Making Turbofan Engines More Energy Efficient

78-GT-198 Copyright 1978 by ASME $3.00 PER COPY $1.50 TO ASME MEMBERS 1 00 I The Society shall not be responsible for statements or opinions advanced in papers or in discussion at meetings of the Society or of its Divisions or Sections, or printed in its publications. Discussion is printed only if the paper is published in an ASME journal or Proceedings. Released for general publication upon presentation. Full credit should be given to ASME, the Technical Division, and the author(s). Making Turbofan Engines More Energy Efficient M. C. HEMSWORTH Mem.ASME M. A. ZIPKIN General Electric Co., Cincinnati, Ohio A review of transport aircraft gas turbine engine development and evolution during the past two decades is presented in terms of energy consumption. The interaction and effects of cycle pressure ratio, firing temperature, bypass ratio, and component efficiencies on installed fuel consumption are reviewed. The possibilities for further substantial improvement in energy efficiency with improved operating economics and with improved environmental characteristics are identified and evaluated. Parametric data are presented showing trade-offs in the areas of efficiency and economics. Environmental considerations are also discussed. The balance of these factors in a cost-effective advanced turbofan is discussed. In conclusion, projections are made for the capability of an advanced turbofan engine compared with the goals established by NASA for their Energy Efficient Engine Program. The characteristics of this more efficient, cost-effective power plant, that can be operational in the late 1980's, are shown in relationship to current turbofan engines. Contributed by the Gas Turbine Division of The American Society of Mechanical Engineers for presentation at the Gas Turbine Conference & Products Show, London, England, April 9-13, 1978. Manuscript received at ASME Headquarters January 13, 1978. Copies will be available until January 1, 1979. THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS, UNITED ENGINEERING CENTER, 345 EAST 47th STREET, NEW YORK, N.Y. 10017

Making Turbofan Engines More Energy Efficient M C. HEMSWORTH, M. A. ZIPKIN ABSTRACT A review of transport aircraft gas turbine engine development and evolution during the past two decades is presented in terms of energy consumption. The interaction and effects of cycle pressure ratio, firing temperature, bypass ratio, and component efficiencies on installed fuel consumption are reviewed. The possibilities for further substantial improvement in energy efficiency with improved operating economics and with improved environmental characteristics are identified and evaluated. Parametric data are presented showing trade-offs in the areas of efficiency and economics. Environmental considerations are also discussed. The balance of these factors in a cost effective advanced turbofan is discussed. In conclusion, projections are made for the capability of an advanced turbofan engine compared with the goals established by NASA for their Energy Efficient Engine Program. The characteristics of this more efficient, cost-effective powerplant, that can be operational in the late 1980's, are shown in relationship to current turbofan engines. INTRODUCTION The fuel consumption per pound of thrust of US commercial jet engines improved in two major steps since the first jet powered US commercial aircraft in the late 1950's. An improvement of 15% was obtained when the first low bypass turbofans were introduced in the early 1960's and the high bypass fans in the early 1970's were in turn 20% better than the low bypass fans. These improvements in engines together with major improvements in aircraft efficiency have been the product of competitive pressures to produce more economical and productive aircraft. Now the urgent realities of the world energy problem require major efforts to further improve engine efficiency to conserve energy in the National interests. This next major step in power plant efficiency improvement must also improve operating costs to provide the incentive for its use in future transports. 1 Mgr - Energy Efficient Engine Program, Fellow ASME 2 Gen Mgr - Advanced Engrg & Technology Programs Following the world fuel crisis in 1973, NASA initiated and funded a succession of important studies to identify advanced engine systems and technology that would provide the basis for another substantial step improvement in aircraft engine fuel efficiency relative to current high bypass turbofans. These studies built on prior company work and concentrated on the payoff of advanced material technologies and on determining the most suitable and efficient engine configurations and cycles for commercial transport applications. Two materials studies were described in References 1 and 2 while cycle and configuration work was described in References 3 and 4. Then followed the NASA Energy Efficient Engine Preliminary Design and Integration Studies, which established a goal of at least 12% reduction in fuel consumption at. 8 Mach, 10, 700M cruise and a reduction of at least 5% in DOC relative to a modern operational high bypass turbofan. The baseline engine relative to which the NASA goals were to be achieved in the General Electric studies was the CF6-50C, which is currently in wide operational use. The more energy efficient turbofan described herein, along with the evaluation ground rules and trends which led to its design selection, was derived from these studies. The report of these studies is in process and is expected to be published in mid-1978 (Reference 5). In this paper, we will examine the component and system design choices that were available and the criteria applied to define an engine to meet the improved efficiency and operating economy goals noted above. ENERGY EFFICIENT ENGINE DEFINITION Since the design of a complete installed engine system is a complex iterative process, the nature and effect of the design choices that are available to improve overall engine performance are most easily understood when they are shown as variations on a basic design. In this way, the effect of the changes inherently has the proper impact on the merit of the end product. The basic engine cycle that is used here to evaluate the design choices for a more fuel efficient turbofan evolved during the NASA Energy Efficient 1

