENCE VALUES

| "- -

I-- "'INCLUDES45 MIN RESERVE " 10-- LAJ t < _+J i 2.9% 0,.

I

3.P/_ • 8.7% 2.3% TAKEOFF DLOCK"" i ! L_J WEIGHT FUEL _ DLOCK A/C

, 0j

5797 KG" TIME.

DRA6 817 KG" CRUISE 2.62 HRS" 112.781 _1 11801I.BI SPEED.

COEFFICIENT.

i -10 0.0252" TAll 472 KM/HR" 1255 KNOTBI _I.h4 " Figure 86.

Projected Reductions in Commuter Parameters, Recuperated Engine Mission Results.

i ' I !

II

f I t_

"i

i

,'1

t

i _° I

!1

-f

tooo..__ _1

mw ,I m.mu _.264/LITER I$I/GALJ _).528/LITEII [_/GALJ .0 .t I _, t Figure 87. Commuter DOC Results for Recuperated Engines.

_7

_o I oo

Y

..................... -- .... ° ..........

J 3.4.1 Tqchnoloqy Identlflcatlon/genefits A number of technologies have been identified by the perfor- mance and DO(: results from _'ask I! and Tack ZZI. They are essen- tially identical to those discussed for the rotorcraft applica- tion in paragraph 2.4.1. These technologies have benefits in terms o_ SFC, weight, si_c, or cost. Several of these technol- ogies, such as metal matrix shafts, are difficult to quantify in terms o_ DOC, but their use is cons!dered beneficial or necessary to meet engine design goals. The identified technologies are: o Component performance (aero) - Compressor n - Turbine n - Co,nbustor _P/P Materials - Ceramics (for turbines, combustors, recuperators) - Polymeric composites (propeller gearbox) - Ni3AI disk [turbines) - Aluminu_ powder metal alloys (compressors) - Cast titanium (compressors) o Combustor - Low pattern factor - High heat release rate o System technologies - Metal matrix shafts - Noncontact face seals/brush seals In order to estimate the benefits derived from these tech- nologies, GTEC isolated each technology using the technical

6&

J

•, % o , -.i_ .

", ,. • • ' , .

II

approach as discussed for the rotorccaft engine in paragraph 2.4

II

and as summarized in Figure 88. This approach involves removing one technology from the year-2000 engine, setting new cycle lim- its as necessary, and generating new engine performance, specifi- cally SFC, weight, diameter, length, and cost data. Finally, trade factors were applied to the ne_ engine parameters, which

'I

result in new DOCs. Comparing the resultant DOC values to the DOC for the baseline year-2000 engine with all technologies shows QI the improvement derived from that technology. The selected tech- nologies are not independent from one another _nd are therefore I not additive.

Qv Of the technologies quantitatively investigated, hot-end materials were found to have the greatest DOC impact, as shown in Figure 89. For example, ceramics for application in turbine air-

.!

foils contribute approximately half of the overall DOC improve- ment projected. Ceramics in the recuperator contributes approxi-

"i

mately one-fourth of the overall DOC improvement.

Compressor and turbine efficiency improvements also show significant DO(: benefit. Combustor technologies for lower pres- sure drop and improved compressor materials resulted in only small improvements in DOC.

o_ Unlike the simple-cycle engines for the rotorcraft applica- tion, Ni3AI did not pay off in DOC benefits. This can be attri- I .I buted to the much lower aerodynamic loading levels in the HP tur- bine (and therefore, lower acceptable blade and disk speeds) for =i the recuperated engine. _e 3.4.2 Technoloqy Plan _4 a_ GTEC's recommended plan for the technologies applicable to year-2000 commuter engines is discussed in this section. The e_

J

, i • .... i

!

,I

CYCLE LIMITS SET NEW ] -- SFC -- WEIGHT -- DIAMETER v TRADE -- LENGTH FACTORS APPLY !

-- COST at4rt | ° Figure 88.

Technical Approach for Estimating Technology Benefits.

Ch

o :7 ,°

!i

o,, eD TOTAL BENEFIT. SO.528/LITER I$2/GALI m SO.2_/LITERISl/OAL) I0 i $0.528/LITER ($2/6ALI RECUPERATEO I ENGINE 1 o4 TOTAL BENEFIT$0.264/LITER I$1/GALi

:7

_e CERAMICS NI3AI CASTALUM COMPONENT Ti PERFORMANCE U-N!.87 e,e bdl w_ .me o* t ¸ Eor Isolated Figure 89. Projected DOC BeneEits (Percent) Technologies.

J!

T_

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I

objective of this plan is to address high payoff technologies in

J

keeping with the technology benefits presented in paragraph 3.4.1. These high-payoff technologies consist of those with quantified DOC benefits and other system technologies consiSered vital to future turboprop engines, as envisioned for year-2000 commuters.

:I

0 key technologies. These include advanced materials, component performance, and system technologies.

I

I iT-

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L II.I 3_',d

M

i 0- ,v

3 e

=J 20- :a I0

i

i ,Z/GAL SI/GAL KEY TECHNOLOGIES • CERAMICS [TURBINE. RECUPEI1ATOR) • IMPROVEO COMPONENT PERFORMANCE [TUROINE. COMOUSTOR.

COMPRESSOR AERO] • SYSTEM TECHNOLOGIES [METAL MATRIX SHAFTS. SEALS) F_guce 90. Commuter Mission Analysis Results.

4.0 SUPERSONIC CRUISE MISSILE ENGINES

(

This section presents Garrett's S_CT study approach, method- ology, and results as established for the cruise miss;le englnee envisioned for the year 2000. This section is organized into Four major tasks as conducted and described in paragraph 1.2 of this report.

4.1 Task I - Selection of Evaluation Procedure_ and Assumptions In Task I, reference missiles are defined to repre=ent year- 2000 vehicle technology, and reference propulsion systems are defined to represent current rocket and near-term gas turbine engine technology. A representative year-2000 mission is estab- lished and the reference airframe/engine is evaluated for missile range for a constrained missile volume.

The impact of environmental constraints on the year-2000 missile is also considered as appropriate, and an economic model is defined to facilitate evaluation of missile costs. Finally, trade factors are computed for primary engine interface param- eters in order to quantify the improvements in missile range and missile cost/range for the advanced turbomachinery cycles/config- urations considered in Task II.

4.1.i Reference Missiles Two reference missiles were configured for the supersonic, tactical mission that is discussed in paragraph 4.1.2. The selection of these configurations was based on discussions held with various missile manufacturers, in particular, Martin a., Marietta.

e- 0.

l p !

The first reference missile discussed is rocket-powered and

is representative of _ current (1985) technology weapon system.

The estimated range and cost of this missile establish the per- formance benchmark for current supersonic, tactical missile sys- tems.

Due to the requirement for increased range, it I_ prnJected that advanced missiles will require the more efficient turbine o4 engine to replace the rocket engine. For this reason, a second missile was configured to represent a turboJet-p_wered missile

T

.e system.

The dimensions of both of these missiles are defined by the packaging limitations of an existing rotary launch envelope: they are confined to a maximum length of 424 cm {Z68 in.), a maximum width of 46 cm (18 in.), and a maximum height of 86 cm (34 in.].

T_ The cross-sectio_ dimensions represe_,t a triangular portion of _e the total payload area.

H

4.1.1.1 Reference Missile (Rocket Propulsion System) o_ The configuration for the rocket-powered missile, shown in

TT

Figure 91, is of conventional cylindrical shape. It is 36 cm (14 in._ in diameter and 427 cm (168 in.) in length. The payload, avionics, and controls are located in the forward section of the _4 missile in a 165 cm (65 in.) long conical housing. The propulsion B_ system, which is in the aft section, accounts for 262 cm (103 in.)

L,,, of the missile's total length.

t_ The estimated weight oE the missile is 729 to 771 kg (1600 to 1700 Ib). The missile's basic aerodynamic characteristics are similar to those of the turbojet-powered missile.

° ° . • •

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:1

I

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t

PAYLOAO AVIONICS ANOCONTAOL| Mi'_UIY LkUNCN INV|LOM \ I_K|T MOTH_ ............ .i ....... . ..............

I O t 3LO Cal ....... " 114Ill !

'_ ml Ill _ !lI| HI 4_1 CU ilOO Iml -" m,MI.72 Figure 91. Reference Missile Configuration with Rocket Propulsion System.

,, .,41' _

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4.1.1.2 Reference Missile _Turbo_et Propulsion System I

I)

The configuration of the turbojet-powered missile, shown in

i'

Figure 92, is generally trapezoid sh&ped. The _orward section o5 the missile contains the payload, avionics, and controls in a 165 cm (65 in.] long conical housing. The a_t section of the missile G. 0 (the trapezoidal section) contains the fuel, the turbojet engine, and the engit_e inlet duct, and accounts for the a_t 262 cm (I03 in.) of the 428-cm (168-1n.) total missile length. The frontal vlew of the missile shows a cross section with a maximum height and width of 53 and 36 cm (21 and 18 in.), respectively. The _e engine inlet is a fixed area Pitot design, which minimizes cost while maintaining adequate inlet pressure recovery. A flush-type .,U inlet design is not being co_sldered in this study due to high- inlet pressure losses associated with supersonic flight condi- oO tions.

The calculated missile weight is displayed in Table 22. The ,,,.t aerodynamics generated Eor the missile are summarized in Table 23.

,Dr i TABLE 22. TURBOJET-POWERED MISSILE WEIGHTS 6e %-- Weight j o,e Component kg (ib) q.

Payload/Avionics/Controls 227 (500) Structure 229 (505) D_ Fuel 129 (284) 96 (2lll Engine Total 680 [1500) .o 4.1.2 Reference Engines 'T" Two reference engines are defined for the supersonic, tacti- _d cal mission that is outlined in paragraph 4.1.3. A two-stage ,.-d.

25O

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• _ ' :J"_l, ,,_..'. -" "_' _" i - ( ,, , • •

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I II.

I Reference Missile Configuration with Turbojet Figure 92.

Propulsion System.

| IP ORIG!NAL PAGE IS

oF POOR QUALITY

)

d .. ii i i • q ¢&BLE 23. TURBOJET-POWERED RISSILE AERODYNAMICS _p bj • i

!l

Mach Ct, . Cn. No..._.

Cn O 0.005 0.0003 0.8 0.00].1 0.015 0.0017 1.0 0.0227 q 0.0297 0.025 0.0037 1.2 o8 0.035 0.0062 1.4 0.0306 0.045 0.0090 t.6 0.0277 Q 0.055 0.0].22 1.8 0.0233 't 0.065 0.0157 2.0 0.0195 0.075 0.0194 2.2 0.0174 2.4 0.0158 !

2.6 0.0146 Missile Reference Area = 1.95 m2 (21 ft 2) _o C L = Lift Coefficient o.

CD = Drag CoefFicient = CDI + CDo CDI = _nduced Drag Coefficient _J CDo = Parasitic Dcag CoefFicient Q.

";d o.

0" °

)

rocket engine that is representative of a current-technology (1985) engine system is first defined. The estimated missile range and missile cost/range for the rocket-_owered missile are used in the final payoff asses_ment in Task IZI.

TABLE 24. REFERENCE TURBOJET COMPONENT MATERIALS Part Material Disk: Titanium Compressor o- Airfoil: Titanium r_ _J Combustor Hastalloy X HP Turbine Disk; Powder-Metal Astroloy o, Airfoil: CMSX-3 Stator: Colombium ShaEt ZNCO 718 _o Enqine Performance Engine performance is predicted using component technology levels representative Of near-term (1989) demonstration engines.

r- .f i 0 i

i

.I o, _ '/ % MIXEDFLOWANO AXIALCOMPhESSOR CPR, 9 SINGLE.STAGE W,,'e/alNLE T AXIALTURBINE MACH 0.8. SEALEVEL 9.1 KG/S IZO LD/S)

i

TRIT (MAXl• 1427C(2600Fi ISA, rTHA M , 0.90 W_'e/aEXIT , 1.92 KG/$14.Z3 L9/SI W _/8/,_INLE T • 3.40 KG/S 17.02LB/S) qPOLY " 0.801 _T " 0,893 r/A 0 ,, 0.924 WCOOL • 4,4 PERCENT OIAMETER • 35.6 CM 114INI LENGTH • 100.8CM (39.7IN) THROUGH-FLOW VOLUME • 100.141CM 3 (6111 iN3J I EXPANSJON/OEFLECTION NOZZLE ANNULAR COMOUSTOR WEIGHT • 95.7 KG 1211L61 _P/P = 0.08 ENGINE COST,, $110,000 ($70K TO$110K) 65.2111.105 Engine ConEiguratLon and ReEerence Turbojet FLgure 93.

Data Sun_.ary.

,t . , ,v m_ . o ...... ........ , LJ ?

Several engine performance conditions must be considered because of the range of operating conditions required by the mission.

l}

p Figure 94 shows the _hrust of the reference (turbojet) engine and missile drag as a function o! flight Math number. Highlighted in this figure are several critical performance levels that are dis- cussed in the following paragraphs.

The engine thermodynamic design point is selected at the missile launch condition (Mach 0.8, sea level). Table 25 summar- ¢ izes the performance at this condition. The selection of this design point assures that adequate thrust is available to sustain i o, and accelerate the missile after launch. It also provides for balanced performance between the missile launch and cruise condi- tions.

:i

Another critical engine condition is max power cruise.

Table 26 summarizes the engine performance at this condition.

The engine performance at Mach 2.5, sea level, is important since ri

_4 the engine operates at this condition for most of the mission duration. The max power condition indicates that excess thrust is available for maneuver requirements. It is also an indicator of engine efficiency. It is at this condition that the TSFCs of

il

the advanced engines were evaluated in Task II.

Other engine performance include_ is the sea-level, static,

n

uninstalled condition. The engine performance at this condition o.

is summarized in Table 27.

Enqine Weight The reference (turbojet) engine weight, including accessor- .-,, ies, has been estimated at 96 kg (211 Ib). This weight was cal- f.

culated in a manner similar to the rotorcrafft and commuter engines, using the WATE-S Computer Program. The WATE-S program &, 6_ was also used to determine advanced engine weights in Task II.

• *. 256

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J _ " • .cl_an_ IP M _ 22,241

t5.ooor

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:HSTALLEDDATA ]REFERENCEIT/Ji _,.

SEA LEVEL MAX THRUST .,,,,,o 17393

14,0001'

MAX. POWER /" d 13,345 ' ,0 -- 13.0001" le t DESIGN POINT _ ,,-" "- I " '" 6.696

/'- I o-ol

12,00Oi" _"_--. SEA LEVEL, / STATIC, UNINSTALLED/s 4,448 ll.m]Ol

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l l 2.0 2.5 O.O 0.5 1.0 i.5 MACH NO.

_1 I!

I Hissile Thrust and Drag.

Figure 94.

TABLE 25. REFERENCE (TURBOJET} ENGINE - DESIGN POINT DATA

:j

Overall EngLne Perl'ormanr.e Component Pertor"L'nce HP ComDreee¢_ 6.0 o PR gnqtne RatLng, N 0194 (lb_) (1842) 02.4 o MAD, t , TSFC, (kq/hr)/N 0.167 o npoly 06.L o [xLt flow 1.92 (4.n) TurbLne Inle_ Temp, C 1427 (r) (2600) Combuetor

;J

gg.5 o q, Overall Cycle PR 6.0 5.0 Inlet W1814, kg/e g,1 o AP, t Q_ (1b/see) (20.0) HP TurbLne PhysLcaL Speed, rad/o ].657 (rpm) (34,925) 3.46 _6 o W/8/4, kg/o (Lb/oec) (7.62) JP-10 Fuel, LHV, kJ/k_ 42,100 (8tu/lb) (10,100) 0g.3 o qAD, 4.4 o CoolLng rXov, t gO o Ram eE_Lc*.ency, em QI , ( e, e_ .° lip i we _ 258 "t

o Ram efficiency, t

t

I

OF F'_.)R QUALrry J TABLE 27. REFERENCE (TURBOJET) ENGINE - SLS, UNINSTALLED DATA . d _4 SLS, Uninstalled PerFormance - Sea Level, Mach 0.0, ISA, Uninstalled With No Production Margin Overaii Engine Performance Component Performance HP Compressor PR 5.98 Enqine Ra%Ln?, N 7099 (Lbt) (L_96) o nADF t 82.3 TSFC, (kq/hr)/N 0.128 86.1 o n ly (ib/hrl/lb (1.253) o E_t _1ow 4.2_ _ombustgr Turbine InLet Tamp, C 1237 (r) (2259) o n, % 99.6 Overall CycLe ?R S.98 o AP, % S.9 inlet W/e/4, kq/t 9.1 (Lb/sec) (20.0) HP Turbine Physical Speed, rad/e 3,443 .8 (rpm) (32,884) o W/O/4, kq/e 3.59 (lb/oec} (7.92) JP-IO Fuel, LHY, kJ/kq 42,100 (Stu/lb) (18,100) O nap, t 89.1 o Cooling Flow, t 4.4 Inle.___._t o Ram efficiency, % 100 2(50 ° JE J m • .nq i ne COs t The manufacturing cost of the reEerence (turboJet_ engine was estimated based on curre.t Eabrication techl;oZogies. The acquisition cost for the reference (turbojet) engine is estimated

'!

¢ to be in the 370,000 to $110,000 range. A conservative cost o[ $110,000 is a_umed for the SECT study. This same method oL engine cost estimating was used to determine manu_acturlng costs oE advanced engines in T_sk If.

