US20090110548A1

US20090110548A1 - Abradable rim seal for low pressure turbine stage

Info

Publication number: US20090110548A1
Application number: US11/928,512
Authority: US
Inventors: Eric Durocher; Guy Lefebvre
Original assignee: Pratt and Whitney Canada Corp
Current assignee: Pratt and Whitney Canada Corp
Priority date: 2007-10-30
Filing date: 2007-10-30
Publication date: 2009-04-30
Also published as: CA2639026A1

Abstract

An abradable rim seal arrangement for a turbine stage of a gas turbine engine, which includes an abradable rim seat attached to a leading edge of an annular inner duct of a turbine exhaust case, causing a trailing edge platform of each rotor blade to overlap the abradable rim seal, and allowing rubbing with the trailing edge platform during engine operation.

Description

TECHNICAL FIELD

The invention relates generally to gas turbine engines and, more particularly, to an improved abradable rim seal for a turbine stage of a gas turbine engine.

BACKGROUND OF THE ART

In a gas turbine engine hot gases produced in a combustion chamber are directed between an outer annular ducting surface formed by a turbine shroud and an annular inner ducting surface formed by a plurality of turbine rotor blade platforms to pass through the turbine rotor blades in one or more turbine stages for powering the engine. The hot gases exhausted from the turbine stages are guided through an annular gas path defined by a turbine exhaust case, to be discharged. Conventionally, a minimum gap is provided between the low pressure turbine rotor and the turbine exhaust case in order to prevent possible rubbing contact between the rotating part of the turbine rotor and the stationary part of the turbine exhaust case, which could damage both compontents. Consequently, rim seal efficiency is poor and may allow some hot gas ingestion in a worst case scenario. When hot gas ingestion occurs, some turbine components in the disc rear cavity are exposed to hot gas temperatures, which has a direct and negative impact on blade/disc durability. Furthermore, such a minimum gap between the turbine exhaust case is typically obtained by accurately machining a leading edge of an inner duct of the turbine exhaust case which is made in a casting or forging process. However, it is difficult to achieve accurate sizing of the leading edge of an inner duct of a turbine exhaust case which is made of sheet metal.
Accordingly, there is a need to provide an improved rim seal for a turbine stage in order to minimize hot gas ingestion that has a direct impact on blade/disc durability.

SUMMARY OF THE INVENTION

In one aspect, the present invention provides an abradable rim seal arrangement for a turbine stage of a gas turbine engine which comprises a trailing edge platform extending axially and rearwardly from each of a plurality of rotor blades, the trailing edge platform having an inner surface facing radially and inwardly; a turbine exhaust case having annular outer and inner ducts located downstream of the rotor blades for ducting hot exhaust gases from the turbine stage; and an abradable rim seal attached to a leading edge of the annular inner duct of the turbine exhaust case, causing the trailing edge platform of each of the rotor blades to overlap the abradable rim seal, and allowing rubbing with the inner surface of the trailing edge platform during engine operation to substantially prevent a rotor disc assembly of the turbine stage from being exposed to the hot gases.
In another aspect, the present invention provides a turbine assembly of a gas turbine engine, which comprises a rotor disc and a plurality of rotor blades extending radially and outwardly from the rotor disc, each of the rotor blades including a trailing edge platform extending axially and rearwardly, the trailing edge platform having an inner surface facing radially and inwardly; a turbine exhaust case having annular outer and inner ducts located downstream of the rotor blades for ducting hot exhaust gases between the ducts; and means for creating an overlapping and contactable condition between a leading edge of the annular inner duct of the turbine exhaust case and the inner surface of the trailing edge platform of the respective rotor blades during engine operation to substantially prevent the rotor disc from being exposed to the hot gases.
In another aspect, the present invention provides a method for substantially reducing hot gas ingestion into a rotor disc rear cavity of a turbine rotor assembly of a gas turbine engine, in which a turbine exhaust case has annular outer and inner ducts located downstream of the turbine rotor assembly for ducting hot exhaust gases having passed through a plurality of turbine rotor blades, the method comprising attaching an abradable rim seal to a leading edge of an annular inner duct of the turbine exhaust case; and machining the abradable rim seal to form a sealing surface thereon for being positioned under and being overlapped by an inner surface of the leading edge platform of each turbine rotor blade when assembling the gas turbine engine, thereby allowing a rubbing condition between the abradable rim seal and the leading edge platform during engine operation
Further details of these and other aspects of the present invention will be apparent from the detailed description and drawings included below.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying drawings depicting aspects of the present invention, in which:

