US20150308351A1 - Fundamental gear system architecture - Google Patents
Fundamental gear system architecture Download PDFInfo
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- US20150308351A1 US20150308351A1 US14/745,802 US201514745802A US2015308351A1 US 20150308351 A1 US20150308351 A1 US 20150308351A1 US 201514745802 A US201514745802 A US 201514745802A US 2015308351 A1 US2015308351 A1 US 2015308351A1
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- gear
- gear system
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- fan drive
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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
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
- F01D25/162—Bearing supports
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/18—Lubricating arrangements
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/18—Lubricating arrangements
- F01D25/20—Lubricating arrangements using lubrication pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/06—Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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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
- F02C3/062—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 the turbine being of the radial-flow type
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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
- 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
- F02C7/06—Arrangements of bearings; Lubricating
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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
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
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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
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D19/00—Axial-flow pumps
- F04D19/02—Multi-stage pumps
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/325—Rotors specially for elastic fluids for axial flow pumps for axial flow fans
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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
- F16H1/48—Special means compensating for misalignment of axes, e.g. for equalising distribution of load on the face width of the teeth
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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
- F16H57/00—General details of gearing
- F16H57/04—Features relating to lubrication or cooling or heating
- F16H57/0412—Cooling or heating; Control of temperature
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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
- F16H57/00—General details of gearing
- F16H57/04—Features relating to lubrication or cooling or heating
- F16H57/048—Type of gearings to be lubricated, cooled or heated
- F16H57/0482—Gearings with gears having orbital motion
- F16H57/0486—Gearings with gears having orbital motion with fixed gear ratio
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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
- 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
- F05D2240/35—Combustors or associated equipment
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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
- 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
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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
- 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
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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
- F05D2260/00—Function
- F05D2260/98—Lubrication
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- a gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion 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 typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
- the high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool
- the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool.
- a speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine.
- the efficiency at which the gear assembly transfers power is a consideration in the development of a gear driven fan. Power or energy not transferred through the gearbox typically results in the generation of heat that is removed with a lubrication system. The more heat generated, the larger and heavier the lubrication system.
- a fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan.
- a lubrication system configures to provide lubricant to the gear system and to remove thermal energy produced by the gear system.
- the lubrication system includes a capacity for removing thermal energy greater than zero and less than about 2% of power input into the gear system during operation of the engine.
- the gear system transfers power input from the fan drive turbine to the fan at an efficiency greater than about 98% and less than 100%.
- the lubrication system includes a capacity for removing thermal energy equal to less than about 1% of power input into the gear system.
- the lubrication system includes a main lubrication system configured to provide lubricant to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.
- the gear system is flexibly supported for movement relative to a static structure of the engine.
- a load limiter is configured to limit movement of the gear system relative to the static structure of the engine responsive to an unbalanced condition.
- the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a non-rotatable carrier, a plurality of star gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of star gears.
- a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the carrier.
- the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a rotatable carrier, a plurality of planet gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.
- a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the ring gear.
- a gas turbine engine includes, among other possible things, a fan including a plurality of fan blades rotatable about an axis, a fan drive turbine, a gear system that provides a speed reduction between the fan drive turbine and the fan, the gear system configured to transfer power input from the fan drive turbine to the fan at an efficiency greater than about 98% and less than 100%, and a lubrication system configured to provide lubricant to the gear system and to remove thermal energy from the gear system produced by the gear system during operation of the engine.
- a second turbine rotor There is a second turbine rotor.
- the lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system during operation of the engine.
- the lubrication system includes a capacity for removing thermal energy equal to less than about 1% of power input into the gear system during operation of the engine.
- the lubrication system includes a main lubrication system configured to provide lubricant to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.
- the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a non-rotatable carrier, a plurality of star gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of star gears.
- the gear system is flexibly supported for accommodating movement relative to an engine static structure.
- a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between the engine static structure and the carrier.
- the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a rotatable carrier, a plurality of planet gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.
- the gear system is flexibly supported for accommodating movement relative to an engine static structure.
- a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between the engine static structure and the ring gear.
- the gear system includes a gear reduction having a gear ratio greater than about 2.3.
- the fan delivers a first portion of air into a bypass duct and a second portion of air into a compressor of the gas turbine engine.
- a bypass ratio which is defined as the first portion divided by the second portion, is greater than about 10.0.
- FIG. 1 is a schematic view of an example gas turbine engine.
- FIG. 2 is a schematic view of an example fan drive gear system including star epicyclical geared architecture.
- FIG. 3 is a schematic view of an example fan drive gear system including planetary epicyclical geared architecture.
- FIG. 4 shows another embodiment.
- FIG. 5 shows yet another embodiment.
- FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- Alternative engines might include an augmenter 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 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26 .
- the combustor section 26 air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24 .
- turbofan gas turbine engine depicts a turbofan gas turbine engine
- the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
- the example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 . It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
- the low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46 .
- the inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48 , to drive the fan 42 at a lower speed than the low speed spool 30 .
- the high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54 .
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
- a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54 .
- the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54 .
- the high pressure turbine 54 includes only a single stage.
- a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
- the example low pressure turbine 46 has a pressure ratio that is greater than about 5.
- the pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- a mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
- the mid-turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46 .
- the core airflow C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46 .
- the mid-turbine frame 58 includes vanes 60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46 . Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58 . Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28 . Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
- the disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine.
- the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10).
- the example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
- the gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
- the fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet.
- TFCT Thrust Specific Fuel Consumption
- Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45.
- Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/518.7) 0.5 ].
- the “Low corrected fan tip speed,” as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
- the example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section 22 includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about 6 turbine rotors schematically indicated at 34 . In another non-limiting example embodiment the low pressure turbine 46 includes about 3 turbine rotors. A ratio between the number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
- the example gas turbine engine includes a lubrication system 98 .
- the lubrication system 98 provides lubricant flow to the rotating components of the gas turbine engine including the bearing assemblies 38 and the geared architecture 48 .
- the lubrication system 98 further provides for the removal of heat generated in the various bearing systems and the geared architecture 48 .
- the example lubrication system 98 includes a main system 80 that provides lubrication during normal operating conditions of the gas turbine engine.
- An auxiliary system 82 is also included to supplement operation of the main lubrication system 80 .
- the size and weight of the lubrication system 90 is directly related to its capacity for removing heat from the geared architecture 48 . The greater the need for removal of heat, the larger and heavier the lubrication system 98 becomes. The amount of heat generated by the geared architecture 48 is therefore an important consideration in the configuration of a fan drive gear system.
- the example geared architecture 48 is part of a fan drive gear system 70 .
- the example geared architecture 48 comprises a gear assembly 65 that includes a sun gear 62 driven by a fan drive turbine 46 .
- the fan drive turbine is the low pressure turbine 46 .
- the sun gear 62 in turn drives intermediate gears 64 mounted on a carrier 74 by journal bearings.
- the carrier 74 is grounded to the static engine structure 36 and therefore the intermediate gears 64 do not orbit about the sun gear 62 .
- the intermediate gears 64 intermesh and drive a ring gear 66 coupled to a fan shaft 68 to drive the fan 42 .
- the gear assembly 65 is flexibly mounted such that it may be isolated from vibrational and transient movement that could disturb alignment between the gears 62 , 64 and 66 .
- flexible mounts 76 support the carrier 74 and accommodate relative movement between the gear assembly 65 and the static structure 36 .
- the example flexible mount 76 includes a spring rate that accommodates deflections that occur during normal operation of the fan drive gear system 70 .
- the flexible coupling 72 also includes a spring rate that allows a defined amount of deflection and misalignment such that components of the gear assembly 65 are not driven out of alignment.
- a load limiting device 78 is provided as part of the gear box mounting structure.
- the load limiter 78 constrains movement of the gear box 65 .
- the limiter 78 further provides a stop that reacts to unbalanced loads on the gear box 65 . Accordingly, the limiter prevents radial unbalanced loads and/or torsional overloads from damaging the gas turbine engine 20 .
- the example fan drive gear system 70 is supported by a lubrication system 98 .
- the lubrication system 98 provides for lubrication and cooling of the gears 62 , 64 and 66 along with bearings supporting rotation of the gears. It is desirable to circulate lubricant as quickly as possible to maintain a desired temperature. Power transmission efficiency through the gear box 65 is detrimentally affected by elevated temperatures.
- the lubricant system 98 includes a main system 80 that provides the desired lubricant flow through a plurality of conduits schematically illustrated by the line 88 to and from the gear box 65 .
- the main oil system 80 also transmits heat, schematically by arrows 92 , away from the gear box 65 to maintain a desired temperature.
- the lubrication system 98 also includes the auxiliary oil system 82 that supplies oil flow to the gear box 65 in response to a temporary interruption in lubricant flow from the main oil system 80 .
- the efficiency of the example gear box 65 and overall geared architecture 48 is a function of the power input, schematically indicated by arrow 94 , through the shaft 40 relative to power output, schematically indicated by arrows 96 , to the fan shaft 68 .
- Power input 94 compared to the amount of power output 96 is a measure of gear box efficiency.
- the example gear box 65 operates at an efficiency of greater than about 98%. In another disclosed example the example gear box 65 operates at an efficiency greater than about 99%.
- the disclosed efficiency is a measure of the amount of power 94 that is specifically transferred to the fan shaft 68 to rotate the fan 42 . Power that is not transmitted through the gear box 65 is lost as heat and reduces the overall efficiency of the fan drive gear system 70 . Any deficit between the input power 94 and output power 96 results in the generation of heat. Accordingly, in this example, the deficit of between 1-2% between the input power 94 and output power 96 generates heat. In other words, between 1% and 2% of the input power 94 is converted to heat energy that must be accommodated by the lubrication system 98 to maintain a working lubricant temperature within operational limits.
- the example lubricant system 98 provides for the removal of thermal energy equal to or less than about 2% of the input power 94 from the low pressure turbine 46 .
- the efficiency of the gear box 65 is greater than about 99% such that only 1% of power input from the low pressure turbine 46 is transferred into heat energy that must be handled by the lubricant system 98 .
- the main oil system includes a heat exchanger 90 that accommodates heat 92 that is generated within the gear box 65 .
- the heat exchanger 90 is an example of one element of the lubrication system 98 that is scaled to the desired capacity for removing thermal energy.
- other elements such as for example lubricant pumps, conduit size along with overall lubricant quantity within the lubrication system 98 would also be increased in size and weight to provide increased cooling capacity. Accordingly, it is desirable to increase power transfer efficiency to reduce required overall heat transfer capacity of lubrication system 98 .
- the high efficiency of the example gear box 65 enables a relatively small and light lubricant system 98 .
- the example lubricant system 98 includes features that can accommodate thermal energy generated by no more than about 2% of the input power 94 .
- the lubrication system 98 has an overall maximum capacity for removing thermal energy equal to no more than about 2% of the input power provided by the low pressure turbine 46 .