4.1.3 Re_erence MissLorl Cruise missile mission types have been consLdored by GTBC from among a number of possLble future requirements. Figure 95 summarizes these requirements into five generic classes that embody different cruise Mach numbers and ranges.

A_ter consultation wlth airframe manufacturers and the Gov- ernment, it was established that the supersonic tactical missile propulsion system should be given the highest priority in this study. This is the result of a perceived need for supersonic speeds with payload/range requirements beyond that which can efficiently be provided by a rocket or a ramjet. Moreover, tur- bomachinery component technology is felt to be lacking _or a low- cost, supersonic turbojet.

The mission objectives for the supersonic tactical missile are tactical, deEense suppression, and/or ship attack (collateral sea control}. For these cases, both speed and low signature are required for successful defense penetration. Based on these mis- sion objectives, it was decided that a low-altitude supersonic flight profile should be evaluated. This mission type, which is illustrated in Figure 96, was used to evaluate the potential payoff of engine technologies as projected in the year 2000.

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T

• ° q _e ILON6.11ANGEi STRAT[OIC MISSILE SUPERSONIC Q i1 SUPERSONIC Zl .,1,10-.30.400M TACTICAL MISSILE {S(A LEV[LI (70.000, rO0.O00 fTi i SECTSTUOY SUBSONIC ILONO-IIANGEi STRATEGIC MISSILE TACTICAL MISSILE 4572M ISEA LEVELI 115.000FTI I- SUBSONIC _45 iMEOIUMRANGEI

-

STRATEGIC MISSILE o, [SEALEVELI 0 I t I I I- ! ' I I I I T!

0 1052 3704 5_ t408 6200 11,112 I2.964 i4.SIS _AJ i10001 120001 [3000i 1401_! i50001 t60OOI 17OOOI 180001 RANGE, KM iNMi 65-281.70 .° Several Missile Scenarios for the Future.

Figure 95.

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i LAUNCH C R--LjtS__.E.E

MACH 0.8 _"'_ MACH 2.5

I SEA LEVEL _ SEA LEVEL

I

REFERENCE (ROCKET_) .-. REFERENCE(

I [_--' ' _-

TARGET TARGET

I • , , _'_, , , __,

0 18.5 37.0 55.6 74.1 92.6 111.1 129.6

RANGE.KM (NM) !

I

J _ ,5.,_, (10) (20) (30) (40l (50) .(60l 170)

I

I Figure 96. Representative R_ference Mission I 263 -| , h •I. __., o ; oJ 4.1.3.1 HlssLon-Reference Missile [Rocket) The mission for the rocket-powered mlss/le was not calcu- lated by GTEC for this study. However, a range of 57 km (30 nm) was suggested by airframer_ to be representative o_ today's rocket technology for a sea level trajectory. This value is useful for comparative purposes in the final payoff evaluation of T_k II. e 0 4.1._.2 M_ssion-Ref,eren, ce Missile (?urbolet) The mission for the turbojet-powered missile was calculated from an _n-house mission analysis program.. It was modeled Ln the following manner, i The missile is launched from sea-level carrier alrcradt o at Mach O.B.

o The mLssile accelerates at sea level at maximum thrust until It reaches Mach 2.5.

e .

o The missile cruises at Mach 2.5, sea level, until fuel .e exhaustlon.

i

Range is the total distance traveled at fuel exhaus- tion.

The range calculated for the turbojet-powered missile is 126 km (68 rim). This mission tvpe was used to determine trade fac- tors for evaluation of year-2000 engines and for the Task III mission analysis.

4.1.4 Economic Model .J The economic model used for the supersonic cruise missile is a modl_ied model that was used in the rotorcraft and commuter applications• * !

®

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The economic calculation addressed the missile elxed costs and variable costs while adhering to a set of assumptions that are summarized in Table 28. The missile 51xed costs consider the airframe and the _uel coats o5 JP-10 selected 5or this study.* The missile variable costs consider thQ enqine and total on-board 5uel costs. The variable costs are dependent on the particular cy( .e/conSiguration that is evaluated in Task II.

These _aJor contributing costs can be added in the following manner to determine total missile costs Total Missile Cost = Airframe Cost + Enq_ne Cost + Fuel Cost Airframe Cost • Cost of Airframe St:ucture, Payload, Avionics, and Controls. These costs are based on discussions with airSramers.

Enqlne Cost • Cost o5 Engine System, Includln9 Accessories Fuel Cost • Cost of Total On-board Fuel (JP-10) I Other costs associated with the acquisition costs of a mls- s_le system, such as englne spares costs, were not considered in I this study for reasons of simplicity.

4.1.5 Environmental Constraints The environmental constraints imposed on supersonic cruise missile systems are generally quite different from those on com- mercial or military aircraft. There are no universal quidelines

I

or standards governing missile environmental constraints; they

I

*Fuel price assumed $2.64/liter ($10/qal) Q

I

than present systems.

(1550) TITANIUM ALUMINIDE LIMIT •

_' 788 ,/"

_ (1450J MAC, Z.5, SEA tzv_t /

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_.I¢.3 .''-" "/RAM 0.7G -- /

Ne{ )

_ 621

(1150 N,IUM LIMIT(CURRENT) , ///y/: _//,/;_////; :////, _/_ y//;

(1050)6 8 10 12 14 16 18

OESIGN CPR

05-Z81.31 6_ I Figure 97. Advanced Compressor High-Temperature Material Requirements.

I

L

ml_,Ui;_%T%_,'_ ¸:'_i" _'_- L( on the existence of a suitable coating. Additional material pro- perties for Ti]A1 and carbon-carbon are summarized in Table 30.

' "I Aerodynamlcs The component performance projections _or the year-2000 cruise missile engine addresses low volume, low cost technolo- gies. These projections a_fect the component efficiency level_ predicted.

Compressor The compressor performance projections are generally dependent on compressor type, flow _ize. and manufacturing cost.

Combustor - The combustor technology projections for year-2000 D cruise missile engines emphasize *in_:eased temperatures in a

I

,9 %-

-..

,q i .

TABLE 30. ADDITIONAL MATERIAL TECHNOLOGIES FOR CRUISE MISSILE ENGINE TitaniL_m Carbon- Aiumlnldes Carbon Application Turbine Vanes, Compressor Vanes, Dl_des, ROtors Blades, ROtOrS, Combus:or, Nozzle 0.00_ Density, kg/cm 3) • (Ib/in) (0.065) 816 (IS00) Material Temperature 2205 (4000) Limit, C(F) (Uncooled] 20.7 34.5-68.9 Design Strength, kN/cm 2 (ksi}* (30) (50-100) Cost Factor** 1.S o *For 1-hour liEe at design temperature **l.0 is 1985 metallic engine part J , 1" ' 8 , T'

Ii

AXIAL COMPRESSOR'_-- • STAGE LOAOING FIRST STAGE I I P/P > 1.85 0.11 • 0.5 EXIT MACH NUMBER + I PERCENT CLEARANCE/HEIGH1 [0.0127 CM 10.005 INJ MIN.i I • EXIT BLADE HEIGHTLIMIT ¢/ >_ 1.09 CM 10.43 INI /, • EFFICIENCYOECHL,41ENT _qR / CASTING2 TO 4 POINTS P O.N 0.91 1.38 I,III 2.27 0.14 0311 023 0,32 0.48 0.09 12.01 13.01 14.0118.01 I0.21 10.31 10.4110.5110.71 II+01 EXIT CORRECTEO +:LOW. KO/SEC ILII/SEC] EFFICIENCY IMPROVEMENT PROJECTIONS • FURTHEROEVELOPMENT OF 3-0 ANALYTICAL METHOOS 160",+I • OESENSITIZE BLADE ROWS TO TIP CLEARANCES 120%1 • REOUCE AIRFOIL AND SECONOARY.FLOW LOSSES 120%1 I OTHER IMPROVEMENTSICOST) INCREASE STAGE LOADINGCAPACITYIN SMALL SIZE MAINTAIN EFFICIENCYWITH THICKEREDGES Pigure 98.

i.!

small volume while maintaining good performance. Specifically, low pattern factors are critical at rotor inlet temperatures of 1927C (]S00r). Also, in order to achieve a s._all volume, a very high • re_erence velocity and a low combustor residence time is _4 desired. A comparison o_ the critical operating parameters o_ the re_erence (turbojet) engine and the year-2000 engine is shown in Figure 99.

Technology advances required to achieve the tabulated per- _ormance and geometry projections include the _urther development o_ ]-D combuetor performance modeling, the evolution o_ advanced di_u_er design tools, and the improvement o_ Euel i nozzle/dome aLr_low interactLon for improved ignition and _,,an b_ow-out characteristics.

Slurry fuel is also an attractive technology _or volume- limited applications due to its inherent high energy-to-volume characte, istics.

Technology advances will address aluminum slurry, carbon °o slurry, and boron slurry. Only boron slurry, however, is eval- I uated in Task _V in order to show the additional improvements possible beyond conventional fuel (JP-IO).

°.

Turbln,_ - The H? turbine performance projections are generally a functien of stage mean work coefficient and turbine flow size.

Figure lO0 shows the turbine efficiency for these parameters. As shown, efficiency is reduced as turbine flow is decreased and turbine loading is increased.

° .

The efficiency curves assume a two-stage axial configura- t tion, no fabrication constraints on turbine blade design, an °o unshruuded blade with 2 percent tip clearance, and an uncooled

i

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OM OJ_lO_ UlUllnllC 0,11 Im I.ZS I.SO 1.711 Z.O0 OTHER IMPROVEMENTS ICOBT)

IJ

|TA|E MEAN WOIIX COEFFICIENT. _AN/UM 2 • IMPROVEO PERFORMANCE/LIFE OPTIMIZATION IN P#ELIMIItARY OESlSN • NO FABCONSTRAINTS • UNSHqOUOEO 8LAOE • Z PERCENT TIP CLEARANCE • UNCOOLEO i • TWO.STAOE AXIAL t m_lll.,41_ 4, Cruise Missile HP Turbine Projections.

100.

FiQure

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O_IG!N_,,L PAGE IS

OF poo_ QUALITY

1' blade. These efficiency curves are the same as shown for the rotor- craft application except that the stage _ork coefficient limit is increased.

Technology advances required to achieve these turbine effi- ciency projections include reduced rotor tip losses and minimized vane/blade interaction losses. Additional technology advances, primarily aimed at reducing costs, and improved performance/lif_ optimization in the preliminary design phas_ -_ m_n_mize the "o,'_rdeslgn" inherent in short lifn ,_ _::_em_able missile sys- tems, resultln_ in reduced costs.

Mechanical The mechanical projections for the year-2000 cruise missile engine are the same as for the rotorcraft and commuter applica- tions discussed in paragraph 2.2.1.4. Refer to that section for the definition and limits of these projections.

System Technoloqies Several additional technology projections were considered for the year-2000 cruise missile engines. These technologies include high-temperature, low-cost accessories, high-_mperature minimum-lubricated bearings, and expansion deflection thrust nozzles. High-temperature, low-cost accessories and high-temper- ature, minimum-lubricated bearings are needed because of the high-temperature operating conditions experienced at supersonic flight conditions.

4.2.2 Enqine Cycle/Configuration Study The cycle/configuration study for the cruise-missile appli- cation consists of a range of cycle parameters and component . _ 277 % m_\_ .... _ u'--_ ..... _....... lk i ', " -- __+ __,"'iii_iI types representing a number of potential engine options for the IL / y ar 2000. A turbojet cycle is considered in this study because _(' o_ the high Mach number (Mach 2.5, sea level) and the importance i oE high speclelc thrust.

_ The thermodynamic cycle study considers a range of turbine _0 rotor inlet temperatures _rom 1427 to 1927C _2600 to 350oF), and f_d cycle-pressure ratios _rom 6:1 to 14:1. Also, several engine configurations are considered, as shown in Pigure I01. The com- L+I , pressor configurations addressed include multistage axial com-.

pressors (stage coun_ depending on pressure ratio) and axlal/cen- -i tri_ugal compressors. Tt_e turbine configuration addresses an unco¢led single-stage axial with carbon-carbon rotor and air- _oils.

+- !

I Three key operating conditions were _ound to be important _or the cruise missile application, sea-level operation at the (1) Mach • 0.B launch, (2) Mach • 1.5 minimum acceleration point, and (3) Mach • 2.5 cruise.

The engine design point was selected at the Mach = 0.8 launch condition. The cycle parametric study assumes a compres-

lj

sor inlet corrected flow of 9.1 kg/s (20 ib/sec) at this condi- tion. and component efficiency levels based on the aerodynamic and mechanical projections as discussed in paragraph 4.2.1.

Also, a fixed-geometry inlet (nram ranging from 0.7 to 0.9) and a

fixed-geometry expansion/deflectlon nozzle were assumed. !!

Designing the engine at the Mach 0.8 operating point avoids overspeed at the higher flight speeds. Although designed at this condition, the engines were sized for the Mach 1.5 operating point. As shown in Figure i02, the Mach 1.5 operating point is • o I a --I ,, In i _ i i

I

T 22,241

150001"

I REFERE_ ,.

INSTALLEO DATA I -'-''; " _ J"

SEA LEVEL | in nu_t.__._.]

1_.7§2.

MAX. POWER " /

f40001

/

ENGINE SIZING h

.. (30001"

o==.,o.= . --,

/ "- \ IMISSILE DRAG I

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"_"(_,"" // MI_IIMUM I ....... I

ACCELERATION

S S 4k,

CONDITION

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, CRUISE _

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2.0 2.5

_i .0 1.5

0.0 0.5

it , MACHNO.

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EngLnes Sized by Minimum Acceleration Condition.

102.

Figure

[

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the missile's minimum accel condition, or the point where th_ difference between engine thrust and missile drag is at a minimum during missile acceleration.

I

pressor configuration for today's technologies because of the

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• eo "0_ .._ e_ . .,dk . ' ......

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r I J,, ?

Q o t 1 MACH 2.5.SEA LEVEL _o ,pnAla • O,70 0.21 ?

UNCOOLEO iUflIINE _,

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12.01

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0,1o IlJl 0.IS It.el

II

2114.2 392,1 490.3 568.4 086.5 704.5 1301 1401 fS)I fOOl 1701 fsol SPECIFIC THRUST. N/IK6/Si, [LB/ILB/SECJI

i

III-NI.IH NOTE:COMPONENT EFFICIENCIEB ARE BASED ON 20 LD/SEC COMPRESSOR INLET FLOW SIZE L Figure 103. HLssile Cycle Study Results.

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Pigure 104. Compressor Inlet and ExLt Corrected Plow VarLat_on.

small last-stage blade heights. Today this would typically

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require the use of axial/centrifugal compressor configurations.

Typical diameter relationships were calculate_ for axial and H

axial-centrifugal compressors. The results are shown in Figure 105. For a diameter-constrained (limited volume) application,

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axial/centrifugal compressor configurations were eliminate_

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because of the large diameter requirements of the centrifugal stage. Axial compressors, however, lie within the diameter con- straint and were selected for further study.

oil There are many axial compressc_ configuration types that can be addressed. The goal of this study was to determin_ a configu- ration type that would give the greatest potential in cycle selection. GTEC first considered a constant outer diameter (COD) compressor configuration with an exit Mach number of 0.5. A high Mach number was chosen in order to allow for higher loading/ stage, thereby reducing stage count. This configuration type resulted in acceptable compressor exit stage blade heights for

I1

pressure ratios of 7 to 9:1 (Figure 106A). To extend the P/P range, a constant mean diameter (CMD) compressor configuration

was considered next. This configuration results in acceptable U

compressor exit stage blade heights for pressure ratios of 9 to ii:I (Figure 106B). Figure I06C shows the results of a COD com- pressor configuration with the exit Mach number decreasing from 0.5 to 0.3 as pressure ratio is increased (stage count increased). This configuration type allows for the greatest pressure ratio selection since more favorable blade heights are achieved. This compressor configuration type was selected and used for the remainder of this study.

Candidate engine volumes, weights, and costs were then cal- culated. As shown in Figures 107A and 107B, engine volume and weight are cycle-dependent. Engine, volume and weight can be

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DIAMETER CONSTRAINT35.6 CM (14 INI

_) AXIAL

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6 8 10 12 14

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65-281.49 Axial and Axla1-Centrifugal Compressor Figure 105.

Diameter Comparison.

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I&al.4_ ol oo Figure 106. Compressor Axial Blade Height Limits.

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<, and Cost Trends.

Figure 107.

Missile Engine Weight, Size,

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minimized, as expected, at higher turbine rotor inlet tempera- tures and, to a degree, at increased cycle pressure ratios.

Engine costs tend to decrease with smaller engine sizes (see Fig- ure 10?C). However, engine costs tend to increase with increasing

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cycle pressure ratio because of increased compressor stage count.

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4.2.3 Cycle/Conflguration Selection The trade factors that were generated in Task I were used to °e calculate the missile _ange and missile cost/range of the year- 2000 missile. This was done for the candidate engines by using .8 their respective TSFC, volume, weight, and cost values. The ?

results of this calculation are shown in Figure IUSA for range, "4 and Figure 108B for missile cost/range. These figures show that ..

the cycle/configuration, with a rotor inlet temperature of •.

(1926C) 3500F and a cycle pressure ratio of 12:1, results in the maximum improvement. This cycle and configuration, as illus- trated in Figure 109, were selected for further evaluation in .4 Task III.

:1

4.3 Task III- System Performance Evaluation ._ In Task IIl, the selected engine from Task II was evaluated in re,ms of its impact on overall missile system performance. A detailed mission analysis was conducted that required extensive off-design performance. Finally, the missile range and cost/ .* range were determined and compared to the reference values calcu- lated in Task I.

