U.S. patent application number 10/868781 was filed with the patent office on 2005-06-16 for rotor blade for a turbo-machine.
This patent application is currently assigned to SNECMA MOTEURS. Invention is credited to Naudet, Jacky.
Application Number | 20050129521 10/868781 |
Document ID | / |
Family ID | 33396863 |
Filed Date | 2005-06-16 |
United States Patent
Application |
20050129521 |
Kind Code |
A1 |
Naudet, Jacky |
June 16, 2005 |
Rotor blade for a turbo-machine
Abstract
The blade of gas turbine engine compressor of the invention
includes a hammer type root to be inserted into a circumferential
groove of the rotor of the compressor, a platform (12) integral
with the root (11) and supporting an airfoil portion (13), the
platform including two edges perpendicular to the axis of the rotor
(20, 21) and two curved flanks (22, 23), the curve of the flanks
being made out of at least one curve defined by an equation, the
curvature centre of the point in the curve whereof the curvature
radius is the smallest, being situated within a band (B) central to
the platform and accounting for 60% of the width (D) of the
platform (12) measured between its parallel rectilinear edges (20,
21), the equation defining the curve being, in the sense of
mathematic functions, continuous and with continuous first
derivative.
Inventors: |
Naudet, Jacky; (Bondoufle,
FR) |
Correspondence
Address: |
OBLON, SPIVAK, MCCLELLAND, MAIER & NEUSTADT, P.C.
1940 DUKE STREET
ALEXANDRIA
VA
22314
US
|
Assignee: |
SNECMA MOTEURS
PARIS
FR
|
Family ID: |
33396863 |
Appl. No.: |
10/868781 |
Filed: |
June 17, 2004 |
Current U.S.
Class: |
416/215 |
Current CPC
Class: |
Y02T 50/673 20130101;
F05D 2240/80 20130101; Y02T 50/60 20130101; F01D 5/143
20130101 |
Class at
Publication: |
416/215 |
International
Class: |
F01D 005/30 |
Foreign Application Data
Date |
Code |
Application Number |
Jun 27, 2003 |
FR |
03 07809 |
Claims
1- A blade (10) of a turbo-machine rotor, including a hammer type
root (11) to be inserted into a circumferential groove (14) of the
rotor (15), a platform (12) integral with the root (11) and
supporting an airfoil portion (13), the platform including two
edges perpendicular to the axis of the rotor (20, 21) and two
curved flanks (22, 23), characterised in that the curve of the
flanks is made out of at least one curve defined by an equation,
the curvature centre of the point in the curve whereof the
curvature radius is the smallest, being situated within a band (B)
central to the platform and accounting for 60% of the width (D) of
the platform (12) measured between its parallel rectilinear edges
(20, 21), the equation defining the curve being, in the sense of
mathematic functions, continuous and with continuous first
derivative.
2- A blade (10) according to claim 1, wherein the curve is made out
of an assembly of tangent circles.
3- A blade (10) according to claim 1, wherein the curve is defined
by a curve such as a spiral, an epicycloids or a circle
involute.
4- A blade according to claim 1, which is a rotor blade of a gas
turbine engine compressor.
Description
[0001] The present invention concerns the fastening of a blade to a
turbo-machine rotor and in particular a rotor blade of an axial
compressor for a gas turbine engine.
[0002] In turbo-jet engines, the high pressure stages of
compressors include generally a large number of blades mounted in a
circumferential groove of the rotor. The blades includes a root
portion, whereon is attached a platform supporting an airfoil
portion. The blades are so-called hammer root type blades with a
shape matching that of the circumferential groove of the rotor,
which exhibits flanks forming a back-up surface in centrifugal
radial direction.
[0003] As can be seen on FIG. 1, in compressors of the previous
art, the platforms 2 were rectangular, their flanks 3 parallel to
the axis 1 of the rotor.
[0004] The improved throughput of compressors has caused a
reduction in the pitches of the blades and an increase in that
tilting angle relative to the axis of the engine. It has therefore
become necessary to tilt the flanks of the platforms, in order to
accommodate a larger number of blades, as can be seen on FIG.
