+

WO2006060005A1 - Ensemble de rotor de turbine-ventilateur avec section d’induction integrale pour moteur de turbine a pression d’entree - Google Patents

Ensemble de rotor de turbine-ventilateur avec section d’induction integrale pour moteur de turbine a pression d’entree Download PDF

Info

Publication number
WO2006060005A1
WO2006060005A1 PCT/US2004/040174 US2004040174W WO2006060005A1 WO 2006060005 A1 WO2006060005 A1 WO 2006060005A1 US 2004040174 W US2004040174 W US 2004040174W WO 2006060005 A1 WO2006060005 A1 WO 2006060005A1
Authority
WO
WIPO (PCT)
Prior art keywords
fan
blade
multitude
inducer
section
Prior art date
Application number
PCT/US2004/040174
Other languages
English (en)
Inventor
Gabriel L. Suciu
James W. Norris
Craig A. Nordeen
Brian Merry
Original Assignee
United Technologies Corporation
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to PCT/US2004/040174 priority Critical patent/WO2006060005A1/fr
Priority to US11/719,854 priority patent/US20090169385A1/en
Publication of WO2006060005A1 publication Critical patent/WO2006060005A1/fr

Links

Classifications

    • 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
    • F01D5/022—Blade-carrying members, e.g. rotors with concentric rows of axial blades
    • 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
    • F02C3/073—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 compressor and turbine stages being concentric
    • 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
    • F02C3/08—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
    • 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
    • F02K3/068—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 being characterised by a short axial length relative to the diameter
    • 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
    • F05D2260/60—Fluid transfer

