Advanced Turbofan Engines for Low Fuel Consumption

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1 78-GT-192 Copyright 1978 by ASME $3.00 PER COPY $1.50 TO ASME MEMBERS 1.00 at Wembley 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). Advanced Turbofan Engines for Low Fuel Consumption WILLIAM SENS Pratt & Whitney Aircraft Group, Commercial Products Division, East Hartford, Conn. The anticipated commercial aircraft fuel usage through the year 2000 is divided into three categories: that which will be consumed by existing engines, new production of current type engines, and new turbofan engines with advanced technology. Means of improving fuel consumption of each of these engine categories will be reviewed and the potential fuel savings identified. The cycle selection and design characteristics of an advanced turbofan engine configuration will be discussed and the potential improvements in fuel consumption and economics identified. 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, Manuscript received at ASME Headquarters January 5, Copies will be available until January 1, THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS, UNITED ENGINEERING CENTER, 345 EAST 47th STREET, NEW YORK, N.Y

2 Advanced Turbofan Engines for Low Fuel Consumption WILLIAM SENS DISCUSSION Fig. 1 shows a projection of jet fuel usage by the free world commercial fleet up through the year This fuel consumption projection is based on an estimated growth rate of the free world fleet of approximately 6 percent per year. This fuel consumption is divided into three parts. 1 The fuel being used by engines that are in the fleet today. 2 The fuel consumed by new production of engine types currently in operation or under development, and their derivatives. 3 The fuel used by new advanced turbofan engines developed from technology advances anticipated in the next six to eight years. The fuel used by the advanced engines would become significant toward the end of this century assuming that the advanced turbofans start entering the fleet by approximately The fuel usage of the free world air fleet over the time period 1980 to the year 2000 is projected to be approximately 4 trillion liters (trillion gallons). As shown in Fig. 2, approximately 90 percent of total will be burned by engines in existence today, or new engines based on existing design technology. Only about 10 percent would be used by advanced turbofan engines designed in the mid eighties or later. This chart illustrates the importance of improving the fuel consumptions of current engine types in order to provide a significant fuel savings in this century. For these reasons, this paper will review improvements that can be made to the fuel consumption of current engines and new production versions of current engines in addition to a discussion of potential fuel saving of advanced turbofan engines. 1 Also the discussion 1 Information updated from "Potential Imrovements in Turbofan Engine ngine Fuel Economy," of performance improvements to current engine types is important as this effort provides a very significant input to the design of improved advanced turbofan engines. There are basically four ways to improve the fuel consumption of engines, as shown in Fig. 3. These are: By improvement in component performance; by improved maintenance procedures that reduce the amount of engine performance deterioration with operational use; by improvement in the powerplant cycle by changes in bypass ratio, overall pressure ratio, and turbine inlet temperature as the major variables; and finally, by designing the engine to be less sensitive to factors causing performance deterioration. The "check marks" indicate where a fuel consumption improvement approach is applicable to a given category of engine. Where there is a limited degree of applicability it is marked 'Some." For instance: an improved cycle can be selected for an advanced engine where we are starting with a clean sheet of paper. However, in growth or derivative versions of current engines (column 2) there are only modest changes in cycle possible due to the constraints of existing hardware and technology. It is a rare occurrence when existing engines are modified to the degree necessary to appreciably affect their cycle. However, performance of engines in operation today can be improved by retrofit of improved components where such a retrofit can be economically justified (this will be discussed further); and by modified maintenance procedures to reduce performance deterioration. A number of component improvements applicable to several current engines are currently under evaluation and development testing as shown in Fig. 4. Use of a single shroud fan blade (instead of the two shrouds on current JT9DIs) along with use of an improved fan blade cross section Hines, R., and Gaffin, W., AIAA Paper No , July 26,

