US20150377123A1 - Turbine section of high bypass turbofan - Google Patents

Turbine section of high bypass turbofan Download PDF

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Publication number
US20150377123A1
US20150377123A1 US14/793,785 US201514793785A US2015377123A1 US 20150377123 A1 US20150377123 A1 US 20150377123A1 US 201514793785 A US201514793785 A US 201514793785A US 2015377123 A1 US2015377123 A1 US 2015377123A1
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United States
Prior art keywords
compressor
turbine
section
fan
ratio
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
US14/793,785
Inventor
Paul R. Adams
Shankar S. Magge
Joseph Brent Staubach
Wesley K. Lord
Frederick M. Schwarz
Gabriel L. Suciu
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United Technologies Corp
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United Technologies Corp
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Filing date
Publication date
Priority to US11/832,107 priority Critical patent/US8256707B2/en
Priority to US201161498516P priority
Priority to US201261593190P priority
Priority to US13/475,252 priority patent/US8844265B2/en
Priority to US13/599,175 priority patent/US9010085B2/en
Priority to US14/692,090 priority patent/US20150345404A1/en
Priority to US14/793,785 priority patent/US20150377123A1/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: MAGGE, SHANKAR S., ADAMS, PAUL R., LORD, WESLEY K., SCHWARZ, FREDERICK M., STAUBACH, JOSEPH B., SUCIU, GABRIEL L.
Publication of US20150377123A1 publication Critical patent/US20150377123A1/en
Priority claimed from EP16178679.3A external-priority patent/EP3115590A1/en
Priority claimed from US16/025,022 external-priority patent/US20190017445A1/en
Priority claimed from US16/025,038 external-priority patent/US20190017446A1/en
Application status is Abandoned legal-status Critical

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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, 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
    • F02C7/36Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24Casings; Casing parts, e.g. diaphragms, casing fastenings
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02Blade-carrying members, e.g. rotors
    • F01D5/06Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/107Gas-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
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, 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
    • F02C7/20Mounting or supporting of plant; Accommodating heat expansion or creep
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16Control of working fluid flow
    • F02C9/18Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02Plants 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
    • F02K3/025Plants 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 by-pass flow being at least partly used to create an independent thrust component
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02Plants 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
    • F02K3/04Plants 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
    • F02K3/06Plants 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
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02Plants 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
    • F02K3/04Plants 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
    • F02K3/075Plants 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 controlling flow ratio between flows
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00Axial-flow pumps
    • F04D19/02Multi-stage pumps
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05BINDEXING SCHEME RELATING TO MACHINES OR ENGINES OTHER THAN NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, TO WIND MOTORS, TO NON-POSITIVE DISPLACEMENT PUMPS, AND TO GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY
    • F05B2250/00Geometry
    • F05B2250/20Geometry three-dimensional
    • F05B2250/28Geometry three-dimensional patterned
    • F05B2250/283Honeycomb
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00Application
    • F05D2220/30Application in turbines
    • F05D2220/32Application in turbines in gas turbines
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00Application
    • F05D2220/30Application in turbines
    • F05D2220/32Application in turbines in gas turbines
    • F05D2220/323Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/20Rotors
    • F05D2240/24Rotors for turbines
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/20Rotors
    • F05D2240/30Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/35Combustors or associated equipment
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/60Shafts
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/40Transmission of power
    • F05D2260/403Transmission of power through the shape of the drive components
    • F05D2260/4031Transmission of power through the shape of the drive components as in toothed gearing
    • F05D2260/40311Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type

Abstract

A turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is defined as air communicated through the bypass path relative to air communicated to the compressor being greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor. The turbine has a first turbine section that includes two or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and rotatable by the turbine.