Definitions

  • the present invention relates to a tip turbine engine, and more particularly to a fan-turbine rotor assembly which includes an inducer formed therein.
  • An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a compressor, a combustor, and an aft turbine all located along a common longitudinal axis.
  • a compressor and a turbine of the engine are interconnected by a shaft.
  • the compressor is rotatably driven to compress air entering the combustor to a relatively high pressure. This pressurized air is then mixed with fuel in a combustor and ignited to form a high energy gas stream.
  • the gas stream flows axially aft to rotatably drive the turbine which rotatably drives the compressor through the shaft.
  • the gas stream is also responsible for rotating the bypass fan.
  • Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor and ignited to form a high energy gas stream which drives the turbine integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490.
  • the tip turbine engine provides a thrust to weight ratio equivalent to conventional turbofan engines of the same class within a package of significantly shorter length.
  • the fan-turbine rotor assembly includes a multitude of components which rotate at relatively high speeds to generate bypass airflow while communicating a core airflow through each of the multitude of hollow fan blades.
  • a large percentage of the expense associated with a tip turbine engine is the manufacture of the fan-turbine rotor assembly and the integration of the inducer with the fan hub.
  • the fan-turbine rotor assembly for a tip turbine engine includes a fan hub which has an outer periphery scalloped by a multitude of elongated openings which extend into a fan hub web. Each elongated opening defines an inducer section and a blade receipt section to retain a hollow fan blade section. The blade receipt section retains each of the hollow fan blade sections adjacent each inducer section.
  • An inner fan blade mount is located adjacent an inducer exhaust section to communicate a core airflow communication path from within each inducer section into the core airflow passage within each fan blade section.
  • the inducer is cast directly into the fan hub which minimizes leakage between each fan blade section and each inducer section to provide increased engine efficiency. Manufacturing and assembly is also readily facilitated.
  • the present invention therefore provides an inducer arrangement for a fan- turbine rotor assembly which is relatively inexpensive to manufacture yet provides a high degree of reliability.
  • Figure 1 is a partial sectional perspective view of a tip turbine engine
  • Figure 2 is a longitudinal sectional view of a tip turbine engine along an engine centerline
  • Figure 3 is an exploded view of a fan-turbine rotor assembly
  • Figure 4 is an assembled view of a fan-turbine rotor assembly
  • Figure 5A is an expanded radial sectional view of an inducer section
  • Figure 5B is a sequential sectional view of the fan hub illustrating the inducer sections therewith;
  • Figure 6 is a schematic view of airflow through the last stage of an axial compressor and into the inducer
  • Figure 7A is an expanded phantom perspective view of a fan blade mounted to a hub of a fan-turbine rotor assembly
  • Figure 7B is an expanded partially sectioned perspective view of a fan blade mounted to a hub of a fan-turbine rotor assembly
  • Figure 7C is an expanded partially sectioned perspective view of a diffuser section of a fan blade.
  • FIG. 1 illustrates a general perspective partial sectional view of a tip turbine engine type gas turbine engine 10.
  • the engine 10 includes an outer nacelle 12, a rotationally fixed static outer support structure 14 and a rotationally fixed static inner support structure 16.
  • a multitude of fan inlet guide vanes 18 are mounted between the static outer support structure 14 and the static inner support structure 16.
  • Each inlet guide vane preferably includes a variable trailing edge 18A.
  • a nose cone 20 is preferably located along the engine centerline A to smoothly direct airflow into an axial compressor 22 adjacent thereto.
  • the axial compressor 22 is mounted about the engine centerline A behind the nose cone 20.
  • a fan-turbine rotor assembly 24 is mounted for rotation about the engine centerline A aft of the axial compressor 22.
  • the fan-turbine rotor assembly 24 includes a multitude of hollow fan blades 28 to provide internal, centrifugal compression of the compressed airflow from the axial compressor 22 for distribution to an annular combustor 30 located within the rotationally fixed static outer support structure 14.
  • a turbine 32 includes a multitude of tip turbine blades 34 (two stages shown) which rotatably drive the hollow fan blades 28 relative a multitude of tip turbine stators 36 which extend radially inwardly from the static outer support structure 14.
  • the annular combustor 30 is axially forward of the turbine 32 and communicates with the turbine 32.