3 80 a, a 70 C) 60 o ' 0 Fig. 1 N C _ C N 100 C 0 50 a, Year Predicted world commercial aircraft jet fuel consumption Component New production Existing current Advanced engines engines engines improvement Some Improved maintenance Improved cycle Designs for lower deterioration Fig. 3 Means of reducing Some Some fuel consumption Typical approaches Existing engines 34% A single shroud fan with improved aerodynamics Reduced HP compressor clearances with abradable rubstrips Increased diameter HP compressor air seals Revised HP turbine air seal Added knife edges Reduced clearances Improved HP turbine active clearance control New production Advanced current engines engines 53-58% `";"::8-13% Fig. 2 Predicted world commercial aircraft jet fuel consumption four trillion liters (multiple circular arc) can reduce block fuel consumption by 1 to 1.5 percent. Fan blade chords must be increased in order to avoid blade flutter and to have satisfactory foreign object damage characteristics. The use of abradable rub strips through the high pressure compressor permits running with tighter tip clearances thereby providing efficiency improvements of up to 2 percent. A sprayed nichrome polyester abradable appears most promising based on cost, durability and abradability considerations. Increasing the diameter of the air seals that seal off the high pressure compressor stator root Thermal barrier coatings on turbine nozzle end walls Fig. 4 Current engines component improvement cavities reduces the parasitic recirculation losses and can potentially improve efficiency by one percent. Revisions to the JT8D high pressure turbine outer air seal by adding additional knife edge seals and honeycomb rub strips with reduced clearances can increase its efficiency by one percent. In the JT9D, lower clearance can be achieved at altitude cruise conditions by use of an improved active clearance control system which incorporates increased coolant air supply and a segmented outer air seal support. Ceramic coatings of the first stage nozzle guide vane end walls provide a "thermal barrier" or insulating effect which allows a reduction in cooling air and a consequent performance improvement of 0.2 percent. Fig. 5 shows the potential reduction in cruise fuel consumption for the JT8D and JT9D engines incorporating the kind of modifications listed in Fig. 4. Both engines require a similar number of component changes in order to achieve the significant performance gains shown. Through its Engine Component Improvement- Performance Improvement Program, NASA is support- 3

4 Potential reduction in cruise fuel consumption Demonstrate SFC benefits of JT8D and JT9D component improvements D Installed SFC reduction High probability of production incorporation 5 fuel savings over engine lifetime Fig. 6 NASA Engine Component Improvement-Performance Improvement Program objectives ME 5 10 Number of modifications Fig. 5 Current engines component improvement ing and participating with the engine manufacturers in the evaluation, selection, and feasibility testing of a number of the current engine component improvements. The objectives of this NASA program are shown in Fig. 6. The goal of a 5 percent fuel savings over the lifetime of the engine appears technically feasible based on the design evaluation studies. However, performance testing and economic evaluation must be carried out to determine how far it is economically feasible to go. An economic evaluation procedure has been developed as a part of the NASA Performance Improvement Program which I believe is somewhat unique to the industry. The general procedure is shown schematically in Fig. 7. The enginemanufacturer, P&WA, and the aircraft companies, Boeing and Douglas, estimate the performance, weight, price, and maintenance cost impact of each of the component improvements. The results of these estimates in terms of airplane performance and economic data are fed into TWA's route and economic simulation. This simulation calculates, for TWA's route system, the impact of each of the component changes on their direct operating cost, investment, and fuel burned. It also calculates the period of time required for the savings in fuel costs to pay back the initial cost of incorporating each of the improved components in the engines. This calculated payback period is then compared with the maximum payback period that the airlines consider acceptable. If the payback period for a given component improvement is greater than the maximum acceptable value, that improvement is rejected. If the payback period is below the maximum acceptable time, it is then evaluated in terms of the cumulative fuel saving over the time period up to 1990 based on the market projection for that particular engine as determined by the team of manufacturers and operators. American Airlines, United Airlines, Pan American Airlines, and Eastern Airlines have been active as consultants in this economic evaluation and are members of the Performance Improvement evaluation team. The component improvements to be pursued in the engine evaluation testing program were selected by P&WA and approved by NASA based on consideration of the fuel saving, economic attractiveness and the evaluation program cost. This general evaluation procedure was used to screen and select the component improvements to be pursued in the engine evaluation testing program. It will be used after evaluation testing to determine which of the performance improvement items should be included in new production engines and offered as a retrofit. As such, it is a useful evaluation tool for the future. The specific fuel consumption of aircraft engines tends to deteriorate with usage for a number of reasons. Fuel savings would be achieved by reducing the amount of this performance deterioration. The general problem is illustrated in Fig. 8. This plot shows historical data on the increase in fuel consumption with flight time for a number of JT9D engines operated and maintained by different airlines. There are several interesting things to be noted. one is that a performance deterioration of 1 to 2 percent occurs in the first several hundred hours of usage. This is called short term deterioration. Our studies to date indicate that the short term deterioration results from an increase in blade tin and seal clearances resulting from rubbing during maximum flight load conditions. Secondly, there is a long term deterioration which takes place over thousands of hours of engine operation. This type of deterioration is a result of foreign object damage, erosion; and warpage of parts in addition to continued seal and tip clearance increases resulting from more severe flight loads which occur relatively infrequently. It is also interesting to note the wide variation in deterioration between the specific engine data points shown. We believe that this 4