Engine studies noted above. Design studies of this engine were carried out for an engine having 15, 900 Kg SLTO thrust. Important cycle parameters for this engine are given in Table 1. Table 1 - Energy Efficient Engine Cycle Fan Pressure Ratio, Max. Climb 1. 71 Bypass Ratio " " 6. 1 Cycle Pressure Ratio " " 38 Mixed Flow Exhaust - Bypass ratio set to achieve desired mixing pressure match Turbine Rotor Inlet Temp 1360 C Hot Day T 0 Turbine Rotor Inlet Temp 1304 C Hot Day Mx Cl Turbine Rotor Inlet Temp 1260 C Hot Day Mx Cr The preliminary design of the installed engine which would operate with this cycle is shown in Figure 1 below. The engine outer fan frame, inlet, and long duct mixed flow nacelle are made of lightweight composite materials and are structurally integrated. Pylon mounted controls and accessories have been considered to reduce nacelle drag. The fan and its quarter stage are directly driven by a 5 stage low pressure turbine. The 10 stage 23:1 pressure ratio core compressor is driven by a 2 stage turbine. The engine system and component design is discussed in more detail later. The currently operational engine used as a baseline for comparative system performance evaluation in the General Electric Studies is the CF6-50C, shown in Figure 2 below (scaled to same Mx Cl Fn at. 8 Mach 10, 700M as Figure 1). Cycle parameters for this engine are shown in Table 2. Table 2 - CF6-50C Cycle Fan Pressure Ratio, Mx. Cl 1.76 Bypass Ratio, " if 4.2 Cycle Pressure Ratio, Mx Cl 32 Turbine Rotor In Temp, C (Hot Day TO) 1316 Separate Flow Exhaust This engine does not have an integrated inlet/ nacelle structure. The fan and 3 stage booster compressor are driven by a 4 stage low pressure turbine. The 14 stage 14:1 pressure ratio core compressor is driven by a 2 stage turbine. Fig. 1 Energy efficient study engine Fig. 2 General Electric CF6-50C 2

ENGINE TRADE FACTORS The design of an aircraft engine requires balance and compromise of characteristics and capabilities to obtain best overall system performance. Engine cost, maintenance cost, weight, life, reliability, safety and fuel consumption must be quantitatively balanced or optimized throughout the engine and system design to define the "right" engine. Engine trade factors provide a rational method of trading off the penalties and benefits resulting from any particular engine characteristics in terms of overall benefit to a particular aircraft and engine system. They are derived for commercial transport operations by creating a mathematical model of the engine/aircraft system, embodying appropriate characteristics of mission and engine/ aircraft technology, and varying each significant design choice in turn to determine its impact on overall system performance, weight and cost. Trade factors for advanced commercial aircraft and engine systems flying particular missions is shown in Table 3. The same trade factors were used in the selection of the more energy efficient engine (Reference 5). Table 3 - Engine Trade Factors Domestic Int'n'l Trijet Trijet Des Range NM 3000 5500 Avg Mission NM 700 2000 Passenger Capacity 225 225 Load Factor % 55 55 Fuel Costs (1977 /Gal) 45 55 Scaled Eng Thrust, Kg 11400 16800 Effect of 1% Increase in SFC @ Constant DOC Engine Weight, Kg 135 235 Fuel Burned, % 1.3 1.7 Engine Price, 1977 $ 150K 330K Engine Maint, 1977 $/Flt Hr 4 7. 5 (Mat'l + 200% OH) Aircraft TOGW, %.6 1. 1 NOTE: +1% SFC = +. 6% DOC Domestic and +1% SFC = 1% DOC Intercontinental The trade factors provide the basis for evaluating component and system design trade-offs such as weight vs cost vs performance. CYCLE SELECTION Fan Size and Pressure Ratio The selection of the fan pressure ratio for a high bypass engine is basically the selection of the propulsion efficiency of the power plant since 75-80% of the engine thrust is produced by the fan flow. For the engine thrust size selected, the relationship of fan airflow and diameter to fan pressure ratio is shown in Figures 3 and 4. 700 2.2 SLTO THRUST = 15875 KGF 7 (35000 LBF) 600 BYPASS RATIO 2.0 500 8 SELECTED P/P 1.8 1.5 1.6 1.7 1.8 FAN P/P AT MAY, CLIMB Fig. 3 Effects of fan pressure ratio/bypass ratio on fan airflow and tip diameter energy efficient study engine 2.6 = 2.2 B 2.5 SLTO THRUST = 15875 KGF. (35000 LBF.) 2.4 BYPASS RATIO 2.3 6 2.1 6 SELECTED P/P 1.5 1.6 1.7 1.8 FAN P/P AT MAX. CLIMB Fig. 4 Effect of fan pressure ratio/bypass ratio on maximum Nacelle diameter energy efficient study engine Figure 5 shows the effect of the fan pressure ratio choice on installed engine cruise SFC. 3 2 INSTALLED EN6INE 1 2^ Q 0 P,,SQN -2 1A SLTO THROS1 35000 LBF K6F. SPV^S?1-3 SELECTED P/P - u 1.5 1.6 1.7 1.8 1.9 FAN P/P a MAX CLIMB Fig. 5 Effects of fan pressure ratio/bypass ratio on cruise sfc energy efficient study engine Note also that LP turbine loading increases with lower fan pressure ratio and resulting higher bypass ratio. This is due to the reduction in LP system rpm that is needed to avoid excessive fan tip speeds as fan diameter increases. ss 5.5 3

