Non-contacting seals for geared gas turbine engine bearing compartments

Abstract

A gas turbine engine includes a fan, a compressor section, a combustor, and a turbine section. The engine also includes a rotating element and at least one bearing compartment including a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. A method and section for a gas turbine engine are also disclosed.

Claims

1. A seal arrangement for a bearing compartment of a gas turbine engine comprising: a rotating element and at least one bearing compartment including a bearing for supporting said rotating element, said rotating element rotatable about an axis; wherein said at least one bearing compartment has a first seal and a second seal each associated with a corresponding one of two opposed axial ends, on either axial side of said bearing relative to said axis, at least one of said first seal and said second seal being a non-contacting seal having a seal face facing a rotating face of said rotating element; and wherein a radially outermost location of said bearing compartment defines a compartment radius with respect to said axis, a radial outermost location of said non-contacting seal establishes a seal radius with respect to the engine axis, and a compartment-seal ratio defined by said compartment radius to said seal radius is less than or equal to 6:1.

2. The seal arrangement as set forth in claim 1, wherein said non-contacting seal is arranged to resist leakage of lubricant outwardly of said at least one bearing compartment and to allow pressurized air to flow from a chamber adjacent said non-contacting seal into said at least one bearing compartment, and a grooved area is formed in one of said faces, with said grooved area having a plurality of circumferentially spaced grooves for generating hydrodynamic lift-off forces and allowing leakage of pressurized air across said faces and into said at least one bearing compartment to resist leakage of lubricant from said at least one bearing compartment.

3. The seal arrangement as set forth in claim 2, wherein said rotating element is configured to rotate at a velocity greater than or equal to 450 feet per second.

4. The seal arrangement as set forth in claim 2, wherein said non-contacting seal being formed with a plurality of passages configured to allow tapping of additional pressurized air to be delivered to said faces at a location in the proximity of said grooved area for generating hydrostatic lift-off forces.

5. The seal arrangement as set forth in claim 4, wherein said grooved area is spaced radially from said plurality of passages at said seal face.

6. The seal arrangement as set forth in claim 5, wherein said rotating element is a shaft rotatable with a rotor having an axial face facing said seal face, and said grooved area is formed in said rotor.

7. The seal arrangement as set forth in claim 2, wherein said non-contacting seal is a controlled gap carbon seal having a full hoop seal and a metal band shrunk fit onto said non-contacting seal, and positioned in a seal carrier.

8. The seal arrangement as set forth in claim 2, wherein said rotating element is a shaft rotatable with a rotor having an axial face facing said seal face.

9. The seal arrangement as set forth in claim 1, wherein said rotating element is configured to rotate at a velocity greater than or equal to 450 feet per second.

10. The seal arrangement as set forth in claim 9, wherein said compartment-seal ratio is between 3:1 to 5:1.

11. The seal arrangement as set forth in claim 9, wherein said rotating element is configured to rotate at a velocity less than or equal to 600 feet per second.

12. A gas turbine engine comprising: a fan section having a plurality of fan blades, a gear arrangement, a compressor section, and a turbine section arranged along an engine axis, said turbine section including a first turbine and a second turbine, said second turbine configured to drive said fan section through said gear arrangement; a seal arrangement comprising: a rotating element and at least one bearing compartment including a bearing for supporting said rotating element; wherein said at least one bearing compartment has a first seal and a second seal each associated with a corresponding one of two opposed axial ends, on either axial side of said bearing relative to said engine axis, at least one of said first seal and said second seal being a non-contacting seal having a seal face facing a rotating face of said rotating element; and wherein a radially outermost location of said fan blades define a fan radius with respect to said engine axis, a radially outermost location of said bearing compartment defines a compartment radius with respect to said engine axis, and a fan-compartment ratio defined by said fan radius to said compartment radius is greater than or equal to 2:1.

13. The gas turbine engine as set forth in claim 12, wherein said first turbine is configured to drive said rotating element.

14. The gas turbine engine as set forth in claim 12, wherein a radially outermost location of said bearing compartment defines a compartment radius with respect to said engine axis, a radial outermost location of at least one of said first and second seals establishes a seal radius with respect to the engine axis, and a compartment-seal ratio defined by said compartment radius to said seal radius is less than or equal to 6:1.

15. The gas turbine engine as set forth in claim 14, wherein said rotating element is configured to rotate at a velocity greater than or equal to 450 feet per second.

16. The gas turbine engine as set forth in claim 15, wherein a grooved area is formed in one of said faces, with said grooved area having a plurality of circumferentially spaced grooves for generating hydrodynamic lift-off forces and allowing leakage of pressurized air across said faces and into said at least one bearing compartment to resist leakage of lubricant from said at least one bearing compartment.