Description

    CROSS-REFERENCE TO RELATED APPLICATION
  • The present disclosure is a continuation-in-part of U.S. patent application Ser. No. 14/692,090, filed Apr. 21, 2015, which was a continuation of U.S. patent application Ser. No. 13/599,175, filed Aug. 30, 2012, which was a continuation of U.S. patent application Ser. No. 13/475,252, filed May 18, 2012, which was a continuation-in-part of U.S. patent application Ser. No. 11/832,107, filed Aug. 1, 2007, and claimed the benefit of U.S. Patent Provisional Application No. 61/593,190, filed Jan. 31, 2012, and U.S. Provisional Application No. 61/498,516, filed Jun. 17, 2011.
  • BACKGROUND
  • The disclosure relates to turbofan engines. More particularly, the disclosure relates to low pressure turbine sections of turbofan engines which power the fans via a speed reduction mechanism.
  • There has been a trend toward increasing bypass ratio in gas turbine engines. This is discussed further below. There has generally been a correlation between certain characteristics of bypass and the diameter of the low pressure turbine section sections of turbofan engines.
  • SUMMARY
  • In a featured embodiment, a turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is defined as air communicated through the bypass path relative to air communicated to the compressor being greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor. The turbine has a first turbine section that includes two or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and rotatable by the turbine.
  • In another embodiment according to the previous embodiment, the compressor includes a first compressor section and a second compressor section. The second compressor section operates at a pressure higher than the first compressor section.
  • In another embodiment according to any of the previous embodiments, the first compressor section includes at least four stages.
  • In another embodiment according to any of the previous embodiments the second compressor section includes at least eight stages.
  • In another embodiment according to any of the previous embodiments the second turbine section comprises a high pressure turbine and the first turbine section comprises a low pressure turbine that operates at a pressure lower than the high pressure turbine section and the low pressure turbine includes an airfoil count below about 1600.
  • In another embodiment according to any of the previous embodiments a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section is between about 0.4 and about 0.5 measured at a maximum Ro axial location within the low pressure turbine.
  • In another embodiment according to any of the previous embodiments the low pressure turbine includes more than four stages.
  • In another embodiment according to any of the previous embodiments the speed reduction mechanism comprises an epicyclic gearbox.
  • In another embodiment according to any of the previous embodiments the epicyclic gearbox provides a speed reduction ratio is between about 2:1 and about 5:1.
  • In another embodiment according to any of the previous embodiments there is a third turbine section.
  • In another featured embodiment, a turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The compressor includes a first compressor section including at least two stages and a second compressor section including at least five stages. The second compressor section is configured to operate at a higher pressure than the first compressor section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor and a bypass ratio defined as air communicated through the bypass path relative to air communicated to the compressor is greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor, and has a first turbine section and a second turbine section. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and is rotatable by the turbine.
  • In another embodiment according to the previous embodiment, the first turbine section includes three or more stages.
  • In another embodiment according to any of the previous embodiments, the first turbine section includes four or more stages.
  • In another embodiment according to any of the previous embodiments, the second turbine section comprises a high pressure turbine and the first turbine section comprises a low pressure turbine that operates at a pressure lower than the high pressure turbine section. The low pressure turbine includes an airfoil count below about 1600.
  • In another embodiment according to any of the previous ,embodiments a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section is between about 0.4 and about 0.5 measured at a maximum Ro axial location within the low pressure turbine.
  • In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 8.0.
  • In another embodiment according to any of the previous embodiments, a fan pressure ratio across the fan is less than about 1.45.
  • In another embodiment according to any of the previous embodiments, the speed reduction mechanism comprises an epicyclic gearbox.
  • In another embodiment according to any of the previous embodiments, the epicyclic gearbox provides a speed reduction ratio between about 2:1 and about 5:1.
  • The details of one or more embodiments are set forth in the accompanying drawings and the description below. Other features, objects, and advantages will be apparent from the description and drawings, and from the claims.
  • BRIEF DESCRIPTION OF THE DRAWINGS
  • FIG. 1 is an axial sectional view of a turbofan engine.
  • FIG. 2 is an axial sectional view of a low pressure turbine section of the engine of FIG. 1.
  • FIG. 3 is transverse sectional view of transmission of the engine of FIG. 1.
  • Like reference numbers and designations in the various drawings indicate like elements.
  • FIG. 4 shows another embodiment.
  • FIG. 5 shows yet another embodiment.
  • DETAILED DESCRIPTION
  • FIG. 1 shows a turbofan engine 20 having a main housing (engine case) 22 containing a rotor shaft assembly 23. An exemplary engine is a high-bypass turbofan. In such an engine, the normal cruise condition bypass area ratio of air mass flowing outside the case 22 (e.g., the compressor sections and combustor) to air mass passing through the case 22 is typically in excess of about 4.0 and, more narrowly, typically between about 4.0 and about 12.0. Via high 24 and low 25 shaft portions of the shaft assembly 23, a high pressure turbine section (gas generating turbine) 26 and a low pressure turbine section 27 respectively drive a high pressure compressor section 28 and a low pressure compressor section 30. As used herein, the high pressure turbine section experiences higher pressures that the low pressure turbine section. A low pressure turbine section is a section that powers a fan 42. Although a two-spool (plus fan) engine is shown, one of many alternative variations involves a three-spool (plus fan) engine wherein an intermediate spool comprises an intermediate pressure compressor between the low fan and high pressure compressor section and an intermediate pressure turbine between the high pressure turbine section and low pressure turbine section.