- Lubrication systems that are required to remove greater than about 2% of input power 94 require larger lubricant systems 98 that can detrimentally impact overall engine efficiency and detract from the propulsion efficiencies provided by the reduction in fan speed.
- another example epicyclical gear box 85 comprises a planetary configuration.
- planet gears 84 are supported on a carrier 86 that is rotatable about the engine axis A.
- the sun gear 62 remains driven by the inner shaft 40 and the low pressure turbine 46 .
- the ring gear 66 is mounted to a fixed structure 36 such that it does not rotate about the axis. Accordingly, rotation of the sun gear 62 drives the planet gears 84 within the ring gear 66 .
- the planet gears 84 are supported on the rotatable carrier 86 that in turn drives the fan shaft 68 .
- the fan shaft 68 and the sun gear 62 rotate in a common direction, while the planet gears 84 individually rotate in a direction opposite to the sun gear 62 but collectively rotate about the sun gear 62 in the same direction as the rotation of the sun gear 62 .
- the example planetary gear box illustrated in FIG. 3 includes the ring gear 66 that is supported by flexible mount 76 .
- the flexible mount 76 allows some movement of the gearbox 85 to maintain a desired alignment between meshing teeth of the gears 62 , 84 , 66 .
- the limiter 78 prevents movement of the planetary gear box 85 beyond desired limits to prevent potential damage caused by radial imbalances and/or torsional loads.
- the example low pressure turbine 46 inputs power 94 to drive the gear box 85 .
- the example gear box 85 transmits more than about 98% of the input power 94 to the fan drive shaft 68 as output power 96 .
- the gear box 85 transmits more than about 99% of the input power 94 to the fan drive shaft 68 as output power 96 .
- the difference between the input power 94 and the output power 96 is converted into heat energy that is removed by the lubrication system 98 .
- the lubrication system 98 has a capacity of removing no more heat 92 than is generated by about 2% of the input power 94 from the low pressure turbine 46 .
- the lubrication system 98 has a capacity of removing no more heat 92 than is generated by about 1% of the input power 94 . Accordingly, the efficiency provided by the example gear box 85 enables the lubrication system 98 to be of size that does not detract from the propulsive efficiency realized by turning the fan section 22 and low pressure turbine 46 at separate and nearer optimal speeds.
- the example fan drive gear system provides for the improvement and realization of propulsive efficiencies by limiting losses in the form of thermal energy, thereby enabling utilization of a lower capacity and sized lubrication system.
- FIG. 4 shows an embodiment 200 , wherein there is a fan drive turbine 208 driving a shaft 206 to in turn drive a fan rotor 202 .
- a gear reduction 204 may be positioned between the fan drive turbine 208 and the fan rotor 202 .
- This gear reduction 204 may be structured and operate like the gear reduction disclosed above.
- a compressor rotor 210 is driven by an intermediate pressure turbine 212
- a second stage compressor rotor 214 is driven by a turbine rotor 216 .
- a combustion section 218 is positioned intermediate the compressor rotor 214 and the turbine section 216 .
- FIG. 5 shows yet another embodiment 300 wherein a fan rotor 302 and a first stage compressor 304 rotate at a common speed.
- the gear reduction 306 (which may be structured as disclosed above) is intermediate the compressor rotor 304 and a shaft 308 which is driven by a low pressure turbine section.
Abstract
A fan drive gear system for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a gear system that provides a speed reduction between a fan drive turbine and a fan. A lubrication system configures to provide lubricant to the gear system and to remove thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy greater than zero and less than about 2% of power input into the gear system during operation of the engine. There is a second turbine rotor. A gas turbine engine is also disclosed.
Description
- This application is a continuation-in-part of U.S. patent application Ser. No. 14/190,159, filed Feb. 26, 2014, which is a continuation of International Application No. PCT/US2013/041761 filed May 20, 2013 that claims priority to U.S. Provisional Application No. 61/653,731 filed May 31, 2012 and U.S. patent application Ser. No. 13/557,614 filed Jul. 25, 2012, now U.S. Pat. No. 8,572,943 granted Nov. 5, 2013.
- A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion 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 typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
- The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. The efficiency at which the gear assembly transfers power is a consideration in the development of a gear driven fan. Power or energy not transferred through the gearbox typically results in the generation of heat that is removed with a lubrication system. The more heat generated, the larger and heavier the lubrication system.
- Although geared architectures can provide improved propulsive efficiency, other factors including heat removal and lubrication can detract from the improved propulsive efficiency. Accordingly, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.
- In a featured embodiment, a fan drive gear system for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a gear system that provides a speed reduction between a fan drive turbine and a fan. A lubrication system configures to provide lubricant to the gear system and to remove thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy greater than zero and less than about 2% of power input into the gear system during operation of the engine. There is a second turbine rotor.
- In a further embodiment of the previous embodiment, the gear system transfers power input from the fan drive turbine to the fan at an efficiency greater than about 98% and less than 100%.
- In a further embodiment of the previous embodiment, the lubrication system includes a capacity for removing thermal energy equal to less than about 1% of power input into the gear system.
- In a further embodiment of the previous embodiment, the lubrication system includes a main lubrication system configured to provide lubricant to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.
- In a further embodiment of the previous embodiment, the gear system is flexibly supported for movement relative to a static structure of the engine.
- In a further embodiment of the previous embodiment, a load limiter is configured to limit movement of the gear system relative to the static structure of the engine responsive to an unbalanced condition.
- In a further embodiment of the previous embodiment, the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a non-rotatable carrier, a plurality of star gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of star gears.
- In a further embodiment of the previous embodiment, a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the carrier.
- In a further embodiment of the previous embodiment, the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a rotatable carrier, a plurality of planet gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.
- In a further embodiment of the previous embodiment, a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the ring gear.
- In another featured embodiment, a gas turbine engine includes, among other possible things, a fan including a plurality of fan blades rotatable about an axis, a fan drive turbine, a gear system that provides a speed reduction between the fan drive turbine and the fan, the gear system configured to transfer power input from the fan drive turbine to the fan at an efficiency greater than about 98% and less than 100%, and a lubrication system configured to provide lubricant to the gear system and to remove thermal energy from the gear system produced by the gear system during operation of the engine. There is a second turbine rotor.
- In a further embodiment of the previous embodiment, the lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system during operation of the engine.
- In a further embodiment of the previous embodiment, the lubrication system includes a capacity for removing thermal energy equal to less than about 1% of power input into the gear system during operation of the engine.
- In a further embodiment of the previous embodiment, the lubrication system includes a main lubrication system configured to provide lubricant to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.
- In a further embodiment of the previous embodiment, the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a non-rotatable carrier, a plurality of star gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of star gears. The gear system is flexibly supported for accommodating movement relative to an engine static structure.
- In a further embodiment of the previous embodiment, a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between the engine static structure and the carrier.
- In a further embodiment of the previous embodiment, the gear system includes a sun gear that is configured to be driven by the fan drive turbine, a rotatable carrier, a plurality of planet gears that are supported on the carrier and that are configured to be driven by the sun gear, and a ring gear circumscribing the plurality of planet gears. The gear system is flexibly supported for accommodating movement relative to an engine static structure.
- In a further embodiment of the previous embodiment, a first flexible coupling is provided between an input shaft that is configured to be driven by the fan drive turbine and the sun gear, and a second flexible coupling between the engine static structure and the ring gear.
- In a further embodiment of the previous embodiment, the gear system includes a gear reduction having a gear ratio greater than about 2.3.
- In a further embodiment of the previous embodiment, the fan delivers a first portion of air into a bypass duct and a second portion of air into a compressor of the gas turbine engine. A bypass ratio, which is defined as the first portion divided by the second portion, is greater than about 10.0.
- In a further embodiment of the previous embodiment, there is a third turbine rotor and said fan drive turbine is a most downstream of the three turbine rotors.
- Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
- These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description.
-
FIG. 1 is a schematic view of an example gas turbine engine. -
FIG. 2 is a schematic view of an example fan drive gear system including star epicyclical geared architecture. -
FIG. 3 is a schematic view of an example fan drive gear system including planetary epicyclical geared architecture. -
FIG. 4 shows another embodiment. -
FIG. 5 shows yet another embodiment. -
FIG. 1 schematically illustrates an examplegas turbine engine 20 that includes afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B while thecompressor section 24 draws air in along a core flow path C where air is compressed and communicated to acombustor section 26. In thecombustor section 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through theturbine section 28 where energy is extracted and utilized to drive thefan section 22 and thecompressor section 24. - Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
- The
example engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing systems 38. It should be understood thatvarious bearing systems 38 at various locations may alternatively or additionally be provided. - The
low speed spool 30 generally includes aninner shaft 40 that connects afan 42 and a low pressure (or first)compressor section 44 to a low pressure (or first)turbine section 46. Theinner shaft 40 drives thefan 42 through a speed change device, such as a gearedarchitecture 48, to drive thefan 42 at a lower speed than thelow speed spool 30. The high-speed spool 32 includes anouter shaft 50 that interconnects a high pressure (or second)compressor section 52 and a high pressure (or second)turbine section 54. Theinner shaft 40 and theouter shaft 50 are concentric and rotate via the bearingsystems 38 about the engine central longitudinal axis A. - A
combustor 56 is arranged between thehigh pressure compressor 52 and thehigh pressure turbine 54. In one example, thehigh pressure turbine 54 includes at least two stages to provide a double stagehigh pressure turbine 54. In another example, thehigh pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. - The example
low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the examplelow pressure turbine 46 is measured prior to an inlet of thelow pressure turbine 46 as related to the pressure measured at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle. - A
mid-turbine frame 58 of the enginestatic structure 36 is arranged generally between thehigh pressure turbine 54 and thelow pressure turbine 46. Themid-turbine frame 58 furthersupports bearing systems 38 in theturbine section 28 as well as setting airflow entering thelow pressure turbine 46. - The core airflow C is compressed by the
low pressure compressor 44 then by thehigh pressure compressor 52 mixed with fuel and ignited in thecombustor 56 to produce high speed exhaust gases that are then expanded through thehigh pressure turbine 54 andlow pressure turbine 46. Themid-turbine frame 58 includesvanes 60, which are in the core airflow path and function as an inlet guide vane for thelow pressure turbine 46. Utilizing thevane 60 of themid-turbine frame 58 as the inlet guide vane forlow pressure turbine 46 decreases the length of thelow pressure turbine 46 without increasing the axial length of themid-turbine frame 58. Reducing or eliminating the number of vanes in thelow pressure turbine 46 shortens the axial length of theturbine section 28. Thus, the compactness of thegas turbine engine 20 is increased and a higher power density may be achieved. - The disclosed
gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, thegas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. - In one disclosed embodiment, the
gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of thelow pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. - A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 22 of theengine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. - “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45.