The fan pressure ratio and hence the diameter of the fan has a major impact on the installed weight of the engine. Figure 6 shows these effects over the range of fan pressure ratios considered. 1000 2 500 7 BYPASS RATIO SLID THRUST = 15675 KGF (35000 LBF,) Engine Pressure Ratio and Turbine Temperature The thermodynamic efficiency of the engine is determined primarily by the overall pressure ratio and the turbine temperature and, of course, by the levels of component efficiency and cooling flow requirements. Figure 9 shows how installed engine SFC is affected by take-off temperature and engine pressure ratio. 0 6 SELECTED P/P < -500. 1 5.5.8SLTT THRUST - 15875 KGF. (35000 LBF.) +6CYCLE PRESSURE RATIO AT MAX. CLIMB z5.4 FAN P/P AT MAX. CLIMB Fig. 6 Effects of fan pressure ratio/b'pass ratio on installed engine weight energy efficient study engine 4 30.2 0 38 45 The effect of fan pressure ratio on system performance is shown in Figure 7 which shows a small and diminishing advantage in fuel burned is obtained as fan pressure ratio is reduced. +2.0 SLTO THRUST = 15875 KGF, (35000 LBF,) +1.0 BYPASS RATIO a 6 0 a 7-1.0 }- 8 SELECTED P/P It 1,5 1.6 1,7 1,8 FAN P/P AT MAX. CLIMB Fig. 7 Effects of fan pressure ratio/bypass ratio on fuel burned energy efficient study_ engine Figure 8 shows an increasing penalty in DOC as fan pressure ratio is reduced from the selected value, based on a transcontinental mission. Fig. 9 Hot day SLTO turbine rotor inlet temperature - deg C effects of turbine temperature and cycle pressure ratio on installed fuel consumption energy efficient study engine Figure 9 shows that the best performance occurs at an overall pressure ratio of 45 and a turbine rotor inlet temperature in excess of 1400 C. However, it is necessary to consider the need for adequate thrust growth without significant degradation in fuel efficiency, mechanical reliability and environmental characteristics. When these factors are considered, the design of the initial energy efficient engine is established at an overall pressure of 38 and a turbine rotor inlet temperature of 1360 C, allowing for further increases in these parameters to be used in achieving a satisfactory growth version. Figure 10 shows the effect of turbine temperature on engine weight. The primary effect is on the core weight; the LP system weight does not change significantly. 750 500 9,20/LITRE FUEL 18.5</LITRE FUEL -2.0 B SETS THRUST = 15875 KGF 250 BYPASS (35000 LBF RATIO X1.0 \ \ 7 4 0 o ^ 0 SELECTED P/P -1.0L 11 1,5 1.6 1. 7 FAN P/P AT MAX CLIMB Fig. 8 Effects of fan pressure ratio/bypass ratio on DOC energy efficient study engine 4 55 -vu 1100 1200 1300 1400 Fig. 10 Ilot day SLTO turbine rotor inlet temperature-deg C effect of SLTO turbine inlet temperature on engine weight energy efficient study engine

As an example, a 56 C reduction in core turbine inlet temperature requires approximately 10% more core airflow to drive the fan. This means that the core weight increases by 12-13% and since the core weight is approximately 40,0 of the engine weight the engine weight increases by approximately 5%. Figure 11 shows fuel burned in a 200 passenger transcontinental mission aircraft vs overall cycle pressure ratio and turbine temperature 20C o,i 1ht.:L:MB Fib,. 11 Hot day SLTO turbine rotor Inlet temperature - deg C effects cf turbine temperature and cycle Dressure ratio on installed fuel consu^ot`on energy efficient stud'. engine These curves are similar to Figure 9, however they are steeper for two reasons: Fuel burned improves more rapidly than SFC --- this factor is as low as 1. 1 on short range aircraft and may be as high as 1. 6 on longer range aircraft. Fuel burned includes the penalty in aircraft fuel consumption caused by the heavier weight of the lower temperature engine. The fuel usage trend, which combines the effects of engine weight and installed SFC, confirms the conclusions derived from Figure 9; the best performance occurs at an overall pressure ratio of 45 and a turbine rotor inlet temperature in excess of 1400 C. The need to provide an attractive growth step drives the selection of these parameters to somewhat lower values for the initial energy efficient engine design. COMPONENT DESIGN SELECTION In selecting component efficiencies to be developed and demonstrated in the early 1980's, aggressive projections of attainable efficiencies were made. The efficiency, weight, cost, and life factors were evaluated using the merit factors from Tables 3 and 6 to find the most effective and efficient combination of characteristics for the future improved efficiency turbofan. In making this projection, improved mechanical and structural design were recognized as necessary and important advanced technology, as well as improvements in aerodynamic design to be developed. Improved control of clearances throughout the engine by improved structure, better control of concentricity, reduced deflection, better transient matching of rotor and stator diameters were important contributing elements of the component performance levels projected. Substantial improvement in engine deterioration characteristics were also a major objective and requirement. The study was conducted using an advanced technology high temperature blade and vane material and advanced cooling technologies with constant maximum metal temperature and with cooling flow varied in the cycle as necessary, to maintain this maximum temperature. The design of each component required a trade-off of characteristics using the trade factors for the overall engine system, given in Table 3, and in addition using the component efficiency/engine performance derivatives listed below. Table 4 shows the effect on engine SFC at.8 Mach 10, 700 M cruise for a 1% change in component efficiency. Table 4 - Component Efficiency/SFC Derivatives Component SFC for +1% Efficiency at Constant Thrust Fan -0. 