EPICYCLIC GEAR TRAIN

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section including a propulsor supported on a propulsor shaft, a turbine section including a turbine shaft, a compressor section having a plurality of compressor hubs with blades driven by the turbine shaft about an engine axis, and an epicyclic gear train interconnecting the propulsor shaft and the turbine shaft. The epicyclic gear train includes a sun gear coupled to the turbine shaft, intermediary gears arranged circumferentially about and meshing with the sun gear, a carrier and a ring gear including first and second portions. The first and second portions have axially opposed faces abutting one another at a radial interface.

Claims

1. A gas turbine engine comprising: a propulsor section including a propulsor supported on a propulsor shaft; a turbine section including a turbine shaft; a compressor section having a plurality of compressor hubs with blades driven by the turbine shaft about an engine axis; and an epicyclic gear train interconnecting the propulsor shaft and the turbine shaft, the epicyclic gear train comprising: a sun gear coupled to the turbine shaft; intermediary gears arranged circumferentially about and meshing with the sun gear; a carrier and support means for supporting each of the intermediate gears relative to the carrier; and a ring gear including first and second portions, the first and second portions each having an inner periphery with teeth intermeshing with the intermediate gears, the first and second portions having axially opposed faces abutting one another at a radial interface, the first and second portions including respective flanges extending along the radial interface radially outward from the teeth, and the ring gear including discharge means for expelling lubricant from the radial interface outwardly of the ring gear.

2. The gas turbine engine as recited in claim 1, wherein the epicyclic gear train defines a gear reduction ratio of greater than or equal to 2.5.

3. The gas turbine engine as recited in claim 2, wherein the epicyclic gear train is a planetary gear system.

4. The gas turbine engine as recited in claim 2, wherein the ring gear includes attachment means for securing the first and second portions of the ring gear to the propulsor shaft.

5. The gas turbine engine as recited in claim 2, wherein the ring gear includes engagement means for forcing the first portion and the second portion toward one another at the radial interface.

6. The gas turbine engine as recited in claim 5, wherein the first and second portions of the gear train includes means for resisting overturning moments.

7. The gas turbine engine as recited in claim 2, wherein the ring gear includes accumulation means for capturing lubrication expelled toward the radial interface.

8. The gas turbine engine as recited in claim 2, wherein the support means includes journal bearings that support the respective intermediate gears, and each journal bearing includes lubrication means for conveying lubricant through the journal bearing to a peripheral journal surface of the journal bearing.

9. The gas turbine engine as recited in claim 2, wherein the gear train includes a torque frame having securement means for fixing the carrier to a fixed housing.

10. The gas turbine engine as recited in claim 2, wherein: the gear train includes collection means for receiving lubricant expelled by the discharge means through the radial interface; the discharge means inhibits an axial flow of lubricant passing along the radial interface prior to being expelled toward the collection means; and the gear train includes return means for communicating lubricant from an outer periphery of the respective first and second portions of the ring gear outwardly to the collection means.

11. The gas turbine engine as recited in claim 2, further comprising: an input shaft that interconnects the sun gear and the turbine shaft, the input shaft including an undulation that extends radially outward relative to the engine axis.

12. The gas turbine engine as recited in claim 2, wherein the gear train includes lubrication means for conveying lubricant from a periphery of the intermediate gears to the radial interface.

13. The gas turbine engine as recited in claim 2, wherein the propulsor is a turbo fan, and the propulsor shaft is a fan shaft supporting the fan.

14. The gas turbine engine as recited in claim 13, further comprising a bypass ratio of greater than 10.5 and a low corrected fan tip speed of less than 1150 ft/second, wherein the fan includes a pressure ratio of less than 1.45 across the fan blade alone at cruise at 0.8 M and 35,000 feet.

15. The gas turbine engine as recited in claim 14, wherein the ring gear includes engagement means for forcing the first portion and the second portion toward one another at the radial interface.

16. The gas turbine engine as recited in claim 15, wherein the ring gear includes attachment means for securing the first and second portions of the ring gear to the fan shaft.

17. The gas turbine engine as recited in claim 16, wherein the gear train includes collection means for receiving lubricant expelled by the discharge means through the radial interface, and the ring gear includes accumulation means for capturing lubrication expelled toward the radial interface.

18. The gas turbine engine as recited in claim 17, wherein the gear train includes return means for communicating lubricant from an outer periphery of the respective first and second portions of the ring gear outwardly to the collection means.

19. The gas turbine engine as recited in claim 18, further comprising: a splitter that defines an entrance to the compressor section; and wherein the radial interface of the ring gear is axially aft of the splitter relative to the engine axis.

20. The gas turbine engine as recited in claim 14, wherein the epicyclic gear train is a planetary gear system.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0020] 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.

[0021] FIG. 2 is an enlarged cross-sectional view of the epicyclic gear train shown in FIG. 1.

[0022] FIG. 3 is an enlarged cross-sectional view of an example ring gear similar to the arrangement shown in FIG. 2.

[0023] 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.

[0024] FIG. 5 shows another embodiment.

[0025] FIG. 6 shows yet another embodiment.

DETAILED DESCRIPTION

[0026] 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.

[0027] 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 a sun gear 30 that is connected to the compressor shaft 24, which provides rotational input, by a 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.

[0028] As shown, each of the star gears 32 is supported on one of the journal bearings 34. Each journal bearing 34 has an internal central cavity 34a that extends between axial ends 35a and 35b. In this example, as shown, the internal central cavity 34a is axially blind in that the axial end 35a is closed. At least one passage 37 extends from the internal central cavity 34a to a peripheral journal surface 39. In the example, the at least one passage 37 includes a first passage 37a and a second passage 37b that is axially spaced from the first passage 37a. As shown, the first and second passages 37a and 37a are non-uniformly spaced with regard to the axial ends 35a and 35b of the internal central cavity 34a.

[0029] In operation, lubricant is provided to the internal central cavity 34a. The lubricant flows through the internal central cavity 34a and then outwardly through the at least one passage 37 to the peripheral journal surface 39. The arrangement of the internal central cavity 34a and at least one passage 37 thereby serves to cool and lubricate the journal bearing 32.

[0030] The gas turbine engine 10 is a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, the engine 10 has a bypass ratio that is greater than about six (6) to ten (10), the epicyclic gear train 22 is a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3 or greater than about 2.5, and a low pressure turbine of the engine 10 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 10 bypass ratio is greater than about ten (10:1) or greater than about 10.5:1, the turbofan 18 diameter is significantly larger than that of the low pressure compressor of the compressor section 14, and the low pressure turbine has a pressure ratio that is greater than about 5:1. In one example, the epicyclic gear train 22 has a gear reduction ratio of greater than about 2.3:1 or greater than about 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.

[0031] A significant amount of thrust is provided by a bypass flow B due to the high bypass ratio. The fan 18 of the engine 10 is designed for a particular flight condition—typically cruise at about 0.8 M and about 35,000 feet. The flight condition of 0.8 M and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise 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. 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 [(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 about 1150 ft/second.

[0032] 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.

[0033] 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.

[0034] 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.

[0035] 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.

[0036] 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.

[0037] Although embodiments of this invention have 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.