  • The engine extends along a longitudinal axis 500 from a fore end to an aft end. Adjacent the fore end, a shroud (fan case) 40 encircles the fan 42 and is supported by vanes 44. An aerodynamic nacelle around the fan case is shown and an aerodynamic nacelle 45 around the engine case is shown.
  • The low shaft portion 25 of the rotor shaft assembly 23 drives the fan 42 through a speed reduction mechanism 46. An exemplary speed reduction mechanism is an epicyclic transmission, namely a star or planetary gear system. As is discussed further below, an inlet airflow 520 entering the nacelle is divided into a portion 522 passing along a core flowpath 524 and a bypass portion 526 passing along a bypass flowpath 528. With the exception of diversions such as cooling air, etc., flow along the core flowpath sequentially passes through the low pressure compressor section, high pressure compressor section, a combustor 48, the high pressure turbine section, and the low pressure turbine section before exiting from an outlet 530.
  • FIG. 3 schematically shows details of the transmission 46. A forward end of the low shaft 25 is coupled to a sun gear 52 (or other high speed input to the speed reduction mechanism). The externally-toothed sun gear 52 is encircled by a number of externally-toothed star gears 56 and an internally-toothed ring gear 54. The exemplary ring gear is coupled to the fan to rotate with the fan as a unit.
  • The star gears 56 are positioned between and enmeshed with the sun gear and ring gear. A cage or star carrier assembly 60 carries the star gears via associated journals 62. The exemplary star carrier is substantially irrotatably mounted relative via fingers 404 to the case 22.
  • Another transmission/gearbox combination has the star carrier connected to the fan and the ring is fixed to the fixed structure (case) is possible and such is commonly referred to as a planetary gearbox.
  • The speed reduction ratio is determined by the ratio of diameters within the gearbox. An exemplary reduction is between about 2:1 and about 13:1.
  • The exemplary fan (FIG. 1) comprises a circumferential array of blades 70. Each blade comprises an airfoil 72 having a leading edge 74 and a trailing edge 76 and extending from an inboard end 78 at a platform to an outboard end 80 (i.e., a free tip). The outboard end 80 is in close facing proximity to a rub strip 82 along an interior surface 84 of the nacelle and fan case.
  • To mount the engine to the aircraft wing 92, a pylon 94 is mounted to the fan case and/or to the other engine cases. The exemplary pylon 94 may be as disclosed in U.S. patent application Ser. No. 11/832,107 (US2009/0056343A1). The pylon comprises a forward mount 100 and an aft/rear mount 102. The forward mount may engage the engine intermediate case (IMC) and the aft mount may engage the engine thrust case. The aft mount reacts at least a thrust load of the engine.
  • To reduce aircraft fuel burn with turbofans, it is desirable to produce a low pressure turbine with the highest efficiency and lowest weight possible. Further, there are considerations of small size (especially radial size) that benefit the aerodynamic shape of the engine cowling and allow room for packaging engine subsystems.
  • FIG. 2 shows the low pressure turbine section 27 as comprising an exemplary three blade stages 200, 202, 204. An exemplary blade stage count is 2-6, more narrowly, 2-4, or 2-3, 3-5, or 3-4. Interspersed between the blade stages are vane stages 206 and 208. Each exemplary blade stage comprises a disk 210, 212, and 214, respectively. A circumferential array of blades extends from peripheries of each of the disks. Each exemplary blade comprises an airfoil 220 extending from an inner diameter (ID) platform 222 to an outer diameter (OD) shroud 224 (shown integral with the airfoil
  • An alternative may be an unshrouded blade with a rotational gap between the tip of the blade and a stationary blade outer air seal (BOAS). Each exemplary shroud 224 has outboard sealing ridges which seal with abradable seals (e.g., honeycomb) fixed to the case. The exemplary vanes in stages 206 and 208 include airfoils 230 extending from ID platforms 232 to OD shrouds 234. The exemplary OD shrouds 234 are directly mounted to the case. The exemplary platforms 232 carry seals for sealing with inter-disk knife edges protruding outwardly from inter-disk spacers which may be separate from the adjacent disks or unitarily formed with one of the adjacent disks.
  • Each exemplary disk 210, 212, 214 comprises an enlarged central annular protuberance or “bore” 240, 242, 244 and a thinner radial web 246, 248, 250 extending radially outboard from the bore. The bore imparts structural strength allowing the disk to withstand centrifugal loading which the disk would otherwise be unable to withstand.
  • A turbofan engine is characterized by its bypass ratio (mass flow ratio of air bypassing the core to air passing through the core) and the geometric bypass area ratio (ratio of fan duct annulus area outside/outboard of the low pressure compressor section inlet (i.e., at location 260 in FIG. 1) to low pressure compressor section inlet annulus area (i.e., at location 262 in FIG. 2). High bypass engines typically have bypass area ratio of at least four. There has been a correlation between increased bypass area ratio and increased low pressure turbine section radius and low pressure turbine section airfoil count. As is discussed below, this correlation may be broken by having an engine with relatively high bypass area ratio and relatively low turbine size.
  • By employing a speed reduction mechanism (e.g., a transmission) to allow the low pressure turbine section to turn very fast relative to the fan and by employing low pressure turbine section design features for high speed, it is possible to create a compact turbine module (e.g., while producing the same amount of thrust and increasing bypass area ratio). The exemplary transmission is a epicyclic transmission. Alternative transmissions include composite belt transmissions, metal chain belt transmissions, fluidic transmissions, and electric means (e.g., a motor/generator set where the turbine turns a generator providing electricity to an electric motor which drives the fan).
  • Compactness of the turbine is characterized in several ways. Along the compressor and turbine sections, the core gaspath extends from an inboard boundary (e.g., at blade hubs or outboard surfaces of platforms of associated blades and vanes) to an outboard boundary (e.g., at blade tips and inboard surfaces of blade outer air seals for unshrouded blade tips and at inboard surfaces of OD shrouds of shrouded blade tips and at inboard surfaces of OD shrouds of the vanes). These boundaries may be characterized by radii RI and RO, respectively, which vary along the length of the engine.