6 High Pressure Compressor -0. 7 High Pressure Turbine -0. 7 Low Pressure Turbine -0. 7 Cooling Flow Increase 1% -0.5 Fan The basic fan aerodynamic requirements in terms of pressure ratio, tip speeds, and stall margin did not require a fan significantly different than the currently operational fan. Improvements in efficiency relative to the best that have been tested to-date were projected as a result of further improvements in aerodynamic design and improved clearance control to be obtained with the much stiffer integrated frame and nacelle structure. A major requirement on the fan mechanical design was the reduction of damage resulting from bird strikes, ice ingestion and other FOD. The projected performance of this future fan included a specific allowance for performance degradation fan required to achieve adequate ruggedness. The quarter stage was used to permit the independent selection of fan tip speed for best fan performance while maintaining proper core boost pressure at high efficiency. General Electric's studies continued to use current fan blade materials to provide assurance that the fan would have the required ruggedness. The use of lightweight fan blades could provide improved performance, weight and possibly cost benefits to the engine, if a design approach can be identified that can be realistically predicted to meet the requirements of commercial use. 5

I Core Compressor The selection of the core compressor technology and configuration is probably the most important design choice because of its major impact on the overall engine arrangement as well as the performance level. Major engine characteristics such as 2 spool or 3 spool, 2 spool with boosters, 2 spool decoupled core, single or 2 stage core turbine are influenced by the basic selection of the core compressor design. Because of this fundamental importance, a NASA program to study Advanced Multi-stage Axial Flow Core Compressors (AMAC) was carried out to provide a thoroughly evaluated basis for selection of a core compressor for an advanced technology high efficiency turbofan engine. It is not appropriate here to present or review the results of these studies in detail, however it is useful to describe the process and show the overall conclusions. The results of this evaluation is described in the summary of the final report of the AMAC study (Reference 6) as follows: "A preliminary design study was conducted to identify an advanced core compressor for use in new high-bypass-ratio turbofan engines to be introduced into commercial service in the 1980's. The initial phase of the study involved a forecast of projected 1985 state-of-the-art technology in compressor and engine system aerodynamic and mechanical design area. The turbine inlet temperature levels projected for use in 1985 vintage engines lead to optimum thermodynamic cycles that require an overall pressure ratio of the order of 40:1. To achieve this overall pressure ratio, two types of core compressor configurations were studied: boosted 14:1 pressure ratio compressors driven by single-stage turbines and unboosted 23:1 pressure ratio compressors driven by two-stage turbines. Based upon the technology projections, a parametric screening study covering a large number of compressor designs was conducted in which the influence of major compressor design features on efficiency, weight, cost, blade life, aircraft direct operating cost and fuel usage was determined. Three high-efficiency, high-economic payoff compressors were developed using the trends observed in the parametric screening studies; these were then studied in detail to better evaluate their aerodynamic and mechanical feasibility. Finally a compressor configuration was selected which demonstrated the best performance potential and good overall system economic payoff. The design selected for development was a 10-stage 23:1 pressure ratio compressor offering the best combination of the following advantages; high efficiency, low operating cost, low fuel usage, and acceptable development risk. " This compressor was selected for use in the high efficiency engine. Figure 12 shows the design trade-offs and Figure 13 shows the economic trade-offs that lead to the selection of the 10 stage compressor. CORPRESSOR SPEED -1 D -15 ENGINE Z INSTALLED REIGNS CO"PREFSOR EFFICIENCY 5 1 9 10 11 12 13 14 NUMBER of COMPRESSOR STAGES Fig. 12 Effect of corlprensor stare nunher INSTALLED PRICE 1 0-1 DIRECT OPERATING COST % 0 FUEL USAGE 0 3 2-1 1 1 9 0 11 12 13 14 NUMBER OF COMPRESSOR STAGES Fi g,. 13 Econoüc results The effect of the compressor characteristics and requirements on high pressure turbine requirements and on the overall system were evaluated as a basis for selecting the 10 stage version. It is interesting to note that the best compressor for an energy efficient engine was not necessarily the configuration with the highest efficiency. High Pressure Turbine The aerodynamic loading of the 2 stage high pressure turbine was relatively conservative and the gains projected from improved aerodynamic design were not large. Important HP turbine performance improvements were projected, however as a result of higher temperature materials, improved cooling technology, improved flowpath sealing and improved concentricity and static part roundness control. Active clearance control to minimize interference during transient operation and maintain clearance throughout the important parts of the operating spectrum were important to performance and to the reduction of deterioration with time. _ 11

