Classifications

• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
  • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
  • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
  • F02C7/06—Arrangements of bearings; Lubricating
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
  • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
  • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
  • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
  • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
  • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
  • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
  • F01D25/162—Bearing supports
• • 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
  • F02K3/077—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 the plant being of the multiple flow type, i.e. having three or more flows
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
  • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
  • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
  • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
  • F01D25/162—Bearing supports
  • F01D25/164—Flexible supports; Vibration damping means associated with the bearing
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2240/00—Components
  • F05D2240/50—Bearings
  • F05D2240/54—Radial bearings
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2260/00—Function
  • F05D2260/40—Transmission of power
  • F05D2260/403—Transmission of power through the shape of the drive components
  • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
  • F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type

Definitions

• FIG. 4 is an enlarged cross-sectional view of a second bearing configuration, taken at detail view 110 of FIG. 3, according to an aspect of the present disclosure.
• first”, “second” or “third” may be used interchangeably to distinguish one component or one air stream from another and are not intended to signify location or importance of the individual components.
• forward and aft may refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operational attitude of the gas turbine engine or vehicle.
• forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.
• upstream and downstream refer to the relative direction with respect to fluid flow in a fluid pathway.
• upstream refers to the direction from which the fluid flows
• downstream refers to the direction to which the fluid flows.
• a ducted secondary fan may be included between the inlet case and the intermediate case of the core turbine engine.
• the ducted secondary fan is driven by the low pressure compressor shaft assembly.
• the core engine includes a splitter that splits the airflow from the ducted secondary fan to partially flow into a core duct, and to partially flow into a compressor bypass duct.
• the separate inlet case and the intermediate case provide support for bearings of the low pressure compressor shaft.
• the separate bearing supports i.e., separated from being part of a single frame structure
• the inclusion of the secondary (ducted) fan, along with the compressor bypass duct present loading challenges to the low pressure compressor shaft during various operating conditions.
• the present disclosure provides a bearing support configuration to address the challenges in handling of loads.
• FIG. 1 is a schematic cross-sectional view of a gas turbine gas turbine engine 10, according to an aspect of the present disclosure.
• the gas turbine engine 10 shown in FIG. 1 is a turbofan engine having a rotor assembly with a single stage of unducted rotor blades, which may be referred to as an “unducted fan.”
• the entire gas turbine engine 10 may be referred to as an “unducted turbofan engine.”
• the gas turbine engine 10 defines an axial direction A, a radial direction R, and a circumferential direction C, for reference purposes.
• the gas turbine engine 10 defines an axial centerline, also referred to as a centerline axis or a longitudinal centerline axis, 12 that extends along the axial direction A.
• the axial direction A extends parallel to the longitudinal centerline axis 12
• the radial direction R extends outward from and inward to the longitudinal centerline axis 12 orthogonal to the axial direction A
• the circumferential direction C extends three-hundred-sixty degrees around the longitudinal centerline axis 12.
• the gas turbine engine 10 extends between a forward end 14 and an aft end 16 along the axial direction A.