  • For low pressure turbine radial compactness, there may be a relatively high ratio of radial span (RO-RI) to radius (RO or RI). Radial compactness may also be expressed in the hub-to-tip ratio (RI:RO). These may be measured at the maximum RO location in the low pressure turbine section. The exemplary compact low pressure turbine section has a hub-to-tip ratio close to about 0.5 (e.g., about 0.4-0.5 or about 0.42-0.48, with an exemplary about 0.46).
  • Another characteristic of low pressure turbine radial compactness is relative to the fan size. An exemplary fan size measurement is the maximum tip radius RTmax of the fan blades. An exemplary ratio is the maximum RO along the low pressure turbine section to RTmax of the fan blades. Exemplary values for this ratio are less than about 0.55 (e.g., about 0.35-55), more narrowly, less than about 0.50, or about 0.35-0.50.
  • To achieve compactness the designer may balance multiple physical phenomena to arrive at a system solution as defined by the low pressure turbine hub-to-tip ratio, the fan maximum tip radius to low pressure turbine maximum RO ratio, the bypass area ratio, and the bypass area ratio to low pressure turbine airfoil count ratio. These concerns include, but are not limited to: a) aerodynamics within the low pressure turbine, b) low pressure turbine blade structural design, c) low pressure turbine disk structural design, and d) the shaft connecting the low pressure turbine to the low pressure compressor and speed reduction device between the low pressure compressor and fan. These physical phenomena may be balanced in order to achieve desirable performance, weight, and cost characteristics.
  • The addition of a speed reduction device between the fan and the low pressure compressor creates a larger design space because the speed of the low pressure turbine is decoupled from the fan. This design space provides great design variables and new constraints that limit feasibility of a design with respect to physical phenomena. For example the designer can independently change the speed and flow area of the low pressure turbine to achieve optimal aerodynamic parameters defined by flow coefficient (axial flow velocity/wheel speed) and work coefficient (wheel speed/square root of work). However, this introduces structural constraints with respect blade stresses, disk size, material selection, etc.
  • In some examples, the designer can choose to make low pressure turbine section disk bores much thicker relative to prior art turbine bores and the bores may be at a much smaller radius RB. This increases the amount of mass at less than a “self sustaining radius”. Another means is to choose disk materials of greater strength than prior art such as the use of wrought powdered metal disks to allow for extremely high centrifugal blade pulls associated with the compactness.
  • Another variable in achieving compactness is to increase the structural parameter AN2 which is the annulus area of the exit of the low pressure turbine divided by the low pressure turbine rpm squared at its redline or maximum speed. Relative to prior art turbines, which are greatly constrained by fan blade tip mach number, a very wide range of AN2 values can be selected and optimized while accommodating such constraints as cost or a countering, unfavorable trend in low pressure turbine section shaft dynamics. In selecting the turbine speed (and thereby selecting the transmission speed ratio, one has to be mindful that at too high a gear ratio the low pressure turbine section shaft (low shaft) will become dynamically unstable.
  • The higher the design speed, the higher the gear ratio will be and the more massive the disks will become and the stronger the low pressure turbine section disk and blade material will have to be. All of these parameters can be varied simultaneously to change the weight of the turbine, its efficiency, its manufacturing cost, the degree of difficulty in packaging the low pressure turbine section in the core cowling and its durability. This is distinguished from a prior art direct drive configuration, where the high bypass area ratio can only be achieved by a large low pressure turbine section radius. Because that radius is so very large and, although the same variables (airfoil turning, disk size, blade materials, disk shape and materials, etc.) are theoretically available, as a practical matter economics and engine fuel burn considerations severely limit the designer's choice in these parameters.
  • Another characteristic of low pressure turbine section size is airfoil count (numerical count of all of the blades and vanes in the low pressure turbine). Airfoil metal angles can be selected such that airfoil count is low or extremely low relative to a direct drive turbine. In known prior art engines having bypass area ratio above 6.0 (e.g. 8.0-20), low pressure turbine sections involve ratios of airfoil count to bypass area ratio above 190.
  • With the full range of selection of parameters discussed above including, disk bore thickness, disk material, hub to tip ratio, and RO/RTmax, the ratio of airfoil count to bypass area ratio may be below about 170 to as low as 10. (e.g., below about 150 or an exemplary about 10-170, more narrowly about 10-150). Further, in such embodiments the airfoil count may be below about 1700, or below about 1600.
  • FIG. 4 shows an embodiment 600, wherein there is a fan drive turbine 608 driving a shaft 606 to in turn drive a fan rotor 602. A gear reduction 604 may be positioned between the fan drive turbine 608 and the fan rotor 602. This gear reduction 604 may be structured and operate like the gear reduction disclosed above. A compressor rotor 610 is driven by an intermediate pressure turbine 612, and a second stage compressor rotor 614 is driven by a turbine rotor 216. A combustion section 618 is positioned intermediate the compressor rotor 614 and the turbine section 616.
  • FIG. 5 shows yet another embodiment 700 wherein a fan rotor 702 and a first stage compressor 704 rotate at a common speed. The gear reduction 706 (which may be structured as disclosed above) is intermediate the compressor rotor 704 and a shaft 708 which is driven by a low pressure turbine section.
  • One or more embodiments have been described. Nevertheless, it will be understood that various modifications may be made. For example, when reengineering from a baseline engine configuration, details of the baseline may influence details of any particular implementation. Accordingly, other embodiments are within the scope of the following claims.