  • the rotationally fixed static inner support structure 16 includes a splitter 40, a static inner support housing 42 and an static outer support housing 44 located coaxial to said engine centerline A.
  • the axial compressor 22 includes the axial compressor rotor 46 from which a plurality of compressor blades 52 extend radially outwardly and a compressor case 50 fixedly mounted to the splitter 40.
  • a plurality of compressor vanes 54 extend radially inwardly from the compressor case 50 between stages of the compressor blades 52.
  • the compressor blades 52 and compressor vanes 54 are arranged circumferentially about the axial compressor rotor 46 in stages (three stages of compressor blades 52 and compressor vanes 54 are shown in this example).
  • the axial compressor rotor 46 is mounted for rotation upon the static inner support housing 42 through a forward bearing assembly 68 and an aft bearing assembly 62.
  • the fan-turbine rotor assembly 24 includes a fan hub 64 that supports a multitude of the hollow fan blades 28.
  • Each fan blade 28 includes an inducer section 66, a hollow fan blade section 72 and a diffuser section 74.
  • the inducer section 66 receives airflow from the axial compressor 22 generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction.
  • the airflow is radially communicated through a core airflow passage 80 within the fan blade section 72 where the airflow is centrifugally compressed. From the core airflow passage 80, the airflow is turned and diffused toward an axial airflow direction toward the annular combustor 30.
  • the airflow is diffused axially forward in the engine 10, however, the airflow may alternatively be communicated in another direction.
  • a gearbox assembly 90 aft of the fan-turbine rotor assembly 24 provides a speed increase between the fan-turbine rotor assembly 24 and the axial compressor 22.
  • the gearbox assembly 90 could provide a speed decrease between the fan-turbine rotor assembly 24 and the axial compressor rotor 46.
  • the gearbox assembly 90 is mounted for rotation between the static inner support housing 42 and the static outer support housing 44.
  • the gearbox assembly 90 includes a sun gear shaft 92 which rotates with the axial compressor 22 and a planet carrier 94 which rotates with the fan-turbine rotor assembly 24 to provide a speed differential therebetween.
  • the gearbox assembly 90 is preferably a planetary gearbox that provides co-rotating or counter-rotating rotational engagement between the fan- turbine rotor assembly 24 and an axial compressor rotor 46.
  • the gearbox assembly 90 is mounted for rotation between the sun gear shaft 92 and the static outer support housing 44 through a forward bearing 96 and a rear bearing 98.
  • the forward bearing 96 and the rear bearing 98 are both tapered roller bearings and both handle radial loads.
  • the forward bearing 96 handles the aft axial loads while the rear bearing 98 handles the forward axial loads.
  • the sun gear shaft 92 is rotationally engaged with the axial compressor rotor 46 at a splined interconnection 100 or the like.
  • the compressed air from the axial compressor 22 enters the inducer section 66 in a direction generally parallel to the engine centerline A and is turned by the inducer section 66 radially outwardly through the core airflow passage 80 of the hollow fan blades 28.
  • the airflow is further compressed centrifugally in the hollow fan blades 28 by rotation of the hollow fan blades 28.
  • From the core airflow passage 80, the airflow is turned and diffused axially forward in the engine 10 into the annular combustor 30.
  • the compressed core airflow from the hollow fan blades 28 is mixed with fuel in the annular combustor 30 and ignited to form a high-energy gas stream.
  • the high-energy gas stream is expanded over the multitude of tip turbine blades 34 mounted about the outer periphery of the fan-turbine rotor assembly 24 to drive the fan-turbine rotor assembly 24, which in turn drives the axial compressor 22 through the gearbox assembly 90.
  • the fan-turbine rotor assembly 24 discharges fan bypass air axially aft to merge with the core airflow from the turbine 32 in an exhaust case 106.
  • a multitude of exit guide vanes 108 are located between the static outer support housing 44 and the rotationally fixed static outer support structure 14 to guide the combined airflow out of the engine 10 to provide forward thrust.
  • An exhaust mixer 110 mixes the airflow from the turbine blades 34 with the bypass airflow through the fan blades 28.
  • the fan hub 64 is the primary structural support of the fan-turbine rotor assembly 24 ( Figure 4).
  • the fan hub 64 is preferably forged and then milled to provide the desired geometry.
  • the fan hub 64 defines a bore 111 and an outer periphery 112.
  • the outer periphery 112 is preferably scalloped by a multitude of elongated openings 111.
  • the fan hub 64 is the primary structural support of the fan- turbine rotor assembly 24.
  • the fan hub 64 supports the multitude of fan blades 28, a diffuser 114, and the turbine 32.