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J Prior to these system evaluations, the selected engine con- figuration was further refined based on the GTEC preliminary design process.

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I -- "I 185 IAI .8LAOE -_ HEIOHT (100111071 _ LIMI_ (eoj, i 170 ,REFERENCe: (T/JI III O 8 I0 12 14 OESIGN CPR 3510 lel (650011 , , , , i REFEKENCE (ROCKET]-- -$8963/KM . I (-$16,600/NMI1 3240 iREFERENCE iT/Ji- $3069/KM i6000iJ- ($7350/HMli __._ 15,500 e,._ eq'

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i 6 8 10 12 14 DESI6N CPR i_.ZlII.71 Figure 108. Missile Engine Range and Cost/Range Trends.

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• COMPRESSOR -- 6-STAGEAXIAL (C.O.O.]

-- CPR = 12:1

• COMBUSTOR -- ANNULARTHROUGHFLOW II

-- CARBON-CARBON • HP TURBINE-- SINGLE-STAGE AXIAL

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-- TRIT = 1926C 13500FI, UNCOOLED -- CARBON-CARBON

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Figure 109. Selected Missile Engine.

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4.3.1 Engine Cycle Refinements Several minor engine refinements were made to the year-2000 engine selected in Task 1£. These refinements were made follow- ing a more detailed design analysis of each engine component.

The identified refineme_ts and the resulting performance effoct_ are summarized in Table 31. The following paragraphs discuss each of these refinements and their effects on component and engine performancc.

As previously discussed, the cycle parametric study oE Task II was conducted with component efflclencies based on a compressor inlet corrected flow (at Macn 0.8, sea level) of 9.1 kg/s (20 Ib/sec). A design point flow size of 6.8 kg/s (15.0 lb/sec) actually results for the engine when it is sized for minimum accel thrust at Math 1.5. This smaller flow size results in a reduction in compressor projected efficiency, as shown in the curve in Figure IIU. This engine sizing effect was not taken into account in the Task ££ study. Also, additional analyses have determined that variable geometry is required for the first two stages of the compressor, at the expense of increased engine diameter and engine cost.

The turbine efficiency was decreas:d by 4.1 points (0.920 to 0.879) as shown in Table 31. This decrease resulted from engine sizing and from turbine diameter and aerodynamic loading changes, as shown in Figure ill. The sizing changes are due to scaling the engine to meet the minimum acceleration condition. The tur- bine diameter (and mean blade radius) was reduced with no change in rotational speed; this increased the stage work coefficient {from 1.65 to 2.65), thereby degrading the efficiency by 2 points. A final correction was made to account for a single tur- bine stage (curves are for two-stage turbines with implicit reheat effects) with exit guide vanes.

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TABLE 31. HZSSTLE ENGINE PERFO_NCE REPINE.qEN?8

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Parameter* Task II Task III

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TSFC, (kglhc)IN 0.161 0.167 o Variable _somatr¥ (Ib/hr}llb (1.575) (1.642) o Refined design 0 Cost rsasiessed 9457 9150 (2110} (2057) OLsmete_, cm 24.4 27.2 (in) (9.6) (10.7) 106.7 Length, cm 101.6 (in) (42.0) (40.0)

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Weight, kg 68.0 (lb) (150) Cost, $ 80,600 87,600 Compressor 6.8 - W/O/4INL_ o EffLclenc¥ projections

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kg/s (lb/sec) (15.0) rss|sessed 0.83 - w/e/4EXIT, 0.83

kg/s (ib/sec) (1.83) (1.83) I}

- CPR 12

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0.874 0.885 - IIPOLY 0.825 0.813 - rlAO

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Turbine - TRIT, C (F) 1927 o Zf_lcLency projections (3500) (3500) reassessed

o Increased aerodynamic II

1.61 1.61 - w/e/41NLET, loadings kg/s (lb/sec) (3.55) (3.55)

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0.920 0.879 - hAD

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*Design points Mach 0.8, sea level, nRAH • 0.90

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ENGINE IZlN6 I O.N 0,84

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O.IlO 1.30 1.81 2,27 0.09 0.14 0.18 0.23 0.32 0,45 0.91 _.01 _.01 15,0l 1.21 1031 1,41 1.51 1.71 II.01 12.01 _IT COHRECTEO FLOW. KG/S IB/SECI • _AGE LO_NG FI_T ST_E PIP > !.85 • 0.5 _IT MACH NUMBER • I PE_ENTCL_4RANCE/HEIGHT 0.013 CMIO,H5INCHIMINIMUM • EXITBLADE HEIGHT LIMIT 1.09 CM _> (0._ INJ L_,_I.I_ Figure 110. Impact of Compressor Size Change.

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Figure 111. Impact of Turbine Size and Loading Change.

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% Other refinements that affect the engine include a better definition of the component geometry and a revised cost assess- ment. The refined component _eometry c_sults in a shorter engine from Task II. Also, an engine cost revision was made for a change in the estimated engine sell cost.

These engine refinements result in a year-2000 engine con- cept that Oisplays engine performance characteristics that are similar to those determined in Task _I. Also, these refinements apply to each of the engine concepLs considered in Task _. For these reasons, it was concluded that these _ngine refinements have no e_fect on the engine/cycle selectAon, and the refinements actually provide E_ a better determination of the optimum engine available in the year 2000.

4.3.2 Mission Analysis The mission analysis for the year-2000 missile system was conducted using the reference missile and mission defined in Task I and the year-2000 engine selected in Task If.

To support the mission analysis, off-design engine perform- ance of the revised year-2000 engine was predicted. Figure 112 summarizes the engi-e thrust and TSFC as a function of Mach num- ber and altitude. As shown in the figure, the year-2000 engine displays similar performance characteristics to the reference (turbojet) engine. The year-2000 engine, however, has a greatly reduced engine envelope. Both the year-2000 engine and the ref- erence (turbojet) engine are sized to the same thrust level at the minimum acceleration condition (Math 1.5, sea level). This • !

t off-design engine performance, in addition to part-power engine performance, is used in the mission analysis model for calcula- tion of the year-2000 missile range and missile cost/range."

,a f I I 0 0,5 1,0 1.5 2.0 2,5 MACHNUMBER HI.IIO Figure 112. Performance Comparison of Reference and Year-2000 Turbojet Engines.

I

The results of the mission analysis are summarized in Figure i13 for both the missile range and missile cost/range payo_ assessment. The missile range calculated for the year-2000

!1

engine is 167 km (90 nm). This compares to 56 km (30 nm) for the reference (rocket) and 126 km (68 nm) for the reference (turbo- Jet) engine. The year-2000 r_n_e exceeds the missile range of the rocket-powered missile by nearly 200 percent and the turbo- Jet-powered missile by more than 30 perce;;t. Similarly, the mis- sile cost/range for the year-2000 engine is $2867/km ($5]10/nm).

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o Component Performance - Compressor aero - Compressor casing treatments - Compresso_ castlng/toollng techniques - Turbine aero - Shaped carbon-carbon ai:foils o Combustor - Low pattern factor - High temperature rise o System technologies - Slurry fuels - Expansion/deflection (E/D) exhaust nozzle - Metal matrix shaft - High-temperature accessories - High-temperature bearings In order to estimate the benefits derived from the above technologies, each was isolated using the technical approach dis- cussed for the rotorcraft engines in paragraph 2.4. This approach involves removing one technology from the year-2000 engine, setting new cycle limits as necessary, and generating new engine performance, specifically, engine TSFC, weight, v3lume, and cost. Finally, the trade factors were applied to the new engine parameters, which resulted in new missile range and cost per mile estimates. Comparing the resultant values to those for the baseline year-2000 missile shows the improvement derived from that technology. The selected technologies evaluated in this manner are not independent from one another and, therefore, are k..

not additive.

Of the technologies quantitatively investigated, material b technologies were found to have the greatest Improvement in mis- i I I For example, most o_ _ sile cost per mile as shown in Figure 114.

II

the cost improvement is due to the use o_ carbon-carbon in the entire hot section o_ the engine. Ti3Al for the compressors was also found to provide a significant unit cost improvement.

Removal o_ Ti3AI limits the achievable compressor discharge tem- perature, which results in greatly reduced cycle pressure ratio

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and mission benefits.

Component aerodynamics were also proven to be important in _ost per mile improvements, specifically the compressor aero- dynamics, where good efficiency with sm_il axial olading i_ critical.

In addition to the technologies examined in the baseline year-2000 engine, several technologies were evaluated to deter- mine further improvements in missile costs. The additional tech- nologies evaluated are slurry fuels (boron), compressor casing treatments, and compressor casting/tooling techniques. As shown in Figures 115 and 116, the use of boron slurry fuels caza signi-

I]

ficantly improve missile cost per mile and increase range to approximately 130 nmi, which is si3ni£icantly greater than that of the baseline year-2000 value. Compressor casing treatments

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and compressor casting/tooling techniques, although important, show a much smaller improvement. The results of this technology evaluation leads to the conclusion that carbon-carbon, slurry fuels, and component performance are key technologies for the cruise missile engine.

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COMPONENT

PERFORMANCE

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Figure 114. Comparison of Technology Benefits to Missile System.

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65-291.112 Benefits of Additional Technologies.

Figure 115.

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4.4.2 Technoloq¥ Plan

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GTEC'S recommended plan for Small Engine Component Technolo- gies for year-2000 cruise missil_ engines is presented in this section. This plan addresses technologies in keeping with the o.

i technology benefits as presented in paragraph 4.4.1. The tech- n .I nologies addressed are as follows: .a 4.4.2.1 Materials 4.4.2.2 Combustor Technologies 4.4.2.3 o.

Compressor IAxie!) Performance Turbine PerEorm_ce 4.4.2.4 °.

Ii 4.4_2.5 Systems Technologies Q_ o, These plans address the high payoff technologies needed to a0 obtain "technology readiness" by the year 2000. Some tech- nologies may require verification and englne demonstration '0 testing prior to commitment to an engine full-scale development.

This activity is not included in the technology plans; it may be

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conducted during the latter program years (11 through 14).

4.4.2.1 Materials This plan consists of two discrete technology programs that

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are identlfie_ as BA and BB in Figure 117. The metal matrix com- B posite shafts program that was previously discussed for rotor- craft engines (paragraph 2.4.2.7, Program X) is shown here for reference and applies to the cruise missile engine as well.

*t I , BA. Carbon-Carbon Combustors and Turbine Rotors/Stators . .

Figure i17. Materials Technology Schedule.

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weight engines. In addition, the ability to operate uncooled

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components will result in additional specific fuel consumption (SFC) improvements. The use of these materials also adds engine

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design flexibility in handling high ram inlet temperatures and the resulting high compressor discharge temperatures.

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Carbon-carbon materials consist of high-strength carbon fibers imbedded in an amorphous carbon m_trix. Since carbon oxi- dizes at significant rates above 427C (800F), a coating is manda- tory to protect th_ composite. Zn an inert atmosphere, C-C main- talns or increases its strength characteristics up To 2205C (4000F). This high temperature capability has led to its use _n rocket nozzles and missile nose tips. Ct is also being widely used in aircraft brake disks and rotors. The successful use of coated C-C as the wing leading edge and the nose cap for the space shuttle has spurred interest in its use as an engine mate- rtal.

"1 The composite nature of C-C means that the fiber architec- ture can be tailored to meet the component stress requirements.

In the case of axial rotors, this primarily means hoop reinforce- ment in the disk and radial reinforcement in the blades. A polar weave fabric layup, which is densi_ied either by pyrolysis of a resin or pitch impregnate, or by chemical vapor deposition (CVD),

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is a natural starting point. The blades would be machined into a densifted billet to form an integral rotor. The coating would be Q,

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the final step, although oxidation tnhibitors would also be in- corporated into the carbon matrix and, ultimately, into the fibers themselves.

o, Combustor liners and exhaust nozzles will also be made as fabric layup with primary and secondary cooling holes drilled into the combustor liner prior to coating. The low stress requirements for the turbine stators result in additional fabri- cation options, including compression molding of a chopped fabric.

I]

I

Based on current programs, tho basic capability appears feasible for C-C substrates to ,Je fabricated in the required com- ponent shapes and to meet the strength requirements. Th:., major question is whether 2-D reinforcement of rotors is adequate or whether 3-D reinforcement is required to achieve improved inter- laminar shear properties.

A number of programs are addressing the coatings technology.

This includes adding oxidation inhibJtors to the substrate as well as usLng overlay coatings. Baseo on current DARPA and USA?

programs, 1371 to 1649C (2500 to 3000F) protection appears feasi- ble in the near term, while 1927C (3500F) coatings are in devel- opment for longer term use. A major hurdle is a reliable oxida- tion protection system that can be cycled repeatedly over a wide temperature range.

Proqram Description This program consists of five major tasks, from the design study through component tests, as scheduled in Figure i17. Three iterations are scheduled to achieve technology readiness.

The initial program task is to select preliminary designs for the coated C-C components. This includes t_e fiber architec- ture, attachment schemes, and fabrication approaches. Prelimi- nary material requirements are thus established. Vendors will be surveyed to determine the state-of-the-art capabilities for com- ponent fabrication approaches. Vendor capabilities will then be evaluated to determine the best long-range approaches and capa- bilities.

Basic fabrication approaches will be evaluated in subcompon- ent trials. Typical questions addressed for the rotor include polar weave versus other fiber architectures, high interlaminar • .

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strength constructions, aerodynamic shape capabilities, shaft

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attachment schemes, and low-cost densification app=oaches. Typi- cal questions for the stator include integral versus segmented

construction, chopped fabric versus laminate, aerodynamic shape 1]

capabilities, attachment schemes, and low-cost denslfication approaches. Subcomponents will be fabricated f_ testing, and NDE techniques will be established.

Under the specimen test activities, vendor material ,_ill be evaluated in terms of strength (before and after coating}, oxida- "!

tion resistance, and mechanical property changes after exposure.

Candidate fabrication approaches will be evaluated by disk spin tests, attachment tests, and mechanical load tests. Substrate, | coating, and fabrication approaches will be selected on the basis b of these tests.

Test specimens will then be fabricated to obtain mechanical

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design property data. Those specimens will have relevant fiber architectures and coatings and will be tested over the appropri-

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ate range of temperature and environment. A significant number of specimens will be tested to compile an adequate data base.

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A detail design for experimental parts is planned, along with two follow-on iterations that will have the benefit of the full design data base. The detail design will consider all mate- b.

rial and fabrication constraints.

ii

Three fabrication cycles are planned. They provide hardware for component and rig tests, hardware for the engine test, and they support the development of NDE technologies for in-process inspections.

Fabricated components will be bench tested individually to assess mechanic_l and aerodynamic perfoLmance. These component

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tests will be used to qualify components for subsequent rig tests. Test results will influence the second detail design effort.

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heat-treated alloy. The purpose of this evaluation is to identi- fy a welding process adequate to weld Ti3AI compressor rotors to one another and to cast conventional titanium alloys. Processes

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to be evaluated include electron beam, laser, and friction weld- ing. Preheating and postheating will be studied as means of eliminating weld cracking.

Based on the properties and the results of the welding study, a single-stage representative of the cruise missile com- pressor will be designed, manufactured, and component-tested.

The component testing will include overspeed burst tests, cyclic _. Jk , J - '1i ,, I Id

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whirlpit tests, and sectioning and testing of finished castings to verify material properties.

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4.4.2.2 Combustor Performance

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This section presents the technology plan for combustors as applicabl_ to year-2000 cruise missile engines.

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PlAN YEJq ° ,

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.I i, BC AOVANCEO OIFFUSIONSYSTEM FOH THROUGH-FLOW COMBUSTOR$ I o1: 2 3 i • MEASURE FLOW FIELO EFFECTSOF PREOIFFUSER. _e OUMP QIFFUSER INTERACTING WITH CURVEO-WALL o't ANO CANTEO COMBUSTOR3 USINGNONINTRUSIVE LASERINSTRUMENTATION

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• ANALYTICAL MOOELiNG/COOE VERIFICATION I • PLEXIGLAS$ MOOELFLOWRIG TESTS VOLUME.LIMITEO THROUGH.FLOW COMOUSTOR EMPLOYING B( NONMETALLIC3 • AOVANCE TECHNOLOGIES FOR HIGH-HEAT-RELEASE, NIGH-REFERENCE.VELOCITY. LOW-RESIOENCE-TIME

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COMOUSTOR3 • ANALYTICAL MOOEUNG/CODE VERIFICATION • COMOUSTION SYSTEM OES/FAB

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• SYSTEM MOOS/RIG TESTS BE SLURRY FUEL COMBUSTION • EVALUATE SLURRYFUEL CANOIOATES INCLUOINO i CARBON. ALUMINUM, ANO BORON • FUEL OELIVERY/ATOMIZATION/COM8USTION.TESTS • COMBUSTION RIG TESTS -i SECT TECHNOLOGY VERIFICATION | i; t., 11,211 .Ill

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-, Figure 118. Combustor Technology Schedule.

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'! Technical Discussion

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The diffusion system length must be minimized fo_ future engines while retaining high effectivene=s. This will require identifying the optimum advanced diffuser types to provide the

least amount of pressure loss in the minimum length. The inter- II

action between combustor and diffuser will require improved analytical models to adequately predict pressure recovery _or :6 advanced diffusers.

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_8 Analytical predictions will be compared with Plexiglas rig test data to identify an optimum advanced diffuser geometry while

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verifying analytical model accuracy. The start of this program will be delayed two years in order to incorporate the results of ,p !

the ongoing Air Force/NASA CDI program for current state-of-the- Qe art diffusers. The advanced diffuser program will employ the 3-O LDV measurement technology being developed in the CDI program and apply it to nonsymmetric advanced, curved-wall, vortex-controlled diffuser geometries.