- “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/518.7)0.5]. The “Low corrected fan tip speed,” as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
- The example gas turbine engine includes the
fan 42 that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, thefan section 22 includes less than about 20 fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about 6 turbine rotors schematically indicated at 34. In another non-limiting example embodiment thelow pressure turbine 46 includes about 3 turbine rotors. A ratio between the number offan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The examplelow pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in thelow pressure turbine 46 and the number ofblades 42 in thefan section 22 disclose an examplegas turbine engine 20 with increased power transfer efficiency. - The example gas turbine engine includes a
lubrication system 98. Thelubrication system 98 provides lubricant flow to the rotating components of the gas turbine engine including thebearing assemblies 38 and the gearedarchitecture 48. Thelubrication system 98 further provides for the removal of heat generated in the various bearing systems and the gearedarchitecture 48. - The
example lubrication system 98 includes amain system 80 that provides lubrication during normal operating conditions of the gas turbine engine. Anauxiliary system 82 is also included to supplement operation of themain lubrication system 80. The size and weight of thelubrication system 90 is directly related to its capacity for removing heat from the gearedarchitecture 48. The greater the need for removal of heat, the larger and heavier thelubrication system 98 becomes. The amount of heat generated by the gearedarchitecture 48 is therefore an important consideration in the configuration of a fan drive gear system. - Referring to
FIG. 2 with continued reference toFIG. 1 , the example gearedarchitecture 48 is part of a fandrive gear system 70. The example gearedarchitecture 48 comprises agear assembly 65 that includes asun gear 62 driven by afan drive turbine 46. In this example, the fan drive turbine is thelow pressure turbine 46. Thesun gear 62 in turn drives intermediate gears 64 mounted on a carrier 74 by journal bearings. The carrier 74 is grounded to thestatic engine structure 36 and therefore the intermediate gears 64 do not orbit about thesun gear 62. The intermediate gears 64 intermesh and drive aring gear 66 coupled to afan shaft 68 to drive thefan 42. - The
gear assembly 65 is flexibly mounted such that it may be isolated from vibrational and transient movement that could disturb alignment between thegears flexible mounts 76 support the carrier 74 and accommodate relative movement between thegear assembly 65 and thestatic structure 36. The exampleflexible mount 76 includes a spring rate that accommodates deflections that occur during normal operation of the fandrive gear system 70. - Power input through the
inner shaft 40 of thefan drive turbine 46 is transmitted through aflexible coupling 72. Theflexible coupling 72 also includes a spring rate that allows a defined amount of deflection and misalignment such that components of thegear assembly 65 are not driven out of alignment. - Although some relative movement is compensated by the
flexible coupling 72 and theflexible mounts 76, movement beyond a desired limitation can detrimentally affect meshing engagement between the gears and therefore aload limiting device 78 is provided as part of the gear box mounting structure. Theload limiter 78 constrains movement of thegear box 65. Thelimiter 78 further provides a stop that reacts to unbalanced loads on thegear box 65. Accordingly, the limiter prevents radial unbalanced loads and/or torsional overloads from damaging thegas turbine engine 20. - The example fan
drive gear system 70 is supported by alubrication system 98. Thelubrication system 98 provides for lubrication and cooling of thegears gear box 65 is detrimentally affected by elevated temperatures. - In this example, the
lubricant system 98 includes amain system 80 that provides the desired lubricant flow through a plurality of conduits schematically illustrated by theline 88 to and from thegear box 65. Themain oil system 80 also transmits heat, schematically byarrows 92, away from thegear box 65 to maintain a desired temperature. - The
lubrication system 98 also includes theauxiliary oil system 82 that supplies oil flow to thegear box 65 in response to a temporary interruption in lubricant flow from themain oil system 80. - The efficiency of the
example gear box 65 and overall gearedarchitecture 48 is a function of the power input, schematically indicated byarrow 94, through theshaft 40 relative to power output, schematically indicated byarrows 96, to thefan shaft 68.Power input 94 compared to the amount ofpower output 96 is a measure of gear box efficiency. Theexample gear box 65 operates at an efficiency of greater than about 98%. In another disclosed example theexample gear box 65 operates at an efficiency greater than about 99%. - The disclosed efficiency is a measure of the amount of
power 94 that is specifically transferred to thefan shaft 68 to rotate thefan 42. Power that is not transmitted through thegear box 65 is lost as heat and reduces the overall efficiency of the fandrive gear system 70. Any deficit between theinput power 94 andoutput power 96 results in the generation of heat. Accordingly, in this example, the deficit of between 1-2% between theinput power 94 andoutput power 96 generates heat. In other words, between 1% and 2% of theinput power 94 is converted to heat energy that must be accommodated by thelubrication system 98 to maintain a working lubricant temperature within operational limits. - The
example lubricant system 98 provides for the removal of thermal energy equal to or less than about 2% of theinput power 94 from thelow pressure turbine 46. In another non-limiting embodiment of the example fandrive gear system 70, the efficiency of thegear box 65 is greater than about 99% such that only 1% of power input from thelow pressure turbine 46 is transferred into heat energy that must be handled by thelubricant system 98. - As appreciated, the larger the capacity for handling and removing thermal energy, the larger and heavier the
lubricant system 98. In this example, the main oil system includes aheat exchanger 90 that accommodatesheat 92 that is generated within thegear box 65. Theheat exchanger 90 is an example of one element of thelubrication system 98 that is scaled to the desired capacity for removing thermal energy. As appreciated, other elements, such as for example lubricant pumps, conduit size along with overall lubricant quantity within thelubrication system 98 would also be increased in size and weight to provide increased cooling capacity. Accordingly, it is desirable to increase power transfer efficiency to reduce required overall heat transfer capacity oflubrication system 98. - In this example, the high efficiency of the
example gear box 65 enables a relatively small andlight lubricant system 98. Theexample lubricant system 98 includes features that can accommodate thermal energy generated by no more than about 2% of theinput power 94. In other words, thelubrication system 98 has an overall maximum capacity for removing thermal energy equal to no more than about 2% of the input power provided by thelow pressure turbine 46. - Greater amounts of capacity for removal of thermal energy results in an overall increase in the size and weight of the
lubrication system 98. Lubrication systems that are required to remove greater than about 2% ofinput power 94 requirelarger lubricant systems 98 that can detrimentally impact overall engine efficiency and detract from the propulsion efficiencies provided by the reduction in fan speed. - Referring to
FIG. 3 with continued reference toFIG. 1 , another exampleepicyclical gear box 85 is disclosed and comprises a planetary configuration. In a planetary configuration, planet gears 84 are supported on acarrier 86 that is rotatable about the engine axis A. Thesun gear 62 remains driven by theinner shaft 40 and thelow pressure turbine 46. Thering gear 66 is mounted to a fixedstructure 36 such that it does not rotate about the axis. Accordingly, rotation of thesun gear 62 drives the planet gears 84 within thering gear 66. The planet gears 84 are supported on therotatable carrier 86 that in turn drives thefan shaft 68. In this configuration, thefan shaft 68 and thesun gear 62 rotate in a common direction, while the planet gears 84 individually rotate in a direction opposite to thesun gear 62 but collectively rotate about thesun gear 62 in the same direction as the rotation of thesun gear 62. - The example planetary gear box illustrated in
FIG. 3 includes thering gear 66 that is supported byflexible mount 76. Theflexible mount 76 allows some movement of thegearbox 85 to maintain a desired alignment between meshing teeth of thegears limiter 78 prevents movement of theplanetary gear box 85 beyond desired limits to prevent potential damage caused by radial imbalances and/or torsional loads. - The example
low pressure turbine 46inputs power 94 to drive thegear box 85. As in the previous example, theexample gear box 85 transmits more than about 98% of theinput power 94 to thefan drive shaft 68 asoutput power 96. In another example, thegear box 85 transmits more than about 99% of theinput power 94 to thefan drive shaft 68 asoutput power 96. - The difference between the
input power 94 and theoutput power 96 is converted into heat energy that is removed by thelubrication system 98. In this example, thelubrication system 98 has a capacity of removing nomore heat 92 than is generated by about 2% of theinput power 94 from thelow pressure turbine 46. In another example. Thelubrication system 98 has a capacity of removing nomore heat 92 than is generated by about 1% of theinput power 94. Accordingly, the efficiency provided by theexample gear box 85 enables thelubrication system 98 to be of size that does not detract from the propulsive efficiency realized by turning thefan section 22 andlow pressure turbine 46 at separate and nearer optimal speeds. - Accordingly the example fan drive gear system provides for the improvement and realization of propulsive efficiencies by limiting losses in the form of thermal energy, thereby enabling utilization of a lower capacity and sized lubrication system.
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FIG. 4 shows anembodiment 200, wherein there is afan drive turbine 208 driving ashaft 206 to in turn drive afan rotor 202. Agear reduction 204 may be positioned between thefan drive turbine 208 and thefan rotor 202. Thisgear reduction 204 may be structured and operate like the gear reduction disclosed above. Acompressor rotor 210 is driven by anintermediate pressure turbine 212, and a second stage compressor rotor 214 is driven by a turbine rotor 216. A combustion section 218 is positioned intermediate the compressor rotor 214 and the turbine section 216. -
FIG. 5 shows yet anotherembodiment 300 wherein afan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction 306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and ashaft 308 which is driven by a low pressure turbine section. - Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.
Claims (21)
1. A fan drive gear system for a gas turbine engine comprising:
a gear system configured to provide a speed reduction between a fan drive turbine and a fan;
a mount flexibly supporting portions of the gear system radially extending from a static structure of the gas turbine engine with respect to a central axis to accommodate radial movement between the gear system and the static structure;
a lubrication system configured to provide lubricant to the gear system and remove thermal energy produced by the gear system, wherein the lubrication system includes a maximum capacity for removing thermal energy from the gear system greater than zero and less than about 2% of power input into the gear system during operation of the engine; and
there being a second turbine rotor.
2. The fan drive gear system as recited in claim 1 , wherein the gear system is configured to transfer power input from the fan drive turbine to the fan at an efficiency greater than about 98% to less 100%.
3. The fan drive gear system as recited in claim 1 , wherein the lubrication system includes a capacity for removing thermal energy equal to less than about 1% of power input into the gear system.
4. The fan drive gear system as recited in claim 1 , wherein the lubrication system comprises a main lubrication system configured to provide lubricant flow to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.
5. The fan drive gear system as recited in claim 1 , wherein the mount includes a load limiter for limiting movement of the gear system responsive to an unbalanced condition.
6. The fan drive gear system as recited in claim 1 , wherein the gear system comprises a sun gear driven by the fan drive turbine, a non-rotatable carrier, a plurality of star gears supported on the carrier and driven by the sun gear and a ring gear circumscribing the plurality of star gears.
7. The fan drive gear system as recited in claim 6 , wherein the mount includes a first flexible coupling between an input shaft driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the carrier.
