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.

Description

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.

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.

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.

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.

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.

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.

This sliding suffers from numerous shortcomings:

Since the size l′ is smaller than the size l, it induces significant clearances at the platforms, which cause leaks.

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.

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.

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.

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.

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.

The present invention intends to remedy these shortcomings.

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.

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.

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.

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:

FIG. 1 represents a schematic view from beneath of blades with flanks parallel to the rotational axis of the compressor of the previous art;

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;

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;

FIG. 4 represents a schematic lateral view of the blade according to the invention;

FIG. 5 represents a schematic view from beneath of three blades of the invention;

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

FIG. 7 represents a schematic view from above of three blades according to the invention after sliding along their flanks.

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.

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.

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.

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.

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.

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.

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.

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.

The curve may also, for exemplification purposes, be defined using curves such as spirals, epicycloids or circle involutes.

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.

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.

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.

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.

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.