FIG. 1 is a schematic cross-sectional view of a bypass gas turbine engine as an exemplary application of the present invention;

FIG. 2 is a partial cross-sectional view of a low pressure turbine stage of the gas turbine engine of FIG. 1, illustrating an abradable rim seal arrangement according to one embodiment of the present invention;

FIG. 3 is a partial cross-sectional view of the low pressure turbine stage showing the details of the rim seal arrangement as circled and indicated by numeral 3 in FIG. 2; and

FIG. 4 is a cross-sectional view of a turbine exhaust case used in the gas turbine engine of FIG. 1.

DETAILED DESCRIPTION OF THE EMBODIMENTS

Referring to FIG. 1, a bypass gas turbine engine presented as an example of the application of the present invention, includes a housing or nacelle 10, a core casing 13, a low pressure spool assembly seen generally at 12 which includes a fan assembly 14, a low pressure compressor assembly 16 and a low pressure turbine assembly 18, and a high pressure spool assembly seen generally at 20 which includes a high pressure compressor assembly 22 and a high pressure turbine assembly 24. The core casing 13 surrounds the low and high pressure spool assemblies 12 and 20 in order to define a main fluid path (not indicated) therethrough. In the main fluid path there is provided a combustor 28 to constitute a gas generator section 26.
Referring to FIGS. 1-4, the bypass gas turbine engine further includes an annular turbine exhaust case 30 located downstream of the low pressure turbine assembly 18 and attached to the core casing 13, as an example of the present invention, which includes an annular inner duct 32 and annular outer duct 34. A plurality of circumferential spaced apart airfoils 36 extend radially between the inner and outer ducts 32, 34, to thereby structurally connect same. A bearing housing 38 is co-axially connected to the inner duct 32 for supporting an aft end of a main shaft (not indicated) of the low pressure spool assembly 12. Optionally, there is a mixer 40 attached to the rear end of the annular outer duct 34. A mounting flange (not indicated) is integrated with the annular outer duct 34 at a front end thereof for securing the annular turbine exhaust case 30 to the engine core casing 13, which in turn is structurally connected to the engine housing 10 through a plurality of radially extending struts (not indicated).
A tail cone indicated by numeral 44 may be attached to a rear end of the annular inner duct 32 of the turbine exhaust case 30 to cover the opening defined by a rear end of the annular inner duct 32, in order to provide an aerodynamic fairing.
The low pressure turbine assembly 18 generally includes a rotor disc 46 and a plurality of rotor blades 48 (only one is shown) extending radially and outwardly from the rotor disc 46. Each of the rotor blades 48 includes a trailing edge platform 50 extending axially and rearwardly from the rotor blade 48. The trailing edge platform 50 may be an integrated portion of an inner blade platform 52 extending laterally from the rotor blade 48. The inner blade platform 52 substantially divides each rotor blade 48 into an airfoil section (not indicated) and a blade root section (not indicated) which is affixed in a corresponding attachment slot in the periphery of the rotor disc 46. The inner blade platforms 52 of the plurality of rotor blades 48 are adjacent one to another and the outer surfaces thereof in combination define an inner ducting surface of the low pressure rotor assembly 18.
A stationary turbine shroud or a rotating turbine shroud ring which may be formed by a plurality of outer blade platforms 54, each being affixed to a top of one of the rotor blades 48, defines an annular outer ducting surface of the low pressure turbine assembly 18. In operation, hot gases produced in the combustion chamber are directed between the annular inner and outer ducting surfaces (defined, for example, by the inner and outer blade platforms 52, 54) to pass through the plurality of rotor blades 48 (the airfoil sections thereof) for powering the low pressure turbine assembly 18, and are then directed by the turbine exhaust case 30 to pass through an annular exhaust passage defined between the annular inner and outer ducts 32, 34 in order to be discharged.
In accordance with one aspect of the present invention, there is means for creating an overlapping and contactable condition between a leading edge (not indicated) of the annular inner duct 32 of the turbine exhaust case 30 and an inner surface 56 of the trailing edge platform 50 of the respective rotor blades 48 during engine operation, to substantially prevent the rotor disc 46 from being exposed to the hot gases. The inner surface 56 of the trailing edge platform 50 of the rotor blade 48 is located under the trailing edge platform 50 and faces radially and inwardly.