• the gas turbine engine 10 includes a core turbine engine 20 and a fan section 50, positioned upstream thereof.
• the core turbine engine 20 includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section.
• the core turbine engine 20 includes a core cowl 22 that defines an annular core inlet 24.
• the core cowl 22 encloses, at least in part, a booster or low pressure (“LP”) compressor 26 that includes at least one low pressure compressor rotor 33 for pressurizing air 25 that enters the core turbine engine 20 through the core inlet 24, and a high pressure (“HP”) compressor 28 that receives pressurized air from the LP compressor 26 and further compresses the air.
• Pressurized air 27 flows downstream to a combustor 30, where fuel is injected into the pressurized air 27 and ignited to raise the temperature and the energy level of the pressurized air 27, thereby forming combustion products 29.
• high/low speed and “high/low pressure” are used with respect to the high pressure/high speed system and the low pressure/low speed system interchangeably. Further, the terms “high” and “low” are used in this same context to distinguish the two systems, and are not meant to imply any absolute speed and/or pressure values.
• the combustion products 29 flow from the combustor 30 downstream to a high pressure turbine 32.
• the high pressure turbine 32 is connected with the high pressure compressor 28 through a high pressure compressor shaft 36 such that the high pressure turbine 32 is drivingly coupled with the high pressure compressor 28.
• the combustion products 29 then flow to a low pressure turbine 34.
• the low pressure turbine 34 is drivingly coupled with the low pressure compressor 26 and components of the fan section 50 via a low pressure compressor shaft 38 that is coaxial with the high pressure compressor shaft 36.
• the core turbine engine 20 defines a core duct 42 defining a core flowpath 42’ that extends between the core inlet 24 and the core turbine engine exhaust nozzle 40.
• the core duct 42 is an annular duct positioned generally inward of the core cowl 22 along the radial direction R.
• the core duct 42 (e.g., the working gas flowpath through the core turbine engine 20) may be referred to herein as a second stream 23 or a second flow path.
• the fan section 50 includes a primary fan assembly 52, which may be referred to as a primary (or a first) fan assembly.
• the primary fan assembly 52 may be an open rotor or an unducted primary fan assembly 52.
• the gas turbine engine 10 may be referred to as an open rotor engine.
• the primary fan assembly 52 is driven by a fan shaft 46 that is coupled to a speed reduction gearbox 55, such as in an indirect-drive or geared-drive configuration.
• the speed reduction gearbox 55 is coupled to the low pressure compressor shaft 38.
• the low pressure compressor shaft 38 may be supported by a plurality of bearings (e.g., four bearings) that are supported by various components parts of the gas turbine engine 10.
• the fan shaft 46 is supported by at least one fan shaft bearing (not shown) that is arranged forward of the speed reduction gearbox 55 and that is supported by a fan frame structure (not shown).
• the primary fan assembly 52 may include an array of fan blades 54 (only one fan blade 54 is shown in FIG. 1). The array of fan blades 54 can be arranged in equal spacing around the longitudinal centerline axis 12.
• the fan blades 54 are rotatable about the longitudinal centerline axis 12 by the primary fan assembly 52 being drivingly coupled to the fan shaft 46 that is coupled with the low pressure compressor shaft 38 via the speed reduction gearbox 55.
• Each fan blade 54 defines a central blade axis 56, and each fan blade 54 of the primary fan assembly 52 is rotatable about its respective central blade axis 56, e.g., in unison with one another.
• One or more fan blade actuators 58 are provided to facilitate the fan blade rotation and, therefore, may be used to change a pitch of the fan blades 54 about the central blade axes 56 of the respective fan blades 54.
• the fan section 50 further includes a fan guide vane array 60 that includes a plurality of circumferentially spaced fan guide vanes 62 (only one fan guide vane 62 is shown in
• FIG. 1 mounted to a fan cowl 70.
• the fan guide vanes 62 may not be rotatable about the longitudinal centerline axis 12.
• the fan guide vanes 62 may be unshrouded as shown in FIG. 1 or, alternatively, may be shrouded, e.g., by an annular shroud spaced outward from the tips of the fan guide vanes 62 along the radial direction R or attached to the fan guide vanes 62.