  • the diffuser 114 defines a diffuser surface 119 formed about the outer periphery of the fan blade sections 72 to provide structural support to the outer tips of the fan blade sections 72 and to turn and diffuse the airflow from the radial core airflow passage 80 ( Figure 3) toward an axial airflow direction.
  • the turbine 32 is mounted to the diffuser surface 119 as one or more turbine ring rotors 118a, 118b which may include a multitude of turbine blade clusters.
  • each inducer section 66 formed by the fan hub 64 is essentially a conduit that defines an inducer passage 118 between an inducer inlet section 120 and an inducer exit section 128 Figures 5 A, 5B).
  • each inducer passage 118 provides separate airflow communication to each core airflow passage 80 when each fan blade section 72 is mounted within each elongated opening 114.
  • each fan blade section 72 includes an attached diffuser section 74 such that the diffuser surface 119 is formed when the fan-turbine rotor assembly 24 is assembled.
  • Figure 6 schematically illustrates the relationship of the angle of the last stage of the compressor rotor blade 52 (one shown) and the last stage of the compressor vanes 54 in the three stage axial compressor 22 ( Figure 2) prior to communication of the airflow from the axial compressor 22 into the inducer sections 66 in the engine 10.
  • the compressor rotor blade 52 is angled relative to the engine centerline A to provide an angle of a relative velocity vector, VrI.
  • the velocity of the counter-rotating compressor blade 52 gives a blade velocity vector, VbI.
  • the resultant vector indicating the resultant core airflow from the compressor blade 52, is the absolute velocity vector, VaI.
  • a stator leading edge 541 of the compressor stator 54 is angled to correspond with the absolute velocity vector, VaI from the compressor rotor blade 52 to efficiently receive and compress the core airflow from the compressor blade 52.
  • the vane trailing edge 54t is angled relative to the engine centerline A to compress and redirect the airflow toward the inducer section 66 (one shown) as the inducer 116 rotates relative thereto at a vane absolute velocity vector, VaI.
  • the inducer inlet 120 of the inducer section 66 is angled to efficiently receive the core airflow from the vane trailing edge 54t which flows toward the inducer section 66 at the absolute velocity vector, VaI from the vane 54.
  • the velocity of the inducer section 66 gives an inducer velocity vector, VbI. Referring to the inducer velocity triangle It, the angle of the inducer 66 is selected such that the sum of the inducer relative velocity vector VrI and the inducer velocity vector VbI match the angle of the core airflow incoming from the compressor vane trailing edge 54t (absolute velocity vector, VaI).
  • the specific angles will depend on a variety of factors, including anticipated blade velocities and the design choices made in the earlier stages of the compressor blades 52 and compressor vanes 54 to provide a length sufficient to turn the core airflow from axial flow to radial flow while decreasing the overall length of the engine 10.
  • the axial compressor 22 may alternatively counter-rotate relative to inducer 116 as disclosed in co-pending application entitled "COUNTER- ROTATING GEARBOX FOR TIP TURBINE ENGINE,” which is assigned to the assignee of the present invention and which is hereby incorporated by reference in its entirety.
  • the fan hub 64 retains each hollow fan blade section 72 through a blade receipt section 122.
  • the blade receipt section 122 preferably forms an axial semi-cylindrical opening formed along the axial length of the elongated openings 111. It should be understood that other retention structures such as a dove-tail, fir-tree, or bulb-type engagement structure will likewise be usable with the present invention.
  • Each hollow fan blade section 72 includes a fan blade mount section 124 that corresponds with the blade receipt section 122 to retain the hollow fan blade section 72 within the fan hub 64.
  • the fan blade mount 124 preferably includes a semi- cylindrical portion to radially retain the fan blade 28.
  • the inner fan blade mount 124 is preferably uni- directionally mounted into the blade receipt section 122 such as from the rear face of the fan hub 64.
  • the fan blade mount section 124 engages the blade receipt section 122 during operation of the fan-turbine rotor assembly 24 to provide a directional lock therebetween. That is, the inner fan blade mount 124 and the blade receipt section 122 may be frustoconical or axially non-symmetrical such that the forward segments form a smaller perimeter than the rear segment to provide a wedged engagement therebetween when assembled.
  • Each inducer section 66 within the fan hub 64 receives core airflow communication from the inducer passages 118 into the core airflow passage 80 and turns and diffuses the airflow through each diffuser section 74 of the diffuser 114 (also illustrated in Figure 7C).