8. The fan drive gear system as recited in claim 1 , wherein the gear system comprises a sun gear driven by the fan drive turbine, a rotatable carrier, a plurality of planet gears supported on the carrier and driven by the sun gear, and a ring gear circumscribing the plurality of planet gears.
9. The fan drive gear system as recited in claim 8 , wherein the mount includes a first flexible coupling between an input shaft driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the ring gear.
10. A gas turbine engine comprising:
a fan including a plurality of fan blades rotatable about an axis;
a compressor section;
a combustor in fluid communication with the compressor section;
a fan drive turbine in communication with the combustor;
a gear system is configured to: (a) provide a speed reduction between the fan drive turbine and the fan; and (b), transfer power input from the fan drive turbine to the fan at an efficiency greater than about 98% to less 100%;
a mount flexibly supporting the gear system radially extending from a static structure of the engine with respect to a central axis to accommodate radial movement between the gear system and the static structure;
a lubrication system configured to provide lubricant to the gear system and remove removing thermal energy from the gear system produced by the gear system; and
there being a second turbine rotor.
11. The gas turbine engine as recited in claim 10 , wherein the lubrication system includes a maximum capacity for removing thermal energy generated by the gear system greater than zero and less than about 2% of power input into the gear system during operation of the engine.
12. The gas turbine engine as recited in claim 10 , wherein the lubrication system includes a capacity for removing thermal energy greater than zero to less than about 1% of power input into the gear system.
13. The gas turbine engine as recited in claim 10 , wherein the lubrication system comprises a main lubrication system configured to provide lubricant flow to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.
14. The gas turbine engine as recited in claim 10 , wherein the gear system comprises a sun gear driven by the fan drive turbine, a non-rotatable carrier, a plurality of star gears supported on the carrier and driven by the sun gear and a ring gear circumscribing the plurality of star gears and the mount includes a first flexible coupling between an input shaft driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the carrier.
15. The gas turbine engine as recited in claim 10 , wherein the gear system comprises a sun gear driven by the fan drive turbine, a rotatable carrier, a plurality of planet gears supported on the carrier and driven by the sun gear, and a ring gear circumscribing the plurality of planet gears and the mount includes a first flexible coupling between an input shaft driven by the fan drive turbine and the sun gear, and a second flexible coupling between a fixed structure and the ring gear.
16. The gas turbine engine as recited in claim 10 , wherein the mount includes a load limiter for limiting movement of the gear system responsive to an unbalanced condition.
17. The gas turbine engine as recited in claim 10 , wherein the gear system comprises a gear reduction having a gear ratio greater than about 2.3.
18. The gas turbine engine as recited in claim 10 , wherein said fan delivers a portion of air into a bypass duct, and a bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section, with the bypass ratio being greater than about 6.0.
19. The gas turbine engine as recited in claim 10 , wherein a fan pressure ratio across the fan is less than about 1.5.
20. The gas turbine engine as recited in claim 10 , wherein said fan has 26 or fewer blades.
21. The gas turbine engine as recited in claim 10 , wherein there is a third turbine rotor and said fan drive turbine is a most downstream of the three turbine rotors.
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US15/379,619 US9840969B2 (en) | 2012-05-31 | 2016-12-15 | Gear system architecture for gas turbine engine |
US15/379,616 US20170096937A1 (en) | 2012-05-31 | 2016-12-15 | Gear system architecture for gas turbine engine |
US15/379,634 US20170096944A1 (en) | 2012-05-31 | 2016-12-15 | Gear system architecture for gas turbine engine |
US15/606,246 US20170260911A1 (en) | 2012-05-31 | 2017-05-26 | Gear system architecture for gas turbine engine |
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US15/872,158 US20180135524A1 (en) | 2012-05-31 | 2018-01-16 | Gear system architecture for gas turbine engine |
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Cited By (23)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20170096937A1 (en) * | 2012-05-31 | 2017-04-06 | United Technologies Corporation | Gear system architecture for gas turbine engine |
EP3415728A1 (en) * | 2017-06-13 | 2018-12-19 | General Electric Company | Gas turbine engine with rotating reversing compound gearbox |
US10221770B2 (en) | 2012-05-31 | 2019-03-05 | United Technologies Corporation | Fundamental gear system architecture |
CN109469724A (en) * | 2017-09-08 | 2019-03-15 | 劳斯莱斯有限公司 | Auxiliary feed-oil equipment for rotary components |
EP3460199A1 (en) * | 2017-09-20 | 2019-03-27 | General Electric Company | Lube system for geared turbine section |
US20190309705A1 (en) * | 2018-04-05 | 2019-10-10 | United Technologies Corporation | Aft counter-rotating boundary layer ingestion engine |
EP3557009A1 (en) * | 2018-04-17 | 2019-10-23 | Rolls-Royce plc | Lubrication system and method of lubrication |
US10590854B2 (en) * | 2016-01-26 | 2020-03-17 | United Technologies Corporation | Geared gas turbine engine |
US11174796B2 (en) * | 2019-04-12 | 2021-11-16 | Rolls-Royce Plc | Accessory gearbox assembly |
US11209164B1 (en) | 2020-12-18 | 2021-12-28 | Delavan Inc. | Fuel injector systems for torch igniters |
US11226103B1 (en) | 2020-12-16 | 2022-01-18 | Delavan Inc. | High-pressure continuous ignition device |
US11286862B1 (en) | 2020-12-18 | 2022-03-29 | Delavan Inc. | Torch injector systems for gas turbine combustors |
US20220186668A1 (en) * | 2020-12-16 | 2022-06-16 | Delavan Inc. | Continuous ignition device exhaust manifold |
US20220195934A1 (en) * | 2020-12-17 | 2022-06-23 | Delavan Inc. | Axially oriented internally mounted continuous ignition device: removable hot surface igniter |
US11428160B2 (en) | 2020-12-31 | 2022-08-30 | General Electric Company | Gas turbine engine with interdigitated turbine and gear assembly |
US11473505B2 (en) | 2020-11-04 | 2022-10-18 | Delavan Inc. | Torch igniter cooling system |
US11608783B2 (en) | 2020-11-04 | 2023-03-21 | Delavan, Inc. | Surface igniter cooling system |
US11635027B2 (en) | 2020-11-18 | 2023-04-25 | Collins Engine Nozzles, Inc. | Fuel systems for torch ignition devices |
US11635210B2 (en) | 2020-12-17 | 2023-04-25 | Collins Engine Nozzles, Inc. | Conformal and flexible woven heat shields for gas turbine engine components |
US11680528B2 (en) | 2020-12-18 | 2023-06-20 | Delavan Inc. | Internally-mounted torch igniters with removable igniter heads |
US11692488B2 (en) | 2020-11-04 | 2023-07-04 | Delavan Inc. | Torch igniter cooling system |
US11754289B2 (en) | 2020-12-17 | 2023-09-12 | Delavan, Inc. | Axially oriented internally mounted continuous ignition device: removable nozzle |
US20230340911A1 (en) * | 2022-04-25 | 2023-10-26 | General Electric Company | Mounting assembly for a gearbox assembly |
Families Citing this family (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9677659B1 (en) * | 2016-01-28 | 2017-06-13 | General Electric Company | Gearbox planet attenuation spring damper |
US10787930B2 (en) | 2018-03-23 | 2020-09-29 | Raytheon Technologies Corporation | Windmill lubrication gear train for lubricant system in a geared gas turbine engine |
US20190323597A1 (en) * | 2018-04-20 | 2019-10-24 | United Technologies Corporation | Electric motor driven auxiliary oil system for geared gas turbine engine |
MX2021003190A (en) * | 2018-09-28 | 2021-05-27 | Kimberly Clark Co | Embossed multi-ply tissue product. |
CN110701288B (en) * | 2019-09-12 | 2020-12-25 | 珠海格力电器股份有限公司 | RV reducer for displaying lubricating states of cycloidal teeth and pin teeth and displaying method |
US11702990B2 (en) | 2021-09-08 | 2023-07-18 | Rolls-Royce North American Technologies Inc. | Redundant electrically driven fuel and oil pumping system for gas turbine with bidirectional pump motor |
US11629640B2 (en) | 2021-09-08 | 2023-04-18 | Rolls-Royce North American Technologies Inc. | Oil pumping control for electrical oil pumping system |
Citations (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6223616B1 (en) * | 1999-12-22 | 2001-05-01 | United Technologies Corporation | Star gear system with lubrication circuit and lubrication method therefor |
US20050257528A1 (en) * | 2004-05-19 | 2005-11-24 | Dunbar Donal S Jr | Retractable afterburner for jet engine |
US20080148708A1 (en) * | 2006-12-20 | 2008-06-26 | General Electric Company | Turbine engine system with shafts for improved weight and vibration characteristic |
US20100105516A1 (en) * | 2006-07-05 | 2010-04-29 | United Technologies Corporation | Coupling system for a star gear train in a gas turbine engine |
US20100296947A1 (en) * | 2009-05-22 | 2010-11-25 | United Technologies Corporation | Apparatus and method for providing damper liquid in a gas turbine |
US8572943B1 (en) * | 2012-05-31 | 2013-11-05 | United Technologies Corporation | Fundamental gear system architecture |
Family Cites Families (257)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3111005A (en) | 1963-11-19 | Jet propulsion plant | ||