• Each fan guide vane 62 defines a central vane axis 64.
• Each fan guide vane 62 of the fan guide vane array 60 may be rotatable about its respective central vane axis 64, e.g., in unison with one another.
• One or more guide vane actuators 66 may be provided to facilitate the rotation of the fan guide vanes 62 and, therefore, may be used to change a pitch of the fan guide vane 62 about its respective central vane axis 64.
• each fan guide vane 62 may be fixed or unable to be pitched about its central vane axis 64.
• a secondary (ducted) fan assembly 84 is included aft of the primary fan assembly 52, such that the gas turbine engine 10 includes both the secondary (ducted) fan assembly 84 (e.g., a secondary or a second fan) and the primary (unducted) fan assembly 52 (e.g., a primary fan assembly or a first fan assembly), which both serve to generate thrust through the movement of air without passage through at least a portion of the core turbine engine 20 (e.g., without passage through the LP compressor 26, the HP compressor 28, and the combustor 30).
• the secondary (ducted) fan assembly 84 e.g., a secondary or a second fan
• the primary (unducted) fan assembly 52 e.g., a primary fan assembly or a first fan assembly
• the secondary (ducted) fan assembly 84 is rotatable about the same axis (e.g., the longitudinal centerline axis 12) as the fan blades 54 of the primary fan assembly 52.
• the secondary (ducted) fan assembly 84 is driven by the low pressure compressor shaft 38.
• the primary fan assembly 52 may be referred to as the primary fan or a first fan
• the secondary (ducted) fan assembly 84 may be referred to as a secondary fan assembly or a second fan assembly.
• the terms “primary” and “secondary” are terms of convenience, and do not imply any particular importance, such as a particular power, or the like.
• the secondary (ducted) fan assembly 84 includes a plurality of fan blades (not separately labeled in FIG. 1; see fan blades 85 in FIG. 2) arranged in a single stage, such that the secondary (ducted) fan assembly 84 may be referred to as a single stage fan.
• the fan blades 85 of the secondary (ducted) fan assembly 84 can be arranged in equal spacing around the
• the fan cowl 70 annularly encases at least a portion of the core cowl 22 and is generally positioned outward of at least a portion of the core cowl 22 along the radial direction R. Particularly, a downstream section of the fan cowl 70 extends over a forward portion of the core cowl 22 to define a compressor bypass duct 72.
• the compressor bypass duct 72 is an annular duct positioned generally outward of the core duct 42 along the radial direction R.
• the compressor bypass duct 72 may be understood as forming at least a portion of a third stream 73 or a third flowpath of the gas turbine engine 10.
• Air 31 from the secondary (ducted) fan assembly 84 may enter the compressor bypass duct 72 through a compressor bypass duct inlet 76 and may exit through a bypass exhaust nozzle 78 to produce propulsive thrust.
• the fan cowl 70 and the core cowl 22 are connected together and supported by a plurality of substantially radially-extending, circumferentially-spaced stationary struts 74 (only one stationary strut 74 is shown in FIG. 1).
• the stationary struts 74 may each be aerodynamically contoured to direct air flowing thereby.
• Other struts (not shown), in addition to the stationary struts 74, may also be used to connect and to support the fan cowl 70 and/or the core cowl 22.
• the compressor bypass duct 72 and the core duct 42 may at least partially coextend (generally axially) on opposite sides (e.g., opposite radial sides) of the core cowl 22.
• the compressor bypass duct 72 and the core duct 42 may each extend directly from a leading edge 44 of the core cowl 22 and may partially co-extend generally axially on opposite radial sides of the core cowl 22.
• FIG. 2 is an enlarged cross-sectional view of a forward portion of the gas turbine engine 10 depicted in FIG. 1.
• the gas turbine engine 10 includes a core flowpath that includes, in serial flow relationship, an inlet case 94, a mid-fan case 108, and an intermediate case 102.
• the inlet case 94 defines an inlet duct 80 having an engine inlet 82.
• the engine inlet 82 of the inlet duct 80 is defined generally at the forward end of the fan cowl 70 and is positioned between the primary fan assembly 52 and the fan guide vane array 60 along the axial direction A.