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

La présente invention concerne un ensemble de rotor de turbine-ventilateur pour un moteur de turbine à pression d’entrée comprenant un moyeu de ventilateur avec une périphérie extérieure festonnée par une multitude d’ouvertures allongées s’étendant dans un tissu de moyeu de ventilateur. Chaque ouverture allongée définit une section d’induction et une section de réception de pale pour retenir une section de pale de ventilateur creuse. La section de réception de pale retient chacune des sections de pale de ventilateur creuse à côté de chaque section d’induction. Les sections d’induction sont moulées directement dans le moyeu de ventilateur ce qui minimise la fuite entre chaque section de pale de ventilateur et chacune des sections d’induction respective, pour minimiser la fuite d’air et accroître l’efficacité du moteur.
PCT/US2004/040174 2004-12-01 2004-12-01 Ensemble de rotor de turbine-ventilateur avec section d’induction integrale pour moteur de turbine a pression d’entree WO2006060005A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
PCT/US2004/040174 WO2006060005A1 (fr) 2004-12-01 2004-12-01 Ensemble de rotor de turbine-ventilateur avec section d’induction integrale pour moteur de turbine a pression d’entree
US11/719,854 US20090169385A1 (en) 2004-12-01 2004-12-01 Fan-turbine rotor assembly with integral inducer section for a tip turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2004/040174 WO2006060005A1 (fr) 2004-12-01 2004-12-01 Ensemble de rotor de turbine-ventilateur avec section d’induction integrale pour moteur de turbine a pression d’entree

Publications (1)

Publication Number Publication Date
WO2006060005A1 true WO2006060005A1 (fr) 2006-06-08

