US20150285155A1 - Method for setting a gear ratio of a fan drive gear system of a gas turbine engine - Google Patents
Method for setting a gear ratio of a fan drive gear system of a gas turbine engine Download PDFInfo
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- US20150285155A1 US20150285155A1 US14/742,954 US201514742954A US2015285155A1 US 20150285155 A1 US20150285155 A1 US 20150285155A1 US 201514742954 A US201514742954 A US 201514742954A US 2015285155 A1 US2015285155 A1 US 2015285155A1
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- fan
- gas turbine
- ratio
- turbine engine
- rotor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/32—Arrangement, mounting, or driving, of auxiliaries
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/06—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
- F02C3/113—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission with variable power transmission between rotors
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H1/00—Toothed gearings for conveying rotary motion
- F16H1/28—Toothed gearings for conveying rotary motion with gears having orbital motion
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H1/00—Toothed gearings for conveying rotary motion
- F16H1/28—Toothed gearings for conveying rotary motion with gears having orbital motion
- F16H1/36—Toothed gearings for conveying rotary motion with gears having orbital motion with two central gears coupled by intermeshing orbital gears
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2210/00—Working fluids
- F05D2210/10—Kind or type
- F05D2210/12—Kind or type gaseous, i.e. compressible
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/40—Use of a multiplicity of similar components
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/40—Transmission of power
- F05D2260/403—Transmission of power through the shape of the drive components
- F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
- F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type
Definitions
- This disclosure relates to a gas turbine engine, and more particularly to a method for setting a gear ratio of a fan drive gear system of a gas turbine engine.
- a gas turbine engine may include a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
- the compressor section can include low and high pressure compressors
- the turbine section can include low and high pressure turbines.
- a high pressure turbine drives a high pressure compressor through an outer shaft to form a high spool
- a low pressure turbine drives a low pressure compressor through an inner shaft to form a low spool.
- the fan section may also be driven by the inner shaft.
- a direct drive gas turbine engine may include a fan section driven by the low spool such that a low pressure compressor, low pressure turbine, and fan section rotate at a common speed in a common direction.
- a speed reduction device which may be a fan drive gear system or other mechanism, may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section. This allows for an overall increase in propulsive efficiency of the engine.
- a shaft driven by one of the turbine sections provides an input to the speed reduction device that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.
- gas turbine engines utilizing speed change mechanisms are generally known to be capable of improved propulsive efficiency relative to conventional engines
- gas turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.
- a gas turbine engine has a fan section including a fan rotatable about an axis.
- a speed reduction device is connected to the fan.
- the speed reduction device includes a planetary fan drive gear system with a planet gear ratio of at least 2.6.
- a bypass ratio is greater than about 11.0.
- the gear ratio is less than or equal to 4.1.
- a fan pressure ratio is below 1.7.
- a fan pressure ratio is below 1.48.
- the fan blade tip speed of the fan section is greater than about 1000 ft/sec and less than about 1400 ft/sec.
- the planetary fan drive gear system includes a sun gear, a plurality of planetary gears, a ring gear, and a carrier.
- each of the plurality of planetary gears includes at least one bearing.
- a low pressure turbine is mechanically attached to the sun gear.
- a low pressure turbine section is in communication with the speed reduction device.
- the low pressure turbine section includes at least three stages.
- the bypass ratio is less than about 22.0.
- a method of improving performance of a gas turbine engine includes determining fan tip speed boundary conditions for at least one fan blade of a fan section. Rotor boundary conditions are determined for a rotor of a low pressure turbine. Stress level constraints are utilized in the rotor of the low pressure turbine and the at least one fan blade to determine if the rotary speed of the fan section and the low pressure turbine will meet a desired number of operating cycles.
- a bypass ratio is greater than about 6.0.
- a speed reduction device connects the fan section and the low pressure turbine and includes a planetary gear ratio of at least about 2.6.
- the planetary gear ratio is less than about 4.1.
- a fan pressure ratio is below 1.7.
- a fan pressure ratio is below 1.48.
- the bypass ratio is greater than about 11 and less than about 22.
