SEALING SYSTEM BETWEEN A BEARING HOUSING AND AN INTERFACE BETWEEN TWO SHAFTS OF A GAS TURBINE DRIVE UNIT

Description

FIELD

[0001] The present invention relates generally to systems for sealing portions of a gas turbine engine and, more specifically, to sealing systems configured to reduce fluid entry into an interface between concentric shafts within a gas turbine engine.BACKGROUND

[0002] Conventional gas turbine engines include a bearing compartment that houses a number of bearings providing support to a number of shafts. Within the bearing compartment, an interface between an inner and outer shaft may be positioned in a separate annulus. This annulus may be sealed from the bearing compartment by an intershaft seal. However, oil or other liquids may pass through the intershaft seal and interact with the interface between the inner and outer shaft. This oil may pose a safety hazard. Accordingly, improved sealing systems may be desirable.

[0003] A prior art sealing system having the features of the preamble to claim 1 is disclosed in US 5,211,535.SUMMARY

[0004] From one aspect, the present invention provides a sealing system in accordance with claim 1. The secondary seal may comprise at least one knife edge seal and / or at least one discourager. The outer shaft may comprise a high pressure compressor shaft, and the inner shaft may comprise a low pressure turbine shaft.

[0005] From another aspect, the present invention provides a gas turbine engine in accordance with claim 4. The secondary seal may include at least one of a knife edge seal and a discourager. The outer shaft comprises a high pressure compressor shaft. The inner shaft may comprise a low pressure turbine shaft.BRIEF DESCRIPTION OF THE DRAWINGS

[0006] The subject matter of the present invention is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements. Figure 1 illustrates a perspective view of an aircraft engine in accordance with the present disclosure; Figure 2 illustrates a cross sectional view of a portion of an aircraft engine in accordance with the present disclosure; Figures 3A-3C illustrate cross sectional views of a portion of an aircraft engine in accordance with the present disclosure. DETAILED DESCRIPTION

[0007] The detailed description of embodiments herein makes reference to the accompanying drawings, which show embodiments by way of illustration. While these embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that logical and mechanical changes may be made without departing from the scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not for limitation. For example, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and / or any other possible attachment option.

[0008] As used herein, "aft" refers to the direction associated with the tail of an aircraft, or generally, to the direction of exhaust of the gas turbine. As used herein, "fore" refers to the direction associated with the nose of an aircraft, or generally, to the direction of flight or motion.

[0009] The present disclosure describes devices and systems for sealing an annulus of a bearing component from an interface of concentric high pressure compressor and low pressure turbine shafts. As will be described in more detail, in various embodiments, systems comprising a secondary seal positioned between an intershaft seal and the shaft interface are provided herein.

[0010] With reference to Figure 1, a gas turbine engine 10 is shown. In general terms, gas turbine engine 10 may comprise a compressor section 24. Air may flow through compressor section 24 and into a combustion section 26, where it is mixed with a fuel source and ignited to produce hot combustion gasses. These hot combustion gasses may drive a series of turbine blades within a turbine section 28, which in turn drive, for example, one or more compressor section blades mechanically coupled thereto. Turbine section 28 may comprise, for example, a high pressure turbine section 40. Compressor section 24 may comprise a high pressure compressor section 36.

[0011] Each of the compressor section 24 and the turbine section 28 may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies may carry a plurality of rotating blades 25, while each vane assembly may carry a plurality of vanes 27 that extend into the core flow path C. The blades 25 create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine 10 along the core flow path C. The vanes 27 direct the core airflow to the blades 25 to either add or extract energy.

[0012] With reference to Figure 2, a portion of gas turbine engine 10 along the axis I-I is illustrated. Gas turbine engine 10 may comprise, for example, a bearing compartment 12, bearings 14 and 14A, a high pressure compressor shaft 16, a low pressure turbine shaft 18, a tower shaft, a cross-over housing 30, and an intershaft seal assembly 32. As depicted in Figure 1, bearings 14 and 14A each comprise a single bearing which extend inwardly and outwardly from the page.

[0013] In various embodiments, bearing compartment 12 may be a mid-bearing compartment of gas turbine engine 10. Bearing compartment 12 may house bearings 14 and 14A, which are disposed around and support a high pressure compressor shaft 16, a low pressure turbine shaft 18, and / or a tower shaft. In particular, bearings 14 may support high pressure compressor shaft 16 and low pressure turbine shaft 18. Bearings 14A may, for example, support a tower shaft. High pressure compressor shaft 16 may, for example, concentrically surround and be spaced apart from low pressure turbine shaft 16 (in such configurations, high pressure compressor shaft 18 is an outer shaft and low pressure turbine shaft 16 is an inner shaft).

[0014] In various embodiments, an outer housing 30 extends annularly around and is disposed around components such as bearings 14. Outer housing 30 may, for example, extend radially from the adjacent casing and may connect to cross-over housing 38. Cross-over housing 38 may extend from outer housing 30 and interface and support bearings 14A disposed around a tower shaft. In various embodiments, intershaft seal assembly 32 is disposed radially inward of cross-over housing 30 and bearing compartment 12.

[0015] With reference to Figures 1-3A, high pressure compressor shaft 16 may, for example, connect high pressure compressor section 36 of gas turbine engine 10 (illustrated in Figure 1) with high pressure turbine section 40 (illustrated in Figure 1). Together, high pressure compressor shaft 16, high pressure compressor section 36, and high pressure turbine section 40 comprise a high pressure spool of engine 10. Similarly, low pressure turbine shaft 18 may connect a fan section with a low pressure compressor section 37 (illustrated in Figure 1) and a low pressure turbine section 41 of engine 10 (illustrated in Figure 1). Together, these sections and components comprise a low pressure spool of engine 10. During operation of gas turbine engine 10, a gas flow may pass through a core flow path to rotate the high pressure and low pressure spools to extract energy from the gas flow and produce thrust. Additionally, gas flow may bypass the core flow path and rotate the fan to extract additional energy and thrust from the gas flow passing through a nacelle portion encasing engine 10.