Family

ID=35447605

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2004/040174 WO2006060005A1 (fr) 2004-12-01 2004-12-01 Ensemble de rotor de turbine-ventilateur avec section d’induction integrale pour moteur de turbine a pression d’entree

Country Status (2)

Country Link
US (1) US20090169385A1 (fr)
WO (1) WO2006060005A1 (fr)

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7607286B2 (en) 2004-12-01 2009-10-27 United Technologies Corporation Regenerative turbine blade and vane cooling for a tip turbine engine
US7874802B2 (en) 2004-12-01 2011-01-25 United Technologies Corporation Tip turbine engine comprising turbine blade clusters and method of assembly
US7878762B2 (en) 2004-12-01 2011-02-01 United Technologies Corporation Tip turbine engine comprising turbine clusters and radial attachment lock arrangement therefor
US7887296B2 (en) 2004-12-01 2011-02-15 United Technologies Corporation Fan blade with integral diffuser section and tip turbine blade section for a tip turbine engine
US7927075B2 (en) 2004-12-01 2011-04-19 United Technologies Corporation Fan-turbine rotor assembly for a tip turbine engine
US8468795B2 (en) 2004-12-01 2013-06-25 United Technologies Corporation Diffuser aspiration for a tip turbine engine
US8672630B2 (en) 2004-12-01 2014-03-18 United Technologies Corporation Annular turbine ring rotor
US8807936B2 (en) 2004-12-01 2014-08-19 United Technologies Corporation Balanced turbine rotor fan blade for a tip turbine engine
US8950171B2 (en) 2004-12-01 2015-02-10 United Technologies Corporation Counter-rotating gearbox for tip turbine engine
US9003768B2 (en) 2004-12-01 2015-04-14 United Technologies Corporation Variable fan inlet guide vane assembly, turbine engine with such an assembly and corresponding controlling method

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8757959B2 (en) * 2004-12-01 2014-06-24 United Technologies Corporation Tip turbine engine comprising a nonrotable compartment
US7840632B2 (en) * 2005-01-05 2010-11-23 New Noah Technology (Shenzhen) Co., Ltd. System and method for portable multimedia network learning machine and remote information transmission thereof
US20130219907A1 (en) * 2012-02-29 2013-08-29 Frederick M. Schwarz Geared turbofan architecture for improved thrust density
US10018119B2 (en) 2012-04-02 2018-07-10 United Technologies Corporation Geared architecture with inducer for gas turbine engine
US10550764B2 (en) * 2013-12-13 2020-02-04 United Technologies Corporation Architecture for an axially compact, high performance propulsion system
US10669946B2 (en) 2015-06-05 2020-06-02 Raytheon Technologies Corporation Geared architecture for a gas turbine engine
US10876407B2 (en) * 2017-02-16 2020-12-29 General Electric Company Thermal structure for outer diameter mounted turbine blades
CN114651120B (zh) * 2019-11-11 2024-12-03 Tns技术公司 燃气涡轮发动机
US11428160B2 (en) 2020-12-31 2022-08-30 General Electric Company Gas turbine engine with interdigitated turbine and gear assembly

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1033849A (fr) * 1951-03-12 1953-07-16 Perfectionnements aux turbines à gaz
US3283509A (en) * 1963-02-21 1966-11-08 Messerschmitt Boelkow Blohm Lifting engine for vtol aircraft
US3496725A (en) * 1967-11-01 1970-02-24 Gen Applied Science Lab Inc Rocket action turbofan engine
WO2004011788A1 (fr) * 2002-07-30 2004-02-05 The Regents Of The University Of California Turbine a rotor simple
US20040025490A1 (en) * 2002-04-15 2004-02-12 Paul Marius A. Integrated bypass turbojet engines for air craft and other vehicles