The goal of this program is to verify the aerodynamic model accuracy for determining pressure loss and flow field in an advanced prediffuser, dump diffuser system. The current CDI pro- gram evaluates combustor-diffuser interaction with current otate- of-the-art prediffusers. The proposed program will extend that study to more aggressive diffuser designs, including curved-wall _m and canted combustor (nonsymmetric) effects.

4_ _o SO. Volume-Limited Through-flow Combustor Employing Nonmetalllcs The use of uncooled nonmetallic combus_ors will significant- ly reduce combustion system problems associated with high-cycle

(

p fabricated. Approximately eight additional liners will be -0 /, partially fabricated to facilitate experimental modifications. A second design-test iteration is scheduled to achieve technology 1 ,lI readiJ_ess.

6e Test results will be compared with analysis predictions of I.

combustor performance to verify accuracy or identify areas where additional modeling development is required.

Pattern factor and combustion efficiency levels are expected to be the limiting parameters in reducing combustor volume.

Ignition and stability characteristics will be determined for

i

each configuration.

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Following initial test evaluation, each combustor will be modified to determine optimum performance for each configuration.

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Iteratlve analysis and testing will allow up to five modifica- tions fOr each configuration during the test program to achieve optimum performance. Analysis prediction and test result compar- isons will be available for all modifications to verify model accuracy for low-residence-time combustors.

At the conclusion of the test program, the impact of reduced "i

combustor volume on pattern factor, combustion efficiency, igni- tion, and stability will be evident. This data will allow more

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Figure 119. Significant Increases in Cruise Missile Range are Possible with High-Energy Fuels•

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characterized for a variety of fuels including the effect of

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carbon loading and additive packages including catalyst. Follow- ing rig test _valuatlon, the most favorable fuel was successfully burned in an existing low-cost disposable turbojet engine, and combustor efflciencLes greater than 95 percent were achieved.

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The atomization of slurry fuels requires designs that address the slurry characteristics to centrifuge the solid part_.- cles out of solution, and its tendency to plug at much lower skin

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temperatures than do liquid fuels.

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w--I I t_ g .P,O _. ,._.'.,,,_._. PAGE OF _P_C_t_. QtJ_LITY Spray tests of aluminum slurry supplied by the Signal Research Cente, were made using the plain Jet atomizer developed during the ELITE program as shown in Figure 121. Test results indicated average droplet size to be less than the boron slurry at comparable conditions.

A comparison of the energy_-per-unit volume of typical slurry fuels is shown in Figure 122, where the lower heating values are in Btu/g_llon. Values for JP-4 and JP-10 are also shown. This _igure demonstrates the significant range incre,_se ootential for the slurries.

Table 32 presents evaluation results for a specific cruise missile application, showing the significant improvements possi- ble with aluminum- and boron-slurry fuels.

4.4.2.3 Compressor {Axial} Performance This plan consists of six discrete technology programs, identified as BF through BK in Figure 123. The plan also repeats the schedule _or a related technology program entitled "Analytical Modeling and Core Verification for Axial/Centrifugal Compressors" as discussed in paragraph 2.4.2.6, Program Q. The technology addressed by this pzogram is vital to the cruise mis- sile engine as well as to the rotorcraEt and commuter engines, and is therefore included in the technology planning for each engine.

BF. Low-Aspect-Ratio Technology for Small Axial Compressors The objective of this program is to reduce airfoil and end- wall losses in low-aspect-ratio axial stages that are required in &_._.--.----_ ....

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CERAMIC INSrRT G5-0082.16A Figure 121. GTEC Pure Alrblast Nozzle.

i Figure 122. Volumetric LHV Advantage Results in Range Improvement Potential.

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MEHTS FOR INCREASED FLOW.RANGE TO ELIMINATE/

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It is probable that airfoil shapes and design parameters that have been shown to be optimum for large axial compressors

I

may not be so for the rear stages of small axial compressors.

Nonseries airfoils c o Solldity and reaction o Spanwise twist or end bends o Spanwise compound leans o Endwall contouring

]

Initial concept testing will be followed by additional ana- ]

lytical design modi_Icatlons and tests to verify the perforn;ance of small low-aspect-ratlo stages.

]

"I BG. Increased Thickness Airfoils for AxLal Compressors =¢41

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The objective of this program is to determine airfoil design concepts incorporating increased thickness (primarily leading and

n

trailing edges) that reduce manufacturing costs and increase ruggedness and resistance to foreign object damage (FOD), while retaining the performance characteristics of thin, scaled air- foils used in larger axial compressors. Thus, the program is aimed at reducing the initial costs and maintenance costs that have been very high £or small axial compressor stages. These tl results are needed before full-stage compressor development in bO order to thoroughly explore compressor design options and ensure that proper deslgn-parameter selections can be made.

The benefits from this program are applicable to the cruise missile engine application that uses small axial stages, and _o secondarily, to enhance axial compressor alternatives for the Tq rotorcraft and commuter engines.

• . d" '. _ _, .'**L w _ Proqram Description This program consists of three major tasks, _rom the design of experimental airEoils t_rough aerodynamic rig tests, as sched- uled in Figure 123.

It is proposed that initial analytic_i, studies be conducted to determine feasibie airfoil design concepts employlno thick blading. The magnitude of thickness levels will be determined by a manufacturing study that considers production methods, toler- ances, and costs. The best concepts resulting f_om this study will be reseed tn a cascade rig to determine design and off- design losses and turning.

The cascade rig results will be examined and incorporated into the axial compressor design system. The _est design will be applied to one stage of a multistage compressor, and perfocma:ce will be evaluated in a single-stage test rig for comparison with a baseline thin-blade design.

Technical Approach The key to this program is to establish methods to retain cur- rent compressor performance with increased blade thickness and chord lengths for reduced cost.

A significant body of knowledge exists concerning optimiza- tion Of propulsion fan blading to cope with the bird-ingestion problem. This information, in conjunction with nonseries airfoil designs, for instance Rechter (reference 8], provides the basis for considering more rugged axial-compressor airfoils for small com- pressors.

i

I: The approach here is to rapidly evaluate a large number oE concept_ o. a cascade rig and choose certain configurations _or full-stage test evaluation. A sy_tematic method will be usud to

tl

screen the candidate configurations, and analytical models will 4_e be developed to be included In the future design system.

BH. Hiqh Stage Loadinq_ for Axial compressors _4 The objective of this program is to determine the e_@ects o_ increased stag_ loadings on compressor _lciency in small axial stages. The obvious benefit is reduced stage count and, in the CME application, significant red_ctlon in engine volume.

The initial CME axial compressor configuration is based on a six-stage configuration that incorporates a 1.75-pressure-ratio _irst stage. This is higher than the corresponding axial config- urations in the rotorcraft application, where volume is a lower i, !

&4 priority parameter. Further advancement in stage loading capa- bility would reduce the stage count even further and would have a significant impact on missile range. The effects are needed before full-stage compressor development in order to thoroughly explore compressor design options and ensure that proper design- parameter selections can be made.

e_ A program to establiah the aerodynamic design technology et required to reduce the number of axial stages from six to four, .I without loss in efficiency potential, would have a significant _° benefit in terms of life-cycle costs.

I e4

i

Program Description ,e° oe This program consists of three major tasks, from the stage design thrpugh aerodynamic rig test, as scheduled in Figure 123.

The program will be initiated with the design o_ the [irst stage of the four-stage axial compressor (including inlet guide vanes). The design would us_ the late_t axial compressor design and analysis methods proven on high-performance tan stages. This design will be tested in a rig capable of running both the Elrst- and _our-st_g_ units. Following the test evaluation, the first stage will be redesigned, _ollowed by design of the three aft stages. Fabrication _nd test o_ the four-stage compressor would then be conducted, followed by an analysis and a final report.

Technical Approach The major emphasis of this program is to determine the load- ing limits _or small axial stages. Existing programs are in place to determine loading limits _or large 22.7 kg/s (50 ib/sec) compressors, such as the Building Block Compressor _or the USAF.

AS the flow size is reduced to 9 kg (20 Ib) and then to less than 2.3 kg (5 Ib) per second, it is not clear that the same loading limits can be used. Current estimates suggest that peak effi- ciency can be achieved with moderately loaded axial stages (first stage P/P = 1.65 to 1.75). Two options are available. The first is to retain the same work coefficient and increase the speed to increase the pressure ratio per stage, thus reducing the stage count. The second option is to retain the existing wheel speed and to increase the loading per stage.

This program addresses the latter option and increases stage loading. The risk involves obtaining the same overall efficiency without losing surge margin. This program will examine methods of designing small, highly loaded blade rows and examining changes in aft stages as a result of increased loading in the front stage.

• This program can be accomplished within four years in a mul- titask effort as shown in Figure 123.

o J BI. Tip-Clearance Loss Minimization for Axial Compressors

II

The objective of this prugram is to establish aerodynamic design methods that desensitize axial compressor blading to tip-

t]

clearance leakage. Direct scaling of large axial compressor designs to small sizes typically leads to practical limits in blade running clearance ratio_. These effects ale needed in advance of full-stage compressor development in order to thor-

.)

-i oughly explore compressor desLgn options and ensure that proper design-parameter selections can be made.

_° .d The proposed program is expected to yietd up to a two-point improvement in efficiency of rear stages where :learance-to-blade height is crucial. These benefits, which are applicable to the CME compressor configurations, secondarily enhance axial compres- sor alternatives for rotorcraft and commuter engines.

t Proqram Description

n

This program consists of three major tasks, from the design of experimental airfoils through aerodynamic rig test, as sched- uled in Figure 123.

In the past, simple flow-rig tests of turbine airfoil tip f; geometry, with the tip-leakage effect simulated, have provided information for development of designs that have reduced airfoil I sensitivity to tip-clearance effects (i.e., AFAPL LART Program).

An initial screening effort of this type to evaluate a large num- o_ ber of concepts will be undertaken, supported by analytical efforts to correlate the results. Following this effort, selec- tion of the best concepts would be made and applied to an interi- _G or stage (rotor and/or stator) row of a multistage axial compres- T sor such as the three-stage rig proposed for the low-aspect-ratio _8 compressor program (Program BF). Rig testing at two clearances

[i

.j m I, will determine the effect of the desiqn modifications on tip- clearance sensitivity relative to the bas_llne configuration.

|,i Technical Approach Attempts to determine tip-clearance effects on axial com- pressors have been more theoretical than experimental. Correla-

"I tions have been suggested using a number of nondirne_sionai param-

eters. Clearance/blade height, clearance/tip chord, and clear- ance/tip exit staggered spacing are the parameters usually men-

I

tioned in the literature, Detailed measurements at universities have provided some basic information on tip leakage and corres-

I

ponding theoretical models.

The proposed approach is to experimentally evaluate a large

I

number of tip configurations in a simplified static flow rig.

The above parameters can easily be modified and tested to obtain

I

both overall losses and detailed flow measurements. Then this data would be used to establish empirical correlations and analyti-

I

cal models of the flow in the tip region. The concepts that appear to reduce losses and secondary flow effects would then be

I

incorporated into a multistage test environment.

This technical approach would permit a systematic evaluation

I

of methods to reduce tip-clearance losses.

stage bleed for stable operation throughout its operating _light / envelope. Casing treatment over the front rotors could poten-

[J

tially eliminate the need for variable vanes by substdnttally increasing the flow range of the _ront-end rotors, at some penal- ty in overall efficiency. The elimination of variable vanes would offer substantial benefits in cost, weight, a_d rellabillty for future crulse missile engines. The results of this study are needed before _ull-stage compressor development in order to thor-

oughly explore compressor design options and ensure that proper il

design-parameter selections can be made.

The primary activity of this program will be to experimen- tally determine the effect o_ v_rlous casing treatment configura-

iJ

tions on the flow range and stall margin from a fixed operating line, and on efficiency for an existing multistage, variable-

Q

vaned compressor. This data, along with analytical modeling, would be used to determine the feasibility and life-cycle-cost advantages of replacing variable vanes with casing treatment on

n

the selected SECT cruise missile engine high-pressure-ratio F_ compressor.

Proqram Description e_ This program consists of two major tasks, from the design of casing modifications through aerodynamic rig tests, as scheduled in Figure 123.

A compressor test vehicle will be selected from existing compressors that is similar in overall pressure ratio and front- end sta_e loading to the SECT high-pressure compressor. Al_er a compressor test vehicle is identified, two or three tip casing J

|

treatment configurations will be designed to fit the front stages of the selected compressor.

Sufficient rig testing of the selected compressor will be conducted to establish the effect of the selected casing treat- ment configurations on flow range, surge margin, and efficien_y for both fixed- and variable-vane operation of the compressor.

This will be done over the entire operating range of the compres- sor, with concentration on the two areas critical to stable tran- sient and steady-state compressor operation, the start region and the midspeed surge knee region.

Using the data from this test and from off-deslgn analytical models for the tested compressor and the SECT cruise missile engine compressor, the feasibility of using casing treatment to replace variable geometry will be evaluated. Estimates will be made for performance, weight, cost, and reliability effects.

Analysis will then be performed to establish the payoff for tip casing treatment.

A final report will document the work done, the data taken, the life-cycle-cost result:, and the final recommendations rela- tive to the SECT cruise missile engine program.

Technical Approach Tip casing treatment has been proven to substantially increase the operating range of axial compressor stages. Pre- sently, front-end compressor stages use variable iGV and stator geometry plus bleed between the front and rear stage blocks in order to get sufficient stable operating range in the critical knee and start regions of compressor operation. This is due to the natural flow capacity mismatch that occurs between the front and rear compressor stages of high-overall-pressure-ratio com- pressors.

J li i

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_'" It it can be demonstrated that tip casing treatment over the

I]

front stages can provide flow range equal to that which results t* from variable-vane geometry, it would be possible to replace the

!i

heavy, complexe expensive, and relatively nonreliable variable- / geomety construction with the simple casing treatment construc-

tion. The technical approach would be to determine, through il

al r, experimentation, the flow range of front s_ages that results from variable vanes and then to determine the flow range for casing ,j treatments with fixed vanes. This could be accomplished in an existing compressor of simlliar overall pressure ratio and front- °: °_ end stage characterlqtlcs (loading and P/P per st_e) a_ a year- - 2000 SECT cruise missile engine compressor. An estimated P

ii

compressor map will be made using the flow range improvements and asscc£ated efficiency effects from the rig tests along with the predicted stage characteristics for the tested compressor and the i; SECT compressor. This map will include the part-speed surge line .° !

for the SECT high-pressure compressor with casing treatment and o_ no variable vanes. The map will be compared with predictions for variable vanes in order to establish life-cycle-cost trade factors and part-speed stability comparison. From these results, the payoffs and feasibility of replacing variable geometry with tip casing treatments in the SECT application will be established.

B_ f_ ee BK. Axial Compressor Casting/Tooling Technology _v To use titanium alloys effectively in compressor applica- _e B" tions, forging processes (as opposed to castings), are conven- tionally used. The forging process is used because of the super- ior fatigue strength of forgings compared to that of castings produced from the same alloy. Design Engineering also requires close tolerance blade profile and blade thicknesses that, to _" date, nave only been achieved through machining.

7i

tb-6 Compressor performance levels projected for the year 2000 require airfoil tolerances that cannot be obtained by current

t

J titanium casting technology. In an attempt to advance the state- of-the-art pertaining to close tolerance titanium investment casting, GTEC Is proposing a three-ysar research and evaluation program. The advantage of casting titanium compressors is the low cost of subsequent machining. Compressor rotors (integral) would be cast to si_e or near net shape, requiring only minimal machining operations. This would el_mlnate many forglngs which currently require extensive and costly machining operations to meet design tolerances.

Proqram Description This program is scheduled as two major tasks, from the cast- ing process development through aerodynamic rig tests, as refer- enced in Figure 123.

The program will be based on an existing compressor rotor as redesigned for integral blading. The rotor will be selected to be representative in size and shape of the SECT requirement for year-2000 cruise missile engines. The program will be initiated by selection of a domestic casting vendor. The process develop- ment and tooling development will be a dual process; tooling can- not be developed without a process, and the process must be adjustable to the tooling. This resQarch and development program will establish the feasibility of advancing the state-of-the-art in close tolerance integrally cast titanium axial compressor rotors.

4.4.2.4 Turbine Performance This section presents the plan for turbine performance improvements, as applicable to year-2000 cruise missile engines.

2he plan consists of four discrete technology programs. Three are shared and one, Program BL !n Figure 124 is new.

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.i _o TURBINE VANE/BLAOE INTERACT!ON • SEEPARAGRAPH 2.4.2.4, P_|OGRAM L TURBINE STATOR HIGH-TURNING CRITERIA • SEEPARAGRAPH 2.4.2.4.PROGRAM M TURBINE TIP.CLEARANCE SENSITIVITY REOUCTION • SEEPARAGRAPH. 2.4.2.4. PARAGRAPH N IlL SRAPEO CARRON-CARRON AIRFOILS FORIMPROVED PERFORMANCE • MECHANICAL/STRESS STUOY OFgLADE OESlGN ANO

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MANUFACTURING PRACTICES TOELIMINATE STRAighT.

LINE.ELEMENT RESTRICTIONS • MECIIANICAL/MANUF;CTURING STUOY SECT TECHNOLOGY VERIFICATION

[i

HI.IN -,, B, b-d Figure 124. Turbine Technology Schedules.