17. The gas turbine engine as set forth in claim 12, wherein said rotating element is configured to rotate at a velocity between 450 feet per second and 600 feet per second.

18. A method of operating a gas turbine engine, the method comprising the steps of: arranging a bearing within a bearing compartment to support a rotating element, said rotating element defining a rotating face, said bearing compartment having a first seal and a second seal each associated with a corresponding one of two opposed axial ends relative to an engine axis, on either axial side of said bearing; rotating said rotating face relative to at least one of said first seal and said second seal; sealing said bearing compartment with said first seal and said second seal, at least one of said first seal and said second seal being a non-contacting seal configured to resist leakage of lubricant outwardly of said bearing compartment and to allow air to flow from a chamber adjacent said non-contacting seal and into said bearing compartment, said non-contacting seal defining a seal face facing said rotating face; and wherein a radially outermost location of said bearing compartment defines a compartment radius with respect to said engine axis, a radial outermost location of said non-contacting seal establishes a seal radius with respect to the engine axis, and a ratio of said compartment radius to said seal radius is less than or equal to 6:1.

19. The method as set forth in claim 18, wherein said rotating element is a shaft rotatable with a rotor having an axial face facing said seal face.

20. The method as set forth in claim 18, comprising: communicating air from a fan to a bypass passage and to a compressor section, wherein a bypass ratio is defined as the volume of air passing into said bypass passage compared to the volume of air passing into said compressor section, said bypass ratio greater than 10 at a cruise condition; and wherein said step of rotating comprises rotating said rotating element at a velocity greater than or equal to 450 feet per second, with said rotating element driving said compressor section.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) FIG. 1 schematically shows a gas turbine engine.

(2) FIG. 2A schematically shows example locations of bearing compartments.

(3) FIG. 2B schematically shows dimensional aspects of the bearing compartments of FIG. 2A.

(4) FIG. 3A is a first embodiment of a non-contact seal according to the present invention.

(5) FIG. 3B shows a second embodiment of a non-contact seal according to the present invention.

(6) FIG. 3C shows a third embodiment of a non-contact seal according to the present invention.

(7) FIG. 3D shows a fourth embodiment of a non-contact seal assembly according to the present invention.

DETAILED DESCRIPTION

(8) 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.

(9) 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.

(10) The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a 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. A radially outermost location of the fan blades of the fan 42 establishes a radius R.sub.F relative to the engine central longitudinal axis A. In some embodiments, radius R.sub.F is between about 28 inches and about 40.5 inches. In an embodiment, the radius R.sub.F is about 36.5 inches. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and 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.

(11) 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 geared architecture 48 may be varied. For example, geared architecture 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 geared architecture 48.

(12) 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 epicyclic 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.

(13) 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 conditiontypically 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 consumptionalso 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 [(Tram 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.

(14) FIG. 2A shows an embodiment of an arrangement of bearing compartments 100 associated with the gas turbine engine, such as the gas turbine engine 20 illustrated in FIG. 1. As shown, a bearing compartment 102 is associated with a low speed shaft 92 at a location associated with the low pressure turbine. Bearings 106 are shown schematically as is a seal 104.

(15) A bearing compartment 108 is associated with a high speed rotor 90 and at the high pressure turbine of FIG. 1. Bearing compartment 108 includes seals 110 at each axial end and a central bearing 112.

(16) Another bearing compartment 114 is also associated with the high speed rotor 90 and the high pressure compressor and includes a bearing 118 and seals 116.

(17) Finally, a bearing compartment is associated with a fan drive gear system 122 at location 120 and with and the fan at location 123. Seals 126 and 128 mechanically seal the axial ends of the bearing compartment 120 and are associated with the fan rotor 127 and the low speed rotor 92. The seals 126, 128 are also respectively associated with the bearings 124 and 130 that are positioned within the bearing compartment 120/123.

(18) Referring to FIG. 2B, the bearing compartments 102, 108, 114, 120/123, establish a dimensional relationship with respect to the engine central longitudinal axis A. A radially outermost location of bearing compartments 102, 108, 114, 120/123 establish corresponding radii R.sub.102, R.sub.108, R.sub.114 and R.sub.120/123.

