FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

Abstract

A gas turbine engine includes a fan section that has fan blades that drive air along a bypass flow path in a bypass duct. A fan shaft is drivingly connected to the fan. A turbine section includes a turbine drive shaft. A gear system is connected to the turbine draft shaft through an input and connected to the fan shaft through an output. At least one of the input and the output include a flexible coupling. A gear system flex mount arrangement accommodates misalignment of the fan shaft and the turbine drive shaft during operation. The gear system flex mount arrangement includes a gear mesh that defines a gear mesh lateral stiffness and a ring gear that defines a ring gear lateral stiffness that is less than 12% of the gear mesh lateral stiffness.

Claims

1. A gas turbine engine, comprising: a fan section having fan blades that drive air along a bypass flow path in a bypass duct; a fan shaft drivingly connected to said fan; a turbine section including a turbine drive shaft; a gear system connected to the turbine draft shaft with a first gear system support on an input side of the gear system and connected to said fan shaft with a second gear system support on an output side of the gear system, wherein at least one of said first gear system support and said second gear system support include a flexible connection; and a gear system flex mount arrangement, wherein said gear system flex mount arrangement accommodates misalignment of said fan shaft and said turbine drive shaft during operation and includes a gear mesh defining a gear mesh lateral stiffness and a ring gear defining a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness.

2. The gas turbine engine of claim 1, wherein the second gear system support on an output side of the gear system defining a second gear system support lateral stiffness and a second gear system support transverse stiffness.

3. The gas turbine engine of claim 2, wherein said input defines an input lateral stiffness and an input transverse stiffness and at least one of said input lateral stiffness and said input transverse stiffness are less than 11% of a respective one of said second gear system support lateral stiffness and said second gear system support transverse stiffness.

4. The gas turbine engine of claim 3, wherein said input lateral stiffness is less than 5% of said gear mesh lateral stiffness.

5. The gas turbine engine of claim 4, wherein said ring gear define a ring gear transverse stiffness and said gear mesh defines a gear mesh transverse stiffness and said ring gear transverse stiffness is less than 12% of said gear mesh transverse stiffness.

6. The gas turbine engine of claim 3, wherein said input lateral stiffness and said input transverse stiffness are less than 11% of a respective one of said second gear system support lateral stiffness and said second gear system support transverse stiffness.

7. The gas turbine engine of claim 3, wherein said bypass duct is at least partially defined by an outer housing.

8. The gas turbine engine of claim 7, wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

9. The gas turbine engine of claim 8, wherein said gear system flex mount arrangement further includes a flexible support which supports said gear system relative to a static structure and defines a flexible support transverse stiffness, and said flexible support transverse stiffness is less than 11% of said second gear system support transverse stiffness.

10. The gas turbine engine of claim 9, further comprising a two stage high pressure turbine.

11. The gas turbine engine of claim 10, wherein said flexible support defines a flexible support lateral stiffness that is less than 11% of said second gear system support lateral stiffness and said second gear system support is a K-frame.

12. The gas turbine engine of claim 2, wherein said gear system flex mount arrangement includes a flexible support which supports said gear system relative to a static structure and defines a flexible support lateral stiffness that is less than 11% of said second gear system support lateral stiffness.

13. The gas turbine engine of claim 12, wherein said input defines an input transverse stiffness, said gear mesh defines a gear mesh transverse stiffness, and said input transverse stiffness is less than 11% of said second gear system support transverse stiffness and less than 5% of said gear mesh transverse stiffness.

14. The gas turbine engine of claim 1, wherein said bypass duct is at least partially defined by an outer housing and said gear system is an epicyclic gear system.

15. The gas turbine engine of claim 14, wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

16. The gas turbine engine of claim 1, further comprising a flexible support which supports said gear system relative to a static structure and defines a flexible support lateral stiffness that is less than 8% of said gear mesh lateral stiffness.

17. The gas turbine engine of claim 16, wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

18. The gas turbine engine of claim 17, wherein said gear mesh defines a gear mesh transverse stiffness and said flexible support defines a flexible support transverse stiffness that is less than 8% of said gear mesh transverse stiffness.

19. The gas turbine engine of claim 18, wherein said ring gear defines a ring gear transverse stiffness that is less than 12% of said gear mesh transverse stiffness.