Low Pressure Turbine The low pressure turbine for the high efficiency engine was initially considered as a 4 stage design recognizing that some efficiency penalty was involved relative to a 5 stage design; however, the lower cost and weight was an important plus for the 4 stage. Table 5 shows the relative merits of the two designs in terms of overall engine system performance and clearly shows that the 5 stage design is the proper choice for this kind of an engine. Table 5 - LP Turbine 4 vs 5 Stage Comparison Original Revised 4 Stage 5 Stage Pitchline Loading Base. 85 (Base) Efficiency, % 90.2 91.7 Length, Meters Base +0. 13 0 SFC - Installed, % -1.0 (10, 700M/. 8M) Weight - Installed (15. 9K +140 Design Size), Kg U Price, $ (15. 9K Des Size) +$32,000 Maintenance Costs +0.22 Fuel Usage - Transcon/ -l.1/-l.4 Intercon, % ii DOC - Transc/Interc, % 2/-. 5 ENGINE INSTALLATION The use of a mixed flow exhaust, in an engine where fan and LP turbine discharge pressures are properly related, has long been recognized as offering potentially improved internal engine performance as well as a significant reduction in exhaust noise. There have been and continue to be questions relative to the net installed engine benefit to the aircraft system when the added weight and cost of the long mixed flow duct are accounted for and especially when the questions of possible installed drag and requirements for possible added pylon length are included. Studies of the use of composite structure in a long duct mixed flow installation for the DC 10 aircraft (Reference 7) support the conclusion that the combination of a short engine and the use of composite/integrated nacelle structure greatly reduces the weight and cost penalties of the long duct and will permit net weight and cost benefits in addition to the performance gain. A major factor in this conclusion is the fact that the fan flow and diameter must be increased for best engine efficiency and thrust if a short cowl nonmixed configuration is used. Table 6 compares the mixed and separated flow engine for a particular set of assumptions (data from Reference 3). Table 6 - Mixed vs Separate Flow Comparison Mixed Baseline Separate High Extract Same Thrust 1 Fan Press Ratio @ Mx Cl 1.71 1.71 Bypass Ratio @ Mx Cl 6.2 7. 0 Fan Diameter - Meters 2. 03 2.16 LP Turb Aero Loading Base 1.2 (Base) Jet Vel@ TO Pri/Fan, M/Sec 357 331 /280 Reverse Thrust- @ 43-50 40-47 4LSL/100 Kts, % Fwd FN A SFC - Same Eng, % Base +1.4 Nacelle Length, Meters 0.71 ASFC - Installed, % +2. 3 Weight - Installed 430 (15. 9K Design Size), Kg Price, $(15.9K Des Size) I +$100, 000 QFuelUsageTransc/Interc,% +3.7/4.8 Q DOC - Transcon/Interc, % +2.4/3.4 The very important question of pylon location and induced wing drag must be resolved with wind tunnel tests which provide direct comparisons of installation requirements and performance for the short vs long duct mixed flow nacelles. Testing to accomplish this is planned. In addition to the use of the mixed flow cycle and composite nacelle, additional performance benefit can be obtained from the use of a thinner cowl inlet and a reduction in frontal area by mounting accessories in the pylon. A major challenge of this last item is, of course, the development of configurations and arrangements that do not significantly increase maintenance time or compromise safety in any of the installation positions anticipated in the fuel efficient aircraft of the 1990's. The combined benefit of these installation features relative to the CF6-50C is estimated as shown in Table 7 (data from Reference 3). Table 7 - Mixed Flow Installation Benefit Installed SFC reduced by 3. 5% Installed weight reduced by 13% Reduced fuel usage in Transcon Trijet 7% ENVIRONMENTAL CONSIDERATIONS An energy efficient engine suitable for future commercial service must be designed with a view to satisfying the anticipated requirements for low noise and emissions. The noise requirements have been reflected in the design of the turbomachinery and engine installation. Emissions requirements have been reflected in the combustor design. 7

Acoustic Design Introduction of an energy efficient engine into commercial service in the late 1980's or early 1990's will require the aircraft to meet the newest acoustic certification requirements. During the past year, these certification requirements have been revised from the 1969 regulations to more stringent levels. The new rules, FAR 3t, (1977), are four to eight EPNdB lower for a typical domestic trijet and reflect the community demands for quieter aircraft. The engine will meet these noise requirements. Of major importance is the basic cycle selection which establishes the fan pressure ratio, fan tip speed, and low pressure turbine discharge pressure. These parameters define the levels of the two most significant sources, fan noise and jet noise (References 3, 5 and 8). Major acoustic features of the engine are: A two-chord spacing has been maintained between the fan rotor and outlet guide vanes to reduce the rotor-stator interaction noise. The exhaust nozzle system contains a. ''daisy'' mixer between the fan and core streams. Although the mixer is used primarily because of its benefit to engine performance, it is also important to the overall acoustic design of the engine nacelle system. Low pressure turbine noise has been reduced by the selection of blade numbers that place the turbine noise in frequency bands beyond the weighted PNL spectrum. This method of lowering turbine noise is a better alternative than using acoustic treatment along the core nozzle flowpath. Nacelle treatment will be located on the fan inlet and fan exhaust duct walls. The treatment will be designed to provide high acoustic efficiency for the specific source characteristics. The inlet treatment will have a length equal to 0. 