(19) Utilizing the seal arrangements disclosed herein, the relative sizes of the bearing compartments 102, 108, 114 and 120/123 can be reduced to provide for a more compact core engine arrangement, which may be utilized with a relatively high bypass ratio. In some embodiments, radii R.sub.120, 123 and R.sub.114 establish a compartment ratio that is between about 1:1 to about 2:1. In embodiments, radii R.sub.120, 123 and R.sub.102 establish a compartment ratio that is between about 2:1 to about 3:1. In some embodiments, radii R.sub.108 and R.sub.102 establish a compartment ratio that is between about 1:1 to about 2:1. For the purposes of this disclosure, the term about is relative to the number of significant digits unless otherwise noted.

(20) The sizes of the bearing compartments 102, 108, 114 and 120/123 relative to the fan 42 can also be reduced. In embodiments, radii R.sub.F and radii R.sub.120/123 establish a fan-compartment ratio that is greater than or equal to about 2:1, such as between about 2.5:1 and about 5:1. In one embodiment, the fan-compartment ratio is about 3:1. In embodiments, radii R.sub.F and radii R.sub.114 establish a fan-compartment ratio that greater than or equal to about 4:1. In an embodiment, radii R.sub.F and radii R.sub.114 establish a fan-compartment ratio that is about 5:1. In embodiments, radii R.sub.F and radii R.sub.108 establish a fan-compartment ratio that greater than or equal to about 8:1, and less than or equal to about 12:1. In an embodiment, radii R.sub.F and radii R.sub.108 establish a fan-compartment ratio that is about 10:1. In embodiments, radii R.sub.F and radii R.sub.102 establish a fan-compartment ratio that is greater than or equal to about 12:1, and less than or equal to about 18:1. In an embodiment, radii R.sub.F and radii R.sub.102 establish a fan-compartment ratio that is about 15:1.

(21) There are challenges with sealing the bearing compartments in a geared turbofan engine. Accordingly, various embodiments disclosed herein relate to the use of non-contacting seals such as lift-off seals at any one or more of the locations of the seals shown in FIG. 2A or in any other bearing compartment on a gas turbine engine. In some embodiments, the seals may be lift-off seals and, more particularly, may be carbon lift-off seals. Of course, in other embodiments other non-contacting seals, including other lift-off seals may be used.

(22) Thus, as shown in FIG. 3A, a shaft 140, which could be any rotating shaft in a gas turbine engine, has a mating rotor 142. An axial face 147 of this mating rotor 142 is sealed relative to a face 145 from a seal 144. The faces 145 and 147 face each other to form a mechanical seal. The seal 144 may be a non-contact seal such as a carbon seal lift-off seal. The interface between faces 145 and 147 experiences high velocities, especially when compared to the prior art. The high velocity is a combination of a high rotational speed of the shaft 140 and a relatively large diameter for the seal 144. Velocities greater than or equal to about 450 feet/second (137.16 meters/second) may be seen.

(23) In the FIG. 3A embodiment seal 144, a set of shallow grooves 152 is provided by cutting into the face 147 of the rotor 142, as shown at circumferentially spaced grooves 154. A spring 146 biases the seal 144 toward the face 147. A higher pressure air is available in a chamber 148, which is on an opposed side of the seal 144 from the bearing compartment 150. The bearing compartment 150 is at a lower pressure than the chamber 148, and this higher pressure air passes through the grooved area 152, such that the air flow levitates (lifts-off) the sealing surface 145 of the non-rotating seal 144 from the sealing surface 147 of the rotor 142. The levitation is a result of hydrodynamic lifting force as the air passes into the bearing compartment 150, preventing oil from escaping the bearing compartment 150.

(24) Another embodiment is illustrated in FIG. 3B. FIG. 3B provides the mechanical sealing between face 168 of a non-rotating seal 162 and a face 170 of a rotor in a manner somewhat similar to the FIG. 3A embodiment. There is a grooved area 172 having circumferentially spaced grooves 174 with the features as described for the first embodiment that generate hydrodynamic lifting force as the gas passes from a high pressure chamber 160 into the bearing compartment 274. Furthermore, the non-rotating seal 162 has an inlet 166, a passage 164, and an outlet 180 which delivers additional high pressure air generating hydrostatic lifting forces at a radial location in the proximity of the grooved area 172, thereby providing a stronger and more stable lift-off seal compared to the first embodiment. As shown, there is a plurality of circumferentially spaced outlets 180. The non-rotating seal 162, which is biased toward the rotor 170 by a spring 161, may be a carbon lift-off seal.

(25) FIG. 3C shows another embodiment 182, wherein the seal 186 has a plurality of circumferentially segmented members biased by spring 184 toward a face 185 of a rotor 142 rotating with the shaft 140. Seal 186 has a radially inwardly facing face 183 providing the seal face with the mating face 185. One of the sealing faces, either 183 or 185, has a set of shallow, circumferentially spaced grooves 210 in a grooved area 211, somewhat similar to those described in the earlier embodiments that generate a hydrodynamic force that levitates (lifts-off) the non-rotating sealing surface 183 from the rotating mating surface 185 when the high pressure chamber 188 delivers pressurized air across the seal 186 to prevent leakage of oil from the bearing compartment 190.

