WO2020240136A1

WO2020240136A1 - Turbine module for an aircraft turbomachine

Info

Publication number: WO2020240136A1
Application number: PCT/FR2020/050894
Authority: WO
Inventors: Charlie Michaël KOUPPER; Yves-Marie Pierre LE BAYON
Original assignee: Safran Helicopter Engines
Priority date: 2019-05-29
Filing date: 2020-05-27
Publication date: 2020-12-03
Also published as: FR3096724B1; FR3096724A1

Abstract

Turbine module (50) for an aircraft turbomachine (10), said module comprising: - arms (52) for connecting annular walls (50a, 50b), said arms comprising upstream leading edges (52a) which are located in a plane P1 perpendicular to an axis, - a nozzle (54) located downstream of the arms and comprising stator vanes (54a) which comprise upstream leading edges (54b) located in a plane P2 perpendicular to the axis and located downstream of the plane P1, and downstream trailing edges (54c) located in a plane P3 perpendicular to the axis and located downstream of the plane P2, the module being characterised in that it is integrally formed, the vanes extending substantially radially between the walls, and the arms extending downstream and each being closely bonded with one of the vanes.

Description

DESCRIPTION

TITLE: TURBINE MODULE FOR A TURBOMACHINE

AIRCRAFT

Technical field of the invention

The present invention relates to a turbine module for an aircraft turbomachine. Technical background

An aircraft turbomachine, for example an airplane or a helicopter, comprises an air inlet supplying a gas generator which comprises from upstream to downstream, with reference to the flow of the gases, to the minus a compressor, an annular combustion chamber, and at least one turbine.

A turbomachine turbine comprises one or more expansion stages comprising a bladed distributor forming a stator, and a bladed wheel forming a rotor. The distributor is fixed to a housing and the wheel comprises a disc carrying vanes at its periphery. The impeller rotates inside the casing and a sealing ring is provided around this impeller in order to limit the passage of gas between the tops of the blades and the casing and therefore to ensure that a maximum of the combustion gases leaving the chamber passes through the wheel to optimize the efficiency of the turbomachine.

It is common practice to provide an annular element located between two turbine stages (high and low pressure), this annular element having a structural function and comprising in particular two annular walls, internal and external respectively, extending one inside. on the other and connected together by radial arms. The walls define between them an annular passage of the combustion gases to the distributor, this passage being crossed by the radial arms.

This type of annular element can for example be used to support rolling bearings for guiding the rotation of a drive shaft of the turbine wheels, and / or to pass easements. In the current art, the annular element is produced independently of the distributor. The distributor comprises an annular row of vanes which extend radially between annular platforms, respectively internal and external. The internal platform extends in the axial extension of the internal wall of the element, and the external platform extends in the axial extension of the external wall. Sealing members are also provided between the downstream ends of the walls and the upstream ends of the platforms to limit gas leaks outside the duct in operation.

Furthermore, this prior art has several aerodynamic drawbacks, among which there may be mentioned:

- losses induced by the presence of arms and their wakes impacting the distributor;

- losses linked to seam dropouts at the interfaces between the walls and the platforms.

The present invention provides an improvement to this prior art.

Summary of the invention

The present invention relates to a turbine module for an aircraft turbomachine, this module comprising:

- two annular walls, respectively external and internal, extending one around the other and around a common axis,

an annular row of connecting arms for said walls, these arms extending substantially radially between these walls, the number of arms being equal to N1 and these arms comprising upstream leading edges which are located in a plane P1 perpendicular to said axis,

a distributor located downstream of said arms and comprising an annular row of stator vanes extending substantially radially relative to said axis, the number of these vanes being equal to N2 which is greater than N1 and at least some of these vanes comprising upstream of the leading edges located in a plane P2 perpendicular to said axis and located downstream of plane P1, and downstream of the trailing edges located in a plane P3 perpendicular to said axis and located downstream of said plane P2,

characterized in that the module is formed in a single piece, the blades extending substantially radially between said walls and being connected to these walls, and in that the arms extend downstream and are each intimately linked with one of said blades.