5 PWA BCAC DACO Manufacturers TWA AA PAA UAL EAL Operators Team Performance Weight Price Maint Required Market Cost I payback I projection period Route & economic simulation Z. " Payback period Out Cumulative fuel saved Fig. 7 Current engine component improvement evaluation procedure "The problem" JT9D-3/7 post repair 5 Short term Long term Percent 4 increase n 3, specific 0 uel 2 f : 0 consumption 0. _ %. ' «0I I I I I i I I I ,000 11,000 Engine flight hours Fig. 8 Current engine performance deterioration scatter is related primarily to the difference in operation and maintenance philosophy between airline operators and to measurement accuracy. We are actively examining data on parts condition, engine performance, and airline maintenance operational procedures to try to better understand the correlation of this long term performance deterioration with the methods of operation and maintenance. The cornerstone of this program is the NASA JT9D Engine Diagnostic program. The objectives of this program are to determine the nature of the JT9D short and long term performance deterioration; to identify and quantify the sources of the JT9D performance deterioration; to determine the sensitivity of component performance to the engine parts deterioration; to develop an analytical model of the JT9D performance deterioration; and based on this model and the analysis of the historical practices of the different airlines, to recommend performance retention techniques for current engines and future engines (see Fig. 9). 5

6 Quantitatively the overall objective is to reduce cruise fuel consumption deterioration at least several percent without adversely affecting the economics of engine operation. The general logic of the short term performance deterioration program is shown in Fig. 10. Operational experience has been determined by detailed measurement of clearances and performance changes in short time 747SP JT9D airlines service engines. An analytical model of the JT- 9D engine and installation have been constructed and used to estimate the change in engine clearances under critical flight load conditions. The operational experience has been compared with the analytically predicted clearance and performance changes. So far, the agreement between analysis and experience has been quite Objectives - NASA JT9D engine diagnostic program Determine short and long term JT9D performance deterioration Identify and quantify sources of short and long term deterioration Determine sensitivity of component performance to engine parts deterioration Develop analytical model of JT9D performance deterioration Recommend performance retention techniques for current and future engines Fig. 9 Current engines reduced deterioration good. It is also planned to measure the changes in critical clearances and performance of the JT9D when it is subject to simulated maximum flight load conditions in a sea level test stand where the critical running clearances can be measured by X-ray. This will provide additional data to calibrate the analytical tools that have been developed. The improved analytical clearance prediction system along with the improved understanding of the causes of deterioration will lead to future engines designed for reduced deterioration. The logic of the long term deterioration program is shown in Fig. 11. Historical data from five airlines on performance trends based on inflight measurements and test stand calibrations taken before and after major repair has been analyzed. Also the mechanical condition of specific parts were examined to determine their relationship with the performance degradation and the performance improvement in the repair process. The effect of specific repairs will be determined based on pre- and post-repair calibration of JT9D7A 747SP engines with expanded instrumentation. These activities will provide an improved understanding of the causes of long term deterioration. This in turn will lead to improved maintenance procedures that will provide a better balance between performance retention and parts replacement costs; and also to design criteria that will provide engines with improved long term deterioration in future JT9D models and new engines. Short term deterioration program approach Predicted flight loads Ope rational e xpe r ience Measured clearance and performance changes n service engines Analytically predicted clearance changes Measured clearance and performance changes under applied loads (X-ray testing) Comparison of arience with analy Improved analytical clearance prediction Improved under of deterioration Designs for improved short term deterioration Fig. 10 Current engines reduced deterioration 6