(26) FIG. 3D shows an embodiment of a controlled gap non-contacting seal assembly 201. The shaft 140 has an outer surface spaced by a small gap 196 from two carbon seals 192. The gap is controlled by design, typically by sizing the sealing diameters of the seals 192 and the rotor 142 such that a small gap is maintained under all conditions. The shaft outer surface 193 and a radially inward facing surfaces 191 of the seals 192 provide the seal faces. In one embodiment, the carbon seals 192 are full hoop members extending around the entire circumference of the shaft 140. A metal band 194 is shrunk fit onto the seal 192. A carrier 195 mounts the seals 192. A high pressure chamber 198 is spaced from the bearing compartment 200, such that high pressure air passes through the gap 196 to prevent the leakage of lubricant.

(27) The seal arrangements disclosed herein can be utilized to reduce the relative sizes of the bearing compartments to provide for a more compact core engine arrangement. The seals 144, 162, 186, 192 establish a dimensional relationship with engine central longitudinal axis A. A radially outermost location of seals 144, 162, 186, 192 establish corresponding radii R.sub.144, R.sub.162, R.sub.186 and R.sub.192 with respect to engine axis A (shown in FIGS. 3A-3D, respectively). For the purposes of this disclosure, the radially outermost location of each seal 144, 162, 186, 192 is measured at an outermost position that the seal 144, 162, 186, 192 is permitted to move relative to the engine central longitudinal axis A during normal operation of the engine 20. A radially outermost location of the bearing compartments 150, 274, 190, 200 establish corresponding radii R.sub.150, R.sub.274, R.sub.190 and R.sub.200 with respect to engine axis A (shown in FIGS. 3A-3D, respectively). For the purposes of this disclosure, the radially outermost location of each bearing compartment 150, 274, 190, 200 is measured at an inner surface of the cavity of the bearing compartment 150, 274, 190, 200 housing the respective bearing(s).

(28) In embodiments, radii R.sub.150/274/190/200 and R.sub.144/162/186/192 establish a compartment-seal ratio that is less than or equal to about 6:1. In some embodiments, the compartment-seal ratio is between about 3:1 and about 5:1.

(29) As previously discussed, the interface between the shaft or rotor and seal(s) may operate at relatively high velocities. In embodiments, the interface between the faces of the shaft 140 or rotor 142 and seal(s) 144/162/186/192 is configured to operate at relative velocities greater than or equal to about 450 feet/second. In some embodiments, the interface operates at relative velocities less than or equal to about 600 feet/second, and more narrowly less than or equal to about 550 feet/second. In embodiments, the shaft 140 corresponds to the high speed rotor 90 (FIG. 2A).

(30) In other embodiments, the interface between the faces of the shaft 140 or rotor 142 and seal(s) 144/162/186/192 is configured to operate at relative velocities less than or equal to about 450 feet/second, such as between about 400 to about 450 feet/second. In embodiments, the shaft 140 corresponds to the low speed rotor 92 (FIG. 2A). In other embodiments, the shaft 140 corresponds to the fan rotor 127 (FIG. 2A). In some embodiments, the interface between the faces of the shaft 140 or rotor 142 and seal(s) 144/162/186/192 is configured to operate at relative velocities between about 200 to about 400 feet/second.

(31) In some embodiments, a velocity of the high speed rotor 90 and the low speed rotor 92 establish a ratio that is between about 1:1 to about 2:1. In embodiments, a velocity of the high speed rotor 90 and the fan rotor 127 establish a ratio that is between about 5:2 to about 3:1, with the fan rotor 127 driven by the fan drive gear system 122.

(32) All of the disclosed embodiments reduce the friction between the seal and the rotating components. This reduces heat generation due to friction, increases the durability of the seals, minimizes loss of oil, and increases the efficiency in fuel consumption of the overall engine. Moreover, as a result of the reduction in friction, less lubricant can be used, thereby also reducing the size of the applicable fluid storage tank (not shown) and the applicable cooling system fluid pumping apparatus (also not shown). Accordingly, the overall weight of the engine may be greatly reduced, thereby increasing the engine's fuel efficiency.

(33) The disclosed embodiments may be useful at any bearing compartment in a gas turbine engine. Although shafts are shown supported by the bearings, the disclosure would extend to other rotating elements supported by a bearing.

(34) Although various 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.