I

The _irst three shared programs address technologles as dis-

!

cussed _or rotorcra_t engines in paragraph 2.4.2.4, Technology Programs L, M, and N. These programs are an integral part o5

!

this plan as well, since the technologies will apply to the year- 2000 cruise mlssle engines. The remaining Program (BL) is dedi-

I

cated to year-2000 cruise missile engines.

BL. Shaped Carbon-Carbon Airfoils For Improved Perform&_ce All axial carbon-carbon (C-C) turbine rotors currently under

Technical Discussion

ii

The most promising C-C architectures intended for use in turbine rotors are the polar weaves. They have both hoop and

n

radial filament bundles (tows) of 3000 to 5000 fibers each in the disk and blade. Toward the bore there are more hoop tows I, than radial, and conversely, more radial tows than hoop in the rim and blade. The radial tows in the blade are needed to take the centrifugal loads while the hoop tows are used only to maintain the radial tow integrity during the fabrication process, and to maintain a more or less uniform ply thickness.

Ideally, the filaments in a fibrous composite should be aligned with the principal stress directions. Unfortunately, this is usually impractical, if not impossible in practice. It J

I

Several of these blades (alrfoil plus root) would be placed in a ring and wound with roving and pitch. This winding would make the hoop lu_d carrying portion of the disk.

The second design concept would use a more conventional integral bladed disk (blts_J approach. The bllsk blank would be fabricated from preimpregnated fibers then cured, pyrolyzed, _nd densified. The airfoils would then be machined from tne blisk blank. This blank could be a conventional 2-D polar weave tamL- hated composite that has been strengthened in the airfoil region by adding a continuous spiral weave insert (see Figure 126).

This spiral weave insert would have fibers In the +45 and -45 degree directions, thus increasing the in-plane shear strength.

The number of hoop fibers in the blade could be reduced, thus making room for the spiral weave insert without sacrificing the strength in the radial direction.

4.4.2.5 System Technoloqles This plan consists of three distinct technology programs, identified as BM through BO in Figure 127. The plan addresses special technologies that are vital to the englne system as envisioned in this study for year-2000 cruise missle engines.

BM. Advanced Supersonic Nozzles Design requirements and constraints for advanced supersonic I cruise missile engine thrust nozzles are functions of the engine -° add airframe designs, the mission (tactical or strategic}, o.

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,o I

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'_ 1 L /i; I Proposed "Spiral" Weave.

°°

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DM AOVANCEO SUFI_RSONIC NOZZLES ,, ICALEO MOOELTEST EVALUATION OF HIgH PNEUUNE.

RATIO NOZZLESSUITABLEFOR SMALL.OIAMETEfl.

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• , ° .

, , .., and the range. In general, however, these requirements and con- straints includes 0 Medium to high nozzle pressure ratios (10 to 30 _or tactical missions, 40 to 100 _or strategic missions)

I

o Light weight/small volume requirements o Maximum diameter constraints ,, 0 High exhaust gas temperatures/little or no cooling _low available 0 FaLrly broad range of operation with high performance requirements o .

o Minimized variable geometry _or low cost and simplicity I i o Requirements Eor nonaxisymmetric geometry ., o Low observables requirements

t!

To satisfy these requirements, an advanced supersonic nozzle is required, Studies in this area have shown that the expan- sion-deflection (E-D) nozzle type would best meet most or all of

i

the above requirements, both for tactical and strategic missions.

t _ _| I i '

[

Design performance close to alternative configurations that are larger and heavier.

0 Superior off-deslgn psrformance due to the aerodynamic adjustment inherent in the E-D nozzle's base region flow field.

!

T o Nozzle performance is Installatlon-lnde_enden_.

!i 0 T Variable geometry is possible via csnterbody transla- tion.

In addition to design studies o_ axisymmetr!c E-.3 nozzles, the ELITE program included studies o_ nonaxlsy._metric (ellipti- cal} Z-D nozzles.

i Current industry and GTEC aerodynamic analysls/deslgn capa- L.

bility for supersonic nozzles in general, and E-D nozzles in par- ticular, includes the following components: I Performance prediction: Theoretical I-D calculations,

L

supported by experimental data base. For E-D nozzles, the existing data base covers only the high end of the CME pressure ratio range.

T

2-D flow analysis: The 2-D, finite-element, viscous tlme-dependent, flow analysis code VNAP2 is used by GTEC for subsonic and supersonic wall contour optimiza-

i

i tion in axisymmetric E-D nozzle design, as well as for other supersonic nozzle types. Other 2-D methods, such

!

as radial equilibrium analysis EOE subsonic ducts and the 2-D method of characteristics for supersonic flows, may also be used.

I

i

speeds and must be mounted internally in the engine to reduce { frontal area and weight associated with tower shafts and redt,c- tion gear boxes.

The objective of this program is to evaluate, through bench i testing, a combined fuel pump/PMG module designed for mountinq In an expendable supersonic engine.

Prcqram Description This program consists of four major tasks, from design through module tests, as shown in Figure 127.

lq L_ Oesiqn - Existing and planned GTEC expendable engines will be studied to determine fuel flow and electrical power require- ments, operating s__:i ranges, environmental considerations, and internal physical space allocations. An- appropriate existing engine will be selected as a test bed.

Analytical calculations and detailed mechanical design tasks will be undertaken. This will support drawings being made for a prototype combined fuel pump/PMG module and associated rig test hardware. The design will permit both components to be function- ally tested as separate units.

e-.

Fabrication - T_'" prototype units and appropriate spare parts will be fabricated. Manufaccuring methods for small quan- ° tities will be employed, but the design will be compatible with L.

mass production technologies.

Bench Test - Initial tests will be conducted at ambient tem- _J !

.! perature on the fuel pump and PMG, as separate components, to T_ validate the basic design and to isolate problems unrelated to system experimentation.

," --_ ., .: ;.., r _ Module Tests - The components will be comoined and subjected to testing over the expected range of fuel flows and electrical lo_da. Environmental tests will then be conducted to expose the module to the expected temperature extremes associated with the candidate engine and mission.

Technical Approach Figure 128 shows the proposed fuel pump/PMG module as con- ceived _oz an existing engine. Fuel enters by means o_ a strut across the aerodynamic flow path. It flows around and through the PMG end turns, slots in the stator, and through the

t

air gap to the pump inlet.

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Pill STILTON T LAMINATIONS 1 ITATOII WlllOIllOII LUIO TUIIUSl !

PMO ROTOR 8L|F.VETYPE

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LIAIIAIIIUM COIALT I|ARINO FUEL.LUIEO O_LI, IMPELLER

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8NAFT rNERAIAL INIULATION

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• ._ BLANKET

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PUMP UlIOUO

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Proposed Fuel Pump/PHG Hodule.

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_ _ .';'_ _ _,,,,, ....

ii Perhaps the most important factor in material selection will be the separator. By definition, externally powder-lubricated bearings lend themselves to use of a conventional steel silver- plated separator, but only up to temperatures less than 538C (1000F). A ceramic retainer is not feasible because it lacks strength in the tensile mode and has no ductility. Even a shrouded ceramic separator would probably not withstand the ball pocket collisions, which would lead to web break-up. A mo_e I io feasible choice for a separator material may be shrouded carbon, although current carbons are limited to approximately 649_ t _d (120OF).

i'

.I

The bearing material with the longest rolling contact fatigue llfe is _ot uecessacily the best candidate because race wear and interaction with the lubricant (to reduce friction and wear) are more important than fatigue life as evaluated under normal high-load, liquld-lubricated conditions. A race candidate

i

Q_ with lower fatigue performance might have better wear life because of improved tribochemical interactions. Similarly, the highest strength retainer material at the operating temperature might not be the best candidate if it forms abrasive surface oxides. A lower-st: _ngth material that forms lubricious oxides might be a better candidate.

Dry-Lubricated Bearings - The design approach for dry- i: lubricated thrust bearings is necessarily different from that for conventional liquid-lubricated bearings. Whereas conventional liquld-lubricated bearings are designed and optimized for skid- ding and fatigue life, solid-lubricated bearings must also con- centrate on the separator design to attain maximum reliability.

Past experience as well as recent GTEC testing have demonstrated that the bearing separator is the limiting component in unconven-

I

tionally lubricated bearings. Failures result from the following situations: O Unbalance failures (inner land-piloted schemes] due to worsening wear of pilot land o Fractures a_ a result of dyna_lcs created from uneven build-up of solid film lubricant 0 Resonance failures (high ball/race traction transmits ball passtng frequency to separator) Aside from the separator, the bearing internal geometry wl_l be optimized for considerations of solid film lubrlca_ion. Race- way curvatures and contact angles will be selected _o produce minimal heat generation, least ball dynamics, and largest in_er- nal clearance. Previous efforts at GTEC have demonstrated that solid lubricated bearings may possess extreme temperature gradi- ents, a factor that must be considered in the design.

Since ceramics are being considered for both the rings and rolling elements, a system of shaft attachment must be incorpo- rated to ensure that _no tensile stress is transmitted into the ceramic inner ring. The method to be used will be simil_r to the one previously discussed.

The sacrificial separator material as well as the type of powder lubricant, will be selected based on past GTEC programs that have researched and surveyed the successes and developments in the field. GTEC has accumulated a significant amount of data

I

on the types and methods of introducing dry film lubricants into the bearing. "

I

p . ._. 'A,. ,T The mission analysis resultu show a significant improvement in both missile _ange and missile cost per mile over the refer- ence missile (Figure 129). for comparison purposes, the approxi- mate range and cost improvement relatlve to a rocket-powered mls- eile al.o is shown. The advanced mi=sile achieves a 22 percent and 91 percent range Increaae using JP-lO _nd bocon slurry _uels, r_spectively, relative to the reference turbojet. When compared to the rocket-powered miesile, a 200 _nd 333 percent ranqe increase is possible with JP-10 and boron slurry, respectively.

Miaaile co_t per mile decreases oE 27.8 anJ 50 percent _or -i JP-lO and boron slurry, respectively, were shown relative to the reEerence turbojet. A 70 to 80 percent cost per mile reduction ia achievable _or JP-10 and boron slurry, when compared to the rocket-powered missile.

These significant improvements are dependent on advanced in materials, advanced _uels (boron slurry], and component oo performance.

I The study assumes an installation weight of two times the APU weight. The aircraft fuel consumption penalty to carry this rK,_ULNG i'AL;E BLANK NO'J: VILMED

-- I

installed weight is assumed to be 0.063 liter fuel/hr/kg (0.00755

!

gal/hr/ib). These factors are based on experience and past studies.

i f The typical installation is in the extreme aft part of the !

fuselage. Since it is aft of the pressure bulkhead, and fuel is • d plumbed to the area, a fireproof enclosure and a flre-suppression system is included. This location has several advantages, such as the followingx Noise and turblne exhaust are directed above and away from passengers, crew, and ground personnel.

° J The fairing structure is aerodynamically designed to reduce drag so that volume is not critical.

° t The location is not convenient for other uses because _d it has limited ground accessibility.

7!

t.J o It is nonpressurized.

For the purposes of this study, the APU volume was not con- sidered as an evaluation factor.

o !

5.1.2 Reference Duty Cycle The operational duty cycle for the year-2000 commercial transport is shown in Table 33. The APU is assumed to be oper- ated one hour per aircraft flight/hour. This is typical usage for many local airlines using current 120- to 1SO-passenger air- .0 craft such as the Boeing 737 and 727, and the McDonnel Douglas DC-9.

T om L

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v s_-+_+,,,,at _ -+._- _ ...... -_+ ._,,.+.

-# • im Jk

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The APU is normally operated only on the ground. A start is

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made during taxi, and the unit is run continuously until taxi out for takeoff. Howeverr the APU also serves as a source for redun- dant electrical and hyra_llc power in _light. It becomes flight

[

critical in two-engine aircraft such as the DC-9 when the air- craft has a propulsion engine generator out o_ service. In this

[

case, the aircraft cannot be dispatched until a backup power f source is available.

l e, The reference engine by definition is a unit that has demon- strated the performance specified but is not in production. A cross section of the 1985 technology uni_, as selected for this study, is shown in Figure 130.

!r The gearbox, inlet _lenum, and the location of the elec-

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tronic control unit are subject to installation requirements established during the air_rame design period. For the purposes

H

of this study, the installed weight was estimated at twice the engine weight. In commercial usage, these engines include con-

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tainment _eatuces which have the demonstrated capability of con- taining a maximum energy hub burst at 125 percent of the design maximum rotational speed. These components on the reference engine have added I0 kg (22 Ib) to the re_erence engine weight.

The APU contains an automatic start system and a control system which will limit the applied load to the selected maximum turbine inlet temperature. As an example of the reliability of this system, many airlines allow completely unskilled cleaning crews to operate the on-board APU when the aircraft is located in U an airport area which is not convenient to commercially produced electrical power.

!

dL_,_lll$( FLOW ANNULAR COMiiUSTOII 81NOUE,.BTAOE

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RAOIAL INLET SCREEN IIAOIAL TURlilN| PLENUM TroT lilAXl • lOS IllOOfl

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Cross Section oE Re_erence APU Without the Gearbox.

Figure 130.

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-.%-.. :... :.,,,,.--..+_-,- , , -. , ....... .. . . ,

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_e_erence engine is glver, in Tables 34

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I,!I and3s.

The reference turbomachlnery estimated weight shown in Fig-

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ure 130 is 42.0 kg {92.6 lb), includlng 10.0 kg {22 Ib) of con- tainment. The total APU estimated weight is 68.9 kg (152 ib), Q4 which includes a gearbox, starter, oil cooler, and five pounds of oil.

The cost of the reference engine in 1985 dollars is e_Li- °i mated at $120,000 each in production quantities.

o_ 5.1.4 Environmental Constraints APU noise has concerned the airlines, aLrpoct authorities, .Q and manuEacturers. To address this concern, ICAO has adopted guidelines for the noise certification of APUs installed in air- craft and operating on the ground. These guidelines establish maximum APU noise levels at work stations and at a 20 m (66 ft) perimeter around an aircraft. The airframe manufacturers have been using these ICAO guidelines as a basis for their APU noise

level requirements. GTBC uses the guidelines as a goal during I!

the design of every APU installation. The guidelines are strin- gent and are difficult to meet without adversely affecting air-

:l

craft weight, centec-of-gravlty limits, compartment env_lope, APU o.

performance, and system cost.

ee APU noise reduction technology in the year 2000 will be o.

focused on improved inlet and exhaust duct and muffler designs and on radial turbine and centrifugal compressor aeroacoustic designs. Even with additional research in these areas over the next 15 years, the existing ICAO guidelines for installed APU noise will be difficult to meet. Therefore, the Gatrett SECT study has used the ICAO Annex 16 Attachment D APU noise guide- $ lines.

O PR 1.97 ,,.a o hAD, % .,!

o npo!y, % O Exit corrected elow, kg/s (lb/sec) ** *No margins e B **Compressor at part load using variable IGV.

ft

O8 i Trade-offs were analyzed between the costs of engine acqui- e# sitlon, _uel, direct labor, and material replac=ment, to ensure the lowest system LCC. Return on investment (ROI) analyses will be made in these tradeoffs.

:i

.¢ The analysis used different modeling approaches for Task II and Task Ill analyses. OurLng Task II, parametric models were primarily used, s_nce results of the evaluations of alternative b, engine configurations could be obtained much mor_ readily than with a bottoms-up approach.

During Task III, a bottoms-up model, based on individual

;I

piece part estimates, was used to more accurately account for major component and llne replaceable unit (LRU) coats.

GTEC used the Air Force 5SC i.i model, which had been used

Ti

successfully in previous cost analyses of commercial APUs. The _4 LSC I.I model allows costs to be evaluated for engine components to the necessary level of indenture as appcoprlate for the trade studies under investigation. Basic assumptions for these studies I are: Bt The aircraft is not all-electrlc, but may be mostly O _f electric. Advanced system studies indicate that some functions can be performed more efficiently using pneu-

!

matics and hydraulics.

o, o The installation will allow operation to 15,240 m (50,000 £t) altitude.

_t o Initial cost and material are based on a 500-unit pro- duction run. APU life Is based on 20,000 hours.

o The installed weight of the APU will be twice that of

oI

the APU weight. The penaAty of transporting this weight will be a composite of current _irfram_ values: 0.063 liter of fuel per flight hour per kg of equip-

I

ment weight (0.00755 gal/hr/ib).

I

Spare parts will be procured at _ rate of 30 percent of initial parts. Associated labor ra_es and buLden rates

I

will be based on 1985 negotiated rates.

o Fuel costs will be based on the duty cycle with the assumption'that the APU is used one hour per aircraft

I

flight hour (3000 hours/year).

T .....................

Z .-- " • 'I complexity can also be seen in the weight factor (either positive + or negative) and typically has only a secondary effect. -I The cost oF fuel has a strong influence on the overall cost oC ownership through both the Fuel and weight factors. It o.

dictates emphasis on SFC and on the cost of transporting the ,e weight of the APU.

• ° _he trade factor values are given _n Table 36 and typify the °, InEluence of a i percent increase Ln cost of these factors on the f_nal cost-of-ownershlp _ummatlon. °e e.

-I TABLE 36. APU TRADE FACTORS ° $0.264/llter $0.528/liter ($1/gal) ($2/gai) i+ [Percent) _Percent) 0.135 0.08 _DOC/1% a Eng/ne Cost e_ 0.20 0.13 aDOC/I% _ Maintenance .4 0.55 0.66 dDOC/1% _ P'tel Burn w 0.10 0.12 _DOC/I% _ W._ight .° As these factors indicate, the fuel and weight factors show a dzrect relationship to the cost of fuel.