54 fan diameter which will allow the overall inlet length to be established by the aerodynamic design. The exhaust duct treatment will be placed on the walls of the flowpath and will extend as far as practical, without compromising the aerodynamic or mechanical designs. Two types of advanced treatment will be considered, two-degree-of-freedom and Kevlar bulk absorber. Particular emphasis has been given to the growth versions of the basic cycle. As growth thrust increases, the fan and core jet velocities increase, thus making the jet noise more significant. The selected growth cycle has predicted jet noise levels which will allow the noise goals to be met with a margin of at least 3 EPNdB. Combustor The combustor in the energy efficient engine will operate with the same pressure loss and high combustor efficiency as current combustors, and therefore it does not contribute to improved fuel efficiency relative to current engines. Meeting emissions requirements in a design that does not penalize the reliability and life of the combustor and has a minimum impact on cost and weight on the engine is a. major technical challenge in this component. The double annular combustor design, Figure 14, selected for this engine was developed in the NASA Experimental Clean Combustor Program discussed in References 9 and 10. It provides the two stage burning necessary to meet emissions requirements in a short compact design. I 1 1 ^ 0.25M Fig. 14 Double annular combustor During start and idle operations ignition and combustion was in the outer annulus only. In this, mode, fuel air ratio was appropriate for good ignition characteristics and combustion efficiency consistent with low carbon monoxide and hydrocarbon emissions. At higher power, combustion took place also in the inner annulus at the low fuel/air ratios required for low NOx emission. As in current technology engines, the NO x emissions levels of these future energy efficient engines will be directly and strongly dependent on the selection of cycle pressure ratio - even though advanced low NOx combustor design technology is used. Extensive development investigations conducted at General Electric and elsewhere clearly show that the NO levels of all combustors, including advanced low NOx combustors, increase rapidly as combustor inlet air temperature and pressure increase. Accordingly, a NO x standard with a variable adjustment for cycle pressure ratio, as is presently being proposed by the EPA for current technology engines, represents an important need in the case of energy efficient engines of the future. A fixed-value NOx standard, with no allowance for pressure ratio effects, could impose a severe limit on the selection of the cycle pressure ratios of these 8

future engines and, therefore, could impose undesirable limits of the attainment of high fuel efficiencies with these engines. In view of expected changes in the nature and availability of petroleum fuel sources, together with the likely usage of some alternate hydrocarbon fuel sources, it is expected that energy efficient engines of the future will be required to accomniodate a much broader range of aviation kerosene fuels. To cope with the higher flame radiation characteristics of such fuels, this latter requirement will necessitate the use of even more advanced combustor liner cooling technology than that used in the modern high bypass turbofan engines of today. In addition, even more advanced smoke abatement features will be needed in the combustors used in these future engines in order to meet the prescribed smoke emission standards with these high aromatic content fuels. Thus, in view of the expected requirement that future energy efficient engines be capable of using broad-specification hydrocarbon fuels, it is expected the development of combustors for use in the future engines will require significant further technology advances. The first of the low bypass fans were not significantly higher in pressure ratio than the turbojets. Pressure ratios in the turbojets remained essentially constant while some of the low bypass fans have continued to increase to values which are 50% higher than they were in the early 1960's. The overall pressure ratio of the high bypass turbofans was much higher; this was a significant factor in the improved fuel efficiency of these engines. The increase in pressure ratio for the more fuel efficient engine is important to efficiency; although increase relative to the highest of the current engines is not large however. Figure 16 shows hot day take-off turbine inlet temperature (rotor inlet) trends. 