US2426792A (en) | 1943-10-29 | 1947-09-02 | Du Pont | Process for preparing a mercury oxide catalyst |
US2608821A (en) | 1949-10-08 | 1952-09-02 | Gen Electric | Contrarotating turbojet engine having independent bearing supports for each turbocompressor |
US2826255A (en) | 1951-06-14 | 1958-03-11 | Gen Motors Corp | Propeller drives |
US3021731A (en) | 1951-11-10 | 1962-02-20 | Wilhelm G Stoeckicht | Planetary gear transmission |
US2748623A (en) | 1952-02-05 | 1956-06-05 | Boeing Co | Orbit gear controlled reversible planetary transmissions |
US2936655A (en) | 1955-11-04 | 1960-05-17 | Gen Motors Corp | Self-aligning planetary gearing |
US3033002A (en) | 1957-11-08 | 1962-05-08 | Fairfield Shipbuilding And Eng | Marine propulsion steam turbine installations |
US3185857A (en) | 1960-02-01 | 1965-05-25 | Lear Siegler Inc | Control apparatus for the parallel operation of alternators |
US3194487A (en) | 1963-06-04 | 1965-07-13 | United Aircraft Corp | Noise abatement method and apparatus |
US3363419A (en) | 1965-04-27 | 1968-01-16 | Rolls Royce | Gas turbine ducted fan engine |
US3287906A (en) | 1965-07-20 | 1966-11-29 | Gen Motors Corp | Cooled gas turbine vanes |
US3352178A (en) | 1965-11-15 | 1967-11-14 | Gen Motors Corp | Planetary gearing |
US3412560A (en) | 1966-08-03 | 1968-11-26 | Gen Motors Corp | Jet propulsion engine with cooled combustion chamber, fuel heater, and induced air-flow |
US3390527A (en) | 1967-07-19 | 1968-07-02 | Avco Corp | High bypass ratio turbofan |
GB1135129A (en) | 1967-09-15 | 1968-11-27 | Rolls Royce | Gas turbine engine |
GB1309721A (en) | 1971-01-08 | 1973-03-14 | Secr Defence | Fan |
GB1350431A (en) | 1971-01-08 | 1974-04-18 | Secr Defence | Gearing |
US3892358A (en) | 1971-03-17 | 1975-07-01 | Gen Electric | Nozzle seal |
US3747343A (en) | 1972-02-10 | 1973-07-24 | United Aircraft Corp | Low noise prop-fan |
GB1418905A (en) | 1972-05-09 | 1975-12-24 | Rolls Royce | Gas turbine engines |
US3861139A (en) | 1973-02-12 | 1975-01-21 | Gen Electric | Turbofan engine having counterrotating compressor and turbine elements and unique fan disposition |
US3988889A (en) | 1974-02-25 | 1976-11-02 | General Electric Company | Cowling arrangement for a turbofan engine |
US3932058A (en) | 1974-06-07 | 1976-01-13 | United Technologies Corporation | Control system for variable pitch fan propulsor |
US3935558A (en) | 1974-12-11 | 1976-01-27 | United Technologies Corporation | Surge detector for turbine engines |
US4020632A (en) * | 1975-07-17 | 1977-05-03 | The United States Of America As Represented By The United States National Aeronautics And Space Administration Office Of General Counsel-Code Gp | Oil cooling system for a gas turbine engine |
US4130872A (en) | 1975-10-10 | 1978-12-19 | The United States Of America As Represented By The Secretary Of The Air Force | Method and system of controlling a jet engine for avoiding engine surge |
SE402147B (en) | 1975-12-05 | 1978-06-19 | United Turbine Ab & Co | GAS TURBINE SYSTEM WITH THREE IN THE SAME GAS PASSAGE ORGANIZED COAXIAL TURBINE ROTORS |
GB1516041A (en) | 1977-02-14 | 1978-06-28 | Secr Defence | Multistage axial flow compressor stators |
US4136286A (en) | 1977-07-05 | 1979-01-23 | Woodward Governor Company | Isolated electrical power generation system with multiple isochronous, load-sharing engine-generator units |
GB2041090A (en) | 1979-01-31 | 1980-09-03 | Rolls Royce | By-pass gas turbine engines |
US4233555A (en) | 1979-04-05 | 1980-11-11 | Dyna Technology, Inc. | Alternating current generator for providing three phase and single phase power at different respective voltages |
US4284174A (en) | 1979-04-18 | 1981-08-18 | Avco Corporation | Emergency oil/mist system |
US4405892A (en) | 1979-07-19 | 1983-09-20 | Brunswick Corporation | Regulator for a generator energized battery |
DE2940446C2 (en) | 1979-10-05 | 1982-07-08 | B. Braun Melsungen Ag, 3508 Melsungen | Cultivation of animal cells in suspension and monolayer cultures in fermentation vessels |
GB2080486B (en) | 1980-07-15 | 1984-02-15 | Rolls Royce | Shafts |
GB2130340A (en) | 1981-03-28 | 1984-05-31 | Rolls Royce | Gas turbine rotor assembly |
FR2506840A1 (en) | 1981-05-29 | 1982-12-03 | Onera (Off Nat Aerospatiale) | TURBOREACTOR WITH CONTRA-ROTATING WHEELS |
US4478551A (en) | 1981-12-08 | 1984-10-23 | United Technologies Corporation | Turbine exhaust case design |
US4660376A (en) | 1984-04-27 | 1987-04-28 | General Electric Company | Method for operating a fluid injection gas turbine engine |
US5252905A (en) | 1985-12-23 | 1993-10-12 | York International Corporation | Driving system for single phase A-C induction motor |
US4696156A (en) | 1986-06-03 | 1987-09-29 | United Technologies Corporation | Fuel and oil heat management system for a gas turbine engine |
GB8630754D0 (en) | 1986-12-23 | 1987-02-04 | Rolls Royce Plc | Turbofan gas turbine engine |
DE3714990A1 (en) | 1987-05-06 | 1988-12-01 | Mtu Muenchen Gmbh | PROPFAN TURBO ENGINE |
GB2207191B (en) | 1987-07-06 | 1992-03-04 | Gen Electric | Gas turbine engine |
US4808076A (en) | 1987-12-15 | 1989-02-28 | United Technologies Corporation | Rotor for a gas turbine engine |
US4879624A (en) | 1987-12-24 | 1989-11-07 | Sundstrand Corporation | Power controller |
US5168208A (en) | 1988-05-09 | 1992-12-01 | Onan Corporation | Microprocessor based integrated generator set controller apparatus and method |
FR2644844B1 (en) | 1989-03-23 | 1994-05-06 | Snecma | SUSPENSION OF THE LOW PRESSURE TURBINE ROTOR OF A DOUBLE BODY TURBOMACHINE |
US4979362A (en) | 1989-05-17 | 1990-12-25 | Sundstrand Corporation | Aircraft engine starting and emergency power generating system |
GB2234035B (en) | 1989-07-21 | 1993-05-12 | Rolls Royce Plc | A reduction gear assembly and a gas turbine engine |
US5081832A (en) | 1990-03-05 | 1992-01-21 | Rolf Jan Mowill | High efficiency, twin spool, radial-high pressure, gas turbine engine |
US5182464A (en) | 1991-01-09 | 1993-01-26 | Techmatics, Inc. | High speed transfer switch |
US5141400A (en) | 1991-01-25 | 1992-08-25 | General Electric Company | Wide chord fan blade |
US5102379A (en) | 1991-03-25 | 1992-04-07 | United Technologies Corporation | Journal bearing arrangement |
US5160251A (en) | 1991-05-13 | 1992-11-03 | General Electric Company | Lightweight engine turbine bearing support assembly for withstanding radial and axial loads |
US5317877A (en) | 1992-08-03 | 1994-06-07 | General Electric Company | Intercooled turbine blade cooling air feed system |
US5447411A (en) | 1993-06-10 | 1995-09-05 | Martin Marietta Corporation | Light weight fan blade containment system |
US5466198A (en) | 1993-06-11 | 1995-11-14 | United Technologies Corporation | Geared drive system for a bladed propulsor |
GB9313905D0 (en) | 1993-07-06 | 1993-08-25 | Rolls Royce Plc | Shaft power transfer in gas turbine engines |
US5307622A (en) | 1993-08-02 | 1994-05-03 | General Electric Company | Counterrotating turbine support assembly |
US5524847A (en) | 1993-09-07 | 1996-06-11 | United Technologies Corporation | Nacelle and mounting arrangement for an aircraft engine |
US5388964A (en) | 1993-09-14 | 1995-02-14 | General Electric Company | Hybrid rotor blade |
RU2082824C1 (en) | 1994-03-10 | 1997-06-27 | Московский государственный авиационный институт (технический университет) | Method of protection of heat-resistant material from effect of high-rapid gaseous flow of corrosive media (variants) |
US5433674A (en) | 1994-04-12 | 1995-07-18 | United Technologies Corporation | Coupling system for a planetary gear train |
JPH07286503A (en) | 1994-04-20 | 1995-10-31 | Hitachi Ltd | Highly efficient gas turbine |
US5625276A (en) | 1994-09-14 | 1997-04-29 | Coleman Powermate, Inc. | Controller for permanent magnet generator |
US5778659A (en) | 1994-10-20 | 1998-07-14 | United Technologies Corporation | Variable area fan exhaust nozzle having mechanically separate sleeve and thrust reverser actuation systems |
JP3489106B2 (en) | 1994-12-08 | 2004-01-19 | 株式会社サタケ | Brushless three-phase synchronous generator |
US5915917A (en) | 1994-12-14 | 1999-06-29 | United Technologies Corporation | Compressor stall and surge control using airflow asymmetry measurement |
US5729059A (en) | 1995-06-07 | 1998-03-17 | Kilroy; Donald G. | Digital no-break power transfer system |
JP2969075B2 (en) | 1996-02-26 | 1999-11-02 | ジャパンゴアテックス株式会社 | Degassing device |
US5734255A (en) | 1996-03-13 | 1998-03-31 | Alaska Power Systems Inc. | Control system and circuits for distributed electrical power generating stations |
US5754033A (en) | 1996-03-13 | 1998-05-19 | Alaska Power Systems Inc. | Control system and circuits for distributed electrical-power generating stations |
US5806303A (en) | 1996-03-29 | 1998-09-15 | General Electric Company | Turbofan engine with a core driven supercharged bypass duct and fixed geometry nozzle |
EP0817350B1 (en) | 1996-06-24 | 2008-03-26 | SANYO ELECTRIC Co., Ltd. | Power-supply system involving system interconnection |
US5857836A (en) | 1996-09-10 | 1999-01-12 | Aerodyne Research, Inc. | Evaporatively cooled rotor for a gas turbine engine |
GB2322914B (en) | 1997-03-05 | 2000-05-24 | Rolls Royce Plc | Ducted fan gas turbine engine |
US5949153A (en) | 1997-03-06 | 1999-09-07 | Consolidated Natural Gas Service Company, Inc. | Multi-engine controller |
US5791789A (en) | 1997-04-24 | 1998-08-11 | United Technologies Corporation | Rotor support for a turbine engine |
US5975841A (en) | 1997-10-03 | 1999-11-02 | Thermal Corp. | Heat pipe cooling for turbine stators |
US5985470A (en) | 1998-03-16 | 1999-11-16 | General Electric Company | Thermal/environmental barrier coating system for silicon-based materials |
US6209311B1 (en) | 1998-04-13 | 2001-04-03 | Nikkiso Company, Ltd. | Turbofan engine including fans with reduced speed |