• the inlet duct 80 is an annular duct that is positioned inward of the fan cowl 70 along the radial direction R.
• the inlet case 94 includes an inlet case bearing support structure 96 that provides support for a first bearing 98, that supports the low pressure compressor shaft 38.
• the first bearing 98 may be a first bearing type, such as a radial load bearing, including a roller bearing, that is assembled with and supported by the inlet case bearing support structure 96 of the inlet case 94, and that engages with the low pressure compressor shaft 38.
• a first bearing type such as a radial load bearing, including a roller bearing, that is assembled with and supported by the inlet case bearing support structure 96 of the inlet case 94, and that engages with the low pressure compressor shaft 38.
• a radial load bearing including a roller bearing
• other types of bearings that provide for radial loading such as a ball bearing, or a double ball bearing, could also be implemented as the first bearing 98.
• the gas turbine engine 10 includes the intermediate case 102 that includes an intermediate duct 112, and the mid-fan case 108 that includes a mid-fan duct 114.
• the mid-fan case 108 connects between the inlet case 94 and the intermediate case 102 such that the mid-fan duct 114 connects the inlet duct 80 and the intermediate duct 112.
• the mid-fan duct 114 also connects the inlet duct 80 and the compressor bypass duct 72.
• the intermediate case 102 includes an intermediate case bearing support structure 104 that provides support for a second bearing 106.
• FIG. 3 is an enlarged cross-sectional view of the mid-fan case 108, according to an aspect of the present disclosure. In FIG.
• a plane Pl is provided as a reference plane that intersects a forwardmost point of a leading edge 87 of the fan blades 85 of the secondary (ducted) fan assembly 84 and a plane P2 is provided as a reference plane that intersects an aftmost point of a trailing edge 89 of the fan blades 85.
• the secondary (ducted) fan assembly 84 has a ducted fan center of gravity 100 and a ducted fan center of gravity reference plane 99 extends radially through, and orthogonal to, the longitudinal centerline axis 12 of the gas turbine engine 10.
• the first bearing 98 may be substantially radially aligned with the ducted fan center of gravity 100 of the secondary (ducted) fan assembly 84 (e.g., substantially radially aligned along the ducted fan center of gravity reference plane 99 through the center of gravity 100 of the secondary (ducted) fan assembly 84).
• “substantially” may mean with an axial offset distance (either axially forward or axially aft along the longitudinal centerline axis 12 of the gas turbine engine 10) from the ducted fan center of gravity 100, such as a one-tenth inch axial offset, such as a one-quarter inch axial offset, or such as a one-half inch axial offset.
• the radial load path of the first bearing 98 may be stiff in the radial direction (R) so as to move the secondary (ducted) fan assembly 84 modal frequency out of the speed range of the low pressure compressor shaft 38.
• the first bearing 98 location helps to control the ducted fan suspension mode and clearance closures under maneuvers and ducted fan unbalance. This provides the ability to reduce the vibration response of the secondary (ducted) fan assembly 84 under nominal and high unbalance conditions.
• the secondary (ducted) fan assembly 84 closures under aircraft maneuvers, specifically angular velocity, can be reduced.
• a reference plane P3 is provided, and the reference plane P3 extends orthogonal to the longitudinal centerline axis 12 through a center of gravity 101 of the low pressure compressor 26.
• a reference plane P4 shown in FIG. 3 extends orthogonal to the longitudinal centerline axis 12 through a center of gravity 103 of the intermediate case 102.
• a reference plane P3’ shown in FIG. 3 extends orthogonal to the longitudinal centerline axis 12 and extends through a center of gravity 105 of the rotating part of a last stage 43 of the low pressure compressor 26.
• the second bearing 106 may be arranged axially between the reference plane P3’ and the reference plane P4. With this location of the second bearing 106, the second bearing 106 helps to control the LP compressor shaft mode position and provides better stability of the low pressure compressor shaft 38.
• a radial stiffness of the inlet case bearing support structure 96 may be at least twice that of a radial stiffness of the intermediate case bearing support structure 104.