20. The gas turbine engine of claim 19, wherein said gear system flex mount arrangement further includes a frame supporting said fan shaft and defining a frame lateral stiffness, wherein said flexible support lateral stiffness is less than 11% of said frame lateral stiffness.

21. The gas turbine engine of claim 20, further comprising a two stage high pressure turbine.

22. The gas turbine engine of claim 2, further comprising a flexible support which supports said gear system relative to a static structure and defines a flexible support lateral stiffness, wherein said flexible support lateral stiffness is less than 11% of said second gear system support lateral stiffness.

23. The gas turbine engine of claim 22, wherein said flexible support defines a flexible support transverse stiffness and said gear mesh defines a gear mesh transverse stiffness and said flexible support transverse stiffness is less than 8% of said gear mesh transverse stiffness.

24. The gas turbine engine of claim 23, wherein said flexible support transverse stiffness is less than 11% of gear second gear system support transverse stiffness.

25. The gas turbine engine of claim 24, wherein said ring gear defines a ring gear transverse stiffness that is less than 20% of said gear mesh transverse stiffness.

26. The gas turbine engine of claim 25, wherein said input defines an input transverse stiffness that is less than 20% of said second gear system support transverse stiffness.

27. The gas turbine engine of claim 26, wherein said bypass duct is at least partially defined by an outer housing.

28. The gas turbine engine of claim 27, wherein said input defines an input lateral stiffness that is less than 11% of said second gear system support lateral stiffness and said second gear system support is a K-frame.

29. The gas turbine engine of claim 28, further comprising a two stage high pressure turbine.

30. The gas turbine engine of claim 29, wherein said gear system has a gear reduction ratio of greater than 2.3, and further comprising a bypass ratio greater than ten (10), a fan pressure ratio of less than 1.45 measured across said fan blades alone, and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows:

[0038] FIG. 1 is a schematic cross-section of a gas turbine engine;

[0039] FIG. 2 is an enlarged cross-section of a section of the gas turbine engine which illustrates a fan drive gear system (FDGS);

[0040] FIG. 3 is a schematic view of a flex mount arrangement for one non-limiting embodiment of the FDGS;

[0041] FIG. 4 is a schematic view of a flex mount arrangement for another non-limiting embodiment of the FDGS;

[0042] FIG. 5 is a schematic view of a flex mount arrangement for another non-limiting embodiment of a star system FDGS;

[0043] FIG. 6 is a schematic view of a flex mount arrangement for another non-limiting embodiment of a planetary system FDGS.

[0044] FIG. 7 is a schematic view of a flex mount arrangement for another non-limiting embodiment of a star system FDGS;

[0045] FIG. 8 is a schematic view of a flex mount arrangement for another non-limiting embodiment of a planetary system FDGS;

[0046] FIG. 9 shows another embodiment; and

[0047] FIG. 10 shows yet another embodiment.

DETAILED DESCRIPTION

[0048] FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

[0049] The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.

[0050] The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

[0051] The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.

[0052] The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

[0053] A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(T.sub.ram °R)/(518.7 °R)].sup.0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.

[0054] With reference to FIG. 2, the geared architecture 48 generally includes a fan drive gear system (FDGS) 60 driven by the low speed spool 30 (illustrated schematically) through an input 62. The input 62, which may be in the form of a coupling, both transfers torque from the low speed spool 30 to the geared architecture 48 and facilitates the segregation of vibrations and other transients therebetween. In the disclosed non-limiting embodiment, the FDGS 60 may include an epicyclic gear system which may be, for example, a star system or a planet system.

[0055] The input coupling 62 may include an interface spline 64 joined, by a gear spline 66, to a sun gear 68 of the FDGS 60. The sun gear 68 is in meshed engagement with multiple planet gears 70, of which the illustrated planet gear 70 is representative. Each planet gear 70 is rotatably mounted in a planet carrier 72 by a respective planet journal bearing 75. Rotary motion of the sun gear 68 urges each planet gear 70 to rotate about a respective longitudinal axis P.