5.2 Engine Configuration and Cycle Evaluation °.

The cycle/configuration studies for the APU application con- sidered a range of potential combinations of turbine rotor inlet temperatures (TRIT), cycle pressure ratios (CPR), component type, ;+ o.

I

J * ,t'_'_qb *4p_Lr_.'%"l=," "..'_'_Q' ;"""_e1" _ * -- • i J i 6d

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tt

°o SINGLE-STAGE COMPRESSORS TWO-STAGE COMPRESSORS -6 LIMIT FOR CORRECTED INLET APU FLOW IN L_/SEC 'T v .'1

3 4 5 6 7 8 9 10 15 20

PRESSURE RATIO

il

65-_0l,I 7 _0 "?

_e Figure 132. Projected Centrifugal Compressor Efficiencies.

• i L ..... ,,;,, ---_. ...........

14.01 (2.oi

z LI,,I P 0.45 LL

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I1.01

0.88 I

W_f8/_ = 0.23 KG/SEC 10.5 LB/SEC]

I 0.86

1 I t.25 1.50 1.75 2.00

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MEAN WORK COEFFICIENT.gJ_ H/UM 2 ASSUMPTIONS: I] "METALLIC TYPE" GEOMETRY 21 UNSHROUDED BLAOE

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3) 2 PERCENTTIP CLEARANCE 4) UNCOOLEO

[

65-261.19

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,G 133. for High-Pressure Projected Efficiency Figure T Axial Turbine Stage.

,[

[

Combustor Technology

[ Combustor improvements projected for the year 2000 are shown

in Figure 136. Discussion of these technologies is found in paragraph 2.2.1 under the rotorcraft application.

Y 5.2.2 APU Cycle/Confi_u;ation Studies Engine Cycle Studies - The engine cycles shown in Figure 137 were studied as applicable to the duty cycle called out in paragraph " 5.1.2.

[

J 0.92 _8 P QQ

o.9o o___

• ° • _ iW,_/'b/OllN -> 0.91 KG/SEC

o88 X

. b bO ,P,

0,"'_ ° , \ '_\ \\\

0.8 1.0 1.2 1.4 1.0 1.8 2.0 f MEANWORK COEFFICIENT gJ AH/UT2 65-2111-78 • REFERENCE ENGINE OE$IGN POINTEFFICIENCY e,e Figure 135. Projected Radlal Inflow Turbine T_ Efficiency in Year 2000.

dt _'b _ _. ',; :'i_,_ .i,.,_ , .

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REFERENCE YEAR-200O ,_GINE ENGINE APU APU APU PARAMETER COMPR($SOR

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:T I; -e id 'T _o

FREEPOWERTURBINE

REGENERATED

G5.26t.20 o, _e Figure 137. APU Cycles Investigated in the SECT Study.

ii

J Each engine was evaluated on the basis of the cycle pressure ratio that generated the minimum fuel consumption at the design point (maximum power on a sea-level, 38C [t0IF} day), using the turbine inlet .temperatures as limited _y material projections.

The cycle analyses of simple-cycle engines using axial turbine wheels are shown on Figure 138. Pertinent points demonstrated by these data include: O Optimum pressure ratios occur at less than 10:l; there- fore, only one stage of compression is necessary.

[

O Uncooled metal turbines at I038C (1900F) produce SFCs lower than the cooled metallic stages at higher temper- atures, due to cooling flow penalties.

O Best SFC produced by ceramic inserted turbine blade, running at 1371C (2500F) (ceramic limiting tempera- Z, ture).

O Blade height limitation of 1 cm (0.4 in.) could affect the operation before optimium pressure ratios are reached.

Zl The free power turbine design with axial components was

L

evaluated and the results are shown in Figure 139. Significant factors apparent from this data are as follows:

L

O Blade height limit restricts this free turbine design to approximately 6.5=i pressure ratio.

: [ O

The free turbine cycle SFC is better than any of the single-shaft axial turbine cycles.

![

Figure 139 also shows a simple-cycle engine with a radiala inflow turbine evaluated at three different pressure ratios. The

;[

SFC from these units is better than the axial turbine designs.

. p • 6, Jb I J • , f 0.,_7

kt

0.450 - I rEFERENCE 0.426 \ • ENGINE °.

-.--.J METAL -- 1371C f2500FI T.I.T. COOLEO 0.365 _ BLADE HEIGHT °o METAL -- 1204C 12200FI TIT, COOLEO 10.601- "_ _ METAL -- 1038C 1190OF)T,I.T, CERAMIC -- t371C I_,,oOF] TIT.

q II-0 10.551' 0.304 !

I0.501.

4 5 6 7 8 9 10 COMPRESSOR PRESSURERATIO ram11 _o ,11 APU Simple Cycles Using Axial Turbine Wheels.

Figure 138.

ated. A schrmatic Of that engine is shown i!, Figure 140. Figu:e 141 shows the projected performance of a regenerated engine scaled from the AGTI01.

i

The performance OE this engine is shown with two control m methods to achieve _rt power. The variable inlet guide vanes e o are effective in maintaining part-load SFC by redczing engine _[CI through-flow. The loss in aerodynamic efficiency i_ the compres- sor and turbine is compensated by the increased eftectlveness in the regenerator. The traditional control method of reducing tur- bine inlet temperature to reduce power output results in a signi-

ii

ficant increase in SFC. As an example, at the 65 percent power point (point 2C in the duty cycle, Table 33, which is used 57.6 percent of the time), the SFC is 25 percent higher for the reduced temperature point (Figure 141), as compared to the engine ° with constant TRIT. Based on these results, the regenerated cycle represents a definite advantage in fuel consumption over all other cycles studied.

Engine Cost Studies - Acquisition costs were estimated for the auxiliary power units studied. The basis for these estimates was I i similarity to existing metallic production components, with factors applied for complexity, machining cost multipliers, number of parts, material costs, and requirements for close tolerances or running clearances. Ceramic component costs were _4 scaled from current experience and projections on the AGTI01 engine program. The results of this analysis are shown on Figure t 142.

Engine Weight Studies - Weight estimates were completed for all configurations using similar techniques. Scaling of components provided most of the hard data. The ceramic AGTI01 components

L

were scaled. The weights used in this study are conservative from the standpoint that projected weight reductions for airborne F J

_Z

I

, "' '"_'.-..,'" _. AGE,

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[

OF POOR QU._Ey

............. ,,- ................ _/'l /

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e-4

0.280

1371C • 0 10.4611

12500F) _=

TURBINE INLETTEIWP

FORPARTLOAD

VARY IGV FOR

ii

PART LOAD

CONSTANT TIT

0.158 (0.26)

4O

PERCENT POWER,PERCENT itS-Z81 .II7 !

i: ..

Figure 141. Projected Performance fo_ APU Based on Scaled AGTIOI.

r: ;_ m

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[

.[

REFERENCE

139.9"

140"

""-'1 ,f

130-

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122.1

I"-'"-"1

120"

,.., llO-

98.9 '_ lO0-

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116-_1 .Z4 NOTE: COSTS ESTIMATEO FOR MINIMUM SFCCONFIGURATIONS *REGENERATOR COST IS 14,800 e_ L Figure 142. APU Cost Comparison.

, . • - _ w_ o.Q_4.___ _" , .o .

,b ............. a"

r

equipment have not been introduced. Moreover, the design con- cepts for brittle materials are still being developed; thus the actual production weights of a scaled AGTI01 or a simple-cycle engine with a ceramic turblne/combustor section may be somewhat :i lower than those estimated for this study. The estimated APU weights are presented in Figure 143.

• 0 _[{_ 5.2.3 Cycle/Enq [ne Selection Using the previous data in the economic model discussed in p_ragraph 5.1.5 resulted in the DOC shown in Figure 144. Th_ composite data represents a family of engines on each line that °_ have been optimized at the pressure ratio_ shown. A simplified ,a engine schematic is assigned to each line to show the cycles ..

studied. All engines, with the exception of the reference (at 1900F) and the metallic turbine (at 2200F) use ceramic turbine °, stages operating at 2500F TRIT. The direct operating cost, in dollars per hour, is shown with fuel costs of $i and $2 per gal- lon. The auxiliary power systems shown are restricted by all of

!!

the study limitations, such as: blade height, turbine tip speed, turbine inlet temperature, costs, fuel consumption, manufacturing ot

:t

capability, and maximum pressure ratio in a single-stage compres- °_ sor, etc.

,o,p From these analyses, the lowest DOC units were the simple- cycle radial component engine and the regenerated design. There- .R fore, these two cycles were selected for detailed off-design analyses in the following study tasks.

6e The simple-cycle engine uses a single-stage centrifugal com- Q, pressor at a pressure ratio of 8:1, a reverse-flow annular burn- er, and a radial inflow turbine running at 1371C (2500F). The "I description of this cycle is found in Table 37. This cycle is selected because it demonstrates the lowest DOC of the simple- 7: j. • ° , :i 1 !.

4_ I _L

]

(152J 115l_ 5J 1153.51

i

45.4 IlOOl.

I

"REGENERATOR COREWEIGHT= 17.7 KG(39 LBJ

I

65.281-79

I

NOTE:APU WEIGHTSESTIMATEDAT MINIMUM SFC

t

Figure 143. APU Weight Comparison.

#- %

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,+ t

It

ii

_4 "i !

.+ *4 FUELAI' SO.528/LITEH II2/OALI o IIEFEA[I_E _:UEL AT $0284/LITEll IIItl2ALI °* FUEL AT $2,00/QAL 80" 1371C[2500Fi oo,, 1371C,25001::1 TiT +, ,EF¢,CE ,_,_ ,Une,NES 7o, ,*+e

i°

++

REGEH REGEM T .f

,i i i fo

PRESSURE RATIO PRESSURE RATIO _1,211 a ,i+..4,

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_4 APU DOC Comparison.

Figure 144.

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I1_'__'_.._..t_:_-:._ - |It _l'_, .,. , ......... _ i I I _-- i i i iii ii i I II -- SECT SIMPLE-CYCLE APU DESIGN POINT PERFORMANCE*, TABLE 37.

6&

._ _ :J _. _. : costs continue at or below these cur-

cycle eng!nes, andiffuel

configuration would be chosen for the applica- eng_.ne at a pressure ratio oE 10 shows a lower " I tion. Although an DO(:, further analysis of compressor technology, risks, and costs 4_ concluded that a pressure ratio of I0 was not feasible for year 2000 f_r the efficiency level quoted for the small flows (0.85 kg/s [1.87 Ib/sec]). Therefore, the pressure ratio of 8 was selected.

.'# The rege,erated engine uses a single-stage centrifugal com- pressor stage at approximately a 5:1 pressure ratio, a single-can combustor, and a radial inflow turbine with a 1371C (2500F) inlet ,!

temperature. The regenerator effectiveness was 94.8 percent; the "I leakage was 3 percent.

5.3 Task III System Performance Evaluation ,bo 5.3.1 Duty Cycle Analysis Further optimization of the cycle pressure ratio, operation- al speed, and other cycle parameters to "fine tune" the two

engines to function with maximum e_ficiency during the specified [J

duty cycle was accomplished. The description of the overall engine and component performance is shown in Tables 37 and 38 for the simple cycle and the regenerated cycles.

" A preliminary analysis on a large number of engines was com- pleted considering design point fuel consumption as a selection criteria. Since the study had now been reduced to the two cycles, the complete duty cycle fuel consumption was calculated.

,1 t As a further investigation, the use of variable geometry on

i

the power section compressor of the simple-cycle configuration was analyzed. The results at part-load indicated that the effi- t- ciency penalties in the aerodynamic components (due to the .J p "b I III_ II I

il

TABLE 38. SECT REGENERATED APU DESIGN POINT PERFORMANCE*,

I

SE_,-LEVEL, STATIC, ISA, MAX UNINSTALLED

I

Overall Engine Pecformance Component Performance Compcessor

[

Engine Rating, kW 315 o WJB/_, Kg/s 1.06 (shp) (423) (iblsec) (2.33)

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!

_lt

!

I

f&

.'i '.-..

reduced flow) produced a fuel consumption that was not signifi- [.'.

U

cantly improved over the conventional temperature reduction con- _J trol concept, Therefore, the _imple-cycle engine final design

i2

did not include variable guide vanes on the compressor.

_4 The regenerated cycle showed definite Improv=d part-load

!i

fuel consumption with variable guide vanes, so this cycle was evaluated using this control concept. The load compressor for this system must also include variable guide vanes to shut off bleed flow when it is not required by the duty cycle. Further- more, a significant reduction in fuel cc;isumption is achieved b F aerodynamic improvements in the guide vanes, impeller, and dif- I .

fuser that will allow complete termination of bleed flow and its shaft power extraction when called for in the duty cycle. (Cur- .

rent systems dump compre3sed air to avoid surge and stall prob- lems).

The SFC reduction (as shown in Figure 139) can result in I significant cost savings for an operator. For example, comparing the reference engine to an advanced simple-cycle (centrifugal compressor and radial inflow turbine with SFC of 0.5 Ib/hp/hr average) APU operating 3000 hours per year, results in a fuel savings of more than 44,800 gallons. This is a savings of $44,800 at $I gallon, or $89,600 at $2 gallon, per engine per year, or a fleet fuel savings of approximately $150 to $300 million over approximately seven years.

A detailed review of initial cost, maintenance cost, and t T_ .!

weight was accomplished on both competitive cycles. In addition,

.!

i: the reference engine described in paragraph 5.1.3 was analyzed s { for fuel consumption for the entire duty cycle. The results of .

• I these analyses are presented on Figure 145. The regenerated _ .J cycle, when compared to the reference engine, can save the opera- i t tot over 47 percent ($40.30 per hour) of the DOC when fuel prices ;i are $0.528/liter ($2/gal). The simple cycle can save 38 percent ($32.60 per hour) under similar conditions.

b- ..

I P I -,-.

• _ &i i m i i

{I

5.4 Small Engine Component Technoloq7 Plan

II

As a result of the analytical effort of this study, it is apparent that in order to provide improved performance from air-

li

borne power units, a number of technologies must be addressed, and design data must be generated and verified. For examplc, several approaches to improved aerodynamic efficiency are _orthy of investigation, rn addition to the creation of component designs that are less sensitive to rotating cleavaflc_s, the pro- _J blem of clearance control for small engines must be i:_vestigated through more refined _haft dynamic analyses, therma_ analyses, dynamic effects on materials, ab:_dable shrouds, boundary layer control, more precise bearings, and static structural systems.

;J J 5.4.1 Advanced Technology Benefits The benefits to be derived from technology are discussed in 7: the following paragraphs and are shown in Figures 146 and I_7.

These figures show the relative contribution of each technology to the "all-technology" payoff for the simple-cycle engine.

n

This study for APUs shows that the largest payoffs can be obtained by establishing ceramic materials technologies for inclusion in radlal inflow turbine stages (Figures 146 and 147).

The capability of these materials to survive exposure to high- _o temperature gas streams for long durations makes them ideal for o.

use in small radial inflow gas turbines, where conventional cool- ! p oo ing techniques are difE1cult, ineffective, and costly. Silicon nitride and silicon carbide ceramic materials can be used in i high-temperature applications at 1371C (2500F). This average temperature can be achieved with a well controlled combustion t process with a maximum hot spot of 1538C (2800F).

The ceramic radial inflow turbine wheel represents a major improvement in component performance. The lower density (approx- imately one-third that of current superalloys) and the strength J

U

+ • +. _ ..... , ............

I

47.0 -q.mmZtml \\J \ \j

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39.0 %" .%_ 38.3 ..... \%!

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"%%, 32.0 31.5 \'N m o %.%, \\ 30" ",."% N\ I ', rv"J,

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I_, L'LI_ m ¢.o I _, _ ._, \\ I,M \\ z \\ 20-

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1,' 2_. \\ {,..1 C,J., I,_ I _L'u z 4:..1 \\

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\\ i,i,i _- 10- %,\ "%\ \\

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° \ \ \\ i -_ `% REGENERATED SIMPLE CERAM;CS TUR61NE COMPRESSOR LOAO

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CYCLE CYCLE EFFICIENCY EFFICIENCY COMPRESSOR EFFICIENCY IMrI-_I.IN

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.I

from APU Technologies.

Figure 146. DOC Benefits

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o .._. ;i ', ..

1]

k l _6 _J 40.3 %\ %% 32.9 '%\ %% %% o lu,.l O $0.2M/LITER 151/6AL) w , f, Jll _----_ $O.5,?.8/LITER I$2/GALi 19.0 '= IP • r"" :Z: 8.0 I--.

i L_ 18.7 X i r,,lO

=_-

m Iz 10"

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i _.,.. %% 9._ 4.4 3.7 .._ %'%

•

0 ! J_ CERAMICS TURBINE COMPRESSOR LOAO REGENERATED SIMPLE o.

[]

point increase in efficiency is projected through the improved impeller performance, use of variable geometry (inlet guide vanes and possibly a variable diffuser), and a design that will allow

{I

discharge flow near zero.

!l

The use of gas-lubricated foil bearings capable of operating in the hot environment are not absolutely necessary but will con- tribute greatly to the reliability, maintainability, a:_d opera- tot's satisfaction in the use of the APU. Problems associated i i with coking of the oil and bearing cooling requirements can be eliminated by the use of foil bearings.

Technologies such as the cold-end material programs and the =1 high-pressure seals will also contribute to the advantages of the _J year-2000 APU over current production units.