Claims (19)

1. A turbofan engine comprising:
a fan having fan blades;
a compressor in communication with the fan section, wherein the fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor and a bypass ratio defined as air communicated through the bypass path relative to air communicated to the compressor is greater than about 6.0;
a combustor in fluid communication with the compressor;
a turbine in communication with the combustor, the turbine having a first turbine section that includes two or more stages and a second turbine section that includes at least two stages, wherein a ratio of airfoils in the first turbine section to the bypass ratio is less than about 170 and the first turbine section includes a maximum gas path radius and a ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50; and
a speed reduction mechanism coupled to the fan and rotatable by the turbine.
2. The turbofan engine as recited in claim 1, wherein the compressor includes a first compressor section and a second compressor section, wherein the second compressor section operates at a pressure higher than the first compressor section.
3. The turbofan engine as recited in claim 2, wherein the first compressor section includes at least four stages.
4. The turbofan engine as recited in claim 3, wherein the second compressor section includes at least eight stages.
5. The turbofan engine as recited in claim 4, wherein the second turbine section comprises a high pressure turbine and the first turbine section comprises a low pressure turbine that operates at a pressure lower than the high pressure turbine section and the low pressure turbine includes an airfoil count below about 1600.
6. The turbofan engine as recited in claim 5, wherein a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section is between about 0.4 and about 0.5 measured at a maximum Ro axial location within the low pressure turbine.
7. The turbofan engine as recited in claim 6, wherein the low pressure turbine includes more than four stages.
8. The turbofan engine as recited in claim 6, wherein the speed reduction mechanism comprises an epicyclic gearbox.
9. The turbofan engine as recited in claim 8, wherein the epicyclic gearbox provides a speed reduction ratio is between about 2:1 and about 5:1.
10. The turbofan engine as recited in claim 1, wherein there is a third turbine section.
11. A turbofan engine comprising:
a fan having fan blades;
a compressor in communication with the fan section wherein the compressor includes a first compressor section including at least two stages and a second compressor section including at least five stages, wherein the second compressor section is configured to operate at a higher pressure than the first compressor section, wherein the fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor and a bypass ratio defined as air communicated through the bypass path relative to air communicated to the compressor is greater than about 6.0;
a combustor in fluid communication with the compressor;
a turbine in communication with the combustor, the turbine having a first turbine section and a second turbine section, wherein a ratio of airfoils in the first turbine section to the bypass ratio is less than about 170 and the first turbine section includes a maximum gas path radius and a ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50; and
a speed reduction mechanism coupled to the fan and rotatable by the turbine.
12. The turbofan engine as recited in claim 11, wherein the first turbine section includes three or more stages.
13. The turbofan engine as recited in claim 12, wherein the first turbine section includes four or more stages.
14. The turbofan engine as recited in claim 13, wherein the second turbine section comprises a high pressure turbine and the first turbine section comprises a low pressure turbine that operates at a pressure lower than the high pressure turbine section and the low pressure turbine includes an airfoil count below about 1600.
15. The turbofan engine as recited in claim 14, wherein a hub-to-tip ratio (Ri:Ro) of the low pressure turbine section is between about 0.4 and about 0.5 measured at a maximum Ro axial location within the low pressure turbine.
16. The turbofan engine as recited in claim 15, wherein the bypass ratio is greater than about 8.0.
17. The turbofan engine as recited in claim 16, wherein a fan pressure ratio across the fan is less than about 1.45.
18. The turbofan engine as recited in claim 17, wherein the speed reduction mechanism comprises an epicyclic gearbox.
19. The turbofan engine as recited in claim 18, wherein the epicyclic gearbox provides a speed reduction ratio between about 2:1 and about 5:1.
US14/793,785 2007-08-01 2015-07-08 Turbine section of high bypass turbofan Abandoned US20150377123A1 (en)