• the radial stiffness of the inlet case bearing support structure 96 may be between two and five times 10 9 m/N and the radial stiffness of the intermediate case bearing support structure 104 may be between five and ten times 10 9 m/N.
• the radial stiffness is defined intrinsically, that is to say by considering the supports as such, outside the gas turbine engine 10.
• the radial stiffnesses values are therefore absolute values, not relative values, and do not depend on the environment in which they are measured, such that it is possible to integrate these supports in any gas turbine engine 10.
• the radial stiffness of a given support may be determined by encasing an end of the support with the other end being free to radially move. This configuration actually reflects the configuration of the inlet case bearing support structure 96 and the intermediate case bearing support structure 104 in the gas turbine engine 10 (the encased end corresponding to the end connected to the engine case and the free end corresponding to the end connected to the corresponding bearing).
• the stiffness of the support is then defined as a ratio between a radial force applied to the free end and the radial displacement of the free end with respect to the encased end that results from the radial force.
• FIG. 4 is an enlarged cross-sectional view of a second bearing mounting configuration, taken at detail view 110 of FIG. 3, according to an aspect of the present disclosure.
• the second bearing 106 may be a second bearing type such as an axial load and a radial load bearing, such as a ball bearing.
• a second bearing 106 is mounted with the intermediate case bearing support structure 104, and may be mounted to include a squirrel cage 115 and squeeze film dampers.
• a squeeze film damper 116 may include at least one squeeze film damper inlet 118 and a squeeze film damper gap 120.
• a squeeze film damping fluid 122 (e.g., oil) is injected into the squeeze film damper inlet 118 and caused to flow to the squeeze film damper gap 120 to provide damping of loads imparted by the low pressure compressor shaft 38 to the intermediate case bearing support structure 104.
• the inclusion of the squirrel cage 115 and the squeeze film damper 116 into the second bearing mounting configuration provides for better rotodynamic stability of the low pressure compressor shaft 38. This is desired because, for the present engine configuration with the speed reduction gearbox 55 (FIG. 1) being arranged between the first (primary) fan assembly 52 (FIG. 1) and the LP compressor 26 (FIG.
• the LP compressor shaft mode is supercritical (in the operating range) and it is desired to ensure that the LP shaft mode has positive damping and is stable under various operating conditions. While the foregoing description described the first bearing 98 (FIG. 3) as constituting a roller bearing, and the second bearing 106 as constituting a ball bearing, these may be reversed in that the first bearing 98 may be a ball bearing and the second bearing 106 may be a roller bearing that is mounted with the squeeze film damper 116 and the squirrel cage 115.
• the modal frequency of the secondary (ducted) fan assembly 84 is governed by two bearings on either side of the secondary (ducted) fan assembly 84 mounted on two different frames (i.e., the inlet case 94 and the intermediate case 102). Therefore, arranging the first bearing 98 substantially aligned with the center of gravity 100 of the secondary (ducted) fan assembly 84 may be preferable, although the first bearing 98 may be somewhat shifted axially off of the ducted fan center of gravity 100 (e.g., shifted up to one inch forward or up to one inch aft of the ducted fan center of gravity 100).
• the two bearing system provides a straddle mounted system for the secondary (ducted) fan assembly 84 with one bearing (i.e., the first bearing 98) being near the ducted fan center of gravity 100 of the secondary (ducted) fan assembly 84 to help achieve the mode margin above redline of the low pressure operating range.
• FIG. 5 is an enlarged view of a low pressure turbine case 35, taken at detail view 128 of FIG. 1, according to an aspect of the present disclosure.
• FIG. 5 depicts an example of locations for a third bearing 130 and a fourth bearing 132 that support the low pressure compressor shaft 38 at the low pressure turbine case 35.
• the third bearing 130 and the fourth bearing 132 are spaced apart to reduce static clearance under maneuvers due to the rotational speed of the LP compressor shaft 38.