In accordance with one embodiment, an abradable rim seal 58 which may be made of a honeycomb material such as Hastelloy® or Inconel® is attached to the leading edge of the annular inner duct 32 of the turbine exhaust case 30. The turbine exhaust case 30 which may be made of sheet metal, for example, by a pressing and welding process, may include a leading edge flange 60 for supporting the abradable rim seal 58 thereon. The leading edge flange 60, for example, may also be made of sheet metal in a continuous ring having a U-shaped cross section and may be welded to the leading edge of the annular inner duct 32 of the turbine exhaust case 30. The leading edge flange 60 has an outer peripheral diameter smaller than both a diameter defined by an outer surface 62 of the annular inner duct 32 and a diameter defined by the inner surface 56 of the trailing edge platform 50 of the respective rotor blades 48, to allow the abradable rim seal 58 which may be brazed to both the annular inner duct 32 and the leading edge flange 60 of the turbine exhaust case 30, to extend under the individual trailing edge platforms 50 of the respective rotor blades 48 to be overlapped by the trailing edge platforms 50, thereby allowing rubbing with the inner surface 56 of the trailing edge platforms 50 during engine operation to substantially avoid hot gas ingestion into a rotor disc rear cavity 64 (see FIG. 2) and to substantially prevent the rotor disc 46 and other attached components such as a disc cover (not shown), etc. from being exposed to the hot gases.
In accordance with another embodiment of the present invention, the abradable rim seal 58 may include a machined annular sealing surface (not indicated) which has a diameter smaller than the diameter defined by the inner surface 56 of the trailing edge platform of the respective rotor blades 48 to form a clearance therebetween (not indicated) when the engine is in a cold condition. The clearance will be substantially closed during engine operation by allowing a rubbing contact between the trailing edge platform 50 and the abradable rim seal 58, due to thermal expansion of engine components. The abradable rim seal 58 may have an external peripheral surface (not indicated) having a diameter substantially equal to the outer diameter (the outer surface 62) of the annular inner duct 32. A step (not indicated) may therefore be formed between the machined sealing surface and the external peripheral surface.
The abradable rim seal arrangement of the present invention advantageously reduces the opening between the hot gas path and the low pressure turbine disc rear cavity 64, thereby substantially reducing or avoiding hot gas ingestion into the disc/blade fixing area (not indicated), and thus improving the durability of turbine components such as the rotor disc, blade fixing parts, disc cover, etc. The present invention is advantageously applicable but not limited to a turbine assembly having a turbine exhaust case made of sheet metal.
The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departure from the scope of the invention disclosed. For example, a bypass gas turbine engine is used as an example of the application of this invention but this invention is applicable to any other type of gas turbine engine. A low pressure turbine stage is illustrated and described as the immediate environment of the abradable rim seal arrangement according to the described embodiments. However, the abradable rim seal arrangement of the present invention may be applicable to other types of turbine assemblies, such as a single stage of a turbine assembly of a gas turbine engine. The abradable rim seal arrangement can also be applied to a turbine exhaust case made of materials other than sheet metal, such as casting or forged products. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.

Claims

1. An abradable rim seal arrangement for a turbine stage of a gas turbine engine, comprising:

a trailing edge platform extending axially and rearwardly from each of a plurality of rotor blades, the trailing edge platform having an inner surface facing radially and inwardly;

a turbine exhaust case having annular outer and inner ducts located downstream of the rotor blades for ducting hot exhaust gases from the turbine stage; and

an abradable rim seal attached to a leading edge of the annular inner duct of the turbine exhaust case, causing the trailing edge platform of each of the rotor blades to overlap the abradable rim seal, and allowing rubbing with the inner surface of the trailing edge platform during engine operation to substantially prevent a rotor disc assembly of the turbine stage from being exposed to the hot gases.