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US11/832,107 US8256707B2 (en) 2007-08-01 2007-08-01 Engine mounting configuration for a turbofan gas turbine engine
US201161498516P true 2011-06-17 2011-06-17
US201261593190P true 2012-01-31 2012-01-31
US13/475,252 US8844265B2 (en) 2007-08-01 2012-05-18 Turbine section of high bypass turbofan
US13/599,175 US9010085B2 (en) 2007-08-01 2012-08-30 Turbine section of high bypass turbofan
US14/692,090 US20150345404A1 (en) 2007-08-01 2015-04-21 Turbine section of high bypass turbofan
US14/793,785 US20150377123A1 (en) 2007-08-01 2015-07-08 Turbine section of high bypass turbofan

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US15/292,249 US20170044978A1 (en) 2007-08-01 2016-10-13 Turbine section of high bypass turbofan
US15/292,472 US10060357B2 (en) 2007-08-01 2016-10-13 Turbine section of high bypass turbofan
US15/292,438 US20170298832A1 (en) 2007-08-01 2016-10-13 Turbine section of high bypass turbofan
US15/292,405 US10371061B2 (en) 2007-08-01 2016-10-13 Turbine section of high bypass turbofan
US16/025,022 US20190017445A1 (en) 2007-08-01 2018-07-02 Turbine section of high bypass turbofan
US16/025,094 US20190048803A1 (en) 2007-08-01 2018-07-02 Turbine section of high bypass turbofan
US16/025,038 US20190017446A1 (en) 2007-08-01 2018-07-02 Turbine section of high bypass turbofan

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US15/292,405 Continuation US10371061B2 (en) 2007-08-01 2016-10-13 Turbine section of high bypass turbofan
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US15/292,405 Active 2028-02-17 US10371061B2 (en) 2007-08-01 2016-10-13 Turbine section of high bypass turbofan
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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3561277A3 (en) * 2018-04-06 2020-01-01 Rolls-Royce plc Geared gas turbine engine
EP3594476A1 (en) * 2018-07-10 2020-01-15 Rolls-Royce plc A geared turbofan gas turbine engine mounting arrangement

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3020658A1 (en) * 2014-04-30 2015-11-06 Snecma Lubricating oil recovery cover for turbomachine equipment
GB201703521D0 (en) * 2017-03-06 2017-04-19 Rolls Royce Plc Geared turbofan

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120251306A1 (en) * 2009-11-20 2012-10-04 United Technologies Corporation Fan Rotor Support