7 Long term deterioration program approach Performance trends In flight Pre/post repair Improved maintenance procedures Parts mechanical condition Improved understanding ^^ of long term deterioration causes Designs for improved long term deterioration Measured performance from specific repair Fig. 11 Current engines reduced deterioration Development of technology by 1985 to permit 12 r reduction in TSFC relative to JT9D-7A 5 improvement in DOC relative to JT9D-7A FAR-36 as amended March 1977 Meet anticipated emissions requirements Fig. 12 Advanced turbofan engines NASA E 3 Program objectives Now let us turn our attention to advanced turbofan engines. The NASA Energy Efficient Engine (E 3 ) Program is the prime focus for our long term turbofan engine technology development. The objectives of the program are shown in Fig. 12. Its primary goal is to achieve at least a 12 percent reduction in cruise specific fuel consumption relative to the JT9D7A powerplant. In addition, there is the objective of obtaining at least a 5 percent improvement in airplane direct operating costs relative to the JT9D7A, to meet the FAR 36 noise requirements as amended in March 1977, and to meet the anticipated future emission requirements. The cycle selection and design optimization of the advanced turbofan was based on the two aircraft requirements shown in Fig. 13. The economics and relative fuel burned for the domestic airplane were calculated for a 1300 km (700 nautical mile) average stage length. The economics and fuel burned for the international aircraft were determined for a 3700 km (2000 nautical Domestic trijet 440 passengers 5600 KM (3000 NM) design range 1300 KM (700 NM) average stage length International quadjet 510 passengers 10,200 KM (5500 NM) design range 3700 KM (2000 NM) average stage length Fig. 13 Advanced turbofan engines evaluation aircraft mile) average stage length. The effect of the major cycle variables on fuel burned for the international range aircraft is shown in Fig. 14. Takeoff turbine inlet temperature and overall pressure ratio are the independent variables, with contours of relative fuel burned. She fan pressure ratio was assumed constant at 1.6 as this fan pressure ratio gives near minimum fuel burned for the range of primary cycle variables covered. 2 The level of component efficiencies and materials technology assumed are 2 "Study of Turbofan Engines Designed for Low Energy Consumption," NASA CR , April

8 International quadjet 0.8M cruise Fan pressure ratio = % +4% +1% Fuel burned a Ia 2800 iann a) I Q) c 6 linimum iel burned L Y Y JT9D ca H H 2400L 1600 k Overall pressure ratio Fig. 14 Advanced turbofan effect of cycle on fuel burned International quadjet 1977 $ /liter (554 /gal) fuel U Y a E C) c I C) ca I- a E 6. I- F_ 0 a) ca L 1600 E Overall pressure ratio Fig. 15 Advanced turbofan effect of cycle on direct operating cost consistent with a mid 801s start f development. With these assumptions, minimum fuel burned occurred at pressure ratio approaching 60 and a takeoff turbine inlet temperature of approximately 1810 K (2800 F). This powerplant cycle would provide about a 10 percent reduction in fuel burned relative to a JT9D cycle engine incorporating the same advanced technology. It is also of interest to note that there is only a small gain to be made in fuel burned by raising turbine inlet temperature above the current JT9D levels. Increasing the overall pressure ratio and increasing the turbine inlet temperature tend to increase initial engine cost and engine maintenance cost, everything else being equal. These effects tend to counterbalance the economic advantages of the lower fuel consumption gained with the higher pressure ratio and temperature. An airplane study was conducted to evaluate the 8