1400 1300 1200 TURBOJETS 8 1100 LOW BYPASS TURBOFANS HIGH BYPASS TU000FA ENERGY EFFICIENT STUDY ENGINE ENERGY EFFICIENT TURBOFAN IN PERSPECTIVE l000 _^. The more energy efficient turbofan described here emerges as an extension and evolution of the high bypass engines of the 1970's with important improvements in every component and a cycle and configuration that takes advantage of technological advances in many areas together with improvements in aircraft/engine system optimization to provide another significant step improvement in turbofan engine efficiency. A technical perspective of this engine is provided by comparing some of the significant design, operating, and performance parameters of this engine with the changes in these factors that have produced the improvements in fuel efficiency in commercial engines to-date. The following series of charts show the trends in specific parameters for commercial turbojet, low bypass turbofans, and for high bypass turbofans up to the present time; the value of each parameter for the more efficient engine is also shown in the appropriate period in which such an engine could be certified. Figure 15 shows trends in overall engine pressure ratio. 900 1955 1960 1965 1970 1975 1980 1985 1990 Fig. 16 Turbine inlet temperature, commercial aircraft engines The turbojet and low bypass fans started service in the 925-990 C range and increased at about 6-8 C per year as these engines were uprated. The high bypass turbofans were first certified in the (1200-1290 C) turbine inlet temperature range which was a large step increase relative to the then current engines. The high bypass engines have also increased in temperature at about the same annual rate. The higher temperature used in these engines was a significant factor in their improved efficiency and a major factor in permitting engine size to more than double at the same time that thrust weight ratio was substantially increased. The turbine temperature of the more efficient engine represents only a small increase relative to currently operational engines. o_ 35 SO r HIGH BYPASS TURBOFANS ENERGY EFFICIENT STUDY ENGINE 1900 RUPTURE STRENGTH 207 IRA, 000 PRS. 1300 1700 ENERGY EFFICIENT STUDY ENGINE z5 k ^ 1600! ::L TURBOJETS LOH BYPASS THRBOFANS 1500 1 1 100 10 1955 1960 1965 1970 1975 1980 1985 1990 Fig. 15 Overall cycle pressure ratio, commercial aircraft engines 1955 1960 1965 1970 1975 1980 1985 1990 Fig. 17 Temperature capability of turbine blade materials

L Figure 17 shows the improvement in high temperature turbine blade materials that has historically provided increased temperature capability of approximately 7-8 C per year. 6ENERGY EFFICIENT HIGH BYPASS STUDY ENGINE TURBOFANS 4-3 2 1 LOG BYPASS TURBOFANS Figures 19 and 20 show changes in thermal and propulsive efficiencies as jet engines have evolved. Thermal efficiency shown in Figure 19 represents the ratio of ideal gas power generated by the power generating system (fan hub, high pressure compressor, combustor, HP turbine and the front portion of the LP turbine that is required to drive the fan hub) to fuel energy supplied. Ideal gas power in this sense is the power that would be generated by means of ideal expansion of the gas generator exhaust to ambient pressure. Propulsive efficiency shown in Figure 20 is the ratio of actual propulsive power (net thrust X aircraft velocity) to ideal gas power from the gas generator. This includes the effects of fan and LP turbine efficiencies, exhaust system pressure losses and exhaust mixer and nozzle performance. 0 1955 1960 1965 1970 1975 1980 1985 Fig. 18 Commercial turbofan bypass ratio Turbine blade material for the more efficient engine has been selected and cooling flows allocated conservatively to provide long design lives for these high cost parts. Figure 18 shows the range of bypass ratios for the low and high bypass ratio engines. ( POWER DELIVERED TO AIRCRAFT ) = PROPULSIVE 65 F- IDEAL GAS POWER EFFICIENCY 65 IDEAL GAS PO4ER THERMAL FUEL ENERGY SUPPLIED EFFICIENCY ( 60 ENERGY EFFICIENT HIGH BYPASS TURBOFANS STUDY ENGINE 55 LOW BYPASS TURBOFANS 55 50 TURBOJETS 8 LOW BYPA HIGH BYPASS TURBOFANS ENERGY EFFICIENT STUDY ENGINE 45 1955 1960 1965 1970 1975 1980 1985 1990 Fig. 19 Thermal efficiency of power generation, commercial aircraft engine 0 50 45 TURBOJETS 1955 1960 1965 1.970 1975 1980 1985 1990 Fig. 20 Commercial aircraft engine propulsive efficiency For the high bypass ratio engines, the relatively large range shown is the result of the fact that growth versions are required to fit the initial nacelle design which drives the growth engine to lower bypass ratio. The more efficient engine is at the high end of the band. 10

'7 The low bypass engines provided a large improvement in fuel efficiency primarily as a result of improved propulsive efficiency; 85% of the improvement came from this source. Cycle pressure ratio and turbine inlet temperature, and th'refore the cycle thermal efficiency, did not change significantly between the turbojets and the first turbofans; only 15 %o of the improvement came from this source. In the high bypass engines, increased cycle thermal efficiency resulting from much higher cycle pressure ratio and temperature, contributes 40 /0 of the improvement provided by these engines; 60% of the gain is the result of higher propulsive efficiency. As seen on these two figures, the improved performance of the more efficient engine is the result of significant improvements in both thermal (45% of the improvement) and propulsive efficiency (55% of the improvement). The record of commercial jet engine fuel efficiency is shown in Figure 21. 