Family Cites Families (120)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2258792A (en) 1941-04-12 1941-10-14 Westinghouse Electric & Mfg Co Turbine blading
US3021731A (en) 1951-11-10 1962-02-20 Wilhelm G Stoeckicht Planetary gear transmission
US2936655A (en) 1955-11-04 1960-05-17 Gen Motors Corp Self-aligning planetary gearing
US3194487A (en) 1963-06-04 1965-07-13 United Aircraft Corp Noise abatement method and apparatus
US3327971A (en) 1964-06-23 1967-06-27 Rolls Royce Mounting arrangement for lift engines
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
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
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
FR2291091B1 (en) 1974-11-13 1977-10-21 Snecma
US3935558A (en) 1974-12-11 1976-01-27 United Technologies Corporation Surge detector for turbine engines
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
US4044973A (en) 1975-12-29 1977-08-30 The Boeing Company Nacelle assembly and mounting structures for a turbofan jet propulsion engine
GB1516041A (en) 1977-02-14 1978-06-28 Secr Defence Multistage axial flow compressor stators
GB1574379A (en) 1977-08-24 1980-09-03 English Electric Co Ltd Turbines and like rotary machines
GB2010969A (en) 1977-12-22 1979-07-04 Rolls Royce Mounting for Gas Turbine Jet Propulsion Engine
US4266741A (en) 1978-05-22 1981-05-12 The Boeing Company Mounting apparatus for fan jet engine having mixed flow nozzle installation
GB2041090A (en) 1979-01-31 1980-09-03 Rolls Royce By-pass gas turbine engines
DE2940446C2 (en) 1979-10-05 1982-07-08 B. Braun Melsungen Ag, 3508 Melsungen, De
US4478551A (en) 1981-12-08 1984-10-23 United Technologies Corporation Turbine exhaust case design
US4696156A (en) 1986-06-03 1987-09-29 United Technologies Corporation Fuel and oil heat management system for a gas turbine engine
US4966338A (en) 1987-08-05 1990-10-30 General Electric Company Aircraft pylon
GB8822798D0 (en) 1988-09-28 1988-11-02 Short Brothers Ltd Ducted fan turbine engine
US4969325A (en) 1989-01-03 1990-11-13 General Electric Company Turbofan engine having a counterrotating partially geared fan drive turbine
US4979362A (en) 1989-05-17 1990-12-25 Sundstrand Corporation Aircraft engine starting and emergency power generating system
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
GB9116986D0 (en) 1991-08-07 1991-10-09 Rolls Royce Plc Gas turbine engine nacelle assembly
US5174525A (en) 1991-09-26 1992-12-29 General Electric Company Structure for eliminating lift load bending in engine core of turbofan
GB9125011D0 (en) 1991-11-25 1992-01-22 Rolls Royce Plc A mounting arrangement for a gas turbine engine
US5275357A (en) 1992-01-16 1994-01-04 General Electric Company Aircraft engine mount
US5320307A (en) 1992-03-25 1994-06-14 General Electric Company Aircraft engine thrust mount
GB2265418B (en) 1992-03-26 1995-03-08 Rolls Royce Plc Gas turbine engine casing
GB2266080A (en) 1992-04-16 1993-10-20 Rolls Royce Plc Mounting arrangement for a gas turbine engine.
US5317877A (en) 1992-08-03 1994-06-07 General Electric Company Intercooled turbine blade cooling air feed system
US5277382A (en) 1992-10-13 1994-01-11 General Electric Company Aircraft engine forward mount
GB2275308B (en) 1993-02-20 1997-02-26 Rolls Royce Plc A mounting for coupling a turbofan gas turbine engine to an aircraft structure
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
US5452575A (en) 1993-09-07 1995-09-26 General Electric Company Aircraft gas turbine engine thrust mount
US5524847A (en) 1993-09-07 1996-06-11 United Technologies Corporation Nacelle and mounting arrangement for an aircraft engine
US5443229A (en) 1993-12-13 1995-08-22 General Electric Company Aircraft gas turbine engine sideways mount
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
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
EP0839285B1 (en) 1994-12-14 2001-07-18 United Technologies Corporation Compressor stall and surge control using airflow asymmetry measruement
GB2303884B (en) 1995-04-13 1999-07-14 Rolls Royce Plc A mounting for coupling a turbofan gas turbine engine to an aircraft structure
JP2969075B2 (en) 1996-02-26 1999-11-02 ジャパンゴアテックス株式会社 Degassing device
GB2312251B (en) 1996-04-18 1999-10-27 Rolls Royce Plc Ducted fan gas turbine engine mounting
US5810287A (en) 1996-05-24 1998-09-22 The Boeing Company Aircraft support pylon
US5857836A (en) 1996-09-10 1999-01-12 Aerodyne Research, Inc. Evaporatively cooled rotor for a gas turbine engine
FR2755942B1 (en) 1996-11-21 1998-12-24 Snecma Suspension redundant prior to turbine engine
FR2755944B1 (en) 1996-11-21 1998-12-24 Snecma Redundant front suspension for a turbine engine
FR2755943B1 (en) 1996-11-21 1998-12-24 Snecma Redundant front suspension for a turbine engine
US5975841A (en) 1997-10-03 1999-11-02 Thermal Corp. Heat pipe cooling for turbine stators
US5921500A (en) 1997-10-08 1999-07-13 General Electric Company Integrated failsafe engine mount
US5927644A (en) 1997-10-08 1999-07-27 General Electric Company Double failsafe engine mount
US6126110A (en) 1997-12-22 2000-10-03 Mcdonnell Douglas Corporation Horizontally opposed trunnion forward engine mount system supported beneath a wing pylon
US5985470A (en) 1998-03-16 1999-11-16 General Electric Company Thermal/environmental barrier coating system for silicon-based materials
US6138949A (en) 1998-10-30 2000-10-31 Sikorsky Aircraft Corporation Main rotor pylon support structure
US6189830B1 (en) 1999-02-26 2001-02-20 The Boeing Company Tuned engine mounting system for jet aircraft
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
GB9927425D0 (en) 1999-11-20 2000-01-19 Rolls Royce Plc A gas turbine engine mounting arrangement
US6315815B1 (en) 1999-12-16 2001-11-13 United Technologies Corporation Membrane based fuel deoxygenator