EP1071870B2 (en) | 1998-04-16 | 2011-06-29 | 3K-Warner Turbosystems GmbH | Turbocharged internal combustion engine |
US6230480B1 (en) | 1998-08-31 | 2001-05-15 | Rollins, Iii William Scott | High power density combined cycle power plant |
US6104171A (en) | 1998-11-23 | 2000-08-15 | Caterpillar Inc. | Generator set with redundant bus sensing and automatic generator on-line control |
US6260351B1 (en) | 1998-12-10 | 2001-07-17 | United Technologies Corporation | Controlled spring rate gearbox mount |
US6196790B1 (en) | 1998-12-17 | 2001-03-06 | United Technologies Corporation | Seal assembly for an intershaft seal in a gas turbine engine |
US6517341B1 (en) | 1999-02-26 | 2003-02-11 | General Electric Company | Method to prevent recession loss of silica and silicon-containing materials in combustion gas environments |
US6410148B1 (en) | 1999-04-15 | 2002-06-25 | General Electric Co. | Silicon based substrate with environmental/ thermal barrier layer |
USH2032H1 (en) | 1999-10-01 | 2002-07-02 | The United States Of America As Represented By The Secretary Of The Air Force | Integrated fan-core twin spool counter-rotating turbofan gas turbine engine |
US6668629B1 (en) | 1999-11-26 | 2003-12-30 | General Electric Company | Methods and apparatus for web-enabled engine-generator systems |
US6315815B1 (en) | 1999-12-16 | 2001-11-13 | United Technologies Corporation | Membrane based fuel deoxygenator |
US6318070B1 (en) | 2000-03-03 | 2001-11-20 | United Technologies Corporation | Variable area nozzle for gas turbine engines driven by shape memory alloy actuators |
US6444335B1 (en) | 2000-04-06 | 2002-09-03 | General Electric Company | Thermal/environmental barrier coating for silicon-containing materials |
US6647707B2 (en) | 2000-09-05 | 2003-11-18 | Sudarshan Paul Dev | Nested core gas turbine engine |
US6631310B1 (en) | 2000-09-15 | 2003-10-07 | General Electric Company | Wireless engine-generator systems digital assistant |
US6555929B1 (en) | 2000-10-24 | 2003-04-29 | Kohler Co. | Method and apparatus for preventing excessive reaction to a load disturbance by a generator set |
US6653821B2 (en) | 2001-06-15 | 2003-11-25 | Generac Power Systems, Inc. | System controller and method for monitoring and controlling a plurality of generator sets |
US6657416B2 (en) | 2001-06-15 | 2003-12-02 | Generac Power Systems, Inc. | Control system for stand-by electrical generator |
JP2003097669A (en) | 2001-09-27 | 2003-04-03 | Jatco Ltd | Torque split type continuously variable transmission with infinite gear ratio |
US6669393B2 (en) | 2001-10-10 | 2003-12-30 | General Electric Co. | Connector assembly for gas turbine engines |
US6708482B2 (en) | 2001-11-29 | 2004-03-23 | General Electric Company | Aircraft engine with inter-turbine engine frame |
US6639331B2 (en) | 2001-11-30 | 2003-10-28 | Onan Corporation | Parallel generator power system |
US6663530B2 (en) | 2001-12-14 | 2003-12-16 | Pratt & Whitney Canada Corp. | Zero twist carrier |
US6735954B2 (en) | 2001-12-21 | 2004-05-18 | Pratt & Whitney Canada Corp. | Offset drive for gas turbine engine |
US6914763B2 (en) | 2002-01-15 | 2005-07-05 | Wellspring Heritage, Llc | Utility control and autonomous disconnection of distributed generation from a power distribution system |
WO2003073312A1 (en) | 2002-02-25 | 2003-09-04 | General Electric Company | Method and apparatus for minimally invasive network monitoring |
US6732502B2 (en) | 2002-03-01 | 2004-05-11 | General Electric Company | Counter rotating aircraft gas turbine engine with high overall pressure ratio compressor |
US6619030B1 (en) | 2002-03-01 | 2003-09-16 | General Electric Company | Aircraft engine with inter-turbine engine frame supported counter rotating low pressure turbine rotors |
US6966174B2 (en) | 2002-04-15 | 2005-11-22 | Paul Marius A | Integrated bypass turbojet engines for air craft and other vehicles |
US20030235523A1 (en) | 2002-06-24 | 2003-12-25 | Maxim Lyubovsky | Method for methane oxidation and, apparatus for use therewith |
US6607165B1 (en) | 2002-06-28 | 2003-08-19 | General Electric Company | Aircraft engine mount with single thrust link |
US6763653B2 (en) | 2002-09-24 | 2004-07-20 | General Electric Company | Counter rotating fan aircraft gas turbine engine with aft booster |
US6814541B2 (en) | 2002-10-07 | 2004-11-09 | General Electric Company | Jet aircraft fan case containment design |
US7021042B2 (en) | 2002-12-13 | 2006-04-04 | United Technologies Corporation | Geartrain coupling for a turbofan engine |
US6847297B2 (en) | 2003-01-06 | 2005-01-25 | General Electric Company | Locator devices and methods for centrally controlled power distribution systems |
US6709492B1 (en) | 2003-04-04 | 2004-03-23 | United Technologies Corporation | Planar membrane deoxygenator |
EP1627145A2 (en) | 2003-04-28 | 2006-02-22 | A. Paul Marius | Turbo rocket with real carnot cycle |
US7055306B2 (en) | 2003-04-30 | 2006-06-06 | Hamilton Sundstrand Corporation | Combined stage single shaft turbofan engine |
US6895741B2 (en) | 2003-06-23 | 2005-05-24 | Pratt & Whitney Canada Corp. | Differential geared turbine engine with torque modulation capability |
US7104918B2 (en) | 2003-07-29 | 2006-09-12 | Pratt & Whitney Canada Corp. | Compact epicyclic gear carrier |
US7019495B2 (en) | 2003-08-28 | 2006-03-28 | C.E. Neihoff & Co. | Inter-regulator control of multiple electric power sources |
GB0321139D0 (en) | 2003-09-10 | 2003-10-08 | Short Brothers Plc | A device |
US7216475B2 (en) | 2003-11-21 | 2007-05-15 | General Electric Company | Aft FLADE engine |
FR2866387B1 (en) | 2004-02-12 | 2008-03-14 | Snecma Moteurs | AERODYNAMIC ADAPTATION OF THE BACK BLOW OF A DOUBLE BLOWER TURBOREACTOR |
US7338259B2 (en) | 2004-03-02 | 2008-03-04 | United Technologies Corporation | High modulus metallic component for high vibratory operation |
DE102004016246A1 (en) | 2004-04-02 | 2005-10-20 | Mtu Aero Engines Gmbh | Turbine, in particular low-pressure turbine, a gas turbine, in particular an aircraft engine |
US7144349B2 (en) | 2004-04-06 | 2006-12-05 | Pratt & Whitney Canada Corp. | Gas turbine gearbox |
US7168949B2 (en) | 2004-06-10 | 2007-01-30 | Georgia Tech Research Center | Stagnation point reverse flow combustor for a combustion system |
US7328580B2 (en) | 2004-06-23 | 2008-02-12 | General Electric Company | Chevron film cooled wall |
DE102004042739A1 (en) | 2004-09-03 | 2006-03-09 | Mtu Aero Engines Gmbh | Fan for an aircraft engine and aircraft engine |
US7269938B2 (en) | 2004-10-29 | 2007-09-18 | General Electric Company | Counter-rotating gas turbine engine and method of assembling same |
US7334392B2 (en) | 2004-10-29 | 2008-02-26 | General Electric Company | Counter-rotating gas turbine engine and method of assembling same |
US7409819B2 (en) | 2004-10-29 | 2008-08-12 | General Electric Company | Gas turbine engine and method of assembling same |
US7458202B2 (en) | 2004-10-29 | 2008-12-02 | General Electric Company | Lubrication system for a counter-rotating turbine engine and method of assembling same |
US7195446B2 (en) | 2004-10-29 | 2007-03-27 | General Electric Company | Counter-rotating turbine engine and method of assembling same |
GB0425137D0 (en) * | 2004-11-13 | 2004-12-15 | Rolls Royce Plc | Blade |
WO2006059981A1 (en) | 2004-12-01 | 2006-06-08 | United Technologies Corporation | Hydraulic seal for a gearbox of a tip turbine engine |
US7309210B2 (en) | 2004-12-17 | 2007-12-18 | United Technologies Corporation | Turbine engine rotor stack |
FR2879720B1 (en) | 2004-12-17 | 2007-04-06 | Snecma Moteurs Sa | COMPRESSION-EVAPORATION SYSTEM FOR LIQUEFIED GAS |
CN1332500C (en) | 2005-02-04 | 2007-08-15 | 艾纯 | Small-sized dipolar single-phase generator |
US20060177302A1 (en) | 2005-02-04 | 2006-08-10 | Berry Henry M | Axial flow compressor |
US7845902B2 (en) | 2005-02-15 | 2010-12-07 | Massachusetts Institute Of Technology | Jet engine inlet-fan system and design method |
GB0506685D0 (en) | 2005-04-01 | 2005-05-11 | Hopkins David R | A design to increase and smoothly improve the throughput of fluid (air or gas) through the inlet fan (or fans) of an aero-engine system |
US7476086B2 (en) * | 2005-04-07 | 2009-01-13 | General Electric Company | Tip cambered swept blade |
US7374403B2 (en) | 2005-04-07 | 2008-05-20 | General Electric Company | Low solidity turbofan |
US7151332B2 (en) | 2005-04-27 | 2006-12-19 | Stephen Kundel | Motor having reciprocating and rotating permanent magnets |
US7500365B2 (en) | 2005-05-05 | 2009-03-10 | United Technologies Corporation | Accessory gearbox |
US7594388B2 (en) | 2005-06-06 | 2009-09-29 | General Electric Company | Counterrotating turbofan engine |
US7513102B2 (en) | 2005-06-06 | 2009-04-07 | General Electric Company | Integrated counterrotating turbofan |
US8220245B1 (en) | 2005-08-03 | 2012-07-17 | Candent Technologies, Inc. | Multi spool gas turbine system |
US9657156B2 (en) | 2005-09-28 | 2017-05-23 | Entrotech, Inc. | Braid-reinforced composites and processes for their preparation |
US7685808B2 (en) | 2005-10-19 | 2010-03-30 | General Electric Company | Gas turbine engine assembly and methods of assembling same |
US7493753B2 (en) | 2005-10-19 | 2009-02-24 | General Electric Company | Gas turbine engine assembly and methods of assembling same |
US7513103B2 (en) | 2005-10-19 | 2009-04-07 | General Electric Company | Gas turbine engine assembly and methods of assembling same |
JP4910375B2 (en) | 2005-11-22 | 2012-04-04 | 日本精工株式会社 | Continuously variable transmission |
DE102006001984A1 (en) | 2006-01-16 | 2007-07-19 | Robert Bosch Gmbh | Method and device for providing a supply voltage by means of parallel-connected generator units |