- a fan blade tip speed of the at least one fan blade is less than 1400 fps.
- a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and the number of stages of the low pressure turbine is increased.
- a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and an annular area of the low pressure turbine is increased.
- a fan drive gear module for a gas turbine engine includes a planetary fan drive gear system with a speed reduction ratio of at least 2.6.
- the speed reduction device is configured to drive a fan section with a bypass ratio greater than about 11.0.
- the speed reduction ratio is less than or equal to 4.1.
- the planetary fan drive gear system is configured to drive a fan section with a fan blade tip speed greater than about 1000 ft/sec and less than about 1400 ft/sec.
- the bypass ratio is less than about 22.0.
- a method of designing a gas turbine engine includes selecting fan tip speed boundary conditions for at least one fan blade of a fan section of a gas turbine engine. Rotor boundary conditions are selected for a rotor of a fan drive turbine of the gas turbine engine. Stress level constraints are determined in the rotor of the fan drive turbine and the at least one fan blade to determine if the rotary speed of the fan section and the fan drive turbine will meet a desired number of operating cycles.
- a bypass ratio is greater than about 6.0.
- the fan section and the fan drive turbine are connected through a speed reduction device that includes a planetary gear ratio of at least about 2.6.
- the planetary gear ratio is less than about 4.1.
- the bypass ratio of the gas turbine engine is greater than about 11 and less than about 22.
- FIG. 1 illustrates a schematic, cross-sectional view of an example gas turbine engine.
- FIG. 2 illustrates a schematic view of one configuration of a low speed spool that can be incorporated into a gas turbine engine.
- FIG. 3 illustrates a fan drive gear system that can be incorporated into a gas turbine engine.
- FIG. 4 shows another embodiment.
- FIG. 5 shows yet another embodiment.
- FIG. 1 schematically illustrates a gas turbine engine 20 .
- the exemplary gas turbine engine 20 is a two-spool turbofan engine that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- Alternative engines might include an augmentor section (not shown) among other systems or features.
- the fan section 22 drives air along a bypass flow path B, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 .
- the hot combustion gases generated in the combustor section 26 are expanded through the turbine section 28 .
- the exemplary gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine centerline longitudinal axis A.
- the low speed spool 30 and the high speed spool 32 may be mounted relative to an engine static structure 33 via several bearing systems 31 . It should be understood that other bearing systems 31 may alternatively or additionally be provided, and the location of bearing systems 31 may be varied as appropriate to the application.
- the low speed spool 30 generally includes an inner shaft 34 that interconnects a fan 36 , a low pressure compressor 38 and a low pressure turbine 39 .
- the inner shaft 34 can be connected to the fan 36 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 45 , such as a fan drive gear system 50 (see FIGS. 2 and 3 ).
- the speed change mechanism drives the fan 36 at a lower speed than the low speed spool 30 .
- the high speed spool 32 includes an outer shaft 35 that interconnects a high pressure compressor 37 and a high pressure turbine 40 .
- the inner shaft 34 and the outer shaft 35 are supported at various axial locations by bearing systems 31 positioned within the engine static structure 33 .
- a combustor 42 is arranged in exemplary gas turbine 20 between the high pressure compressor 37 and the high pressure turbine 40 .
- a mid-turbine frame 44 may be arranged generally between the high pressure turbine 40 and the low pressure turbine 39 .
- the mid-turbine frame 44 can support one or more bearing systems 31 of the turbine section 28 .
- the mid-turbine frame 44 may include one or more airfoils 46 that extend within the core flow path C.
- gear system 50 may be located aft of combustor section 26 or even aft of turbine section 28
- fan section 22 may be positioned forward or aft of the location of gear system 50 .
- the inner shaft 34 and the outer shaft 35 are concentric and rotate via the bearing systems 31 about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes.
- the core airflow is compressed by the low pressure compressor 38 and the high pressure compressor 37 , is mixed with fuel and burned in the combustor 42 , and is then expanded over the high pressure turbine 40 and the low pressure turbine 39 .