-y, / As shown in the DOC p_ojections (Figure 146), the regener- ated cycle for this application has the lowest DOC. As expected, the regenerated cycle demonstrated greater advantages in DOC as fuel price increased.

Since the ceramic materials considered for use in the year 2000 are not subject to normal fatigue-type failures, long lives can be predicted for these components. Therefore, the hot sec- tion of the engine (all ceramic) can be run at maximum tempera- _J ture continuously with the flow varied to match the required T' power output. Variable inlet guide vanes performance can provide this function. The loss in performance of the aerodynamic components at part speed is compensated for by the increased heat transfer of the regenerator over the part-power range of interest .° in this application. SFC is reduced dramatically at lower 7: powers, as shown in Figure 141. This results in a regenerator

t

-o DOC benefit of 2.3 percent at $1/gal and 8.770 at $2/gal fuel cost.

7j

o .. q- ., - - J 5.4.2 Technoloqy Plan GTEC's recommended plan for small engine component technolo- gies _or year-2000 APUs is presented in this section. This plan addresses seven technologies in keeping with the technology bene-

¥

i

fits presented in paragraph 5.4.1. The technology categories are tabulated below, along with their respective paragraph numbers.

will result in uncooled operating capability with turbine inlet

;I

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I, 011 Z 7 0

-?

m

I

CERAA_IC$ FOR COMBUSTORS ANO TUROINEOLAOI:$ ANO VANES • SEE PARAGRAPH Z.4.Z.I, TECHNOLOGYPHOGIIAN il CA CERAMICRAOiALTUrBiNE WHEEI/STATOR o_ • EXPERIMENTAL MATERIAL/HARDWARE PROGRAM FOR INTEGRAL MONOLITHIC CERAMICWHEELSIZEDFOR SECTSIMPLE CYCLEAPU ,;A • MATERIAL/PROCESS OEV. OESIGNDATA • FABRICATION TECHNOLOGY VERIFICATION • COMPONENT FABRICATION !, t, • COMPONENT OENCH/HOT RIG TESTING C8 ENGINEENVIRONMENT TEST OF CERAMICRAOIALWHEEL _e • EVALUATE TURBINEON TEST-BEO ENGINE • ENGINEAOAPTIVE HARDWARE OES/FAO

• EVALUATION TURBINETESTS 11

It

I NASA/DOECERAMICS :NITIATIVE I _8 i _8 SECT TECHNOLOGY VERIFICATION _., 18.ml.lill Ceramics Technology Schedule.

Figure 148.

6&

L.s i '| temperatures up to 1371C (2500F). By eliminating the wheel and _ stator cooling requirements and operating at 1371C (2500F), sig- " '_ nificant improvement in SFC and power density can be achieved " over current APUs Additionally, because the candidate ceramic I I material, SI3N 4, is approximately one-thlrd the density of super- / ! alloys typically used for turbines, some weight reduction can be j realized. Finally, by eliminating the need for complex cooling _ schemes and fabrication approaches that would be required for metal turbine components, a potential for cost savings exists with the use of ceramic parts.

Program OescrL, pttcn This program consists of four major tasks, from material development through hot rig testing, as scheduled in Figure 148.

The initial activity will be to configure a turbine wheel and matching stator that are representative in size and shape to

II

the SECT APU engine as conceptualized for the year 2000. These designs will be based on known property values of the candidate ceramic materials. Baseline material testing will also be per- formed to provide additional design data as required. Design of the adaptive hardware and the rotor attachment scheme will be connected as important facets to assure minimal stressing of ceramic components due to distortion and contact loading.

Ceramic component fabrication process development will emphasize the rotor development, addressing the size and volume influence, dimensional tolerances, and component material proper- ties. Stator fabrication will be performed in parallel, but is

I anticipated to require significantly less effort due to existing

I

ceramic (monolithic and/or composite) materials fabrication pro- gram as described previously for Program CA. This program will I also build on advanced turbine aerodynamic design technologies.

Program Description

,|

This program consists of two major tasks, as scheduled on

I

Figure 148. The major activities of this program consist of the design and fabrication of test-bed engine parts, and engine

-I

environment testing at 1371C (2500F).

.!

'I

O

J i ,, , : : : °; -- , k Design - Specifications for the test rotor will be selected,

il

and aerodynamic and mechanical design analyses will be conducted to configure the rotor and a matching ceramic stat0r. Design- board layout and detail design will be conducted through prelimi- nary and final design phases. Design activities will include integration oE the ceramic stage into a suitable turbine test rig 4J and into a test-bed engine.

Fabr[catlon - Test parts and spares will be fabricated/pro- cured for test-bed engine parts.

o_ Enqine Environment Tests

;I

_0 The ceramic turbine stage and adaptive parts will be installed in the selected test-bed engine and will be run at 1371C (2500F) test conditions. Performance and mechanical data will be obtained and compared with design predictions.

Technical Approach Qe Specifications Err a turbine stage will be established that are representative of the SECT APU requirements (simple cycle) as envisioned for the year 2000. The specification will also be compatible with the selection of a suitable engine test bed for _d e_gine environment tests.

oy ,j Analytical design of the integral ceramic radial turbine rotor will draw from the latest technologies available for aero- dynamic and mechanical design, materials properties, and fabrica- tion/manufacturing processes. The mechanical design will accept risks commensurate with performance benefits and will be based on i the available material properties as established for the pilot rotor.

J J

!

5.4.2.2 Regenerated Enqine TechnoloqLes

$

These technologies will impact APU design options in the year-2000 time frame. For SECT planning purposes, it is envi- sioned that the regenerator technology will be advanced by separate initiatives such as the NASA/DOE-sponsored programs for the AGTZ01 and/or related programs.

J

5.4.2.3 Radial Turbine Performance This section presents the plan for radial turbine perfor- mance technologies as applicable for year-2000 APUs. The plan Ls

I

made up of two discrete technology programs identified as Pro- grams CC aod CD in ?_qure 149.

CC. Radial Turbine Inlet/Stator ;-rEormance Improvement

I

Current radial turbin÷ fJPslgn practices concentrate design efforts to optimize the stator and rotor, while less attention is paid to the combustor/turbine transition duct. Historically, an aerodynamically configured scroll has been used upstream oE the stator to achieve uniform flow. Future designs will be more size-llmited due to the desire to increase APU power density. As the acceptable size of the diameter of the turbine decreases, the transition duct must become more compact. Decreasing duct size f implies tighter turning (Erom axial-to-radial) and, perhaps, higher Mach number levels. Thus, the performance of the transi- tion duct becomes a critical issue, both from a total-pressure loss standpoint and from a stator inlet flow uniformity concern.

Radial turbine performance could be improved from current levels by designing the inlet duct and stator as a system using .,4 advanced 3-D, viscous analyses. Moreover, the possibility of reducing envelope size at current performance levels is present.

J

1 i.| i

I1

ii

t Ibe

:!

_s PLAN YEAR 45'67891 10111 I_ 14 U II2 3 ,!

flAOiAl,TURBINEINI.ET/STATQRPERFORMANCE _a CC IMPROVEMENT • ESTAilLISX ANO VERIFY AN AOVANCEO S-O VISCOUS 41 FLOWCOOE FOR OESIGNANO OPTIMIZATION OF INLET I.

_J DUCT/STATOR CONFIGURATIQNS • COOEMOOELING/VERIFICATIQN _ _ • INLET/STATOR DES/FAR '_/_

ii

• AERO RIG TESTS CD i PERFORMANCE TECHNOLOGY FOR BACKWAIIO_URVED RAOIAL TURBINEWHEELS _4 • AERODESIGN ALTERNATIVE TQ ROTOR ENTRANCE iNCIOENCE LOSSESFOR TIP-SPEEO-LIMITEO TURBINES • TURBINEOES/FA8 _ • AERO ICOLO)RIG TEST ._ SECTTECHNOLOGY VERIFICATION .-_o, f, t_ wt A.

i.s -,- 'T,T !

APU Radial Turbine Technology Schedule.

Figure 149.

|

J ,.... _'_ c _:._ :_i _': "__'i¸_ ................ = c:_ ...... ...... " " i" _ This SECT program will create the tools necessary to design optimized inlet/stator conflgura_ions. Actual designs will be I generated and tested to calibrate the new codes. The result will be a design system capable of producing radial tu=bines having enhanced performance levels.

Proqram Description This program consist_ of three major tasks, from code model- ing through areodynamic rig tests, as scheduled in Figure 149.

Two computer codes will be generated to analyzp the inlet/ stator system. Since an iterative process is anticipated, these codes will be written so that file transfers will be easy and

I

thorough.

I

The first code will be the stator geometry generator. This code must be capable of designing stator sections along stream-

I

lines having arbitrary orientation. In addition, the code will have stacking capability to create the full stator 3-D geometry from the several design sections.

I

The second code will be the 3-D, viscous geometry analyzer.

Its input will be a file of geometry from the geometry generator code, plus the applicable flow-field boundary conditions. At this point it is assumed that one or two stator passages, but not all of them, will be analyzed. The code is envisioned as being !

something llke the Denton 3-D code modified to include an appro- priate boundary layer analysis.

AnalTsls and Desiqn - Analysis and design will be conducted to choose a suitable existing system and analyze it with the new codes. This will constitute the baseline configuration, which will then be redesigned to obtain the maximum performance

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ll,41 • o possible. The only constraint will be that it does not violate

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the maximum diameter Of the baseline design.

\, A second redesign will be acccmpllshed to minimize the sys-

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tem maximum diameter. A Judgment will be made as to defining i , this diameter; for example, whether it is minimized when the sys-

I1

tem losses are equal to the baseline system losses. It could also be based on equivalent engine power densities.

The choice of these two redesigns is due to the desire to bracket a reasonable range of design constraints and parameters.

i

eJ Fabrication - To validate/calibrate the design codes, test-

:j

ing of the system redesigns is essential. The test vehicle will be the baseline system cold air test rig modi[ied to accept the redesigned systems. Adequate instrumentation will be included ,ll fOr obtaining detailed flow-field characteristics. The requiced new rig hardware will be procured.

Rig Tests - Two types of activities are performed in this

II

will be three-dimensional. It is likely that, for axial flow at the duct inlet, the stator meridional flow path will nc. longer be

[ purely radial. It may be beneficial to do some of the meridional

turning, presently done in the duct, in the accelerating flow field of the stator. This would lead to a "mixed flow" type of

[

stator design.

t CD. Performance Technoloqy for Backward-Curved Radial Turbine Wheels

[

As the work requirement of a radial turbine increases, so does the inducer tip speed. When the tip speed requirement

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-- ° ) b surpasses the wheel material capablllties, the designer is faced wlth trading high Mach number and exLt swirl (downstream losses) for flow entering the rotor at- positive relative angles (inci- dence losses). Typically, the lesser penalty is due to inci- dence, so the design becomes a tip-speed-limited configuration.

These incidence losses, however, can be two points (or higher) for high-work turbines.

-0 The SECT APUs of this study did not suffer from this trait because oE the projected (year 2000) ceramic capability oE tip speeds of up to 9144 m/s (3000 ft/sec). This tip speed is ade- *i quate for turbines having pressure ratios up to lO:l. An

l

approach to improve turbine performance prior to that time, which °4 could also be applied to the SECT ceramic turbine, is presented here.

"4 This program will advance the concept of the backward-curved

.I

"radial" turbine blade, a concept that relaxes the radial con- straint and allows the inlet portion of the blade to have back-

!

ward curvature (this is routinely done in compressor impellers).

,,a The nonradial inlet alleviates the incidence problem but creates higher stress levels. Considering the progress being made in dual-alloy wheels and the improvements in ceramic materials, these higheL stress levels may soon be tolerable. This means higher performance for the majority of recent radial turbine designs.

..4 -T Program Description This program consists of two major tasks, as scheduled i_ Figure 149.

potential advancement for improved perfor_.ance.

Looking farther ahead, this concept is also a possibility with ceramic radial turbine wheels. An important parameter that allows the use of backward-curved blades is the material strength-to-weight ratio. Advanced ceramic material property projections imply higher ratios than current metallic properties.

Thus, the concept could be proven and used for immediate perfor- mance improvements and be applied in future ceramic designs as well.

5.4.2.4 Combustor Technologies These technologies must be advanced for year-2000 APUs.

GTEC envisions a requirement for ceramic reverse-flow combustor technologies as discussed for the rotorcraft engines in paragraph 2.4.2.5, Technology Programs O and P.

5.4.2.5 Compressor (Centrifugal) Performance " This section presents the plan for performance improvements to centrifugal compressors as applicable for year-2000 APUs. The 8- &.

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................. .. ,, ,,'- ............ . _ ....T . J[ ;t,r RL m ussed _or the cctorcraft and commuter engines, and two ' ' III grams Q, R, T, and U,

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i" tm Program CE, whLch addresses desLgn t_chniques for sn_all APU

!

_. compressors, is planned to follow the first-stage Performance i' Optimization Program (R) to maximize the benefits for thi_ plan.

t I Similarily, the load compressor program (CF) is scheduled to tak_ full advantage of the basic compressor programs and Program CE for best plan results.

CE. Small _APU} Compressor Design Techniques

!

I The one-stage centrifugal compressors proposed for the SECT

APUs are characterized by high-pressure ratios (8 to i0:i) and low flow rates (0.45 to 0.91 kg/s) (I to 2 ib/sec). A major

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challenge in meeting the performance goals of these compressors is overcoming the restrictions that current materials and manu-

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facturing techniques impose on the aerodynamic design of compres- sors of this size. GTEC expe_ience with scaling an 8:1 pressure

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ratio from 11.34 to 0.91 kg/s (25 to 2 ib/sec) produced greater performance decrement than could be accounted for by Reynold's

[

number effects and increased parasitic losses. This shortfall is apparently because minimum blade thicknesses do not scale, which causes increased blade blockage. This is a critical performance

, (

parameter with the high Mach numbers that high-pressure-ratio machines require. Also, current manufacturing techniques dictate that blade count be lower than that for a full-sized compressor.

This SECT program will examine the design tools and techniques used for centrifugal compressors, and will modify or develop new +_,._ ! l .

ones for use in small, high-pressure-ratio compressors.

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ANALY'---'--_ICAL MODELING AN VE I ICA ION AXIAL/CENTRIFUGAL COMPRESSORS • Ill PAAA_IRAi_N Z4.Z,ll. TECNNOLOtlY PHOGRAM 0 PERFORMANCE _PTIMIZATISN OF IMPELLER/OIFFUSERI -6 X-OVER FOR A FIRST STA2E CENTRIF • SEEPARAGRAPH 2.4.2.8,TECHNOLOGY PROORAM II CENTRIFUGAL COMPRESSOR CASINO TREATMENTS • lEE PARAGRAPH Z,4.Z.G, TECHNOLOGY PROGRAM T IMPROVEO OIFFUSION SYSTEMS • I1[| PIIRAGRAPH 1,4.ZJI, TECHNOLOGY PROGRAM U ¢E SMALL (APUI COMPRESSOROESIGN TECHNIQUES

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• MOOEL ANO VERIFY COMPUTER COOE TO OESIGN/ OPTIMIZE SMALL/HIGH.PRESSURE,RATIO COMPRESSOR "t STAGE FOIl flIGH 6LAOE BLOCK, AGE ANO LOW SLAOE COUNT

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• COOE MOOELING/VERIFICATION • BAli COMPRESSOR RIG TESTS • COMPRESSOR OE|/FAil

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• AFRO RIO TESTS CF VARIABLE GEOMETRY FOR APU LOAO COMPRESSOR • EVALUATE THE CONCEPT OF VARIABLE INLET GUIOE VANES + VARIABLE OIFFUSER VANES FOR SURGE-FREE RANGE • STAGE OES/FAD • AFRO RIG TEST °!

SECT TECHNOLOGYVERIFICATION # ml,141 !

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Figure 150.

based on the new model and baseline compressor data, will be designed.

Aero Rig Tests - A test similar to that run on the baseline compressor will be conducted with the redesigned compressor.

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Extensive Elow-field measurements using the LDV will be obtained.

These will also be compared to the model results as a proof-of- concept.

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_ .... ,+,¢n+.+,_,.,_-.,t-+..,,+._.m, ..- :-t ,i, - , ' - , .... " . ,+¢ 'i.v ,:"J_l,"_l_+.g,--'"--"'+' _,P+_'._,-,,e ?,,_-+ -_,.. ,, :.+..+,,.+,.+.+ , .. + , ++,, ++ .+, •. ..

I Approach ,2:I Although centrifugal compressor flow fields have bep.n m.-.,a- .-!

,_ I sured before, they have usually been for large lower-speed impel- t lets where the viscous and Hach number effects that will be pre- i I P valant in the SECT APU compressor are minimal. The data obtained I in Task I will be important to the development of the new model.

t This model mu_t be able to account for the 3-D compressible vis- cous flow. Since the thick vanes and high Mach numbers of small high-pressur_ .ratio compressors are conducive to shock loss,

]

attempts will be made to apply some of _:he shock-free design techniques for axlal-flow compressors to centrifugal.

CF. Variable Geometry for APU Load Compressor

]

One reason that SFC is compromised in gas turbine APUs is that typical load compressors do not possess sufficient range to

]

allow low-load operation without employing an inefficient surge valve. To circumvent this problem, the load compressor proposed for the SECT APU will employ variable geometry in the form of variable inlet guide vanes and variable radial diffuser vanes in order to provide additional compressor range and eliminate the need for a surge valve.

This technology program covers the design and evaluation of the variable geometry centrifugal compressor stage to be used as the load compressor for the SECT APU.

Program Description This program is scheduled as two major tasks, as shown in Figure 150.