Making the module in one piece simplifies its design and manufacture, this realization preferably being made by additive manufacturing. It is then no longer necessary to provide fixing and / or sealing members between the walls and the distributor, which is particularly advantageous.

Furthermore, the integration of the arms and the distributor reduces the aforementioned losses. Indeed, this design naturally eliminates the wake losses associated with the arms, both the losses by mixing of the boundary layers in the downstream wakes of the arms, and also the losses associated with the transport of the wakes in the downstream distributor.

The module according to the invention may include one or more of the following characteristics, taken in isolation from one another or in combination with one another:

- each of the arms comprises in cross section an aerodynamic profile and has a thickness, preferably normal to the skeleton of the profile, which varies continuously and regularly from the leading edge of the arm to the trailing edge of the blade at which arm is associated with; the skeleton of an aerodynamic profile of an arm or a blade can be identified by a person skilled in the art,

- each of the arms has a maximum thickness Emax just downstream of the plane P1, an intermediate thickness Emoy in the plane P2 and a minimum thickness Pmin in the plane P3,

- each of the arms has a general orientation inclined at an angle a with respect to said axis, and each of the blades has an orientation inclined at an angle b with respect to said axis, b being possibly different from a,

- b is equal to or greater than a, - angles a and b are adapted to the angle of incidence of the flow; in the present application, the main flow or flow corresponds to the gas flow which flows from upstream to downstream; the thickness may be substantially parallel to this flow,

- Said annular walls extend continuously from upstream ends located upstream of plane P1, up to downstream ends located downstream of plane P3.

The present invention also relates to an aircraft turbomachine, comprising at least one module as described above

The present invention also relates to a method of manufacturing a module as described above, characterized in that it is obtained by additive manufacturing.

Brief description of the figures

Other characteristics and advantages of the invention will become apparent on reading the detailed description which follows, for the understanding of which reference is made to the appended drawings in which:

[Fig. 1] Figure 1 is a schematic half view in axial section of a part of an aircraft turbomachine,

[Fig. 2] Figure 2 is a very schematic cross-sectional and top view of an arm and distributor vanes, according to a technique prior to the invention

[Fig. 3] FIG. 3 is a very schematic view in axial section and from the side of the arm and of the blades of FIG. 2,

[Fig. 4] FIG. 4 is a very schematic view in cross section and from above of an arm and of distributor vanes, according to an embodiment of the module of the invention,

[Fig. 5] FIG. 5 is a very schematic view in axial section and from the side of the arm and of the blades of FIG. 4.

Detailed description of the invention

FIG. 1 represents a part of an aircraft turbomachine 10 such as a helicopter turbojet. The turbomachine 10 comprises from upstream to downstream, with reference to the direction of gas flow (see arrows), an air inlet 12, at least one compressor 14, here centrifugal, an annular combustion chamber 16, and minus one turbine 18.

The air which enters the engine through the air inlet 12 is compressed in the compressor 14 which is here a centrifugal compressor. The compressed air exits radially outwards and feeds the combustion chamber 16 via an annular assembly forming a rectifier 20 and a diffuser 22.

The combustion chamber 16 comprises two annular walls, respectively external 16b and internal 16a which extend one around the other and which are themselves arranged inside an external casing 24 of the combustion chamber. 16.

This casing 24 comprises at its upstream end an annular flange 24a for fixing to annular flanges of the rectifier-diffuser assembly 20-22 as well as a casing 25 of the compressor 14 and the air inlet 12.

The compressed air is mixed with fuel and then burned in the combustion chamber 16, which generates combustion gases which are then injected into the turbines 18.

A high pressure turbine stage 18a is located just downstream of the outlet of the combustion chamber 16 and includes a stator distributor 28 and a rotor wheel 26. A low pressure turbine stage 18b is located downstream of the stage 18a and also includes a distributor 28 and a rotor wheel 26.

A turbine nozzle includes an annular row of fixed gas flow straightening vanes, and a turbine wheel includes an annular row of vanes carried by a rotor disc.

