US20080050238A1

US20080050238A1 - Disc firtree slot with truncation for blade attachment

Info

Publication number: US20080050238A1
Application number: US11/508,999
Authority: US
Inventors: Eric Durocher; Minh Loc Tran; Othmane Leghzaouni
Original assignee: Pratt and Whitney Canada Corp
Current assignee: Pratt and Whitney Canada Corp
Priority date: 2006-08-24
Filing date: 2006-08-24
Publication date: 2008-02-28
Also published as: CA2595849A1

Abstract

A turbine rotor disc includes a plurality of attachment slots for receiving firtree-profiled blade roots. Each attachment slot is formed in an undulating profile having substantially smoothly curved and laterally, alternately recessed and projecting portions. At least one pair of substantially smoothly curved projecting portions being truncated at a tip thereof, respectively.

Description

TECHNICAL FIELD

The present invention relates generally to gas turbine engines, and more particularly to an improved blade root retaining system for attachment of a turbine blade to a turbine disc of a gas turbine engine.

BACKGROUND OF THE ART

A conventional gas turbine engine includes various rotor blades in the fan, compressor and turbine sections thereof, which are removably mounted to respective rotor discs. Each of the rotor blades includes a blade root at the radially innermost end thereof. Each of the blade roots conventionally includes one or more pairs of lobes which slide axially into and be retained in one of a plurality of axially extending attachment slots in the periphery of the rotor disc. The disc and blade fixings of a rotor assembly of gas turbine engines, particularly of the high pressure turbine rotor assembly, conventionally requires a complicated undulating or firtree profile in order to meet the requirements of engine performance, weight reduction, secondary air consumption, disc/blade life considerations, etc. Nevertheless, the undulating or firtree profile of the disc and blade fittings, particularly the profiled attachment slots of rotor discs, present a challenge in the manufacturability thereof. Efforts have been made in disc and blade fitting configurations, to provide an optimized compromise among engine performance, weight, secondary air consumption, manufacturability, manufacturing costs and disc/blade life expectancy.
Accordingly, there is a need to provide an improved disc and blade fixing configuration of rotor assemblies of gas turbine engines.

SUMMARY OF THE INVENTION

It is therefore an object of the present invention to provide an improved disc and blade fitting configuration of rotor assemblies for gas turbine engines.
In one aspect, the present invention provides a disc of a turbine rotor adapted to support a plurality of blades attached thereto, which comprises a plurality of attachment slots for receiving a root of the respective blades, the attachment slots being circumferentially spaced apart one from another and axially extending through a periphery of the disc, each slot including a pair of opposed side walls, each side wall being in an undulating profile having substantially smoothly curved and laterally, alternately recessed and projecting portions extending along a length of the respective slots, the recessed and projecting portions of one side wall substantially and circumferentially aligning with the respective recessed and projecting portions of the other side wall to thereby provide a profiled space defined between the opposed side walls, at least one pair of the circumferentially aligning and substantially smoothly curved projecting portions being truncated at a tip thereof, respectively.
In another aspect, the present invention provides a rotor assembly of a gas turbine engine, which comprises a rotor disc defining a plurality of attachment slots circumferentially spaced apart one from another and extending axially through a periphery thereof; an array of rotor blades extending radially outwardly from the periphery of the rotor disc, each of the rotor blades including an airfoil section, a blade root and platform segments extending laterally from sides of the airfoil section into opposing relationship with corresponding platform segments of adjacent blades, each of the blade roots including a series of smoothly curved lateral projections in pairs on opposite sides thereof extending along an axial length of the blade root to form a firtree profile; and wherein each slot of the rotor disc includes a pair of opposed side walls having substantially smoothly curved and laterally, alternately recessed and projecting portions extending along a length of the respective slots, the recessed and projecting portions of one side wall substantially and circumferentially aligning with the recessed and projecting portions of the other side wall to thereby provide a profiled space defined between the opposed side walls substantially in accordance with the firtree profile of the respective blade roots, at least one pair of the circumferentially aligning and substantially smoothly curved projecting portions being truncated at a tip thereof, respectively, thereby creating a small clearance between the truncated tip of the respective projecting portions of the attachment slot and the blade root of a rotor blade attached thereto.
In a further aspect, the present invention provides a disc for a gas turbine engine comprising a plurality of firtree slots provided through the disc around a periphery of the disc, the slots defined by a plurality of opposed lobes pairs extending through the disc to define sidewalls of the slot, at least one opposed lobe pair each having opposed rounded apexes, and at least one opposed lobe pair each having opposed truncated apexes.
Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below.

