U.S. patent application number 11/508999 was filed with the patent office on 2008-02-28 for disc firtree slot with truncation for blade attachment.
This patent application is currently assigned to Pratt & Whitney Canada Corp.. Invention is credited to Eric Durocher, Othmane Leghzaouni, Minh Loc Tran.
Application Number | 20080050238 11/508999 |
Document ID | / |
Family ID | 39113640 |
Filed Date | 2008-02-28 |
United States Patent
Application |
20080050238 |
Kind Code |
A1 |
Durocher; Eric ; et
al. |
February 28, 2008 |
Disc firtree slot with truncation for blade attachment
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.
Inventors: |
Durocher; Eric; (Vercheres,
CA) ; Tran; Minh Loc; (Mont-Royal, CA) ;
Leghzaouni; Othmane; (Longueull, CA) |
Correspondence
Address: |
OGILVY RENAULT LLP (PWC)
1981 MCGILL COLLEGE AVENUE, SUITE 1600
MONTREAL
QC
H3A 2Y3
US
|
Assignee: |
Pratt & Whitney Canada
Corp.
|
Family ID: |
39113640 |
Appl. No.: |
11/508999 |
Filed: |
August 24, 2006 |
Current U.S.
Class: |
416/219R |
Current CPC
Class: |
F01D 5/3007
20130101 |
Class at
Publication: |
416/219.R |
International
Class: |
F01D 5/30 20060101
F01D005/30 |
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.
Description
TECHNICAL FIELD
[0001] 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
[0002] 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.
[0003] 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
[0004] 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.
[0005] 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.
[0006] 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.
[0007] 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.
[0008] 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
[0009] Reference is now made to the accompanying figures depicting
aspects of the present invention, in which:
[0010] FIG. 1 is a schematic cross-sectional view of a turbofan gas
turbine engine, as an example illustrating an application of the
present invention;
[0011] FIG. 2 is a schematic partial cross-sectional view of a
turbine rotor assembly of the engine of FIG. 1;
[0012] 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;
[0013] 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
[0014] 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
[0015] 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.
[0016] 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.
[0017] 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.
[0018] 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.
[0019] 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.
[0020] 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 42a, 44a,
46a and 52, 54, 56. The recessed and projecting portions 42a, 44a,
46a and 52, 54, 56 of one side wall substantially and
circumferentially align with the respective recessed and projecting
portions 42a, 44a, 46a 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.
[0021] 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 42a, 44a, 46a.
[0022] 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.
[0023] 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).
[0024] Each of the truncated projecting portions 52, 54 or 56
preferably defines a small flat surface 52a, 54a or 56a (at the tip
thereof), which extends across the entire width of the projecting
portion. The small flat surface 52a, 54a or 56a 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 56a, but also exist in relation to the
respective small flat surfaces 52a and 54a.)
[0025] 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.
[0026] 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.
[0027] 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.
* * * * *