US7591754B2 (en) | 2006-03-22 | 2009-09-22 | United Technologies Corporation | Epicyclic gear train integral sun gear coupling design |
US7610763B2 (en) | 2006-05-09 | 2009-11-03 | United Technologies Corporation | Tailorable design configuration topologies for aircraft engine mid-turbine frames |
US7600370B2 (en) | 2006-05-25 | 2009-10-13 | Siemens Energy, Inc. | Fluid flow distributor apparatus for gas turbine engine mid-frame section |
US20080003096A1 (en) | 2006-06-29 | 2008-01-03 | United Technologies Corporation | High coverage cooling hole shape |
US7926260B2 (en) | 2006-07-05 | 2011-04-19 | United Technologies Corporation | Flexible shaft for gas turbine engine |
US7704178B2 (en) | 2006-07-05 | 2010-04-27 | United Technologies Corporation | Oil baffle for gas turbine fan drive gear system |
US7765788B2 (en) | 2006-07-06 | 2010-08-03 | United Technologies Corporation | Cooling exchanger duct |
US7658060B2 (en) * | 2006-07-19 | 2010-02-09 | United Technologies Corporation | Lubricant cooling exchanger dual intake duct |
US7594404B2 (en) | 2006-07-27 | 2009-09-29 | United Technologies Corporation | Embedded mount for mid-turbine frame |
US7594405B2 (en) | 2006-07-27 | 2009-09-29 | United Technologies Corporation | Catenary mid-turbine frame design |
US7694505B2 (en) | 2006-07-31 | 2010-04-13 | General Electric Company | Gas turbine engine assembly and method of assembling same |
US8939864B2 (en) | 2006-08-15 | 2015-01-27 | United Technologies Corporation | Gas turbine engine lubrication |
WO2008105815A2 (en) | 2006-08-22 | 2008-09-04 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine with intermediate speed booster |
US7815417B2 (en) | 2006-09-01 | 2010-10-19 | United Technologies Corporation | Guide vane for a gas turbine engine |
US7632064B2 (en) | 2006-09-01 | 2009-12-15 | United Technologies Corporation | Variable geometry guide vane for a gas turbine engine |
US7816813B2 (en) | 2006-09-28 | 2010-10-19 | Asco Power Technologies, L.P. | Method and apparatus for parallel engine generators |
US8313280B2 (en) | 2006-10-12 | 2012-11-20 | United Technologies Corporation | Method and device to avoid turbo instability in a gas turbine engine |
WO2008045068A1 (en) * | 2006-10-12 | 2008-04-17 | United Technologies Corporation | Turbofan engine with variable area fan nozzle and low spool generator for emergency power generation and method for providing emergency power. |
US7662059B2 (en) | 2006-10-18 | 2010-02-16 | United Technologies Corporation | Lubrication of windmilling journal bearings |
US7832193B2 (en) | 2006-10-27 | 2010-11-16 | General Electric Company | Gas turbine engine assembly and methods of assembling same |
US7841165B2 (en) | 2006-10-31 | 2010-11-30 | General Electric Company | Gas turbine engine assembly and methods of assembling same |
US7966806B2 (en) | 2006-10-31 | 2011-06-28 | General Electric Company | Turbofan engine assembly and method of assembling same |
US8205427B2 (en) * | 2006-11-09 | 2012-06-26 | United Technologies Corporation | Interdependent lubrication systems in a turbine engine |
US7841163B2 (en) | 2006-11-13 | 2010-11-30 | Hamilton Sundstrand Corporation | Turbofan emergency generator |
US8215454B2 (en) | 2006-11-22 | 2012-07-10 | United Technologies Corporation | Lubrication system with tolerance for reduced gravity |
US8020665B2 (en) | 2006-11-22 | 2011-09-20 | United Technologies Corporation | Lubrication system with extended emergency operability |
US7882693B2 (en) | 2006-11-29 | 2011-02-08 | General Electric Company | Turbofan engine assembly and method of assembling same |
US7797946B2 (en) | 2006-12-06 | 2010-09-21 | United Technologies Corporation | Double U design for mid-turbine frame struts |
US20080148881A1 (en) | 2006-12-21 | 2008-06-26 | Thomas Ory Moniz | Power take-off system and gas turbine engine assembly including same |
US7716914B2 (en) | 2006-12-21 | 2010-05-18 | General Electric Company | Turbofan engine assembly and method of assembling same |
US7791235B2 (en) | 2006-12-22 | 2010-09-07 | General Electric Company | Variable magnetic coupling of rotating machinery |
FR2912181B1 (en) | 2007-02-07 | 2009-04-24 | Snecma Sa | GAS TURBINE WITH HP AND BP TURBINES CONTRA-ROTATIVES |
US7721549B2 (en) | 2007-02-08 | 2010-05-25 | United Technologies Corporation | Fan variable area nozzle for a gas turbine engine fan nacelle with cam drive ring actuation system |
US7942079B2 (en) | 2007-02-16 | 2011-05-17 | Hamilton Sundstrand Corporation | Multi-speed gearbox for low spool driven auxiliary component |
US8015828B2 (en) | 2007-04-03 | 2011-09-13 | General Electric Company | Power take-off system and gas turbine engine assembly including same |
US8017188B2 (en) | 2007-04-17 | 2011-09-13 | General Electric Company | Methods of making articles having toughened and untoughened regions |
FR2915175B1 (en) | 2007-04-20 | 2009-07-17 | Airbus France Sa | ENGINE ATTACHING MACHINE FOR AN AIRCRAFT HAVING A REAR ENGINE ATTACHMENT BEAM DEPORTEE FROM THE HOUSING |
US7557544B2 (en) | 2007-04-23 | 2009-07-07 | Cummins Power Generation Ip, Inc. | Zero crossing detection for an electric power generation system |
US20080318066A1 (en) | 2007-05-11 | 2008-12-25 | Asml Holding N.V. | Optical Component Fabrication Using Coated Substrates |
US8262817B2 (en) | 2007-06-11 | 2012-09-11 | Honeywell International Inc. | First stage dual-alloy turbine wheel |
US7950237B2 (en) | 2007-06-25 | 2011-05-31 | United Technologies Corporation | Managing spool bearing load using variable area flow nozzle |
US8104265B2 (en) | 2007-06-28 | 2012-01-31 | United Technologies Corporation | Gas turbines with multiple gas flow paths |
US7882691B2 (en) | 2007-07-05 | 2011-02-08 | Hamilton Sundstrand Corporation | High to low pressure spool summing gearbox for accessory power extraction and electric start |
US20120124964A1 (en) | 2007-07-27 | 2012-05-24 | Hasel Karl L | Gas turbine engine with improved fuel efficiency |
US8347633B2 (en) | 2007-07-27 | 2013-01-08 | United Technologies Corporation | Gas turbine engine with variable geometry fan exit guide vane system |
US8256707B2 (en) | 2007-08-01 | 2012-09-04 | United Technologies Corporation | Engine mounting configuration for a turbofan gas turbine engine |
US7942635B1 (en) | 2007-08-02 | 2011-05-17 | Florida Turbine Technologies, Inc. | Twin spool rotor assembly for a small gas turbine engine |
US8074440B2 (en) | 2007-08-23 | 2011-12-13 | United Technologies Corporation | Gas turbine engine with axial movable fan variable area nozzle |
US9957918B2 (en) | 2007-08-28 | 2018-05-01 | United Technologies Corporation | Gas turbine engine front architecture |
US8973374B2 (en) * | 2007-09-06 | 2015-03-10 | United Technologies Corporation | Blades in a turbine section of a gas turbine engine |
US8075261B2 (en) | 2007-09-21 | 2011-12-13 | United Technologies Corporation | Gas turbine engine compressor case mounting arrangement |
US8205432B2 (en) | 2007-10-03 | 2012-06-26 | United Technologies Corporation | Epicyclic gear train for turbo fan engine |
US8104289B2 (en) | 2007-10-09 | 2012-01-31 | United Technologies Corp. | Systems and methods involving multiple torque paths for gas turbine engines |
US7656060B2 (en) | 2007-10-31 | 2010-02-02 | Caterpillar Inc. | Power system with method for adding multiple generator sets |
US8015798B2 (en) | 2007-12-13 | 2011-09-13 | United Technologies Corporation | Geared counter-rotating gas turbofan engine |
US8485783B2 (en) | 2007-12-20 | 2013-07-16 | Volvo Aero Corporation | Gas turbine engine |
JP5419251B2 (en) | 2008-02-20 | 2014-02-19 | 富士フイルム株式会社 | Fine particle production method and fine particle production apparatus |
US7762086B2 (en) | 2008-03-12 | 2010-07-27 | United Technologies Corporation | Nozzle extension assembly for ground and flight testing |
US20090265049A1 (en) | 2008-04-22 | 2009-10-22 | Honeywell International, Inc. | Aircraft system emissions and noise estimation mechanism |
GB0807775D0 (en) | 2008-04-29 | 2008-06-04 | Romax Technology Ltd | Methods for model-based diagnosis of gearbox |
DE102008023990A1 (en) | 2008-05-16 | 2009-11-19 | Rolls-Royce Deutschland Ltd & Co Kg | Two-shaft engine for an aircraft gas turbine |
US8128021B2 (en) | 2008-06-02 | 2012-03-06 | United Technologies Corporation | Engine mount system for a turbofan gas turbine engine |
US8807477B2 (en) | 2008-06-02 | 2014-08-19 | United Technologies Corporation | Gas turbine engine compressor arrangement |
US8210800B2 (en) | 2008-06-12 | 2012-07-03 | United Technologies Corporation | Integrated actuator module for gas turbine engine |
US8973364B2 (en) | 2008-06-26 | 2015-03-10 | United Technologies Corporation | Gas turbine engine with noise attenuating variable area fan nozzle |
US20100005810A1 (en) | 2008-07-11 | 2010-01-14 | Rob Jarrell | Power transmission among shafts in a turbine engine |
US8266889B2 (en) | 2008-08-25 | 2012-09-18 | General Electric Company | Gas turbine engine fan bleed heat exchanger system |
US7997868B1 (en) | 2008-11-18 | 2011-08-16 | Florida Turbine Technologies, Inc. | Film cooling hole for turbine airfoil |
US8166748B2 (en) | 2008-11-21 | 2012-05-01 | General Electric Company | Gas turbine engine booster having rotatable radially inwardly extending blades and non-rotatable vanes |
US8061969B2 (en) | 2008-11-28 | 2011-11-22 | Pratt & Whitney Canada Corp. | Mid turbine frame system for gas turbine engine |
US20100132377A1 (en) | 2008-11-28 | 2010-06-03 | Pratt & Whitney Canada Corp. | Fabricated itd-strut and vane ring for gas turbine engine |
US8091371B2 (en) | 2008-11-28 | 2012-01-10 | Pratt & Whitney Canada Corp. | Mid turbine frame for gas turbine engine |
US8177488B2 (en) * | 2008-11-29 | 2012-05-15 | General Electric Company | Integrated service tube and impingement baffle for a gas turbine engine |