2.
[0005] Because of the loads perpendicular to the rotational axis 1,
as for example the inertia loads and aerodynamic loads exerted
thereon, the blades are caused to pivot round the longitudinal axis
of the airfoil portion 4. The platforms 2 slide with respect to one
another in order to adopt a position as represented on FIG. 3. The
platforms 2 then tend to pile up according to the shortest
circumferential dimension, i.e. along a width l', with respect to
the plane of the rotor 5, smaller than the initial width l. In
other words, in duty, the platforms 2 appear with their shortest
width relative to the plane of the rotor 5.
[0006] This sliding of the platforms 2 is allowed by the clearance
existing between the roots of the blades and their housing as well
as between the platforms 2 and their housing.
[0007] This sliding suffers from numerous shortcomings:
[0008] Since the size l' is smaller than the size l, it induces
significant clearances at the platforms, which cause leaks.
[0009] It promotes the rotation of the blades in the direction of
increase in the setting angle of the airfoil portion 4, which is
detrimental to the throughput of the compressor.
[0010] The roots do not rest correctly in their housing on the
surfaces designed to that effect, which translates in surface
hammering and an increase in the local load levels in the disc and
the blade root.
[0011] It can also be noted during the operation of adjusting the
length of the end 2 of the airfoil portion, that the centrifugal
load is not large enough to bring the blades back to their correct
position. At low speed, the blades pivot and lock in the wrong
position by friction, and cannot resume their correct position,
even at higher rotational speed.
[0012] One has therefore attempted to confer to the flanks of the
platforms, such a profile that for the same rotation of each blade
caused by a tangential load, they do slip over one another and such
that the contact loads oppose the rotation.
[0013] The American patent U.S. Pat. No. 4,878,811 provides
platforms whereof the flanks include two rectilinear portions,
parallel to the rotational axis and offset, connected by an oblique
portion. The purpose of this solution is to reduce the rotation of
the airfoil portion and to avoid the leaks between the platforms by
limiting the slippage of the platforms with respect to one another.
It involves, however, uneasy machining of the platforms, since each
flank entails several machining entities.
[0014] The present invention intends to remedy these
shortcomings.
[0015] To this effect, the invention concerns a rotor blade of a
turbo-machine, including a root inserted in an longitudinal annular
groove of the rotor, a platform integral with the root and
supporting a airfoil portion, the platform including two
longitudinal edges and two bent-in flanks forming a curve,
characterised in that the curve is made out of at least one curve
defined by an equation, the curvature centre of the point in the
curve whereof the curvature radius is the smallest being situated
inside a band central to the platform and accounting for 60% of the
width of the platform measured between its parallel rectilinear
edges, the equation defining the curve being, in the sense of
mathematic functions, continuous and with continuous first
derivative.
[0016] Thanks to this definition of the shape of the curved flanks
of the platforms, if the blades are not placed correctly, they
resume their right position naturally as the rotor rotates.
Moreover, the flanks of the platforms can be machined in a single
machining entity.
[0017] The invention relates in particular to a rotor blade for a
gas turbine engine compressor, but the applicant does not intend to
limit the extent of its rights to that application.
[0018] The present invention will be understood better using the
following description of the preferred embodiment of the blade
according to the invention, with reference to the appended
drawings, whereon:
[0019] FIG. 1 represents a schematic view from beneath of blades
with flanks parallel to the rotational axis of the compressor of
the previous art;
[0020] FIG. 2 represents a schematic view from above of blades with
flanks parallel relative to the rotational axis of the compressor
of the previous art;
[0021] FIG. 3 presents a schematic view from above of the sliding
of the blades with parallel flanks tilted relative to the
rotational axis of the compressor of the previous art;
[0022] FIG. 4 represents a schematic lateral view of the blade
according to the invention;
[0023] FIG. 5 represents a schematic view from beneath of three
blades of the invention;
[0024] FIG. 6 represents a schematic view from above of three
blades of the invention after rotation along the longitudinal axis
of the airfoil portion, and
[0025] FIG. 7 represents a schematic view from above of three
blades according to the invention after sliding along their
flanks.