2. The abradable rim seal arrangement as defined in claim 1 wherein the abradable rim seal comprises a honeycomb material.

3. The abradable rim seal arrangement as defined in claim 2 wherein the honeycomb material is brazed to the annular inner duct of the turbine exhaust case.

4. The abradable rim seal arrangement as defined in claim 1 wherein the abradable rim seal comprises a first annular surface having a first diameter, allowing the abradable rim seal to be overlapped by the inner surface of the rotor blades to create a clearance between the trailing edge platform and the abradable rim seal during a cold condition of the engine while allowing the inner surface of the trailing edge platform to rub against the abradable rim seal during engine operation.

5. The abradable rim seal arrangement as defined in claim 4 wherein the abradable rim seal comprises a second annular surface having a diameter substantially equal to a diameter of an inner surface of the inner duct of the turbine exhaust case.

6. A turbine assembly of a gas turbine engine, comprising:

a rotor disc and a plurality of rotor blades extending radially and outwardly from the rotor disc, each of the rotor blades including a trailing edge platform extending axially and rearwardly, the trailing edge platform having an inner surface facing radially and inwardly;

a turbine exhaust case having annular outer and inner ducts located downstream of the rotor blades for ducting hot exhaust gases between the ducts; and

means for creating an overlapping and contactable condition between a leading edge of the annular inner duct of the turbine exhaust case and the inner surface of the trailing edge platform of the respective rotor blades during engine operation to substantially prevent the rotor disc from being exposed to the hot gases.

7. The turbine assembly as defined in claim 6 wherein the means comprises an abradable rim seal attached to the leading edge of the annular inner duct of the exhaust case.

8. The turbine assembly as defined in claim 7 wherein the turbine exhaust case is made of sheet metal.

9. The turbine assembly as defined in claim 8 wherein the annular inner duct of the turbine exhaust case comprises a leading edge flange for supporting the abradable rim seal thereon.

10. The turbine assembly as defined in claim 9 wherein the leading edge flange comprises an outer surface for attachment with the abradable rim seal, the outer surface of the leading edge flange having a diameter smaller than an outer diameter of the annular inner duct of the turbine exhaust case.

11. The turbine assembly as defined in claim 9 wherein the abradable rim seal comprises a machined annular sealing surface being overlapped by the inner surface of the trailing edge platform of the rotor blades in order to allow rubbing with same during engine operation.

12. The turbine assembly as defined in claim 11 wherein the machined annular sealing surface of the abradable rim seal has a diameter smaller than a diameter defined by the inner surface of the trailing edge platform of the respective rotor blades to form a clearance therebetween, the clearance being substantially closed by allowing a rubbing contact between the trailing edge platform and the abradable rim seal due to thermal expansion of engine components during engine operation.

13. The turbine assembly as defined in claim 12 wherein the abradable rim seal comprises an external peripheral surface having a diameter substantially equal to the outer diameter of the inner duct.

14. The turbine assembly as defined in claim 7 wherein the abradable rim seal comprises a honeycomb material.

15. The turbine assembly as defined in claim 6 wherein the trailing edge platform is an integral portion of a blade platform of each rotor blade, the blade platform substantially dividing each rotor blade into a blade section and a root section.

16. A method for substantially reducing hot gas ingestion into a rotor disc rear cavity of a turbine rotor assembly of a gas turbine engine, in which a turbine exhaust case has annular outer and inner ducts located downstream of the turbine rotor assembly for ducting hot exhaust gases having passed through a plurality of turbine rotor blades, the method comprising:

attaching an abradable rim seal to a leading edge of an annular inner duct of the turbine exhaust case; and

machining the abradable rim seal to form a sealing surface thereon for being positioned under and being overlapped by an inner surface of the leading edge platform of each turbine rotor blade when assembling the gas turbine engine, thereby allowing a rubbing condition between the abradable rim seal and the leading edge platform during engine operation.

17. The method as defined in claim 16 comprising attaching an annular leading edge flange to a leading end of the inner duct of the turbine exhaust case for supporting the abradable rim seal thereon.

18. The method as defined in claim 17 wherein the abradable rim seal is made of a honeycomb material and is brazed to the leading edge flange.