• a reference plane P6 extends orthogonal to the longitudinal centerline axis 12 and extends through a center of gravity 134 of a low pressure turbine rotor assembly 37.
• the low pressure turbine rotor assembly 37 may be connected to the low pressure compressor shaft 38 via a shaft connection 39.
• the low pressure turbine rotor assembly 37 may include a plurality of turbine rotors, including a first turbine rotor 136, a second turbine rotor 138, a third turbine rotor 140, and a fourth turbine rotor 142.
• a reference plane P5 extends orthogonal to the longitudinal centerline axis 12 and extends through a forwardmost point of the second turbine rotor 138, which is a turbine rotor forward of the reference plane P6.
• a reference plane P5’ extends orthogonal to the longitudinal centerline axis 12 and extends through an aft-most point of the fourth turbine rotor 142, which is aft of the reference plane P6.
• the third bearing 130 may be axially located close to the reference plane P6.
• This configuration of the location of the third bearing 130 reduces the length portion of the LP compressor shaft 38 that affects the LP shaft modes, such that the LP shaft length between the second bearing 106 (FIG. 3) and the third bearing 130 may be reduced.
• the rotational speed of the LP compressor shaft 38 may therefore be increased, such that the theoretical unstable speed range of the LP compressor shaft 38 is reduced.
• Locating the third bearing 130 axially aligned with the reference plane P6 i.e., the gravity center of the low pressure turbine rotor assembly 37
• the third bearing 130 affects the LP shaft modes and stability, and helps to control clearance closures under inertial loads.
• the fourth bearing 132 may be positioned axially as far as possible from the third bearing 130, taking into account the available space below the LP turbine case 35 and a turbine rear frame 144.
• a fourth bearing 132' may be supported by a fourth bearing support structure 150' and may be located under the turbine rear frame 144, such as axially aligned with a reference plane P7 that extends orthogonal to the longitudinal centerline axis 12 and extends through a center of gravity 146 of the turbine rear frame 144.
• the fourth bearing 132 may be axially located aft of the reference plane P7.
• the fourth bearing 132 helps to control static clearances under dynamic loads (e.g., angular velocity maneuvers).
• static clearances under dynamic loads e.g., angular velocity maneuvers.
• the third bearing 130 may be connected to a third bearing support structure 148 that is connected to the turbine rear frame 144, and the fourth bearings 132 may be connected to a fourth bearing support structure 150 that is also connected to the turbine rear frame 144.
• FIG. 6 depicts an alternate arrangement of the third bearing and the fourth bearing to that depicted in FIG. 5, according to another aspect of the present disclosure.
• a reference plane P8 extends orthogonal to the longitudinal centerline axis 12 and extends through a forward portion of an inter-turbine case 152
• a reference plane P9 extends orthogonal to the longitudinal centerline axis 12 and extends through a center of the first turbine rotor 136.
• the third bearing 130 may therefore be axially located between the reference plane P8 and the reference plane P9, and may be located axially upstream of the shaft connection 39.
• the third bearing 130 may be connected to the inter-turbine case 152 via a third bearing support structure 154, rather than being connected to the turbine rear frame 144 as shown in FIG.
• the FIG. 6 location of the third bearing 130 reduces the length portion of the LP compressor shaft 38 that affects the LP shaft modes, such that the LP shaft length between the second bearing 106 and the third bearing 130 may be reduced, thereby reducing the unstable speed range (between the LP shaft mode and the maximum speed range of the LP rotor).
• axial location of the third bearing 130 helps to control the LP shaft mode location and instability of the low pressure compressor shaft 38.
• the fourth bearing 132 is positioned as far as possible from the third bearing 130. However, in the FIG.
• the fourth bearing 132 is not be located under the turbine rear frame 144 so as to limit the loads applied to the third bearing 130 and to the fourth bearing 132. Rather, the fourth bearing 132 may be axially located between the reference plane P6 and a reference plane PIO that intersects a forwardmost portion of the turbine rear frame 144.