US8106633B2 (en) | 2008-12-18 | 2012-01-31 | Caterpillar Inc. | Generator set control system |
US8191352B2 (en) | 2008-12-19 | 2012-06-05 | General Electric Company | Geared differential speed counter-rotatable low pressure turbine |
US8216108B2 (en) * | 2009-02-05 | 2012-07-10 | Friede & Goldman, Ltd. | Gear assembly with tapered flex pin |
US8307626B2 (en) | 2009-02-26 | 2012-11-13 | United Technologies Corporation | Auxiliary pump system for fan drive gear system |
US8181441B2 (en) | 2009-02-27 | 2012-05-22 | United Technologies Corporation | Controlled fan stream flow bypass |
GB0903423D0 (en) | 2009-03-02 | 2009-04-08 | Rolls Royce Plc | Variable drive gas turbine engine |
FR2944558B1 (en) | 2009-04-17 | 2014-05-02 | Snecma | DOUBLE BODY GAS TURBINE ENGINE PROVIDED WITH SUPPLEMENTARY BP TURBINE BEARING. |
US8246503B2 (en) | 2009-06-10 | 2012-08-21 | United Technologies Corporation | Epicyclic gear system with improved lubrication system |
US8172716B2 (en) | 2009-06-25 | 2012-05-08 | United Technologies Corporation | Epicyclic gear system with superfinished journal bearing |
US8375695B2 (en) | 2009-06-30 | 2013-02-19 | General Electric Company | Aircraft gas turbine engine counter-rotatable generator |
US8176725B2 (en) | 2009-09-09 | 2012-05-15 | United Technologies Corporation | Reversed-flow core for a turbofan with a fan drive gear system |
US8457126B2 (en) | 2009-10-14 | 2013-06-04 | Vss Monitoring, Inc. | System, method and apparatus for distributing captured data packets including tunneling identifiers |
US8672801B2 (en) | 2009-11-30 | 2014-03-18 | United Technologies Corporation | Mounting system for a planetary gear train in a gas turbine engine |
JP5282731B2 (en) | 2009-12-22 | 2013-09-04 | 株式会社安川電機 | Power converter |
US9170616B2 (en) | 2009-12-31 | 2015-10-27 | Intel Corporation | Quiet system cooling using coupled optimization between integrated micro porous absorbers and rotors |
US8905713B2 (en) | 2010-05-28 | 2014-12-09 | General Electric Company | Articles which include chevron film cooling holes, and related processes |
US9006930B2 (en) | 2010-07-08 | 2015-04-14 | Delta Electronics Inc. | Power supply having converters with serially connected inputs and parallel connected outputs |
US8366385B2 (en) | 2011-04-15 | 2013-02-05 | United Technologies Corporation | Gas turbine engine front center body architecture |
US8172717B2 (en) | 2011-06-08 | 2012-05-08 | General Electric Company | Compliant carrier wall for improved gearbox load sharing |
US8297916B1 (en) | 2011-06-08 | 2012-10-30 | United Technologies Corporation | Flexible support structure for a geared architecture gas turbine engine |
US9498823B2 (en) | 2011-11-07 | 2016-11-22 | United Technologies Corporation | Metal casting apparatus, cast work piece and method therefor |
BR102012027097B1 (en) | 2011-11-23 | 2022-01-04 | United Technologies Corporation | GAS TURBINE ENGINE |
US8935913B2 (en) | 2012-01-31 | 2015-01-20 | United Technologies Corporation | Geared turbofan gas turbine engine architecture |
US20130192266A1 (en) | 2012-01-31 | 2013-08-01 | United Technologies Corporation | Geared turbofan gas turbine engine architecture |
US20150308351A1 (en) * | 2012-05-31 | 2015-10-29 | United Technologies Corporation | Fundamental gear system architecture |
US8756908B2 (en) * | 2012-05-31 | 2014-06-24 | United Technologies Corporation | Fundamental gear system architecture |
-
2015
- 2015-06-22 US US14/745,802 patent/US20150308351A1/en not_active Abandoned
-
2016
- 2016-12-15 US US15/379,634 patent/US20170096944A1/en not_active Abandoned
- 2016-12-15 US US15/379,619 patent/US9840969B2/en active Active
- 2016-12-15 US US15/379,616 patent/US20170096937A1/en not_active Abandoned
-
2017
- 2017-05-26 US US15/606,246 patent/US20170260911A1/en not_active Abandoned
-
2018
- 2018-01-16 US US15/872,158 patent/US20180135524A1/en not_active Abandoned
- 2018-01-16 US US15/872,136 patent/US20180238241A1/en not_active Abandoned
Patent Citations (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6223616B1 (en) * | 1999-12-22 | 2001-05-01 | United Technologies Corporation | Star gear system with lubrication circuit and lubrication method therefor |
US20050257528A1 (en) * | 2004-05-19 | 2005-11-24 | Dunbar Donal S Jr | Retractable afterburner for jet engine |
US20100105516A1 (en) * | 2006-07-05 | 2010-04-29 | United Technologies Corporation | Coupling system for a star gear train in a gas turbine engine |
US20080148708A1 (en) * | 2006-12-20 | 2008-06-26 | General Electric Company | Turbine engine system with shafts for improved weight and vibration characteristic |
US20100296947A1 (en) * | 2009-05-22 | 2010-11-25 | United Technologies Corporation | Apparatus and method for providing damper liquid in a gas turbine |
US8572943B1 (en) * | 2012-05-31 | 2013-11-05 | United Technologies Corporation | Fundamental gear system architecture |
Cited By (37)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US11773786B2 (en) | 2012-05-31 | 2023-10-03 | Rtx Corporation | Fundamental gear system architecture |
US9840969B2 (en) | 2012-05-31 | 2017-12-12 | United Technologies Corporation | Gear system architecture for gas turbine engine |
US20170096937A1 (en) * | 2012-05-31 | 2017-04-06 | United Technologies Corporation | Gear system architecture for gas turbine engine |
US10221770B2 (en) | 2012-05-31 | 2019-03-05 | United Technologies Corporation | Fundamental gear system architecture |
US10590854B2 (en) * | 2016-01-26 | 2020-03-17 | United Technologies Corporation | Geared gas turbine engine |
US10663036B2 (en) | 2017-06-13 | 2020-05-26 | General Electric Company | Gas turbine engine with rotating reversing compound gearbox |
EP3415728A1 (en) * | 2017-06-13 | 2018-12-19 | General Electric Company | Gas turbine engine with rotating reversing compound gearbox |
CN109469724A (en) * | 2017-09-08 | 2019-03-15 | 劳斯莱斯有限公司 | Auxiliary feed-oil equipment for rotary components |
EP3453924B1 (en) * | 2017-09-08 | 2022-06-29 | Rolls-Royce plc | Auxiliary oil supply apparatus for a rotating component |
US11085521B2 (en) | 2017-09-08 | 2021-08-10 | Rolls-Royce Plc | Auxiliary oil supply apparatus for a rotating component |
EP3460199A1 (en) * | 2017-09-20 | 2019-03-27 | General Electric Company | Lube system for geared turbine section |
US11255221B2 (en) | 2017-09-20 | 2022-02-22 | General Electric Company | Lube system for geared turbine section |
US20190309705A1 (en) * | 2018-04-05 | 2019-10-10 | United Technologies Corporation | Aft counter-rotating boundary layer ingestion engine |
US11098678B2 (en) * | 2018-04-05 | 2021-08-24 | Raytheon Technologies Corporation | Aft counter-rotating boundary layer ingestion engine |
US11313454B2 (en) | 2018-04-17 | 2022-04-26 | Rolls-Royce Plc | Lubrication system |
EP3557009A1 (en) * | 2018-04-17 | 2019-10-23 | Rolls-Royce plc | Lubrication system and method of lubrication |
CN110388448A (en) * | 2018-04-17 | 2019-10-29 | 劳斯莱斯有限公司 | Lubricating system |
US11174796B2 (en) * | 2019-04-12 | 2021-11-16 | Rolls-Royce Plc | Accessory gearbox assembly |
US11719162B2 (en) | 2020-11-04 | 2023-08-08 | Delavan, Inc. | Torch igniter cooling system |
US11473505B2 (en) | 2020-11-04 | 2022-10-18 | Delavan Inc. | Torch igniter cooling system |
US11692488B2 (en) | 2020-11-04 | 2023-07-04 | Delavan Inc. | Torch igniter cooling system |
US11608783B2 (en) | 2020-11-04 | 2023-03-21 | Delavan, Inc. | Surface igniter cooling system |
US11635027B2 (en) | 2020-11-18 | 2023-04-25 | Collins Engine Nozzles, Inc. | Fuel systems for torch ignition devices |
US11891956B2 (en) | 2020-12-16 | 2024-02-06 | Delavan Inc. | Continuous ignition device exhaust manifold |
US20220186668A1 (en) * | 2020-12-16 | 2022-06-16 | Delavan Inc. | Continuous ignition device exhaust manifold |
US11226103B1 (en) | 2020-12-16 | 2022-01-18 | Delavan Inc. | High-pressure continuous ignition device |
US11421602B2 (en) * | 2020-12-16 | 2022-08-23 | Delavan Inc. | Continuous ignition device exhaust manifold |
US20220195934A1 (en) * | 2020-12-17 | 2022-06-23 | Delavan Inc. | Axially oriented internally mounted continuous ignition device: removable hot surface igniter |
US11635210B2 (en) | 2020-12-17 | 2023-04-25 | Collins Engine Nozzles, Inc. | Conformal and flexible woven heat shields for gas turbine engine components |
US11486309B2 (en) * | 2020-12-17 | 2022-11-01 | Delavan Inc. | Axially oriented internally mounted continuous ignition device: removable hot surface igniter |
US11754289B2 (en) | 2020-12-17 | 2023-09-12 | Delavan, Inc. | Axially oriented internally mounted continuous ignition device: removable nozzle |
US11680528B2 (en) | 2020-12-18 | 2023-06-20 | Delavan Inc. | Internally-mounted torch igniters with removable igniter heads |
US11286862B1 (en) | 2020-12-18 | 2022-03-29 | Delavan Inc. | Torch injector systems for gas turbine combustors |
US11209164B1 (en) | 2020-12-18 | 2021-12-28 | Delavan Inc. | Fuel injector systems for torch igniters |
US11913646B2 (en) | 2020-12-18 | 2024-02-27 | Delavan Inc. | Fuel injector systems for torch igniters |
US11428160B2 (en) | 2020-12-31 | 2022-08-30 | General Electric Company | Gas turbine engine with interdigitated turbine and gear assembly |
US20230340911A1 (en) * | 2022-04-25 | 2023-10-26 | General Electric Company | Mounting assembly for a gearbox assembly |
Also Published As
Publication number | Publication date |
---|---|
US9840969B2 (en) | 2017-12-12 |
US20170260911A1 (en) | 2017-09-14 |
US20180135524A1 (en) | 2018-05-17 |
US20180238241A1 (en) | 2018-08-23 |
US20170096937A1 (en) | 2017-04-06 |
US20170096943A1 (en) | 2017-04-06 |
US20170096944A1 (en) | 2017-04-06 |
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