The casing 24 further comprises at its downstream end an annular flange 24b for fixing to sealing ring support flanges 36, 38.

A housing 32 extends inside the wall 16a and carries at its upstream end the sealing ring 36 which extends around the wheel 26 of the stage 18a. The casing 24 has at its downstream end a flange 32a for fixing to the flange 24b. A crown 34 carries the sealing ring 38 which extends around the wheel 26 of stage 18b. This ring 34 comprises a flange 34a for fixing to the flanges 32a, 24b.

The wheels 26 are interconnected by a shaft 40 which is furthermore connected to the impeller of the centrifugal compressor 14. The shaft 40 is guided in rotation about an axis A by rolling bearings 41 which are carried by an annular support 42 interposed between the two floors 18a, 18b.

The bearing support 42 comprises two annular walls, internal 42a and external 42b respectively, interconnected by an annular row of arms 44 extending substantially radially with respect to the axis A of rotation of shaft 40. The arms 44 are tubular and can be used for the passage of easements 46 such as fluid conduits or electric cables.

The bearing support 42 is mounted inside the housing 32 and carries a bearing housing which comprises a ring 48 for supporting the outer rings 41 has bearings 41. The bearings 41 are here two in number, one upstream bearing to rollers and a ball bearing, the internal rings 41 b of which are directly mounted on the shaft 40

Figures 2 and 3 very schematically show the current state of the art in the manufacture and assembly of the bearing support 42, of the arm 44 and of the distributor 28. As mentioned in the above, it can be seen that these parts are produced independently of each other.

The distributor 28 comprises an annular row of vanes 28a which extend radially between annular platforms, respectively internal 28b and external 28c. The internal platform 28b extends in the axial extension of the internal wall 42a and the external platform 49 extends in the axial extension of the external wall 42b. Sealing members or a shape design to aid in the sealing may also be provided between the downstream ends of the walls 42a, 42b and the upstream ends of the platforms 28b, 28c to limit gas leaks outside the stream. operation.

The arms 44 comprise upstream leading edges situated in a plane P1 perpendicular to the axis A. The blades 28a comprise upstream leading edges situated in a plane P2 perpendicular to the axis A and situated downstream of the plan P1. These blades have trailing edges downstream located in a plane P3 perpendicular to the axis A and located downstream from plane P2.

It can be seen in FIG. 2 that the plane P2 is downstream and away from the trailing edges of the arms 44. Losses are induced by the presence of the arms 44 and their wakes impacting the distributor 28. Losses are also linked to the stalls vein at the interfaces between the walls 42a, 42b and the platforms 28b, 28c, as well as the mixture of intergrid leaks at the hub and casing.

An embodiment of a turbine module 50 according to the invention is shown in Figures 4 and 5.

This module 50 comprises two annular walls, respectively external 50b and internal 50a, extending one around the other and around a common axis A.

An annular row of arms 52 extends substantially radially between the walls 50a, 50b.

The number of arms 52 is equal to N1 and these arms have upstream leading edges 52a which are located in a plane P1 perpendicular to the axis A.

A distributor 54 is located between P2 and P3 and has an annular row of stator vanes 54a extending substantially radially with respect to the axis A. The number of such vanes 54a is equal to N2 which is greater than N1. Some of these blades (their number is equal to N2 - N1) have upstream leading edges 54b located in a plane P2 perpendicular to the axis A and located downstream of the plane P1. These blades have trailing edges 54b downstream located in a plane P3 perpendicular to axis A and located downstream from plane P2.

Unlike the prior art, the module 50 is formed in one piece. The blades 54a extend here substantially radially between the walls 50a, 50b and are connected to these walls.

The arms 52 extend downstream and are each intimately linked with one of the blades 54a. This concerns the other blades numbering N2, namely those which are directly located downstream of the arms with respect to the direction of flow of the gases in the stream (FIG. 4). Figure 4 shows that each of the arms 52 comprises in cross section an aerodynamic profile and has a thickness normal to the skeleton of the profile, which varies continuously and regularly from the leading edge 52a of the arm to the trailing edge. 54c of the blade 54a with which this arm is associated.