Family Cites Families (62)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2221685A (en) * 1939-01-18 1940-11-12 Gen Electric Elastic fluid turbine bucket unit
NL69078C (fr) * 1944-01-31
US2595829A (en) * 1946-12-19 1952-05-06 Benson Mfg Company Axial flow fan and compressor
US2620554A (en) * 1948-09-29 1952-12-09 Westinghouse Electric Corp Method of manufacturing turbine blades
US2698711A (en) * 1951-02-06 1955-01-04 United Aircraft Corp Compressor air bleed closure
US2753140A (en) * 1951-07-28 1956-07-03 United Aircraft Corp Engine mount
US2874926A (en) * 1954-12-31 1959-02-24 Gen Motors Corp Compressor air bleed-off
US2936978A (en) * 1957-03-29 1960-05-17 United Aircraft Corp Rear engine mount
US3023998A (en) * 1959-03-13 1962-03-06 Jr Walter H Sanderson Rotor blade retaining device
US3042349A (en) * 1959-11-13 1962-07-03 Gen Electric Removable aircraft engine mounting arrangement
US3132842A (en) * 1962-04-13 1964-05-12 Gen Electric Turbine bucket supporting structure
US3267667A (en) * 1964-06-25 1966-08-23 Gen Electric Reversible flow fan
US3269120A (en) * 1964-07-16 1966-08-30 Curtiss Wright Corp Gas turbine engine with compressor and turbine passages in a single rotor element
US3363419A (en) * 1965-04-27 1968-01-16 Rolls Royce Gas turbine ducted fan engine
BE755508A (fr) * 1966-05-16 1971-02-01 Gen Electric Rotor pour moteurs a turbine a gaz
US3404831A (en) * 1966-12-07 1968-10-08 Gen Electric Turbine bucket supporting structure
GB1113087A (en) * 1967-02-27 1968-05-08 Rolls Royce Gas turbine power plant
GB1294898A (fr) * 1969-12-13 1972-11-01
GB1309721A (en) * 1971-01-08 1973-03-14 Secr Defence Fan
US4563875A (en) * 1974-07-24 1986-01-14 Howald Werner E Combustion apparatus including an air-fuel premixing chamber
GB1484898A (en) * 1974-09-11 1977-09-08 Rolls Royce Ducted fan gas turbine engine
US4271674A (en) * 1974-10-17 1981-06-09 United Technologies Corporation Premix combustor assembly
US3979087A (en) * 1975-07-02 1976-09-07 United Technologies Corporation Engine mount
GB2044358B (en) * 1979-03-10 1983-01-19 Rolls Royce Gas turbine jet engine mounting
US4251987A (en) * 1979-08-22 1981-02-24 General Electric Company Differential geared engine
GB2098719B (en) * 1981-05-20 1984-11-21 Rolls Royce Gas turbine engine combustion apparatus
US4460316A (en) * 1982-12-29 1984-07-17 Westinghouse Electric Corp. Blade group with pinned root
US4505640A (en) * 1983-12-13 1985-03-19 United Technologies Corporation Seal means for a blade attachment slot of a rotor assembly
US4817382A (en) * 1985-12-31 1989-04-04 The Boeing Company Turboprop propulsion apparatus
GB2195712B (en) * 1986-10-08 1990-08-29 Rolls Royce Plc A turbofan gas turbine engine
US4785625A (en) * 1987-04-03 1988-11-22 United Technologies Corporation Ducted fan gas turbine power plant mounting
DE3714990A1 (de) * 1987-05-06 1988-12-01 Mtu Muenchen Gmbh Propfan-turbotriebwerk
US4834614A (en) * 1988-11-07 1989-05-30 Westinghouse Electric Corp. Segmental vane apparatus and method
US5010729A (en) * 1989-01-03 1991-04-30 General Electric Company Geared counterrotating turbine/fan propulsion system
GB2234035B (en) * 1989-07-21 1993-05-12 Rolls Royce Plc A reduction gear assembly and a gas turbine engine
EP0431800B1 (fr) * 1989-12-05 1994-08-31 ROLLS-ROYCE plc Dispositif d'accrochage de sécurité pour moteurs
FR2661213B1 (fr) * 1990-04-19 1992-07-03 Snecma Moteur d'aviation a tres grand taux de dilution et du type dit contrafan avant.
GB9009588D0 (en) * 1990-04-28 1990-06-20 Rolls Royce Plc A hydraulic seal and method of assembly
FR2677953B1 (fr) * 1991-06-19 1993-09-10 Snecma Structure de suspension arriere d'un turboreacteur.
US5267397A (en) * 1991-06-27 1993-12-07 Allied-Signal Inc. Gas turbine engine module assembly
GB9116986D0 (en) * 1991-08-07 1991-10-09 Rolls Royce Plc Gas turbine engine nacelle assembly
GB2262313B (en) * 1991-12-14 1994-09-21 Rolls Royce Plc Aerofoil blade containment
US5466198A (en) * 1993-06-11 1995-11-14 United Technologies Corporation Geared drive system for a bladed propulsor
GB2303884B (en) * 1995-04-13 1999-07-14 Rolls Royce Plc A mounting for coupling a turbofan gas turbine engine to an aircraft structure
GB2332024B (en) * 1997-12-03 2000-12-13 Rolls Royce Plc Rotary assembly
DE19828562B4 (de) * 1998-06-26 2005-09-08 Mtu Aero Engines Gmbh Triebwerk mit gegenläufig drehenden Rotoren
DE19844843B4 (de) * 1998-09-30 2006-02-09 Mtu Aero Engines Gmbh Planetengetriebe
US6095750A (en) * 1998-12-21 2000-08-01 General Electric Company Turbine nozzle assembly
US6102361A (en) * 1999-03-05 2000-08-15 Riikonen; Esko A. Fluidic pinch valve system
GB9922619D0 (en) * 1999-09-25 1999-11-24 Rolls Royce Plc A gas turbine engine blade containment assembly
US6223616B1 (en) * 1999-12-22 2001-05-01 United Technologies Corporation Star gear system with lubrication circuit and lubrication method therefor
GB0019533D0 (en) * 2000-08-10 2000-09-27 Rolls Royce Plc A combustion chamber
US6430917B1 (en) * 2001-02-09 2002-08-13 The Regents Of The University Of California Single rotor turbine engine
GB0107973D0 (en) * 2001-03-30 2001-05-23 Rolls Royce Plc A gas turbine engine blade containment assembly
US6708482B2 (en) * 2001-11-29 2004-03-23 General Electric Company Aircraft engine with inter-turbine engine frame
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
GB0206163D0 (en) * 2002-03-15 2002-04-24 Hansen Transmissions Int Gear unit lubrication
US20030192303A1 (en) * 2002-04-15 2003-10-16 Paul Marius A. Integrated bypass turbojet engines for aircraft and other vehicles
US6910854B2 (en) * 2002-10-08 2005-06-28 United Technologies Corporation Leak resistant vane cluster
US7021042B2 (en) * 2002-12-13 2006-04-04 United Technologies Corporation Geartrain coupling for a turbofan engine
US7927075B2 (en) * 2004-12-01 2011-04-19 United Technologies Corporation Fan-turbine rotor assembly for a tip turbine engine
WO2006060003A2 (fr) * 2004-12-01 2006-06-08 United Technologies Corporation Pale de ventilateur comprenant une section de diffuseur integrale et une section de pale de turbine a aube destine a une moteur a turbine a aube