- the high pressure turbine 40 and the low pressure turbine 39 rotationally drive the respective high speed spool 32 and the low speed spool 30 in response to the expansion.
- the gas turbine engine 20 is a high-bypass geared aircraft engine.
- the gas turbine engine 20 bypass ratio is greater than about six (6:1).
- the geared architecture 45 can include an epicyclic gear train, such as a planetary gear system, a star gear system, or other gear system.
- the geared architecture 45 enables operation of the low speed spool 30 at higher speeds, which can enable an increase in the operational efficiency of the low pressure compressor 38 and low pressure turbine 39 , and render increased pressure in a fewer number of stages.
- the pressure ratio of the low pressure turbine 39 can be pressure measured prior to the inlet of the low pressure turbine 39 as related to the pressure at the outlet of the low pressure turbine 39 and prior to an exhaust nozzle of the gas turbine engine 20 .
- the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 38
- the low pressure turbine 39 has a pressure ratio that is greater than about five (5:1).
- the bypass ratio is greater than 11 and less than 22, or greater than 13 and less than 20.
- the low pressure turbine 39 includes at least one stage and no more than eight stages, or at least three stages and no more than six stages. In another non-limiting embodiment, the low pressure turbine 39 includes at least three stages and no more than four stages.
- TSFC Thrust Specific Fuel Consumption
- Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system.
- the low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45. In another non-limiting embodiment of the example gas turbine engine 20 , the Fan Pressure Ratio is less than 1.38 and greater than 1.25. In another non-limiting embodiment, the fan pressure ratio is less than 1.48. In another non-limiting embodiment, the fan pressure ratio is less than 1.52. In another non-limiting embodiment, the fan pressure ratio is less than 1.7.
- Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram ° R)/(518.7°R)] 0.5 , where T represents the ambient temperature in degrees Rankine.
- the Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
- the Low Corrected Fan Tip Speed according to another non-limiting embodiment of the example gas turbine engine 20 is less than about 1400 fps (427 m/s).
- the Low Corrected Fan Tip Speed according to another non-limiting embodiment of the example gas turbine engine 20 is greater than about 1000 fps (305 m/s).
- FIG. 2 schematically illustrates the low speed spool 30 of the gas turbine engine 20 .
- the low speed spool 30 includes the fan 36 , the low pressure compressor 38 , and the low pressure turbine 39 .
- the inner shaft 34 interconnects the fan 36 , the low pressure compressor 38 , and the low pressure turbine 39 .
- the inner shaft 34 is connected to the fan 36 through the fan drive gear system 50 .
- the fan drive gear system 50 provides for counter-rotation of the low pressure turbine 39 and the fan 36 .
- the fan 36 rotates in a first direction D 1
- the low pressure turbine 39 rotates in a second direction D 2 that is opposite of the first direction D 1 .
- FIG. 3 illustrates one example embodiment of the fan drive gear system 50 incorporated into the gas turbine engine 20 to provide for counter-rotation of the fan 36 and the low pressure turbine 39 .
- the fan drive gear system 50 includes a star gear system with a sun gear 52 , a ring gear 54 disposed about the sun gear 52 , and a plurality of star gears 56 having journal bearings 57 positioned between the sun gear 52 and the ring gear 54 .
- a fixed carrier 58 carries and is attached to each of the star gears 56 . In this embodiment, the fixed carrier 58 does not rotate and is connected to a grounded structure 55 of the gas turbine engine 20 .
- the sun gear 52 receives an input from the low pressure turbine 39 (see FIG. 2 ) and rotates in the first direction D 1 thereby turning the plurality of star gears 56 in a second direction D 2 that is opposite of the first direction D 1 . Movement of the plurality of star gears 56 is transmitted to the ring gear 54 which rotates in the second direction D 2 opposite from the first direction D 1 of the sun gear 52 .
- the ring gear 54 is connected to the fan 36 for rotating the fan 36 (see FIG. 2 ) in the second direction D 2 .
- a star system gear ratio of the fan drive gear system 50 is determined by measuring a diameter of the ring gear 54 and dividing that diameter by a diameter of the sun gear 52 .
