RING GEAR MOUNTING ARRANGEMENT WITH OIL SCAVENGE SCHEME

Abstract

A gas turbine engine includes a bypass ratio greater than about ten (10). A fan is supported on a fan shaft and has a plurality of fan blades. A gear system is connected to the fan shaft. A plurality of planetary gears and a first set of opposed angled ring gear teeth are separated from a second set of opposed angled ring gear teeth. A lubricant flow path is located axially between the first set of opposed angled ring gear teeth and the second set of opposed angled ring gear teeth. A gear system support is relative to a fixed housing facilitating segregation of vibrations. An annular channel is axially aligned with the lubricant flow path. A low pressure turbine has an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1.

Claims

1. A gas turbine engine comprising: a bypass ratio greater than about ten (10); a fan supported on a fan shaft and having a plurality of fan blades; a gear system connected to the fan shaft, a plurality of planetary gears, a first set of opposed angled ring gear teeth separated from a second set of opposed angled ring gear teeth, wherein a lubricant flow path is located axially between the first set of opposed angled ring gear teeth and the second set of opposed angled ring gear teeth; a gear system support relative to a fixed housing facilitating segregation of vibrations; an annular channel axially aligned with the lubricant flow path; and a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1.

2. The gas turbine engine of claim 1, further comprising a torque frame at least partially supporting the gear system, wherein the torque frame facilitates segregation of vibrations.

3. The gas turbine engine of claim 2, wherein the gear system includes a sun gear having a sun gear splined connection.

4. The gas turbine engine of claim 3, wherein said gear system includes a gear reduction ratio of greater than 2.5.

5. The gas turbine engine of claim 4, wherein the annular channel is at least partially defined by a gutter.

6. The gas turbine engine of claim 5, wherein the fan shaft is supported by at least one tapered roller bearing.

7. The gas turbine engine of claim 5, further comprising a fan tip speed less than 1150 ft/second.

8. The gas turbine engine of claim 7, further comprising a low fan pressure ratio of less than 1.45 across the fan blades alone.

9. The gas turbine engine of claim 5, wherein said gutter is for collecting oil expelled from the gear system.

10. The gas turbine engine of claim 9, further comprising a fan tip speed less than 1150 ft/second.

11. The gas turbine engine of claim 10, further comprising a low fan pressure ratio of less than 1.45 across the fan blades alone.

12. The gas turbine engine of claim 11, wherein the gutter is formed from a soft material.

13. The gas turbine engine of claim 12, wherein the soft material is aluminum.

14. The gas turbine engine of claim 8, wherein the first set of opposed angled ring gear teeth and the second set of opposed angled ring gear teeth each include a radially extending flange that extends radially outward away from a corresponding set of opposed angled ring gear teeth, and the fan shaft includes a radially extending flange connected to the radially extending flange on the first set of opposed angled ring gear teeth and the radially extending flange on the second set of opposed angled ring gear teeth.

15. The gas turbine engine of claim 14, further comprising a seal attached to at least one of the radially extending flanges on the first and second set of opposed angled ring gear teeth or the radially extending flange on the fan shaft.

16. The gas turbine engine of claim 15, wherein the seal includes an oil return passage.

17. The gas turbine engine of claim 16, wherein the oil return passage includes a slot.

18. The gas turbine engine of claim 17, wherein the slot is located in the seal.

19. The gas turbine engine of claim 17, wherein the slot is located in the radially extending flange on the fan shaft.

20. The gas turbine engine of claim 19, further comprising an input to the gear system, wherein the input facilitates segregation of vibrations.

21. A gas turbine engine comprising: a fan supported on a fan shaft by at least one roller bearing and having a plurality of fan blades; a bypass ratio greater than about ten (10); an epicyclic gear system connected to the fan shaft, the epicyclic gear system having a gear reduction ratio of greater than 2.5, a plurality of intermediate gears, and a ring gear with an aperture that is, wherein the ring gear includes a first portion separated from a second portion, the first portion and the second portion each include a radially extending flange that extends radially outward away from a corresponding set of opposed angled teeth; a torque frame as least partially supporting the gear system with respect to a housing to facilitate segregation of vibrations; and an annular channel axially aligned with the aperture and located radially outward from the aperture.