US6223616B1 (en) 1999-12-22 2001-05-01 United Technologies Corporation Star gear system with lubrication circuit and lubrication method therefor
GB0002257D0 (en) 2000-02-02 2000-03-22 Rolls Royce Plc Rotary apparatus for a gas turbine engine
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
GB2375513B (en) 2001-05-19 2005-03-23 Rolls Royce Plc A mounting arrangement for a gas turbine engine
US6517027B1 (en) 2001-12-03 2003-02-11 Pratt & Whitney Canada Corp. Flexible/fixed support for engine cowl
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
US6607165B1 (en) 2002-06-28 2003-08-19 General Electric Company Aircraft engine mount with single thrust link
US6652222B1 (en) 2002-09-03 2003-11-25 Pratt & Whitney Canada Corp. Fan case design with metal foam between Kevlar
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
US6899518B2 (en) 2002-12-23 2005-05-31 Pratt & Whitney Canada Corp. Turbine shroud segment apparatus for reusing cooling air
US6709492B1 (en) 2003-04-04 2004-03-23 United Technologies Corporation Planar membrane deoxygenator
FR2856656B1 (en) 2003-06-30 2006-12-01 Snecma Moteurs Aircraft engine rear suspension with boomerang shaft and boomerang shaft
US7055330B2 (en) 2004-02-25 2006-06-06 United Technologies Corp Apparatus for driving an accessory gearbox in a gas turbine engine
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
US7328580B2 (en) 2004-06-23 2008-02-12 General Electric Company Chevron film cooled wall
US7121802B2 (en) 2004-07-13 2006-10-17 General Electric Company Selectively thinned turbine blade
US7134286B2 (en) 2004-08-24 2006-11-14 Pratt & Whitney Canada Corp. Gas turbine floating collar arrangement
US7409819B2 (en) 2004-10-29 2008-08-12 General Electric Company Gas turbine engine and method of assembling same
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
US7374403B2 (en) 2005-04-07 2008-05-20 General Electric Company Low solidity turbofan
US7500365B2 (en) 2005-05-05 2009-03-10 United Technologies Corporation Accessory gearbox
FR2887521B1 (en) 2005-06-28 2007-08-17 Airbus France Sas Engine assembly for an aircraft comprising an engine and a device for hitching such an engine
WO2007038673A1 (en) 2005-09-28 2007-04-05 Entrotech Composites, Llc Linerless prepregs, composite articles therefrom, and related methods
GB0520850D0 (en) 2005-10-14 2005-11-23 Rolls Royce Plc Fan static structure
US7591754B2 (en) 2006-03-22 2009-09-22 United Technologies Corporation Epicyclic gear train integral sun gear coupling design
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
US7694505B2 (en) 2006-07-31 2010-04-13 General Electric Company Gas turbine engine assembly and method of assembling same
US7841165B2 (en) 2006-10-31 2010-11-30 General Electric Company Gas turbine engine assembly and methods of assembling same
US8017188B2 (en) 2007-04-17 2011-09-13 General Electric Company Methods of making articles having toughened and untoughened regions
US7950237B2 (en) 2007-06-25 2011-05-31 United Technologies Corporation Managing spool bearing load using variable area flow nozzle
US20120124964A1 (en) 2007-07-27 2012-05-24 Hasel Karl L Gas turbine engine with improved fuel efficiency
US8256707B2 (en) 2007-08-01 2012-09-04 United Technologies Corporation Engine mounting configuration for a turbofan gas turbine engine
US8844265B2 (en) 2007-08-01 2014-09-30 United Technologies Corporation Turbine section of high bypass turbofan
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
US7997868B1 (en) 2008-11-18 2011-08-16 Florida Turbine Technologies, Inc. Film cooling hole for turbine airfoil
US8181441B2 (en) 2009-02-27 2012-05-22 United Technologies Corporation Controlled fan stream flow bypass
US8172716B2 (en) 2009-06-25 2012-05-08 United Technologies Corporation Epicyclic gear system with superfinished journal bearing
US8689538B2 (en) 2009-09-09 2014-04-08 The Boeing Company Ultra-efficient propulsor with an augmentor fan circumscribing a turbofan
US8439637B2 (en) 2009-11-20 2013-05-14 United Technologies Corporation Bellows preload and centering spring for a fan drive gear system
US9170616B2 (en) 2009-12-31 2015-10-27 Intel Corporation Quiet system cooling using coupled optimization between integrated micro porous absorbers and rotors
US8636195B2 (en) 2010-02-19 2014-01-28 General Electric Company Welding process and component formed thereby
US8905713B2 (en) 2010-05-28 2014-12-09 General Electric Company Articles which include chevron film cooling holes, and related processes
US8834099B1 (en) 2012-09-28 2014-09-16 United Technoloiies Corporation Low noise compressor rotor for geared turbofan engine
US8807916B2 (en) 2012-09-27 2014-08-19 United Technologies Corporation Method for setting a gear ratio of a fan drive gear system of a gas turbine engine
US10190496B2 (en) 2013-03-15 2019-01-29 United Technologies Corporation Turbofan engine bearing and gearbox arrangement
US8869504B1 (en) 2013-11-22 2014-10-28 United Technologies Corporation Geared turbofan engine gearbox arrangement

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120251306A1 (en) * 2009-11-20 2012-10-04 United Technologies Corporation Fan Rotor Support

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
Rauch, "Design Study of an Air Pump and Integral Lift Engine ALF-504 Using the Lycoming 502 Core", NASA/CR-120,992, July 1972 *

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3561277A3 (en) * 2018-04-06 2020-01-01 Rolls-Royce plc Geared gas turbine engine
EP3594476A1 (en) * 2018-07-10 2020-01-15 Rolls-Royce plc A geared turbofan gas turbine engine mounting arrangement

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US20170298832A1 (en) 2017-10-19
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US10060357B2 (en) 2018-08-28
US20190048803A1 (en) 2019-02-14
US20170044978A1 (en) 2017-02-16
US10371061B2 (en) 2019-08-06

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