Each arm 52 has a maximum thickness Emax just downstream of the plane P1, an intermediate thickness Emoy in the plane P2 and a minimum thickness Emin in the plane P3. These thicknesses can be measured according to normals to the gas flow in the vein, as mentioned above.

In addition, each arm 52 has a general orientation inclined at an angle a with respect to the axis A. The vanes 54a each have an orientation inclined at an angle b with respect to the axis A. The angles a and b are adapted to the angle of incidence of the upstream flow b is equal to or greater than a.

The annular walls 50a, 50b extend continuously from upstream ends located upstream of plane P1, to downstream ends located downstream of plane P3 (Figure 5). It is therefore understood that these walls 50a, 50b result from the fusion of the platforms 28b, 28c and the aforementioned walls 42a, 42b of the prior art.

The integration of the arms 52 into the distributor 54 and vice versa makes it possible to eliminate the losses by wake of the prior art, that is to say the losses by mixing of the boundary layers in the downstream wake of the arms, and also the losses linked to the transport of the wake in the distributor.

By taking advantage of the possibilities offered by additive manufacturing, the structural arm and the distributor can be made in one piece. These two profiles can thus be aerodynamically designed as a single piece, thus reducing the overall aerodynamic losses per wake. In addition, the realization of a single and continuous vein, eliminates the effects of unhooking or walking in the assembly of the parts, which reduces the pressure losses associated with both this shape accident for the flow of vein, and mixing the leak with the flow air. All problematic assembly issues are thus removed by additive manufacturing which produces a single piece. In a particular embodiment of the invention, the digital simulations carried out on a target configuration indicate an isentropic efficiency gain of 0.3 points for the arm and turbine assembly.

The one-piece module also makes it possible to significantly reduce (of the order of 25 to 30% in the example shown) its mass compared to the prior art.

Additive manufacturing helps meet these manufacturing and optimization goals.

Claims

1. Turbine module (50) for an aircraft turbomachine (10), this module comprising:

- two annular walls, respectively external (50b) and internal (50a), extending one around the other and around a common axis (A),

- an annular row of arms (52) for connecting said walls, these arms extending substantially radially between these walls, the number of arms being equal to N1 and these arms comprising upstream leading edges (52a) which are located in a plane P1 perpendicular to said axis,

- a distributor (54) located downstream of said arms and comprising an annular row of stator vanes (54a) extending substantially radially with respect to said axis, the number of these vanes being equal to N2 which is greater than N1 and to at least some of these blades comprising upstream leading edges (54b) located in a plane P2 perpendicular to said axis and located downstream of plane P1, and downstream of trailing edges (54c) located in a plane P3 perpendicular to said axis and located downstream of said plane P2, characterized in that the module is formed in one piece, the vanes extending substantially radially between said walls and being connected to these walls, and in that the arms extend downstream and are each intimately linked with one of said blades.

2. Module (50) according to claim 1, wherein each of the arms (52) comprises in cross section an aerodynamic profile and has a thickness normal to the skeleton which varies continuously and regularly from the leading edge (52a) of the. arm up to the trailing edge (54c) of the blade (54a) with which this arm is associated.

3. Module (50) according to claim 2, wherein each of the arms (52) has a maximum thickness Emax just downstream of the plane P1, an intermediate thickness Emoy in the plane P2 and a minimum thickness Emin in the plane P3.

4. Module (50) according to one of the preceding claims, wherein each of the arms (52) has a general orientation inclined at an angle a by relative to said axis (A), and each of the vanes (54a) has an orientation inclined at an angle b with respect to said axis.

5. Module (50) according to claim 4, wherein b is equal to or greater than a.

6. Module (50) according to one of the preceding claims, wherein said annular walls (50a, 50b) extend continuously from upstream ends located upstream of the plane P1, to downstream ends located downstream of the plan P3.

7. Aircraft turbomachine, comprising at least one module (50) according to one of the preceding claims.

8. A method of manufacturing a module (50) according to one of claims 1 to 6, characterized in that it is obtained by additive manufacturing.