WO2006060005A1 - Fan-turbine rotor assembly with integral inducer section for a tip turbine engine - Google Patents

Fan-turbine rotor assembly with integral inducer section for a tip turbine engine Download PDF

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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
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WO
WIPO (PCT)
Prior art keywords
fan
blade
multitude
inducer
section
Prior art date
Application number
PCT/US2004/040174
Other languages
French (fr)
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/en
Priority to US11/719,854 priority patent/US20090169385A1/en
Publication of WO2006060005A1 publication Critical patent/WO2006060005A1/en

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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).

Abstract

A fan-turbine rotor assembly for a tip turbine engine includes a fan hub with 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. The inducer sections are cast directly into the fan hub which minimizes leakage between each fan blade section and each of the respective inducer sections to minimize airflow leakage and increase engine efficiency.

Description

FAN-TURBINE ROTOR ASSEMBLY WITH INTEGRAL INDUCER SECTION FOR A TIP TURBINE ENGINE
BACKGROUND OF THE INVENTION This invention was made with government support under Contract No.:
F33657-03-C-2044. The government therefore has certain rights in this invention.
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. In some instances, there are multiple shafts or spools. In such instances, there is a separate turbine connected to a separate corresponding compressor through each shaft. In most instances, the lowest pressure turbine will drive the bypass fan. Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable longitudinal length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications. A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. 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.
One significant rotational component of a tip turbine engine is the fan- turbine rotor assembly. 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.
Accordingly, it is desirable to provide an inducer arrangement for a fan- turbine rotor assembly, which is relatively inexpensive to manufacture yet provides a high degree of reliability.
SUMMARY OF THE INVENTION
The fan-turbine rotor assembly for a tip turbine engine according to the present invention 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. BRIEF DESCRIPTION OF THE DRAWINGS
The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows:
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; and Figure 7C is an expanded partially sectioned perspective view of a diffuser section of a fan blade.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
Figure 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.
Referring to Figure 2, 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. Preferably 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. Alternatively, 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. In operation, air enters the axial compressor 22, where it is compressed by the three stages of the compressor blades 52 and compressor vanes 54. 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. Concurrent therewith, 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.
Referring to Figure 3, the fan-turbine rotor assembly 24 is illustrated in an exploded view. 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.
Referring to Figure 4, the fan hub 64 itself forms the multitude of inducer sections 66. 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).
Referring to Figures 5A and 5B, the inducer sections 66 together form the inducer 116 of the fan-turbine rotor assembly 24. The inducer inlet section 120 of each inducer passage 118 extends forward of the fan hub 64 and is canted toward a rotational direction of the fan hub 64 such that inducer inlet 120 operates as an air scoop during rotation of the fan-turbine rotor assembly 24. 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. Preferably, 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. Referring to the compressor blade velocity triangle Bt, 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.
Referring to the vane velocity vector St, 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).
It should be understood that 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. It should be understood that 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.
Referring to Figure 7A, 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.
Referring to Figure 7B, 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).
It should be understood that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The preferred embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.

Claims

1. A fan hub assembly for a tip turbine engine comprising: a fan hub defining a hub axis of rotation, said fan hub defining a multitude of elongated openings through an outer periphery of said fan hub; a blade receipt section defined by each of said elongated openings; and an inducer section defined within each of said elongated openings to turn an airflow from a generally axial direction to a generally radial direction.
2. The fan hub assembly as recited in claim 1, wherein said inducer section is cast within said fan hub.
3. The fan hub assembly as recited in claim 1, further comprising a multitude of fan blades, each of said multitude of fan blade receivable within each of said blade receipt sections to receive an airflow through a core airflow passage defined within each of said fan blades.
4. The fan hub assembly as recited in claim 3, wherein each of said multitude of fan blades include a fan blade mount section receivable within each of said blade receipt sections.
5. The fan hub assembly as recited in claim 4, wherein each of said fan blade mount sections are semi-cylindrical to radially lock said fan blade sections within said fan hub.
6. A fan-turbine rotor assembly for a tip turbine engine comprising: a fan hub defining a hub axis of rotation, said fan hub defining a multitude of elongated openings through an outer periphery of said fan hub; an inducer defined by each of said elongated openings to turn an airflow from a generally axial direction to a generally radial direction; a blade receipt section defined by each of said elongated openings; a multitude of fan blade sections which each define a core airflow passage therethrough; and a fan blade mount section extending from each of said multitude of fan blade sections, each of said fan blade mount sections receivable within one of said multitude of blade receipt sections for retention therein to communicate said airflow from said inducer to each of said multitude of core airflow passages.
7. The fan-turbine rotor assembly as recited in claim 6, further comprising a diffuser about said multitude of fan blade sections, said diffuser in communication with each of said multitude of core airflow passages to turn said airflow from said radial direction to a second axial airflow direction.
8. The fan-turbine rotor assembly as recited in claim 7, further comprising a turbine which extends from said diffuser.
9. The fan-turbine rotor assembly as recited in claim 8, wherein said turbine includes a first row of shrouded turbine blades and a second row of shrouded turbine blades.
PCT/US2004/040174 2004-12-01 2004-12-01 Fan-turbine rotor assembly with integral inducer section for a tip turbine engine WO2006060005A1 (en)

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US11/719,854 US20090169385A1 (en) 2004-12-01 2004-12-01 Fan-turbine rotor assembly with integral inducer section for a tip turbine engine

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US7878762B2 (en) 2004-12-01 2011-02-01 United Technologies Corporation Tip turbine engine comprising turbine clusters and radial attachment lock arrangement therefor
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