- the star system gear ratio of the geared architecture 45 is between 1.5 and 4.1.
- the system gear ratio of the fan drive gear system 50 is between 2.6 and 4.1.
- the star system gear ratio is below 1.5, the sun gear 52 is relatively much larger than the star gears 56 .
- This size differential reduces the load the star gears 56 are capable of carrying because of the reduction in size of the star gear journal bearings 57 .
- the sun gear 52 may be much smaller than the star gears 56 .
- This size differential increases the size of the star gear 56 journal bearings 57 but reduces the load the sun gear 52 is capable of carrying because of its reduced size and number of teeth.
- roller bearings could be used in place of journal bearings 57 .
- Improving performance of the gas turbine engine 20 begins by determining fan tip speed boundary conditions for at least one fan blade of the fan 36 to define the speed of the tip of the fan blade.
- the maximum fan diameter is determined based on the projected fuel burn derived from balancing engine efficiency, mass of air through the bypass flow path B, and engine weight increase due to the size of the fan blades.
- Boundary conditions are then determined for the rotor of each stage of the low pressure turbine 39 to define the speed of the rotor tip and to define the size of the rotor and the number of stages in the low pressure turbine 39 based on the efficiency of low pressure turbine 39 and the low pressure compressor 38 .
- FIG. 4 shows an embodiment 100 , wherein there is a fan drive turbine 108 driving a shaft 106 to in turn drive a fan rotor 102 .
- a gear reduction 104 may be positioned between the fan drive turbine 108 and the fan rotor 102 . This gear reduction 104 may be structured and operate like the geared architecture 45 disclosed above.
- a compressor rotor 110 is driven by an intermediate pressure turbine 112 , and a second stage compressor rotor 114 is driven by a turbine rotor 116 .
- a combustion section 118 is positioned intermediate the compressor rotor 114 and the turbine section 116 .
- FIG. 5 shows yet another embodiment 200 wherein a fan rotor 202 and a first stage compressor 204 rotate at a common speed.
- the gear reduction 206 (which may be structured as the geared architecture 45 disclosed above) is intermediate the compressor rotor 204 and a shaft 208 which is driven by a low pressure turbine section.
Abstract
Description
- This disclosure is a continuation of U.S. application Ser. No. 14/705,577, filed May 6, 2015, which is a continuation in part of PCT/US2013/061115 filed on Sep. 23, 2013, which claims priority to U.S. Provisional Patent Application No. 61/706,212 filed on Sep. 27, 2012.
- This disclosure relates to a gas turbine engine, and more particularly to a method for setting a gear ratio of a fan drive gear system of a gas turbine engine.
- A gas turbine engine may include a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. Among other variations, the compressor section can include low and high pressure compressors, and the turbine section can include low and high pressure turbines.
- Typically, a high pressure turbine drives a high pressure compressor through an outer shaft to form a high spool, and a low pressure turbine drives a low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the inner shaft. A direct drive gas turbine engine may include a fan section driven by the low spool such that a low pressure compressor, low pressure turbine, and fan section rotate at a common speed in a common direction.
- A speed reduction device, which may be a fan drive gear system or other mechanism, may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section. This allows for an overall increase in propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the speed reduction device that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.
- Although gas turbine engines utilizing speed change mechanisms are generally known to be capable of improved propulsive efficiency relative to conventional engines, gas turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.
- In a featured embodiment, a gas turbine engine has a fan section including a fan rotatable about an axis. A speed reduction device is connected to the fan. The speed reduction device includes a planetary fan drive gear system with a planet gear ratio of at least 2.6. A bypass ratio is greater than about 11.0.
- In another embodiment according to the previous embodiment, the gear ratio is less than or equal to 4.1.
- In another embodiment according to any of the previous embodiments, a fan pressure ratio is below 1.7.
- In another embodiment according to any of the previous embodiments, a fan pressure ratio is below 1.48.
- In another embodiment according to any of the previous embodiments, the fan blade tip speed of the fan section is greater than about 1000 ft/sec and less than about 1400 ft/sec.