[0056] Each planet gear 70 is also in meshed engagement with rotating ring gear 74 that is mechanically connected to a fan shaft 76. Since the planet gears 70 mesh with both the rotating ring gear 74 as well as the rotating sun gear 68, the planet gears 70 rotate about their own axes to drive the ring gear 74 to rotate about engine axis A. The rotation of the ring gear 74 is conveyed to the fan 42 (FIG. 1) through the fan shaft 76 to thereby drive the fan 42 at a lower speed than the low speed spool 30. It should be understood that the described geared architecture 48 is but a single non-limiting embodiment and that various other geared architectures will alternatively benefit herefrom.

[0057] With reference to FIG. 3, a flexible support 78 supports the planet carrier 72 to at least partially support the FDGS 60A with respect to the static structure 36 such as a front center body which facilitates the segregation of vibrations and other transients therebetween. It should be understood that various gas turbine engine case structures may alternatively or additionally provide the static structure and flexible support 78. It should be understood that lateral as defined herein is generally transverse to the axis of rotation A and the term “transverse” refers to a pivotal bending movement with respect to the axis of rotation A which typically absorbs deflection applied to the FDGS 60. The static structure 36 may further include a number 1 and 1.5 bearing support static structure 82 which is commonly referred to as a “K-frame” which supports the number 1 and number 1.5 bearing systems 38A, 38B. Notably, the K-frame bearing support defines a lateral stiffness (represented as Kframe in FIG. 3) and a transverse stiffness (represented as Kframe.sup.BEND in FIG. 3) as the referenced factors in this non-limiting embodiment.

[0058] In this disclosed non-limiting embodiment, the lateral stiffness (KFS; KIC) of both the flexible support 78 and the input coupling 62 are each less than about 11% of the lateral stiffness (Kframe). That is, the lateral stiffness of the entire FDGS 60 is controlled by this lateral stiffness relationship. Alternatively, or in addition to this relationship, the transverse stiffness of both the flexible support 78 and the input coupling 62 are each less than about 11% of the transverse stiffness (Kframe.sup.BEND). That is, the transverse stiffness of the entire FDGS 60 is controlled by this transverse stiffness relationship.

[0059] With reference to FIG. 4, another non-limiting embodiment of a FDGS 60B includes a flexible support 78′ that supports a rotationally fixed ring gear 74′. The fan shaft 76′ is driven by the planet carrier 72′ in the schematically illustrated planet system which otherwise generally follows the star system architecture of FIG. 3.

[0060] With reference to FIG. 5, the lateral stiffness relationship within a FDGS 60C itself (for a star system architecture) is schematically represented. The lateral stiffness (KIC) of an input coupling 62, a lateral stiffness (KFS) of a flexible support 78, a lateral stiffness (KRG) of a ring gear 74 and a lateral stiffness (KJB) of a planet journal bearing 75 are controlled with respect to a lateral stiffness (KGM) of a gear mesh within the FDGS 60.

[0061] In the disclosed non-limiting embodiment, the stiffness (KGM) may be defined by the gear mesh between the sun gear 68 and the multiple planet gears 70. The lateral stiffness (KGM) within the FDGS 60 is the referenced factor and the static structure 82′ rigidly supports the fan shaft 76. That is, the fan shaft 76 is supported upon bearing systems 38A, 38B which are essentially rigidly supported by the static structure 82′. The lateral stiffness (KJB) may be mechanically defined by, for example, the stiffness within the planet journal bearing 75 and the lateral stiffness (KRG) of the ring gear 74 may be mechanically defined by, for example, the geometry of the ring gear wings 74L, 74R (FIG. 2).

[0062] In the disclosed non-limiting embodiment, the lateral stiffness (KRG) of the ring gear 74 is less than about 12% of the lateral stiffness (KGM) of the gear mesh; the lateral stiffness (KFS) of the flexible support 78 is less than about 8% of the lateral stiffness (KGM) of the gear mesh; the lateral stiffness (KJB) of the planet journal bearing 75 is less than or equal to the lateral stiffness (KGM) of the gear mesh; and the lateral stiffness (KIC) of an input coupling 62 is less than about 5% of the lateral stiffness (KGM) of the gear mesh.

[0063] With reference to FIG. 6, another non-limiting embodiment of a lateral stiffness relationship within a FDGS 60D itself are schematically illustrated for a planetary gear system architecture, which otherwise generally follows the star system architecture of FIG. 5.

[0064] It should be understood that combinations of the above lateral stiffness relationships may be utilized as well. The lateral stiffness of each of structural components may be readily measured as compared to film stiffness and spline stiffness which may be relatively difficult to determine.