9 International quadjet 1977 $ 294; liter (S1.10 gal) fuel 3000,^ Y 1900 a a E E m c c a ; o o a) a) Y Y f0 f Overall pressure ratio Fig. 16 Advanced turbofan effect of cycle on direct operating cost effects of these design variables on aircraft economics. The results of the international quadjet evaluation are shown in Fig. 15 where contours of constant aircraft direct operating cost are plotted as a function of the primary cycle variables. These results are based on the use of 14.5 cent/liter (55 cent/gal) fuel and 1977 economic assumptions for the rest of the airplane. In each case, the size of the aircraft and its powerplant were adjusted to give the same payload and range characteristics and priced accordingly. The replacement life of the hot section parts was assumed constant (as metal temperature was held constant). However, the price of the hot section parts was increased with increasing pressure ratio and temperature to account for increased complexity of the cooled parts. The optimum direct operating costs (Fig. 16) occur at close to the current JT9D engine cycle variables. The turbine inlet temperatures and overall pressure ratios could be varied over a fairly wide range without significantly affecting the direct operating cost computed with the stated assumptions. However, the assumptions on engine cost and engine maintenance cost become more questionable as temperature and pressure ratios are increased beyond current commercial operating experience. So there seems to be little incentive to push up to very high overall pressure ratios or turbine inlet temperatures with fuel priced near current levels. The next chart shows what would happen if Advanced JT9D-7A turbofan Bypass ratio Fan pressure ratio Fan tip speed - m sec (ft sec) 410 (1350) 475 (1550) Overall pressure ratio Max. combustor exit temp - K ( F) 1644 (2500) 1700 (2600) Turbine cooling air Base -35 Turbine exhaust/ fan duct mixer No Yes Fig. 17 Advanced turbofan preliminary design parameters the price of fuel were to double while the other airplane operating costs were to remain constant. Under these conditions, the cycle pressure ratio for optimum direct operating costs moves all the way up to 50 and the direct operating cost penalty for using today's JT9D cycle becomes significant. However the gains due to increasing turbine inlet temperature above the JT9D values are still small. The results of the domestic aircraft evaluation show somewhat less incentive to increase overall engine pressure ratio. The design parameters selected for the advanced engine are shown in Fig. 17. The JT9D-7A engine is shown for comparison. An overall pressure ratio of 38 and a maximum combustor exit temperature of 1700 K (2600 F) have been selected 9

10 % change in Fuel burned International airplane Domestic airplane -4 tiii Advanced separate flow turbofan ' Advanced mixed flow turbofan % change in DOC +2 / JI J 85 % mixer efficiency- Fig. 18 Advanced turbofan potential mixer benefits as being a reasonable compromise between improved performance and the anticipated capabilities of materials and cooling technology in the projected time period. The fan bypass ratio of 6.5 and fan pressure ratio slightly above 1.7 have been selected on the basis of achieving the best possible combination of installed weight and performance. Also, it was desired to reduce the primary jet velocity to the point where noise levels 4 db below the FAR 36 March 1977 amendment levels could be achieved if advances in fan noise reduction were achieved. Fan tip speed has been set at 475 m/sec (1550 fps) in order to achieve acceptable fan stability margin at the desired fan pressure ratio level. The design objective is to reduce the turbine cooling air by approximately 1/3 relative to the JT9D -7A in spite of the higher pressure ratio and turbine inlet temperature by use of single crystal moterials in the high pressure turbine blade airfoils and improved airfoil cooling effectiveness. The advanced turbofan incorporates a short length high effectiveness mixer in the exhaust stream in order to further improve fuel consumption. The exhaust stream mixer transfers energy from the turbine exhaust gas to the fan exit air by a mixing process. If done efficiently, this can provide a 4 to 5 percent improvement in specific fuel consumption relative to a non-mixed engine of the same diameter. As shown in Fig. 18, use of a mixer could provide a 4 percent reduction in fuel burned and 2 percent reduction in direct operating cost for the international range airplane and somewhat smaller gain for the +2 Direct drive fan! % change in Fuel burned 2 % change in Geared fan ;-- International airplane --- Domestic airplane J i Direct drive Geared fan fan Fan bypass ratio Fig. 19 Advanced turbofan geared fan evaluation domestic aircraft. It should be emphasized that these are preliminary estimates. Significant design study, mixer model tests, and nacelle/wing model tests are required to determine whether these benefits can be made a reality. The potential benefits of the mixer are sufficient to warrant such an evaluation program. One possibility that was seriously considered was the use of gearing between the low pressure compressor and fan. Use of a gear would permit lower fan tip speeds and higher bypass ratios (both of which favor lower fuel consumption 10