- In another embodiment according to any of the previous embodiments, the planetary fan drive gear system includes a sun gear, a plurality of planetary gears, a ring gear, and a carrier.
- In another embodiment according to any of the previous embodiments, each of the plurality of planetary gears includes at least one bearing.
- In another embodiment according to any of the previous embodiments, a low pressure turbine is mechanically attached to the sun gear.
- In another embodiment according to any of the previous embodiments, a low pressure turbine section is in communication with the speed reduction device. The low pressure turbine section includes at least three stages.
- In another embodiment according to any of the previous embodiments, the bypass ratio is less than about 22.0.
- In another embodiment according to any of the previous embodiments, there are three turbine rotors with a first fan drive turbine rotor driving the fan through the speed reduction device, and an intermediate turbine rotor and a high pressure turbine rotor each driving a compressor rotor.
- In another featured embodiment, a method of improving performance of a gas turbine engine includes determining fan tip speed boundary conditions for at least one fan blade of a fan section. Rotor boundary conditions are determined for a rotor of a low pressure turbine. Stress level constraints are utilized in the rotor of the low pressure turbine and the at least one fan blade to determine if the rotary speed of the fan section and the low pressure turbine will meet a desired number of operating cycles. A bypass ratio is greater than about 6.0. A speed reduction device connects the fan section and the low pressure turbine and includes a planetary gear ratio of at least about 2.6.
- In another embodiment according to the previous embodiment, the planetary gear ratio is less than about 4.1.
- In another embodiment according to any of the previous embodiments, a fan pressure ratio is below 1.7.
- In another embodiment according to any of the previous embodiments, a fan pressure ratio is below 1.48.
- In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 11 and less than about 22.
- In another embodiment according to any of the previous embodiments, a fan blade tip speed of the at least one fan blade is less than 1400 fps.
- In another embodiment according to any of the previous embodiments, if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and the number of stages of the low pressure turbine is increased.
- In another embodiment according to any of the previous embodiments, if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and an annular area of the low pressure turbine is increased.
- In another featured embodiment, a fan drive gear module for a gas turbine engine includes a planetary fan drive gear system with a speed reduction ratio of at least 2.6. The speed reduction device is configured to drive a fan section with a bypass ratio greater than about 11.0.
- In another embodiment according to the previous embodiment, the speed reduction ratio is less than or equal to 4.1.
- In another embodiment according to any of the previous embodiments, the planetary fan drive gear system is configured to drive a fan section with a fan blade tip speed greater than about 1000 ft/sec and less than about 1400 ft/sec.
- In another embodiment according to any of the previous embodiments, the bypass ratio is less than about 22.0.
- In another featured embodiment, a method of designing a gas turbine engine includes selecting fan tip speed boundary conditions for at least one fan blade of a fan section of a gas turbine engine. Rotor boundary conditions are selected for a rotor of a fan drive turbine of the gas turbine engine. Stress level constraints are determined in the rotor of the fan drive turbine and the at least one fan blade to determine if the rotary speed of the fan section and the fan drive turbine will meet a desired number of operating cycles. A bypass ratio is greater than about 6.0. The fan section and the fan drive turbine are connected through a speed reduction device that includes a planetary gear ratio of at least about 2.6.
- In another embodiment according to the previous embodiment, the planetary gear ratio is less than about 4.1.
- In another embodiment according to any of the previous embodiments, the bypass ratio of the gas turbine engine is greater than about 11 and less than about 22.
- In another embodiment according to any of the previous embodiments, lowering the speed reduction ratio of the speed reduction device and increasing a number of stages of the fan drive turbine responsive to determine that a stress level in the rotor or the at least one fan blade is outside a predefined number of operating cycles.
- In another embodiment according to any of the previous embodiments, lowering the speed reduction ratio of the speed reduction device and increasing an annular area of the fan drive turbine responsive to determine that a stress level in the rotor or the at least one fan blade is outside a predefined number of operating cycles.
- These and other features may be best understood from the following drawings and specification.