DESCRIPTION OF THE DRAWINGS

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

FIG. 1 is a schematic cross-sectional view of a turbofan gas turbine engine, as an example illustrating an application of the present invention;

FIG. 2 is a schematic partial cross-sectional view of a turbine rotor assembly of the engine of FIG. 1;

FIG. 3 is a partial rear side elevational view of the turbine rotor assembly of FIG. 2, showing an undulating or firtree profile of a disc and blade fitting configuration incorporating one embodiment of the present invention;

FIG. 4 is a partial cross-sectional view of the disc of the turbine rotor assembly of FIG. 3, showing the details of the attachment slots of the disc; and

FIG. 5 is a view similar to that of FIG. 3, showing an undulating or firtree profile of a disc and blade fitting configuration incorporating another embodiment of the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring to FIG. 1, a turbofan gas turbine engine, presented as an example of the application of the present invention, includes a housing or a 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. The compressor assemblies 16, 22 drive a main air flow (not indicated) along the main fluid path and provide bleed air flow as a cooling air source for cooling the combustor 28 and the turbine assemblies 18 and 24.
Referring to FIG. 1-5, a rotor assembly, for example a turbine rotor assembly 30 in one rotor state of the high pressure turbine assembly 24, is described herein according to one embodiment of the present invention. The turbine rotor assembly 30 includes a turbine rotor disc 32 mounted on a rotating shaft (not indicated) of the high pressure spool assembly 20 and is rotatable about a longitudinal axis 29 of the engine, which is also the longitudinal axis of the turbine rotor assembly 30. An array of rotor blades 34 (only one shown in FIG. 2) extend radially outwardly from the periphery of the turbine rotor disc 32. Each of the rotor blades 34 includes an airfoil section 36, a root section 38 and platform segments 40 extending laterally from opposed sides of the airfoil section 36 into opposing relationship with corresponding platform segments 40 of adjacent rotor blades 34.
The rotor assembly 30 will now be described in greater detail with reference, in particular, to FIGS. 2-5. The root section 38 of each turbine rotor blade 34 includes a series of smoothly curved lateral projections preferably referred to as lobes 42, 44 and 46 in pairs on opposite sides thereof, extending along the axial length of the blade root 38. The pairs of lobes 42, 44 and 46 have circumferential widths decreasing from the radially outermost lobes 42 (“top lobe”), to the radially innermost lobes 46 (“bottom lobe”), with the radially central lobes 44 (“mid lobes”) disposed therebetween having an intermediate lobe width. The root section 38 of such a multi-lobed type is often referred to as a firtree, because of this characteristic shape. Although three pairs of lobes are illustrated as an example of this invention, the number of lobes may vary in different embodiments.
The platform segments 40 of turbine rotor blades 34, in combination form an inner section of an inner annular wall of the main fluid path of the engine, as shown in FIG. 1. The platform segments 40 of the turbine rotor blades 34 are preferably shaped to provide a flared gas path in order to achieve high levels of efficiency in engine performance.
The turbine rotor disc 32 includes a web section 33 extending radially outwardly from a hub (not shown) which is mounted to the rotating shaft (not indicated) of the high pressure spool assembly 20 of FIG. 1, and a rim section 50 extending radially outwardly from the web section 33. Rim section 50 has an axial thickness defined by respective front and rear sides thereof (not indicated), and also defines an outer periphery 55.
The turbine rotor disc 32 further includes a plurality of attachment slots 48 (only one shown in FIGS. 3 and 4), circumferentially spaced apart one from another and axially extending through the periphery 55 of the turbine rotor disc 32 which in this embodiment, is the entire axial thickness of the rim section 50. Each of the axial attachment slots 48 includes a pair of opposed side walls (not indicated) each being defined in an undulating profile having substantially smoothly curved and laterally, alternately recessed and projecting portions 42 a, 44 a, 46 a and 52, 54, 56. The recessed and projecting portions 42 a, 44 a, 46 a and 52, 54, 56 of one side wall substantially and circumferentially align with the respective recessed and projecting portions 42 a, 44 a, 46 a and 52, 54, 56 of the other side wall to thereby provide a profiled space defined between the opposed side walls, substantially in accordance with the firtree profile of the root section 38 of the respective turbine rotor blades 34. The axial attachment slot 48 is thus substantially complimentary in both shape and size to the firtree profile of the root sections 38 of a turbine rotor blade 34, so as to form abutting retaining surfaces of the respective root sections and attachment slot 48 for radially retaining blade 34 in the turbine rotor assembly 30 against centrifugal forces represented by arrow 58 (see FIG. 3) cause by high speed rotation of the turbine rotor assembly 30. Radial retaining forces represented by arrows 60 (see FIG. 3) occur between the abutting retaining surfaces which extend substantially along both axial lengths of the turbine rotor blade 34 and the axial thickness of the rim section 50 of the turbine rotor disc 32.