• the axial location of the fourth bearing 132 in the FIG. 6 aspect helps to control unbalanced loads and clearance closures (due to unbalances and static clearance under inertial loads).
• the primary fan assembly 52 in operation of the gas turbine engine 10, the primary fan assembly 52 generates a first stream 124 of air flowing in a downstream direction external to the core turbine engine 20.
• a portion of the first stream 124 enters the engine inlet 82 of the core turbine engine 20 as an inlet airflow 125, where the inlet airflow 125 passes through the inlet duct 80 to the mid-fan duct 114.
• the secondary (ducted) fan assembly 84 imparts work to the inlet airflow 125 to generate a secondary fan airflow 126.
• the pressurized air 27 flows into the combustor 30, whereby combustion products 29 then flow through the high pressure turbine 32 to drive the high pressure compressor shaft 36, and then through the low pressure turbine 34 to drive the low pressure compressor shaft 38, thereby sustaining operation of the primary fan assembly 52 and the secondary (ducted) fan assembly 84, as well as the low pressure compressor 26.
• the flow of air 25 through the core turbine engine 20 may be referred to as a second stream 25' of air.
• the gas turbine engine 10 located downstream of the secondary (ducted) fan assembly 84 and upstream of the compressor bypass duct inlet 76, the gas turbine engine 10 includes an array of outlet guide vanes 90.
• the array of outlet guide vanes 90 are not rotatable about the longitudinal centerline axis 12.
• the array of outlet guide vanes 90 are configured as fixed-pitch outlet guide vanes.
• the bypass exhaust nozzle 78 of the compressor bypass duct 72 is further configured as a variable geometry exhaust nozzle.
• the gas turbine engine 10 includes one or more exhaust nozzle actuators 92 for modulating the variable geometry exhaust nozzle.
• the variable geometry exhaust nozzle may be configured to vary a total cross-sectional area (e.g., an area of the nozzle in a plane perpendicular to the longitudinal centerline axis 12) to modulate an amount of thrust generated based on one or more engine operating conditions (e.g., temperature, pressure, mass flowrate, etc., of an airflow through the compressor bypass duct 72).
• a fixed geometry exhaust nozzle may also be adopted.
• the combination of the array of inlet guide vanes 86 located upstream of the secondary (ducted) fan assembly 84, the array of outlet guide vanes 90 located downstream of the secondary (ducted) fan assembly 84, and the bypass exhaust nozzle 78 may result in a more efficient generation of third stream thrust during one or more engine operating conditions. Further, by introducing a variability in the geometry of the inlet guide vanes 86 and the bypass exhaust nozzle 78, the gas turbine engine 10 may be capable of generating more efficient third stream thrust across a relatively wide array of engine operating conditions, including takeoff and climb (where a maximum total engine thrust is generally needed) as well as cruise (where a lesser amount of total engine thrust is generally needed).
• air passing through the compressor bypass duct 72 may be relatively cooler (e.g., lower temperature) than one or more fluids utilized in the core turbine engine 20.
• one or more heat exchangers 48 may be positioned in thermal communication with the compressor bypass duct 72.
• one or more heat exchangers 48 may be disposed within the compressor bypass duct 72 and utilized to cool one or more fluids from the core turbine engine 20 with the air 31 passing through the compressor bypass duct 72, as a resource for removing heat from a fluid, e.g., compressor bleed air, oil, or fuel.
• the heat exchanger 48 may be an annular heat exchanger extending substantially three-hundred-sixty degrees in the compressor bypass duct 72 (e.g., at least three hundred degrees, such as at least three-hundred-thirty degrees). In such a manner, the heat exchanger 48 may effectively utilize the air 31 passing through the compressor bypass duct 72 to cool one or more systems of the gas turbine engine 10 (e.g., lubrication oil systems, compressor bleed air, electrical components, etc.). The heat exchanger 48 uses the air 31 passing through the compressor bypass duct 72 as a heat sink and correspondingly increases the temperature of the air 31 downstream of the heat exchanger 48 and exiting the bypass exhaust nozzle 78.
• the gas turbine engine may be implemented in various environments.