-
FIG. 1 illustrates a schematic, cross-sectional view of an example gas turbine engine. -
FIG. 2 illustrates a schematic view of one configuration of a low speed spool that can be incorporated into a gas turbine engine. -
FIG. 3 illustrates a fan drive gear system that can be incorporated into a gas turbine engine. -
FIG. 4 shows another embodiment. -
FIG. 5 shows yet another embodiment. -
FIG. 1 schematically illustrates agas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C for compression and communication into thecombustor section 26. The hot combustion gases generated in thecombustor section 26 are expanded through theturbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to two-spool turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures. - The exemplary
gas turbine engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centerline longitudinal axis A. Thelow speed spool 30 and thehigh speed spool 32 may be mounted relative to an enginestatic structure 33 viaseveral bearing systems 31. It should be understood that other bearingsystems 31 may alternatively or additionally be provided, and the location of bearingsystems 31 may be varied as appropriate to the application. - The
low speed spool 30 generally includes aninner shaft 34 that interconnects afan 36, alow pressure compressor 38 and alow pressure turbine 39. Theinner shaft 34 can be connected to thefan 36 through a speed change mechanism, which in exemplarygas turbine engine 20 is illustrated as a gearedarchitecture 45, such as a fan drive gear system 50 (seeFIGS. 2 and 3 ). The speed change mechanism drives thefan 36 at a lower speed than thelow speed spool 30. Thehigh speed spool 32 includes anouter shaft 35 that interconnects ahigh pressure compressor 37 and ahigh pressure turbine 40. In this embodiment, theinner shaft 34 and theouter shaft 35 are supported at various axial locations by bearingsystems 31 positioned within the enginestatic structure 33. - A
combustor 42 is arranged inexemplary gas turbine 20 between thehigh pressure compressor 37 and thehigh pressure turbine 40. Amid-turbine frame 44 may be arranged generally between thehigh pressure turbine 40 and thelow pressure turbine 39. Themid-turbine frame 44 can support one ormore bearing systems 31 of theturbine section 28. Themid-turbine frame 44 may include one ormore airfoils 46 that extend within the core flow path C. It will be appreciated that each of the positions of thefan section 22,compressor section 24,combustor section 26,turbine section 28, and fandrive gear system 50 may be varied. For example,gear system 50 may be located aft ofcombustor section 26 or even aft ofturbine section 28, andfan section 22 may be positioned forward or aft of the location ofgear system 50. - The
inner shaft 34 and theouter shaft 35 are concentric and rotate via the bearingsystems 31 about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by thelow pressure compressor 38 and thehigh pressure compressor 37, is mixed with fuel and burned in thecombustor 42, and is then expanded over thehigh pressure turbine 40 and thelow pressure turbine 39. Thehigh pressure turbine 40 and thelow pressure turbine 39 rotationally drive the respectivehigh speed spool 32 and thelow speed spool 30 in response to the expansion. - In a non-limiting embodiment, the
gas turbine engine 20 is a high-bypass geared aircraft engine. In a further example, thegas turbine engine 20 bypass ratio is greater than about six (6:1). The gearedarchitecture 45 can include an epicyclic gear train, such as a planetary gear system, a star gear system, or other gear system. The gearedarchitecture 45 enables operation of thelow speed spool 30 at higher speeds, which can enable an increase in the operational efficiency of thelow pressure compressor 38 andlow pressure turbine 39, and render increased pressure in a fewer number of stages. - The pressure ratio of the
low pressure turbine 39 can be pressure measured prior to the inlet of thelow pressure turbine 39 as related to the pressure at the outlet of thelow pressure turbine 39 and prior to an exhaust nozzle of thegas turbine engine 20. In one non-limiting embodiment, the bypass ratio of thegas turbine engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 38, and thelow pressure turbine 39 has a pressure ratio that is greater than about five (5:1). In another non-limiting embodiment, the bypass ratio is greater than 11 and less than 22, or greater than 13 and less than 20. It should be understood, however, that the above parameters are only exemplary of a geared architecture engine or other engine using a speed change mechanism, and that the present disclosure is applicable to other gas turbine engines, including direct drive turbofans. In one non-limiting embodiment, thelow pressure turbine 39 includes at least one stage and no more than eight stages, or at least three stages and no more than six stages. In another non-limiting embodiment, thelow pressure turbine 39 includes at least three stages and no more than four stages. - In this embodiment of the exemplary