11 Fig. 20 Advanced turbofan preliminary layout Performance improvements: Shroudless hollow titanium fan blade Contoured compressor airfoils Single crystal turbine airfoils Graded ceramic turbine outer air seals Reduced clearances - structurally integrated fan ducts - active clearance control designs Short efficient mixer Weight/cost reductions: Increased rotor speeds/bearing ON's Improved compressor and turbine disk materials Oxide dispersion strengthened combustor liners Composite nacelle Emissions improvement: Vorbix combustor Fig. 21 Advanced turbofan potential design features and lower noise) without compromising the design of the low pressure compressor and low pressure turbine. Fig. 19 shows the results of the geared engine study in terms of fuel burned and direct operating cost. The results indicate the possibility of a 4 percent improvement in fuel burned for both types of aircraft and 1 percent in direct operating costs for the international aircraft. No change in DOC was indicated for the domestic aircraft. However, the geared fan was not selected for the advanced turbofan engine for several reasons: (a) The durability and reliability characteristics of lightweight high power gearing under flight load conditions is open to question, (b) a very extensive program would be required to substantiate the gearing durability under flight operating conditions, (c) the geared fan engine has low pressure compressor and low pressure turbine components that are quite different aerodynamically from conventional turbofan components. Thus, test program results would not directly support our ongoing current product line improvement in these component areas, and (d) the evaluation studies conducted did not recognize all of the secondary effects that a large geared fan might have on the airplane configuration. The foregoing considerations were felt to outweigh the potential advantages shown by the preliminary study. A preliminary layout of the advanced turbofan engine is shown in Fig. 20. It incorporates two spools and a two stage high pressure turbine. The fan ducts are structurally integrated with the engine in order to stiffen the engine and thereby minimize deflection and clearance changes under flight load conditions. The engine incorporates a long duct to provide adequate fan noise suppression and mixing of the turbine exit gas to improve performance. The potential design features for this engine are listed in Fig. 21. For improved performance, it features shroudless hollow titanium fan blades; contoured compressor airfoils similar in nature to supercritical aircraft wing airfoils, single crystal turbine airfoils which permit higher metal temperature and therefore reduce cooling air requirements; turbine outer air seals of a graded ceramic material which permits running with reduced tip clearances and reduced shroud cooling. Clearances are minimized by use of the structually integrated fan ducts to minimize deflection and clearance changes under flight loads, and by the use of active clearance control in the high pressure turbine and compressor. The active clearance control is used to open up clearances at operating conditions where maximum flight loads and critical transients occur. As discussed 11

12 Relative to JT9D-7A International aircraft Reduction Design feature (%) Thermodynamic cycle 6 Component aero refinement 1 6 Improved materials Improved structure/clearance control 4 Mixed flow nacelle4 Total reduction 20% Fig. 22 Advanced turbofan potential fuel burned reduction previously the design also incorporates a short high efficiency exhaust mixer. Improved weight and cost are achieved through the higher rotor bearing and speeds. The higher rotor speed allows the same amount of work to be done with a fewer number of stages and airfoils. Improved compressor and turbine disk materials are utilized to achieve the higher tip speeds without increase in weight or cost. Oxide dispersion strengthened material is used for the combustion liners to provide the required burner liner life at the higher compressor exit pressure and temperatures. Composite materials will be used extensively in the nacelle to minimize weight. A vorbix type combustor derived from the NASA Clean Combustor Program will be utilized in order to minimize the emissions. This is particularly important in order to reduce the NO emissions which otherwise would increase at the higher levels of combustor inlet temperature of the advanced engine. The advanced turbofan engine has a potential reduction in fuel burned of 20 percent relative to the JT9D-7A. As shown in Fig. 22, approximately 30 percent of this improvement is due to the thermodynamic cycle change, about 30 percent due to refinement in the component aerodynamics and improved materials, about 20 percent due to improvements in the structure and clearance control, and 20 percent due to the use of a mixed flow nacelle. This potential improvement is impressive. However, it will take many years of diligent design and component research to reach the point where we have the necessary know-how to proceed with the design and development of such an engine. In summary, we have identified means of improving the fuel consumption of current engines by as much as 5 percent through improved component performance and reduced engine deterioration. The advanced turbofan configuration described herein has the potential of providing a 20 percent reduction in fuel burned relative to today's high bypass ratio turbofan by Programs funded by NASA and the engine manufacturers are in being to make such progress possible. Successful completion of these efforts would reduce the fuel consumption of the free world transport fleet by over 10 percent by the year