[0026] With reference to FIG. 4, the blade 10 of the invention
comprises a root 11, so-called hammer root type, because of its
oblong base tapering upward, integral with a platform 12 supporting
a blade 13.
[0027] The root 11 is inserted into an annular groove 14 of the
rotor 15 of the compressor, its upper surface 11' resting against
the internal wall of the groove when the rotor 15 is rotating,
because of the centrifugal forces.
[0028] The lower portion 16 of the platform 12, of width smaller
than that of its upper section 17, supporting the blade 13, rests
laterally against a rim 18 of the rotor 15, with a clearance
enabling, on the one hand, the assembly of the blades 10 in the
groove 14, on the other hand, the elevation of the blade 10 until
the upper surface 11' of the root 11 contacts the internal wall of
the groove 14 when the rotor 15 rotates.
[0029] It is the duty of the man of the art to define the geometry
of the root 11 and of the airfoil portion 13 of the blade 10, the
invention residing in the form of the platform 12.
[0030] With reference to FIG. 5, the platform 12 of the invention
comprises, as a planar view from above, two rectilinear transversal
edges 20, 21 perpendicular to the axis of the rotor. It also
includes flaks 22, 23 connecting both edges, which are curvilinear
in shape.
[0031] One of the objects of the invention is to be able to machine
the flanks 22, 23 of the platform 12 without changing the angle of
attack of the milling cutter, i.e. using a single machining entity.
Thus, and in the perspective according to which the blades 12
should not pivot along the longitudinal axis 6 of the blades 13,
the curve delineating the flanks 22, 23 of the platform 12 of the
invention meets certain conditions.
[0032] Thus, the curve delineating the flanks 22, 23 of the
platform 12 must be built from a curve defined by an equation, or a
set of curves defined by equations, with the following condition:
the curvature centre of the most bent-in portion of the curve, i.e.
the curvature centre corresponding to the smallest curvature
radius, must be contained within the band B central to the platform
12 accounting for 60% of the width D of the platform 12, measured
between its rectilinear parallel edges 20, 21. Moreover, the curve
must be, in the sense of mathematic functions, continuous and with
continuous first derivative.
[0033] In particular, the curve delineating the flanks 22, 23 of
the platform may be defined by an assembly of tangent circles,
whereas the centre of the circle with the smallest radius should
lie within the band B defined above.
[0034] The curve may also, for exemplification purposes, be defined
using curves such as spirals, epicycloids or circle involutes.
[0035] With reference to FIG. 6, where the three platforms 12 of
FIG. 5 have been simulated after pivoting around the longitudinal
axis 6 of their blade 13, it can be seen that the width L' then
exhibited by the platforms 12 relative to the plane of the blade
assembly 5, is greater than the length L exhibited by the blades 12
when they are arranged correctly, i.e. with the flanks 22, 23 of
the neighbouring platforms adjacent, and their edges 20, 21
co-linear. The width L' being greater than the width L, and
according to what has been said in the preamble, the platform will
tend to resume their correct position, shown on FIG. 5, when the
rotor rotates.
[0036] With reference to FIG. 7, where the three platforms 12 of
FIG. 5 have been simulated, after sliding relative to one another
along their adjacent flanks 22, 23, i.e. by offsetting their edges
20, 21 without keeping them aligned among neighbours, it can be
seen that the width L" then exhibited by the platforms 12, with
respect to the plane of the blade assembly 5, is greater than the
length L exhibited by the blades when they are arranged correctly.
Similarly, in such a case, the blades will therefore tend to resume
their correct position when the rotor rotates.
[0037] The arrangement of the blades 10 of the invention around the
rotor 12 is conventional, since the blades 10 are inserted one by
one into the groove 14, and blocked circumferentially by a certain
number of locks.
[0038] The lower portion 16 of the platform 12 of the invention is
adjacent to its upper portion 17 at the flanks 22, 23, and of
smaller width at its upper portion 17 at the edges 20, 21.
* * * * *