gas turbine engine 20, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. Thefan section 22 of thegas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with thegas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. - Fan Pressure Ratio is the pressure ratio across a blade of the
fan section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. In another non-limiting embodiment of the examplegas turbine engine 20, the Fan Pressure Ratio is less than 1.38 and greater than 1.25. In another non-limiting embodiment, the fan pressure ratio is less than 1.48. In another non-limiting embodiment, the fan pressure ratio is less than 1.52. In another non-limiting embodiment, the fan pressure ratio is less than 1.7. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram ° R)/(518.7°R)]0.5, where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the examplegas turbine engine 20 is less than about 1150 fps (351 m/s). The Low Corrected Fan Tip Speed according to another non-limiting embodiment of the examplegas turbine engine 20 is less than about 1400 fps (427 m/s). The Low Corrected Fan Tip Speed according to another non-limiting embodiment of the examplegas turbine engine 20 is greater than about 1000 fps (305 m/s). -
FIG. 2 schematically illustrates thelow speed spool 30 of thegas turbine engine 20. Thelow speed spool 30 includes thefan 36, thelow pressure compressor 38, and thelow pressure turbine 39. Theinner shaft 34 interconnects thefan 36, thelow pressure compressor 38, and thelow pressure turbine 39. Theinner shaft 34 is connected to thefan 36 through the fandrive gear system 50. In this embodiment, the fandrive gear system 50 provides for counter-rotation of thelow pressure turbine 39 and thefan 36. For example, thefan 36 rotates in a first direction D1, whereas thelow pressure turbine 39 rotates in a second direction D2 that is opposite of the first direction D1. -
FIG. 3 illustrates one example embodiment of the fandrive gear system 50 incorporated into thegas turbine engine 20 to provide for counter-rotation of thefan 36 and thelow pressure turbine 39. In this embodiment, the fandrive gear system 50 includes a star gear system with asun gear 52, aring gear 54 disposed about thesun gear 52, and a plurality of star gears 56 havingjournal bearings 57 positioned between thesun gear 52 and thering gear 54. Afixed carrier 58 carries and is attached to each of the star gears 56. In this embodiment, the fixedcarrier 58 does not rotate and is connected to a groundedstructure 55 of thegas turbine engine 20. - The
sun gear 52 receives an input from the low pressure turbine 39 (seeFIG. 2 ) and rotates in the first direction D1 thereby turning the plurality of star gears 56 in a second direction D2 that is opposite of the first direction D1. Movement of the plurality of star gears 56 is transmitted to thering gear 54 which rotates in the second direction D2 opposite from the first direction D1 of thesun gear 52. Thering gear 54 is connected to thefan 36 for rotating the fan 36 (seeFIG. 2 ) in the second direction D2. - A star system gear ratio of the fan
drive gear system 50 is determined by measuring a diameter of thering gear 54 and dividing that diameter by a diameter of thesun gear 52. In one embodiment, the star system gear ratio of the gearedarchitecture 45 is between 1.5 and 4.1. In another embodiment, the system gear ratio of the fandrive gear system 50 is between 2.6 and 4.1. When the star system gear ratio is below 1.5, thesun gear 52 is relatively much larger than the star gears 56. This size differential reduces the load the star gears 56 are capable of carrying because of the reduction in size of the stargear journal bearings 57. When the star system gear ratio is above 4.1, thesun gear 52 may be much smaller than the star gears 56. This size differential increases the size of thestar gear 56journal bearings 57 but reduces the load thesun gear 52 is capable of carrying because of its reduced size and number of teeth. Alternatively, roller bearings could be used in place ofjournal bearings 57. - Improving performance of the
gas turbine engine 20 begins by determining fan tip speed boundary conditions for at least one fan blade of thefan 36 to define the speed of the tip of the fan blade. The maximum fan diameter is determined based on the projected fuel burn derived from balancing engine efficiency, mass of air through the bypass flow path B, and engine weight increase due to the size of the fan blades. - Boundary conditions are then determined for the rotor of each stage of the