22. The gas turbine engine of claim 21, wherein the gear system includes a sun gear having a sun gear connection.

23. The gas turbine engine of claim 22, wherein the annular channel includes a gutter for collecting oil expelled from the gear system.

24. The gas turbine engine of claim 23, wherein the gutter is formed from aluminum.

25. The gas turbine of claim 24, further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, and a low fan pressure ratio of less than 1.45 and the low fan pressure ratio is measured across the fan blades alone.

26. The gas turbine of claim 24, further comprising a low pressure turbine with an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5:1, a low fan pressure ratio of less than 1.45 and the low fan pressure ratio is measured across the fan blades alone, and a gear system support relative to a fixed housing facilitating segregation of vibrations.

27. The gas turbine engine of claim 26, wherein the fan shaft includes a radially extending flange connected to the radially extending flange on the first portion of the ring gear and the radially extending flange on the second portion of the ring gear.

28. The gas turbine engine of claim 27, further comprising a seal attached to at least one of the radially extending flanges on the ring gear or the radially extending flange on the fan shaft, wherein the seal includes an oil return passage.

29. The gas turbine engine of claim 28, further comprising a fan tip speed less than 1150 ft/second.

30. The gas turbine engine of claim 29, further comprising an input to the gear system and a torque frame as least partially supporting the gear system with respect to a housing, wherein the input and the torque frame each facilitates segregation of vibrations.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0034] FIG. 1 is a partial cross-sectional view of a front portion of a gas turbine engine illustrating a turbo fan, epicyclic gear train and a compressor section.

[0035] FIG. 2 is an enlarged cross-sectional view of the epicyclic gear train shown in FIG. 1.

[0036] FIG. 3 is an enlarged cross-sectional view of an example ring gear similar to the arrangement shown in FIG. 2.

[0037] FIG. 4 is a view of the ring gear shown in FIG. 3 viewed in a direction that faces the teeth of the ring gear in FIG. 3.

[0038] FIG. 5 shows another embodiment.

[0039] FIG. 6 shows yet another embodiment.

DETAILED DESCRIPTION

[0040] A portion of a gas turbine engine 10 is shown schematically in FIG. 1. The turbine engine 10 includes a fixed housing 12 that is constructed from numerous pieces secured to one another. A compressor section 14 having compressor hubs 16 with blades are driven by a turbine shaft 25 about an axis A. A turbo fan 18 is supported on a turbo fan shaft 20 that is driven by a compressor shaft 24, which supports the compressor hubs 16, through an epicyclic gear train 22. The engine 10 is a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, the engine 10 bypass ratio is greater than about six (6) to ten (10), the gear train 22 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 18 has a pressure ratio that is greater than about 5. The engine 10 in one non-limiting embodiment is a high-bypass geared architecture aircraft engine. In one disclosed embodiment, the engine 10 bypass ratio is greater than ten (10:1), the turbofan diameter is significantly larger than that of the low pressure compressor 16, and the low pressure turbine 27a (shown schematically) has a pressure ratio that is greater than 5:1. As understood, the low pressure turbine 27a is downstream of at least one upstream, or high pressure turbine. The gear train 22 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5:1. 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.

[0041] A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan 18 of the engine 10 is designed for a particular flight conditiontypically cruise at about 0.8M and about 35,000 feet. The flight condition of 0.8 M and 35,000 ft, with the engine at its best fuel consumptionalso known as bucket cruise TSFCis 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 the FEGV system 36. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than 1.45. Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tambient deg R)/518.7){circumflex over ()}0.5]. The Low corrected fan tip speed as disclosed herein according to one non-limiting embodiment is less than 1150 ft/second.

[0042] The gear train 22 generally includes a fan drive gear system (FDGS) 100 driven by the compressor shaft 24 through an input coupling 102. The input coupling 102 both transfers torque from the compressor shaft 24 to the gear train 22 and facilitates the segregation of vibrations and other transients therebetween.

[0043] The input coupling 102 may include an interface spline 102 joined, by a gear spline 106, to the sun gear 30. The sun gear 30 is in meshed engagement with multiple star gears 32. Each star gear 32 is also in meshed engagement with rotating ring gear 38 that is mechanically connected to the fan shaft 20. Since the star gears 32 mesh with both the rotating ring gear 38 as well as the rotating sun gear 30, the star gears 32 rotate about their own axes to drive the ring gear 38. The rotation of the ring gear 38 is conveyed to the fan 20 through the fan shaft 20 to thereby drive the fan 18 at a lower speed than the turbine shaft 25.