[0065] By flex mounting to accommodate misalignment of the shafts under design loads, the FDGS design loads have been reduced by more than 17% which reduces overall engine weight. The flex mount facilitates alignment to increase system life and reliability. The lateral flexibility in the flexible support and input coupling allows the FDGS to essentially ‘float’ with the fan shaft during maneuvers. This allows: (a) the torque transmissions in the fan shaft, the input coupling and the flexible support to remain constant during maneuvers; (b) maneuver induced lateral loads in the fan shaft (which may otherwise potentially misalign gears and damage teeth) to be mainly reacted to through the number 1 and 1.5 bearing support K-frame; and (c) both the flexible support and the input coupling to transmit small amounts of lateral loads into the FDGS. The splines, gear tooth stiffness, journal bearings, and ring gear ligaments are specifically designed to minimize gear tooth stress variations during maneuvers. The other connections to the FDGS are flexible mounts (turbine coupling, case flex mount). These mount spring rates have been determined from analysis and proven in rig and flight testing to isolate the gears from engine maneuver loads. In addition, the planet journal bearing spring rate may also be controlled to support system flexibility.

[0066] FIG. 7 is similar to FIG. 5 but shows the transverse stiffness relationships within the FDGS 60C (for a star system architecture). The transverse stiffness (KIC.sup.BEND) of the input coupling 62, a transverse stiffness (KFS.sup.BEND) of the flexible support 78, a transverse stiffness (KRG.sup.BEND) of the ring gear 74 and a transverse stiffness (KJB.sup.BEND) of the planet journal bearing 75 are controlled with respect to a transverse stiffness (KGM.sup.BEND) of the gear mesh within the FDGS 60.

[0067] In the disclosed non-limiting embodiment, the stiffness (KGM.sup.BEND) may be defined by the gear mesh between the sun gear 68 and the multiple planet gears 70. The transverse stiffness (KGM.sup.BEND) within the FDGS 60 is the referenced factor and the static structure 82′ rigidly supports the fan shaft 76. That is, the fan shaft 76 is supported upon bearing systems 38A, 38B which are essentially rigidly supported by the static structure 82′. The transverse stiffness (KJB.sup.BEND) may be mechanically defined by, for example, the stiffness within the planet journal bearing 75 and the transverse stiffness (KRG.sup.BEND) of the ring gear 74 may be mechanically defined by, for example, the geometry of the ring gear wings 74L, 74R (FIG. 2).

[0068] In the disclosed non-limiting embodiment, the transverse stiffness (KRG.sup.BEND) of the ring gear 74 is less than about 12% of the transverse stiffness (KGM.sup.BEND) of the gear mesh; the transverse stiffness (KFS.sup.BEND) of the flexible support 78 is less than about 8% of the transverse stiffness (KGM.sup.BEND) of the gear mesh; the transverse stiffness (KJB.sup.BEND) of the planet journal bearing 75 is less than or equal to the transverse stiffness (KGM.sup.BEND) of the gear mesh; and the transverse stiffness (KIC.sup.BEND) of an input coupling 62 is less than about 5% of the transverse stiffness (KGM.sup.BEND) of the gear mesh.

[0069] FIG. 8 is similar to FIG. 6 but shows the transverse stiffness relationship within the FDGS 60D for the planetary gear system architecture.

[0070] FIG. 9 shows an embodiment 200, wherein there is a fan drive turbine 208 driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction 204 may be positioned between the fan drive turbine 208 and the fan rotor 202. This gear reduction 204 may be structured, mounted and operate like the gear reduction disclosed above. A compressor rotor 210 is driven by an intermediate pressure turbine 212, and a second stage compressor rotor 214 is driven by a turbine rotor 216. A combustion section 218 is positioned intermediate the compressor rotor 214 and the turbine section 216.

[0071] FIG. 10 shows yet another embodiment 300 wherein a fan rotor 302 and a first stage compressor 304 rotate at a common speed. The gear reduction 306 (which may be structured, mounted and operate as disclosed above) is intermediate the compressor rotor 304 and a shaft 308, which is driven by a low pressure turbine section.

[0072] 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.

[0073] It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

[0074] Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.

[0075] The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason, the appended claims should be studied to determine true scope and content.