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i Various variable diffuser concepts will be evaluated and a

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final concept will be detailed. Compressor performance and range estimates, including the effects of variable geometry, will be

|

made. The compressor will be mapped at various IGV and radial diffuser vane angles to measure the variable geometry concept.

!

Technical Approach In order to obtain the gains in SFC that are available from

i

a variable-geometry load compressor, two go_Is must be accom- plished. First, the compressor must have considerable range and high efficiency at its nominal setting. This will allow the compressor to be matched at peak efficiency at full load and will minimize the efficiency loss that will occur at part-load and at

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no-load conditions. Even though the load compressor will produce considerably lower pressure ratios than the engine compressor (4 to 5:1, as opposed to 8 to I0:i), design tools based oa engine compressors should assure that this goal is met.

The second goal is that a reliable, practical, producible mechanical actuation system must be developed to vary the radial

[

diffuser vane angle through sufficient range with a minimum end- wall leakage. R&D efforts are currently under way at GTEC to

[

evaluate and optimize several candidate concepts.

Load compressors have unique complexities that engine com- pressors do not have. This includes the requirement for some sort of scroll to transfer bleed air to its intended area of use.

Also, load compressor power absorption must be reduced to as low [ a level as possible for quick starting of the power section of the gas turbine with the smallest, lightest starting system pos- . range of conditions, load compressors also challenge the _ [ sible. Because of the requirement to operate over such a wide 7" 7- - "_ ' _{'_

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• o r r mechanical design to provide for vibration resistance under situ- ations where significant rotating stall can be present. The technology program that addre;ses the problems of high-perfor- mance centrifugal load compressors with variable geometry must address all of these operating characteristics.

5.4.2.6 Materials for "Cold" Parts These technologies can benefit APUs in much the same manner as discussed for the rotorcraft engines in paragraph 2.4.2.?, Programs V, W, X, and Y.

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ol 5.4.2.7 System Technologies

:]

This section presents the plan for system technologies as applicable to year-2000 APUs. The plan consists of two discrete

j

technology programs, one as previously discussed for the rotor- craft engines, and one new program identified as CG on Figure 151.

The first program addresses the requirement for noncontact face seals as discussed for rotorcraft engines in paragraph 2.4.2.8, Technology Program Z. This program is an integral part of this plan as well, since the requirement is projected for year-2000 APUs. The second Program (CG) is dedicated to future

;!

le APUs.

CG. Gas Lubricated Foil Journal Bearing (for APU) Future trends for high-performance and high-power-density APUs will impose severe requirements on the turbine end bearing.

Projected APU engine cycles will feature 1371C (2500F) turbine inlet temperatures through extensive use of ceramic components.

High rotational speeds and compact packaging are also envisioned for these units. The gas-lubricated foil Journal bearing offers N the potential to meet these future gas turbine requirements.

J 'T ,, 'II i II I I - I II

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PlAN YEAR NONCONTACT FACE SEALS • |EE PARAGRAPH 2.4.2,8, TECHNOLOGY PROGRAM Z CG GAS-LUBRICATED FOILJOURNALGEARING [FORAPUI • FOIL BEARING MATERIALS ANOTEST EVALUATION FOR HIGH.TEMPERATURE SECTENGINEWITH HIGH ROTATIONAL SPEEDSANO COMPACTPACKAGING • ANALYTICAL MOOELING/CODE VERIFICATION • COATINGEVALUATIONS/TESTS • BEARING OES/FAB • GEARING RIG TESTS • ENGINEENVIRONMENT TESTS SECT TECHNOLOGY VERIFICATION g1.11 151. Systems Technology Scheme Eor APU.

Figure

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The anticipated extended operating conditions, under hlgh distortion and misalignment, are: a O 816C (1500F) Coil temperature O 10,470 rad/s (100,000) rpm speed O 34.3 N/cm 2 (50 psi) unit load capacity These requirements can be compared to current EoLI Journal bearing load capacity limits under ideal rig test conditions of 20.7 N/cm2 (30 psi) ,:nit load capacity for temperatures up to 649C (1200_) and 34.3 N/cm 2 (50 psi) for temperatures up to 37_C

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{700F). Rotor speeds fo_ theno tests were less than 5235 rad/s {50,000 rpm).

A schematic o_ the current GTEC loll journal bearing design is shown in Figure 152. The design features overlapping foil leaves supported by flexible backino springs. A thin antlfric- tlon coating on the foils allows self-startlng capability. While

]

this basic design provides a baseline configuration from which design iterations can proceed, it is subject to modification.

]

Proqram Desclptlon This program consists of four major tasks, from analytical modeling through engine environment tests, as scheduled in Figure

;]

151.

The requirement for high load capacity at high temperatures t will require a fully coordinated effort between bearing and coat- e_ ing development. The foil bearing materials and coatings will also require development to withstand the higher operating tem- peratures achievable with ceramics.

_j _A The requirement to define alternate materials may necessi- tate changes to the basic geometry of the bearing. While a m" (4 Figure 152. GTEC Self-Acting Foil Journal Bearing Concept With Backing Spring.

I

J " *| _ 0 coupled elastohydrodynamic analysis is still expected to be valid, program modifications will be required to reflect the new geometry of _he bearing.

Corresponding to the trend for higher power density engines, rotor speeds and temperatures will increase. While the high sur- face speed of the bearing Journal must be taken into account in the bearing design, the tendency of rotor configurations toward supercritical operation and overhung bearing supports is of more _8 concern. The support characteristics of the foil bearing will be _g evaluated to ensure proper bearing operation under these extreme _e dynamic operating conditions.

Due to the complexity of the foil bearing configuration, these dynamic performance parameters can best be obtained by r_g testing. A foil bearing test rig specifically designed to evalu- ate bearing dynamic parameters will be used to assess the follow- ing quantities: Cross-coupling stifEness coefficients Bearing damping characteristics Power dissipation Dynamic orthogonal stifEness coeEEicients The bearing design will consider the load degradation influ- .we ence of thermal and mechanical distortion and bearing misalign- o.

ment. An integrated design analysis will be conducted to supple- ment the bearing design and optimization procedures. The analy- sis will consist of: Bearing elastohydrodynamic analysis i Rotor dynamics analysis Secondary flow analysis _t Thermal analysis oe Structural analysis

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! These individual analyses will be fully iterated with comprehen-

sive rig testing to evolve a viable foil bearing design.

!

Technical Approach To achieve high load capacity at elevated temperatures, the appropriate foil/Journal coating combination must be identified.

A dedicated coating evaluation program will be initiated at the outset of the program to enhance the possibility of defining an

[

acceptable coating pair. rt is expected that, b[ the start of this program, a considerable body of knowledge regarding F_e- viously tested coatings will exist. This program will no_ be limited to screening existing coatings, but will include the development of new coatings.

As coatings and materials for this application are selected,

I

bearing design activities will begin. In concurrence with these materials, compatible foil bearing designs will be synthesized.

This will require modifications to the elastohydrodynamic analy.

sis to treat these new bearing geometries. Bearing design activ- ity will continue to be iterated with bearing performance test results.

L

Test rigs will be required to evaluate the various parame- ters required for complete evaluation of the foil bearing. Due to the complexity of attaining some of the dynamic bearing char- acteristics, more than one test rig may be required. The design of these test rigs constitutes an important step in ensuring the proper evaluation of the iterated bearing designs. The test rigs will evaluate the following bearing characteristics: 0 Static load-deflection curve o Static breakaway torque o o Actual sway space

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"I _i;:_&Z%__-.." .... _.

B o Bearing bottoming load

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o Bearing hysteresis o Dynamic orthogonal stiffness o Dynamic cross-coupling stiffness ¢ o Rotor bearing dynamic response o Load capacity

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o Power dissipation Many oF the preceding characteristics are routinely evalu- ated on existing Fell Journal bearing tes_ rigs. The complete z i itemization ot desired bearing characteristics are the result of extensLve testing and evaluation of bearing/rotor system dynam- LCS.

The foil Journal bearing application to the hostile turbine- end environment of the advanced APU will be a success-oriented program based on extensive GTEU experience. The first demonstra- tor JFSI00 jet fuel starter was successfully operated in 1973.

That unit used both fell journal and thrust bearings to complete- ly eliminate the oil lubrication system. Since then, turbine-end foil bearings have been successfully operated in the JFSIg0 jet fuel starter, the GTCPI65 APU, and are incorporated in the AGTI01 advanced gas turbine.

5.5 Summary As determined by the mission analysis, the advanced engines achieve fuel burn reductions of 43.2 and 70.8 percent, respec- tively, for the simple and regenerated cycles (Figure 153).

.r 'I Despite the fuel burn advantage of the regenerated engine, how- ever, the correspondipg size, weight, and cost advantage of the 4m_ simple cycle result in nearly equal DOCs at the low fuel price.

At the high fuel price, however, the lower mission fuel require- ments of the regenerated engin_ translate into a significant DOC

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by 47 _ercent relative to the re_erence engine, compared to 38.J percent _or the simple _ycle.

i These stgn£_Icant Impcovements depend on advanced materLals, compone_wt performance, and system technologLes.

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*Low fuel price = $0.264/liter ($1/gal) ; k. [ High fuel price = $0.528/liter ($2/gal) **JP-1O fuel price = $2.64/liter ($10/gal) .(

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,,L O At low fuel prices ($1/gal), both simple and heat- recovery cycles will be competitive in terms of b.

DOC.

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At high fuel prices ($2/gal), heat-recovery cycles will have a definite DOC advantage over simple cycles.

. _J Significant DOC reductions are possible from the pro- Jected year-200 engine technologies.

O Rotorcraft - Fuel burn reductions of 22 and 42 percent are projected for simple and recuperated engines, respectively. This translates to system DOC reductions of 7.0 and 7.4 percent at $1/gal.

At $2/gal, DOC reductions of 8.7 and 11.4 percent are predicted for the simple and recuperated engine_ respectively.

O Commuter - Projected commuter benefits will be similar to those for the rotorcraft.

_e _e O AP._UU - Simple and regenerated engines are projected to achieve fuel burn reductions of 43 and 71 per- cent, respectively. At $1/gal thiz translates to DOC reductions of 37 and 39 percent. At $2/gal, -!

oa simple and regenerated engines are projected to reduce DOC by 38 and 47 percent, respectively.

_J Cruise Missile Enqines e Limited-diameter, low-volume engines are required for this application to achieve interface compatibility ,b4 with launchers and to facilitate increased missile fuel r | fraction for increased mission range. There are no known gas turbine engines of current technology that q are suitable for the cruise mLssile of this study.

A year-2000 mlsslle-tu£toJet system is projected to increase the mission range, compared to current rocket-powered systems, by appreximately 200 per- cent (3i5 percent with slurry fuels). This system is projected to Increase the _Issile range by 32 percent, compared to a near-term (1989) reference turbojet-powered missile.

| 6.2 High _ayotf Technoloq[e_

il

The results 0_ this study show that technology planning for year-2000 engines should support both slmple-cycle and heat- recovery-cycle engines for the rotorcraft, commuter, and APU applications. Technology planning for cruise missile engines should address extremely compact, high-temperature turbojets.

I

Rotorcraft r Commuter t and APU Engines I , Ceramics for high-temperature combustors, t-rbine blad- ing, and/or rotors will dominate engine cycles and con- figurations.

i'

Ceramic heat recovery devices will be needed to meet the probability of rising fuel costs.

t

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0 Ceramic recuperators are envisioned for rotorcraft and commuter engines.

D I

k i .

o Ceramic regenerators, similar to the disk type incorporated in the NASA/DOE-sponsored AGTI01 vehicular engine, are envisioned for APUs.

Metallic technology advances will be required for tur- bine disks and blades and shafting to facilitate increased blade speeds, rotational speeds, and engLne weight and cost reductions.

, Combustor technologies will be required f_r high tem- peratures (_427C [2600F]) and heat release rates, with emphasis on pattern factors _0.12.

Advanced turbomachinery components are vital to these

q

high cycle-pressure-ratio (simple cycle) and tem- perature-ratio engines.

o System technologies will be required to meet the mechanical and packaging challenges of these compact engines with high rotational speeds. Advanced seals and lubricants will be required. Foil journal bearings will b" 5eneficial for small, compact APUs.

Cruise Missile Engines 7_ Coated carbon-carbon materials for hot-section parts show the most promise.

Advanced metallics for high-temperature, axial-compres- sor stages will be required to meet the requirements of high flight Mach number and high cycle-pressure-ratio.

Combustor technologies will be needed for extremely high temperatures" (1927C [3500F]) and heat- D w k ' , ,,. I '.P r%:

!

release-rates, with emphasis on Improved pattern fac- tors.

10. Advanced turbine and compressors will be required.

Multistage axial compressor technologies _or low weight/cost, with minimum variable geometry, will be important. Non-straight-llne-element turbine bladli_

,j

(coated carbon-carbon) will be important for acceptable

!

turbine performance levels.

ii. System technologies will be required for compact thrust nozzle systems. High-temperature accessories of low cost/weight will requi_e new advanced technologies.

I

.I

i

mQ.-- ................................

.o | I'4 REFERENCES "User's Manual _or HESCOMP - Helicopter Sizing and Computer Program," Boeing Vertol, NASA CR-152018, September 1973, Revised 1979.

a - " GTEC NASA CR- "Weight Analysis of Turbine Engine Small, 165U49.

W , "Parametric _tudy of Helicopter Aircraft Systems Costs and

j

Weights," NASA Report No. 0-22305.

J • "Application oE Advanced Technoloo[es to Small, Short Haul

: J

Transport Aircraft" (STAT), NASA CR-16561C.

, "Advanced Material Technology Candidates Eor the 1980s," NASA CR-165176, August 1980.

I

, Humphreys, J.R., "Why So Few All-New General Aviation Air- craft," Society Of Experimental Test Pilots Technical Review, Vol. 12, No. 3, Spring 1975, pp 43-50.

Wisler, D.C., "Core Compressor Exit Stage Study," Volume II, _ j 7 NASA CR-159498, November 1980.

Rechter, M., P. Schimming, M. Starken, "Design and Testing of Two Supercritical Compressor Cascades," ASME Paper 79-GT- iI, March 1979.

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Next: APPENDIX A →

APPENDIX A

i . ,j'] , L I APPENDIX A LIST OF SYMBOL5 .,,, ._: -,A_,'._ p'_.,%NK I_IOT FILM.I_I ° PRE"CED;f_G PAGE BLANK NOT FILMED t_'_,_,t_ .... _.'T_.:_"_, -¸ . ' "_ .,.

LIST OF SYMBOLS ,,. Symbol AN 2 Turbine Annulus Area (in 2) Multiplied by Rotor RPM Squared C-C Carbon-Carbon Combustor Diffuser Interaction -0 Constant Mean Diameter CME Cruise Missile Engine -a COD Constant Outside Diameter CPR Compressor Pressure Ratio CST Constitutional Solution Treatment Rotor Corrected Thrust Coefficient Ct/Sigma DN Bearing Mean Diameter Times RPM DOC Direct Operating Cost DS Directionally Solidified E/0 Elastic Modulus to Density FOD Foreign Object Damage gJ_H/UM2 Turbine Stage Mean Work Coefficient HESCOMP Helicopter Sizing and Performance Computer Program ,oo HHR High Heat Release HIP Hot Isostatic Pressing ,,a HOGE Hover Out of Ground Effect IGV Inlet Guide Vanes ..a IPS Inlet Particle Separator IRP Intermediate Rated Power r__. ' "o , J i i i -- LIST OF SYMBOLS (Contd) Symbol ISA International Standard Atmosphere K Constant LART Low Aspect Ratio Turbine LDV Laser Doppler Velocimeter L/O Li_t/Drag Ratio L2F Laser. Two Focus LHV Lower Heating Value

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i LPT-NOZ _P/P Low-Pressure Turbine Percent Pressure Drop

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MMC Metal Matrix Composite

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MN Mach Numb3r MPa MegaPascals - N/m 2 x 10 6 Nondestructive Evaluation NDE High-Pressure Spool Physical RPM NHp Low-Pressure Spool Physical RPM NLp NR_,M Inlet Recovery PF Pattern Factor PFC Planar Flow Casting High Pressure PH LOW Pressure PL PM Powder Metal

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PMG Permanent Magnet Generator P/p Pressure Ratio Across the Axial Compressor

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LIST OF SYMBOLS (Contd) Symbol, Pressure Ratio PR RCS Radar Cross Section

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RSP Rapid Solidification Process SFC Specific Fuel Consumption SLE Straight Line Element SLS Sea Level, _tatic QJ Rotor Blade Area/D_sc Area Sol idity TOGW Takeoff Gross Weight T/J Turbojet TRrT Turbine Rotor Inlet Temperature TSFC Thrust Specific Fuel Consumption T-T Total to Total VBest Range Velocity at Which the Aircraft's Range is at a Maximum Reference Velocity VREF Corrected Airflow w/e/6 WATE-S Weight Analysis of Turbine Engine - Small Turbine Cooling Flow (Percent of Core Flow) Wcool Turbine Efficiency nT P ADOC Change in D1rect Operating Cost K Heat Exchanger E_fectlveness Turbine Corrected Work _x/e

Overview

Part--Power Performance Comparison for the View in PDF

Reference Missile Configuration with Turbojet i

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I technologies for year-2000 engines and to provide companion tech-

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I Year-2000 aircraft and missions as well as performance

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I tical cruise missile_, and auxiliary power units (APU). Parallel

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I Operating costs of aircraft with selected year-2000 engines

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Part--Power Performance Comparison for the