[0044] In the example arrangement shown, the epicyclic gear train 22 is a star gear train. Referring to FIG. 2, the epicyclic gear train 22 includes the sun gear 30 that is connected to the compressor shaft 24, which provides rotational input, by the splined connection. A carrier 26 is fixed to the housing 12 by a torque frame 28 using fingers (not shown) known in the art. The carrier 26 supports star gears 32 using journal bearings 34 that are coupled to the sun gear 30 by meshed interfaces between the teeth of sun and star gears 30, 32. Multiple star gears 32 are arranged circumferentially about the sun gear 30. Retainers 36 retain the journal bearings 34 to the carrier 26. A ring gear 38 surrounds the carrier 26 and is coupled to the star gears 32 by meshed interfaces. The ring gear 38, which provides rotational output, is secured to the turbo fan shaft 20 by circumferentially arranged fastening elements, which are described in more detail below.

[0045] The torque frame 28 supports the carrier 26 with respect to the housing 12 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 be provided.

[0046] The fixed housing 12 may further include a number 1 and 1.5 bearing support frame 108 which is commonly referred to as a K-frame which supports the number 1 and number 1.5 bearing systems 110A, 110B to support the fan shaft 20 (FIG. 1). The number 1 and number 1.5 bearing systems 110A, 110B may include tapered roller bearings which provide a line contact.

[0047] Referring to FIGS. 3 and 4, the ring gear 38 is a two-piece construction having first and second portions 40, 42. The first and second portions 40, 42 abut one another at a radial interface 45. A trough 41 separates oppositely angled teeth 43 (best shown in FIG. 4) on each of the first and second portions 40, 42. The arrangement of teeth 43 forces the first and second portions 40, 42 toward one another at the radial interface 45. The back side of the first and second portions 40, 42 includes a generally S-shaped outer circumferential surface 47 that, coupled with a change in thickness, provides structural rigidity and resistance to overturning moments. The first and second portions 40, 42 have a first thickness T1 that is less than a second thickness T2 arranged axially inwardly from the first thickness T1. The first and second portions 40, 42 include facing recesses 44 that form an internal annular cavity 46.

[0048] The first and second portions 40, 42 include flanges 51 that extend radially outward away from the teeth 43. The turbo fan shaft 20 includes a radially outwardly extending flange 70 that is secured to the flanges 51 by circumferentially arranged bolts 52 and nuts 54, which axially constrain and affix the turbo fan shaft 20 and ring gear 38 relative to one another. Thus, the spline ring is eliminated, which also reduces heat generated from windage and churning that resulted from the sharp edges and surface area of the splines. The turbo fan shaft 20 and ring gear 38 can be rotationally balanced with one another since radial movement resulting from the use of splines is eliminated. An oil baffle 68 is also secured to the flanges 51, 70 and balanced with the assembly.

[0049] Seals 56 having knife edges 58 are secured to the flanges 51, 70. The first and second portions 40, 42 have grooves 48 at the radial interface 45 that form a hole 50, which expels oil through the ring gear 38 to a gutter 60 that is secured to the carrier 26 with fasteners 61 (FIG. 2). The direct radial flow path provided by the grooves 48 reduces windage and churning by avoiding the axial flow path change that existed with splines. That is, the oil had to flow radially and then axially to exit through the spline interface. The gutter 60 is constructed from a soft material such as aluminum so that the knife edges 58, which are constructed from steel, can cut into the aluminum if they interfere. Referring to FIG. 3, the seals 56 also include oil return passages 62 provided by first and second slots 64 in the seals 56, which permit oil on either side of the ring gear 38 to drain into the gutter 60. In the example shown in FIG. 2, the first and second slots 64, 66 are instead provided in the flange 70 and oil baffle 68, respectively.

[0050] FIG. 5 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 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.

[0051] FIG. 6 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 as disclosed above) is intermediate the compressor rotor 304 and a shaft 308 which is driven by a low pressure turbine section.

[0052] Although a preferred embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.