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1033849A (fr) * 1951-03-12 1953-07-16 Perfectionnements aux turbines à gaz
US3283509A (en) * 1963-02-21 1966-11-08 Messerschmitt Boelkow Blohm Lifting engine for vtol aircraft
US3496725A (en) * 1967-11-01 1970-02-24 Gen Applied Science Lab Inc Rocket action turbofan engine
US20040025490A1 (en) * 2002-04-15 2004-02-12 Paul Marius A. Integrated bypass turbojet engines for air craft and other vehicles
WO2004011788A1 (fr) * 2002-07-30 2004-02-05 The Regents Of The University Of California Turbine a rotor simple

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7607286B2 (en) 2004-12-01 2009-10-27 United Technologies Corporation Regenerative turbine blade and vane cooling for a tip turbine engine
US7874802B2 (en) 2004-12-01 2011-01-25 United Technologies Corporation Tip turbine engine comprising turbine blade clusters and method of assembly
US7878762B2 (en) 2004-12-01 2011-02-01 United Technologies Corporation Tip turbine engine comprising turbine clusters and radial attachment lock arrangement therefor
US7887296B2 (en) 2004-12-01 2011-02-15 United Technologies Corporation Fan blade with integral diffuser section and tip turbine blade section for a tip turbine engine
US7927075B2 (en) 2004-12-01 2011-04-19 United Technologies Corporation Fan-turbine rotor assembly for a tip turbine engine
US8468795B2 (en) 2004-12-01 2013-06-25 United Technologies Corporation Diffuser aspiration for a tip turbine engine
US8672630B2 (en) 2004-12-01 2014-03-18 United Technologies Corporation Annular turbine ring rotor
US8807936B2 (en) 2004-12-01 2014-08-19 United Technologies Corporation Balanced turbine rotor fan blade for a tip turbine engine
US8950171B2 (en) 2004-12-01 2015-02-10 United Technologies Corporation Counter-rotating gearbox for tip turbine engine
US9003768B2 (en) 2004-12-01 2015-04-14 United Technologies Corporation Variable fan inlet guide vane assembly, turbine engine with such an assembly and corresponding controlling method

Also Published As

Publication number Publication date
US20090169385A1 (en) 2009-07-02

Similar Documents

Publication Publication Date Title
US7887296B2 (en) Fan blade with integral diffuser section and tip turbine blade section for a tip turbine engine
EP1825112B1 (fr) Moteur a turbine de bout en porte-a-faux
EP1828568B1 (fr) Ensemble rotor de soufflante-turbine pour moteur a turbine de bout
US8950171B2 (en) Counter-rotating gearbox for tip turbine engine
US7976273B2 (en) Tip turbine engine support structure
EP1888905B1 (fr) Agencement d etancheite pour un ensemble de rotor de soufflante de turbine d' extremite
US7883315B2 (en) Seal assembly for a fan rotor of a tip turbine engine
EP1828574B1 (fr) Ensemble de boite de vitesses a couplage proximal pour moteur de turbine a pression d'entree
US20090169385A1 (en) Fan-turbine rotor assembly with integral inducer section for a tip turbine engine
EP1947297A2 (fr) Ensemble joint pour rotor de soufflante-turbine d'un moteur à turbine en bout
EP1841959B1 (fr) Ailettes de rotor de soufflante pour moteur a turbine en bout
EP1834071B1 (fr) Inducteur de pale de ventilateur de moteur de turbine a pression d'entree
US7845157B2 (en) Axial compressor for tip turbine engine
EP1831520B1 (fr) Moteur à turbine d'extremité et procédé d'opération associé
US7883314B2 (en) Seal assembly for a fan-turbine rotor of a tip turbine engine
WO2006060002A1 (fr) Pale de ventilateur avec une multitude de canaux d’ecoulement internes

Legal Events

Date Code Title Description
AK Designated states

Kind code of ref document: A1

Designated state(s): AE AG AL AM AT AU AZ BA BB BG BR BW BY BZ CA CH CN CO CR CU CZ DE DK DM DZ EC EE EG ES FI GB GD GE GH GM HR HU ID IL IN IS JP KE KG KP KR KZ LC LK LR LS LT LU LV MA MD MG MK MN MW MX MZ NA NI NO NZ OM PG PH PL PT RO RU SC SD SE SG SK SL SY TJ TM TN TR TT TZ UA UG US UZ VC VN YU ZA ZM ZW

AL Designated countries for regional patents

Kind code of ref document: A1

Designated state(s): GM KE LS MW MZ NA SD SL SZ TZ UG ZM ZW AM AZ BY KG KZ MD RU TJ TM AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LT LU MC NL PL PT RO SE SI SK TR BF BJ CF CG CI CM GA GN GQ GW ML MR NE SN TD TG

121 Ep: the epo has been informed by wipo that ep was designated in this application
DPE1 Request for preliminary examination filed after expiration of 19th month from priority date (pct application filed from 20040101)
WWE Wipo information: entry into national phase

Ref document number: 11719854

Country of ref document: US

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 04822060

Country of ref document: EP

Kind code of ref document: A1

点击 这是indexloc提供的php浏览器服务，不要输入任何密码和下载