low pressure turbine 39 to define the speed of the rotor tip and to define the size of the rotor and the number of stages in thelow pressure turbine 39 based on the efficiency oflow pressure turbine 39 and thelow pressure compressor 38. - Constraints regarding stress levels in the rotor and the fan blade are utilized to determine if the rotary speed of the
fan 36 and thelow pressure turbine 39 will meet a desired number of operating life cycles. If the stress levels in the rotor or the fan blade are too high, the gear ratio of the fandrive gear system 50 can be lowered and the number of stages of thelow pressure turbine 39 or annular area of thelow pressure turbine 39 can be increased.FIG. 4 shows anembodiment 100, wherein there is afan drive turbine 108 driving ashaft 106 to in turn drive afan rotor 102. Agear reduction 104 may be positioned between thefan drive turbine 108 and thefan rotor 102. Thisgear reduction 104 may be structured and operate like the gearedarchitecture 45 disclosed above. Acompressor rotor 110 is driven by anintermediate pressure turbine 112, and a secondstage compressor rotor 114 is driven by aturbine rotor 116. Acombustion section 118 is positioned intermediate thecompressor rotor 114 and theturbine section 116. -
FIG. 5 shows yet anotherembodiment 200 wherein afan rotor 202 and afirst stage compressor 204 rotate at a common speed. The gear reduction 206 (which may be structured as the gearedarchitecture 45 disclosed above) is intermediate thecompressor rotor 204 and ashaft 208 which is driven by a low pressure turbine section. - Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.
- It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure.
- The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claim should be studied to determine the true scope and content of this disclosure.
Claims (28)
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US14/705,577 US20180073439A9 (en) | 2012-09-27 | 2015-05-06 | A fan drive gear system for driving a fan in a gas turbine engine having a high bypass ratio |
US14/742,954 US20150285155A1 (en) | 2012-09-27 | 2015-06-18 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
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US14/705,459 Abandoned US20150233301A1 (en) | 2012-09-27 | 2015-05-06 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US14/705,577 Abandoned US20180073439A9 (en) | 2012-09-27 | 2015-05-06 | A fan drive gear system for driving a fan in a gas turbine engine having a high bypass ratio |
US14/742,954 Abandoned US20150285155A1 (en) | 2012-09-27 | 2015-06-18 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US14/746,910 Abandoned US20150308335A1 (en) | 2012-09-27 | 2015-06-23 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US15/181,102 Abandoned US20160281610A1 (en) | 2012-09-27 | 2016-06-13 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US15/875,656 Abandoned US20180156135A1 (en) | 2012-09-27 | 2018-01-19 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US17/217,754 Abandoned US20210215101A1 (en) | 2012-09-27 | 2021-03-30 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
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US14/705,459 Abandoned US20150233301A1 (en) | 2012-09-27 | 2015-05-06 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US14/705,577 Abandoned US20180073439A9 (en) | 2012-09-27 | 2015-05-06 | A fan drive gear system for driving a fan in a gas turbine engine having a high bypass ratio |
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US14/746,910 Abandoned US20150308335A1 (en) | 2012-09-27 | 2015-06-23 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US15/181,102 Abandoned US20160281610A1 (en) | 2012-09-27 | 2016-06-13 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US15/875,656 Abandoned US20180156135A1 (en) | 2012-09-27 | 2018-01-19 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
US17/217,754 Abandoned US20210215101A1 (en) | 2012-09-27 | 2021-03-30 | Method for setting a gear ratio of a fan drive gear system of a gas turbine engine |
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2015
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US8807916B2 (en) | 2014-08-19 |
US20210215101A1 (en) | 2021-07-15 |
US20160281610A1 (en) | 2016-09-29 |
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US20150308335A1 (en) | 2015-10-29 |
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US20140193238A1 (en) | 2014-07-10 |
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US20150233303A1 (en) | 2015-08-20 |
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