• the engine may be implemented in an aircraft, but may also be implemented in non-aircraft applications, such as power generating stations, marine applications, or oil and gas production applications.
• non-aircraft applications such as power generating stations, marine applications, or oil and gas production applications.
• present disclosure is not limited to use in aircraft.
• a gas turbine engine having a centerline axis including a primary fan assembly, and a core turbine engine including (a) an inlet case defining an inlet duct and including an inlet case bearing support structure, (b) an intermediate case defining an intermediate duct and a compressor bypass duct, and including an intermediate case bearing support structure, (c) a low pressure compressor shaft drivingly coupled with the primary fan assembly, and including a secondary fan assembly and at least one low pressure compressor rotor, each axially arranged along the centerline axis between the inlet case and the intermediate case, (d) a first bearing supported by the inlet case bearing support structure and engaging with the low pressure compressor shaft, and (e) a second bearing supported by the intermediate case bearing support structure and engaging with the low pressure compressor shaft.
• the primary fan assembly is an unducted primary fan assembly and provides a first stream of air external to the core turbine engine, a portion of the first stream of air being provided to the inlet duct.
• the secondary fan assembly provides a second stream of air from the inlet duct to the at least one low pressure compressor rotor, and a third stream of air from the inlet duct to the compressor bypass duct.
• first bearing is one of a first bearing type or a second bearing type
• second bearing is the other of the first bearing type or the second bearing type.
• second bearing is mounted with squeeze film dampers and a squirrel cage.
• the core turbine engine further comprises (f) a low pressure turbine case, (g) an inter-turbine case, (h) a turbine rear frame, (i) a third bearing arranged axially downstream of the second bearing and engaging with the low pressure compressor shaft, and (j) a fourth bearing arranged axially downstream of the third bearing and engaging with the low pressure compressor shaft.
• the turbine rear frame includes a third bearing support structure supporting the third bearing and a fourth bearing support structure supporting the fourth bearing.
• the interturbine case includes a third bearing support structure supporting the third bearing
• the turbine rear frame includes a fourth bearing support structure supporting the fourth bearing.
• the third bearing is axially positioned along the low pressure compressor shaft forward of a center of gravity of a low pressure turbine rotor assembly of the low pressure turbine case.
• a gas turbine engine having a centerline axis including a first fan assembly providing a first stream of air external to a core turbine engine, and the core turbine engine including (a) a core flowpath defined by an inlet case and an intermediate case, (b) a shaft drivingly coupled with the first fan assembly, and including a second fan assembly and at least one compressor rotor, each axially arranged with respect to the centerline axis between the inlet case and the intermediate case, (c) a first bearing supported by the inlet case and engaging with the shaft, and (d) a second bearing supported by the intermediate case and engaging with the shaft.
• the inlet case includes an inlet case bearing support structure supporting the first bearing
• the intermediate case includes an intermediate case bearing support structure supporting the second bearing
• the first fan assembly is an unducted fan assembly and provides a portion of the first stream of air to the inlet case.
• first bearing is one of a first bearing type or a second bearing type
• second bearing is the other of the first bearing type or the second bearing type.
• second bearing is mounted with squeeze film dampers and a squirrel cage.

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)
• Turbine Rotor Nozzle Sealing (AREA)

Applications Claiming Priority (1)

Publications (1)

Family

ID=93014773

Family Applications (1)

Country Status (3)

Family Cites Families (5)

• 2023
  • 2023-04-11 EP EP23932299.3A patent/EP4695149A2/de active Pending
  • 2023-04-11 WO PCT/US2023/018162 patent/WO2024228694A2/en not_active Ceased
  • 2023-04-11 CN CN202380097018.5A patent/CN121152749A/zh active Pending

Also Published As

Similar Documents

Legal Events