[0016] In various embodiments, bearing compartment 12 circumscribes shafts 16 and 18 and bearings 14 and 14A and acts to contain a lubricant oil that is supplied to bearing compartment 12 to lubricate and cool the bearings 14 and 14A during operation of the gas turbine engine 10. In various embodiments, bearings 14 support high pressure compressor shaft 16 and low pressure turbine shaft 18 and allow shafts 16 and 18 to rotate relative to a stator engine casing. Similarly, bearings 14A may support a tower shaft and allow the tower shaft to rotate relative to the stator engine casing.

[0017] Cross-over housing 30 may, for example, support intershaft seal assembly 32. In various embodiments, intershaft seal assembly 32 controls a gas circulated between shafts 16 and 18 from to leak into the lubricating oil supplied portions of bearing compartment 12. Further, intershaft seal assembly 32 prevents the lubricating oil from leaking past an interface into the volume between the shafts 16 and 18. To accomplish this, intershaft seal assembly 32 comprises seals that are biased to ride against seats. In various embodiments, intershaft seal assembly 32 comprises a metallic material. In other embodiments, intershaft seal assembly 32 comprises a carbon material. Although described with reference to specific embodiments, intershaft seal assemblies of any suitable material are within the scope of the present disclosure.

[0018] With initial reference to Figures 3A-3C, intershaft seal assembly 32 comprises a buffer air inlet 60. Buffer air inlet 60 allows pressurized air to enter intershaft seal assembly 32. Air entering intershaft seal assembly 32 through buffer air inlet 60 provides back pressure to annulus 46.

[0019] In various embodiments, bearing compartment 12 further comprises a secondary seal 50. For example, lubricating oil from bearing compartment 12 may, undesirably, pass through intershaft seal assembly 32 and into annulus 46. Secondary seal 50 is configured to prevent lubricating oil that has passed into annulus 46 from entering interface 48. Lubricating oil entering interface 48 may present a safety risk, as it may pass through interface 48, into the space between high pressure compressor shaft 16 and low pressure turbine shaft 18, and eventually accumulate in the turbine section.

[0020] In various embodiments, secondary seal 50 may comprise one or more knife edge seals 52. For example, in the illustrated embodiment, secondary seal 50 comprises two knife edge seals 52. However, any number of knife edge seals, including a single knife edge seal, is within the scope of the present disclosure.

[0021] Secondary seal 50 may further comprise, for example, one or more discouragers 54. For example, in the illustrated embodiment, secondary seal 50 comprises one discourager 54. However, any number of discouragers (including zero) is within the scope of the present disclosure. Further, any number of knife edge seals and discouragers may be combined to form secondary seal 50.

[0022] With reference to Figure 3B, annulus 46 comprises an oil drain. For example, as illustrated in Figure 3B, oil drain may be in or close to the same axial location as buffer air inlet 60 and offset in the circumferential direction. In various embodiments, when sufficient lubricating oil has accumulated in annulus 46, lubricating oil may be removed from gas turbine engine 10 via oil drain. Buffer air from buffer air inlet 60 assists in removal of lubricating oil from annulus 46 via oil drain by providing back pressure to annulus 46 (as discussed previously).

[0023] It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the invention is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean "one and only one" unless explicitly so stated, but rather "one or more." Moreover, where a phrase similar to "at least one of A, B, or C" is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials.

[0024] Systems, methods and apparatus are provided herein. In the detailed description herein, references to "one embodiment," "an embodiment," "an example embodiment," etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.

Claims

1. A sealing system for a gas turbine engine (10) comprising: an intershaft seal (32) separating a bearing compartment (12) from an annulus (46) , the intershaft seal being configured to prevent a lubricating oil in the bearing compartment (12) from entering the annulus (46); an outer shaft (16) concentrically surrounding at least a portion of an inner shaft (18), wherein an interface (48) between the outer shaft (16) and the inner shaft (18) is located within the annulus (46), and the intershaft seal (32) is configured to prevent the lubricating oil from leaking past the interface (48) into a volume defined between the outer shaft (16) and the inner shaft (18); a secondary seal (50) positioned within the annulus (46) such that it is configured to substantially prevent a lubricating oil in the annulus (46) from entering the interface (48) and passing through to the volume; and characterised by an oil drain located in the annulus (46); and in that the intershaft seal (32) is radially outward of the annulus (46) and the secondary seal (50) is radially inward of the intershaft seal (32), and the intershaft seal (32) comprises a buffer air inlet (60) to allow pressurized air to enter through the intershaft seal (32) and provide a back pressure to the annulus (46).

2. The sealing system of claim 1, wherein the secondary seal (50) comprises at least one knife edge seal (52).

3. The sealing system of claim 1 or 2, wherein the secondary seal (50) comprises a discourager (54).

4. A gas turbine engine (10) comprising the sealing system of claim 1, wherein the secondary seal (50) comprises at least one knife edge seal (52).

5. The gas turbine engine of claim 4, wherein the secondary seal (50) comprises at least one discourager (54).

6. The sealing system or gas turbine engine of any preceding claim, wherein the outer shaft (16) comprises a high pressure compressor shaft (16).

7. The sealing system or gas turbine engine of any preceding claim, wherein the inner shaft (18) comprises a low pressure turbine shaft (18).