1.0 0.9 F F LOW 0,8 TURBOJETS BYPASS TU FA HIGH BYPASS TURBOFANS 0.6 ENERGY EFFICIENT -UDY ENGINE 0.5i I I i 1955 1960 1965 1970 1975 1980 1985 1990 Fig. 21 Specific fuel consumption trends, commercial aircraft engines Altitude cruise SFC is shown for each of the three engine classes vs time. The curve passes thru the representative fuel efficiency for each class at the time that it first entered commercial service, In each class, the shaded area shows the changes in fuel efficiency with time as these engines continued in use; the slope at the top of the band in general reflects the fact that the fuel efficiency of the engines worsened somewhat as growth versions of the basic engines were introduced. This chart also shows that the goal established by NASA of a 12% SFC improvement for the more fuel efficient engine can be met. CONCLUDING REMARKS an engine with this improved fuel efficiency could be certified in the late 1980's. The advances in performance, cost, and weight to be achieved by the energy efficient engine result from the incorporation of advanced technology features in structural design, installation packaging and turbomachinery performance. Most importantly, these include the following: high pressure ratio compressor (23:1 on one shaft); highly loaded low pressure turbine; active clearance control on both high pressure compressor and turbine; advanced high temperature materials and cooling design technology; composite/integrated engine and nacelle structure with mixed flow exhaust; thin cowl inlet and pylon-mounted accessory package. In addition, environmental considerations of noise and emissions are benefitted by careful attention to the acoustic aspects of turbomachinery design, the incorporation of the double-annular combustor and the mixed flow exhaust system. In the definition of this engine, it was recognized that engine first cost, life and maintenance cost must be realistically traded off against fuel efficiency in defining a cost effective efficient engine. The cycle selected is not the optimum that could have been constructed; an overall pressure ratio of 45 with a turbine rotor inlet temperature of 1430 C would have offered approximately 1. 5%/. better SFC. The very great importance of reliability and lower maintenance costs combined with the need to provide an attractive growth path led to the selection of a pressure ratio 38 and turbine inlet temperature of 1360 C as a better balance because it requires only a small increase over the pressure and temperature levels of some of todays high bypass engine. With this philosophy, it is realistic to project lower maintenance costs for this engine. Improvements in fuel efficiency beyond these engines can and will be made but they will necessarily be smaller. Further improvements will come from somewhat better component efficiencies, improved materials and improved cooling technology which will permit operation at higher engine pressure ratio and turbine temperature. In turbine engines, as in any other maturing technology, the opportunities for step function improvements are smaller as time goes on but the engine described here is not the end of the line. ACKNOWLEDGEMENT The authors thank R. F. Patt Manager of Energy Efficient Engine Aerothermodynamics Systems for his assistance in the preparation of this paper. The more energy efficient engine described here will meet the objective of a 12% altitude cruise SFC reduction and a 5% DOC reduction established by NASA. Environmental requirements will also be met. A 5 year program of design, technology development and system demonstration has started; 11

REFERENCES SI - U S CUSTOMARY CONVERSION FACTORS 1 Ross, E. W., Johnston, R.P., and Neitzel, R. E., "Cost Benefit Study of Advanced Materials Technology for Aircraft Turbine Engines", prepared for NASA under Contract NAS 3-17805, NASA CR-134702, November, 1974. 2 Hillery, R. V., and Johnston, R.P., "Cost Benefit Study of Advanced Materials Technology for Aircraft Turbine Engines ", prepared for NASA under Contract NAS 3-20074, NASA CR-135235, September, 1977. 3 Neitzel, R. E., Hirschkron, R. and Johnston, R.P., "Study of Turbofan Engines Designed for Low Energy Consumption", prepared for NASA under Contract NAS 3-19201, NASA CR-135053, August, 1976. 4 Hirschkron, R., Johnston, R.P. and Neitzel, R. E., "Study of Unconventional Aircraft Engines Designed for Low Energy Consumption", prepared for NASA under Contract NAS 3-19519, NASA CR-135136, December, 1976. 5 Final report to be issued in Mid 1978 on NASA Contract NAS 3-20627, "Energy Efficient Engine Preliminary Design and Integration Studies". 6 Wisler, D.C., Koch, C., and Smith, Jr., L. H., "Preliminary Design Study of Advanced Multi-stage Axial Flow Core Compressor", prepared for NASA under Contract NAS 3-19444, NASA CR-135133, February, 1977. 7 Nordstrom, K. (Douglas), Marsh, A. (Douglas) and Sargisson, D. F., "Conceptual Design Study of Advanced Acoustic Composite Nacelles", preparedfor NASA under Contract NAS 1-13356, NASA CR-132703, July, 1975. 8 Kohn, A. 0., "Noise Considerations in High Bypass Ratio Fan Engine Design", ASME paper 70-WA/GT-14, November, 1970. 9 Bahr, D.W., and Gleason, C.C., "Experimental Clean Combustor Program - Phase I Final Report'; prepared for NASA under Contract NAS 3-16830, NASA CR-134737, June, 1975. 10 Bahr, D. W., Gleason, C.C., and Rogers, D. W., "Experimental Clean Combustor Program - Phase II Final Report", prepared for NASA under Contract NAS 3-18551, CR-134971, August, 1976. Length Volume Mass Force Stress Temperature Conversion K Meters to inches 2. 7340E-01 Meters to feet 3.2808E+00 Litres to gallons 2. 6417E-01 =^= Kgm to Lbm 2. 2046E+00 "' Kgf to Lbf 2. 2046E+00 * MPa to psi 1.4504E+02 oc to of TC +32 1.8 U S Customary = (SI) X (K) 12