It should be noted that the firtree-profiled blade root section 38 and attachment slot 48 are sized so as to provide a desired radial play therebetween such that the firtree of root section 38 fits loosely into attachment slot 48 to allow the rotor blade 34 to self-adjust in position under the centrifugal forces 58 during operation, in order to significantly reduce or eliminate stresses on the root section 38 caused by inappropriate attachment. Therefore, during operation, there are gaps between the root section 38 of the rotor blade 34 and the rotor disc 32. In particular, there are gaps 62 between the bottom surface of top, mid and bottom lobes 42, 44, 46 and the respective adjacent surfaces of recessed portions 42 a, 44 a, 46 a.
According to one embodiment the present invention, one pair of the circumferentially aligning and substantially smoothly curved projecting portions, for example, the radially innermost projecting portions 56, are truncated at a tip thereof, respectively, as illustrated in FIG. 5.
FIGS. 3 and 4 illustrate another embodiment of the present invention in which each pair of the circumferentially aligning and substantially smoothly curved projecting portions 52, 54, 56 are truncated at a tip thereof, respectively. (The broken lines in FIG. 4 show the conventional tip of the projecting portions 52, 54 and 56.) Thereby, there is a small clearance between the truncated tip of the respective projecting portions 52, 54, 56 of the attachment slot 48 and the blade root 38 of the rotor blade 34 attached thereto (see FIG. 3).
Each of the truncated projecting portions 52, 54 or 56 preferably defines a small flat surface 52 a, 54 a or 56 a (at the tip thereof), which extends across the entire width of the projecting portion. The small flat surface 52 a, 54 a or 56 a is preferably smoothly connected to an adjacent surface of the truncated projecting portion 52, 54 or 56 through a curved transitional surface 64 or 66 at each side of the small flat surface. (The curved transitional surfaces 64, 66 are only indicated in relation to the small flat surface 56 a, but also exist in relation to the respective small flat surfaces 52 a and 54 a.)
The firtree or undulating profile with truncation of the attachment slot configuration of the turbine rotor disc, advantageously reduces difficulties in the manufacturing of turbine rotor discs, particularly in the formation of the attachment slots thereof for high volume engines. Truncation of the attachment slots provides the possibility of designing better and stronger slot cutting tools used in the disc slot machining process, thereby reducing cutting tool wear and/or risk of tool breakage. Less wear and breakage of cutting tools result in cost savings in the production phase of engine manufacturing.
FIGS. 3-5 are schematic illustrations and as such, the truncations shown are exaggerated and do not represent the proportional dimensions of the slot configuration. It is preferable to determine the dimensions of truncation as an optimum compromise among engine performance, cost, weight, secondary air consumption and manufacturability that meet disc/blade life requirements. In general, the truncation is small enough to not significantly affect the undulating or firtree profile of the attachment slots of the rotor disc and thus not to significantly affect the load stress distribution between abutting surfaces of the respective disc slots and blade roots during engine operation. The orientation of the small flat surface defined by the truncation may vary within a limited angular range, depending on consideration of a better slot cutting tool design. It should also be noted that in consideration of a better slot cutting tool design, the truncation at the tip of the pair of radially innermost projecting portions 56 of the slot 48 is preferable if only one pair of projections are selected to be truncated, as illustrated in FIG. 5.
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, although a turbofan gas turbine engine is taken as an example to illustrate an application of the present invention, this invention is applicable to gas turbine engines of other types. 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. A disc of a gas turbine rotor adapted to support a plurality of blades attached thereto, the disc comprising a plurality of attachment slots for receiving a root of the respective blades, the attachment slots being circumferentially spaced apart one from another and axially extending through a periphery of the disc, each slot including a pair of opposed side walls, each side wall being in an undulating profile having substantially smoothly curved and laterally, alternately recessed and projecting portions extending along a length of the respective slots, the recessed and projecting portions of one side wall substantially and circumferentially aligning with the respective recessed and projecting portions of the other side wall to thereby provide a profiled space defined between the opposed side walls, at least one pair of the circumferentially aligning and substantially smoothly curved projecting portions being truncated at a tip thereof, respectively.

2. The disc as defined in claim 1 wherein the truncated pair of projecting portions each define a flat surface at the tip thereof extending along an entire width of the respective projecting portions.

3. The disc as defined in claim 2 wherein the flat surface is smoothly connected to an adjacent surface of each truncated projecting portion through a curved transitional surface at each side of the flat surface.

4. The disc as defined in claim 1 wherein each pair of the circumferentially aligning and substantially smoothly curved projecting portions are truncated at a tip thereof, respectively.

5. A rotor assembly of a gas turbine engine, comprising:

a rotor disc defining a plurality of attachment slots circumferentially spaced apart one from another and extending axially through a periphery thereof;

an array of rotor blades extending radially outwardly from the periphery of the rotor disc, each of the rotor blades including an airfoil section, a blade root and platform segments extending laterally from sides of the airfoil section into opposing relationship with corresponding platform segments of adjacent blades, each of the blade roots including a series of smoothly curved lateral projections in pairs on opposite sides thereof extending along an axial length of the blade root to form a firtree profile; and

wherein each slot of the rotor disc includes a pair of opposed side walls having substantially smoothly curved and laterally, alternately recessed and projecting portions extending along a length of the respective slots, the recessed and projecting portions of one side wall substantially and circumferentially aligning with the recessed and projecting portions of the other side wall to thereby provide a profiled space defined between the opposed side walls substantially in accordance with the firtree profile of the respective blade roots, at least one pair of the circumferentially aligning and substantially smoothly curved projecting portions being truncated at a tip thereof, respectively, thereby creating a small clearance between the truncated tip of the respective projecting portions of the attachment slot and the blade root of a rotor blade attached thereto.

6. The rotor assembly as defined in claim 5 wherein the firtree-profiled blade roots and attachment slots are sized so as to provide a desired radial play therebetween.

7. The rotor assembly as defined in claim 5 wherein the truncated pair of projecting portions define a flat surface at the tip thereof extending along an entire width of the respective projecting portion.

8. The rotor assembly as defined in claim 7 wherein the flat surface is smoothly connected to an adjacent surface of the truncated projecting portion through a curved transitional surface at each side of the small flat surface.

9. The rotor assembly as defined in claim 5 wherein each pair of the circumferentially aligning and substantially smoothly curved projecting portions are truncated at a tip thereof, respectively.

10. A disc for a gas turbine engine comprising a plurality of firtree slots provided through the disc around a periphery of the disc, the slots defined by a plurality of opposed lobes pairs extending through the disc to define sidewalls of the slot, at least one opposed lobe pair each having opposed rounded apexes, and at least one opposed lobe pair each having opposed truncated apexes.