FR3096724A1 - TURBINE MODULE FOR AN AIRCRAFT TURBOMACHINE - Google Patents

Abstract

Turbine module (50) for an aircraft turbomachine (10), this module comprising: - arms (52) for connecting annular walls (50a, 50b), 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 stator vanes (54a) which have upstream leading edges (54b) located in a plane P2 perpendicular to said axis and located downstream of the plane P1, and downstream of the 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 the walls, and the arms extending downstream and each being intimately linked with one of the vanes. Figure for abstract: Figure 4

Description

TURBINE MODULE FOR AN AIRCRAFT TURBOMACHINE

Technical field of the invention

The present invention relates to a turbine module for an aircraft turbine engine.

Technical background

An aircraft turbomachine, for example of 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 gases, to the at least one compressor, one 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 casing and the impeller comprises a disc carrying blades on 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 tips of the blades and the casing and therefore to ensure that a maximum of the combustion gases leaving the chamber passes through the impeller to optimize the performance of the turbomachine.

It is common 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, respectively internal and external, extending one inside on the other and connected together by radial arms. The walls define between them an annular stream for the passage of the combustion gases to the distributor, this stream being crossed by the radial arms.

This type of annular element can for example be used to support roller bearings for guiding in rotation a drive shaft of the turbine wheels, and/or to pass easements.

In the current technique, the annular element is produced independently of the distributor. The distributor comprises an annular row of blades 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 devices are also provided between the downstream ends of the walls and the upstream ends of the platforms to limit gas leaks outside the section in operation.

Furthermore, this prior art has several aerodynamic drawbacks, among which we can cite:

• the losses induced by the presence of the arms and their wakes impacting the distributor;
• losses related to seam stalls at the interfaces between the walls and the platforms.

The present invention proposes an improvement to this prior art.

The present invention relates to a turbine module for an aircraft turbine engine, 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 of 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 with respect 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 the 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 one piece, the vanes extending substantially radially between the said walls and being connected to these walls, and in that the arms extend downstream and are each intimately connected with one of said blades.

The production of the module in one piece makes it possible to simplify its design and its manufacture, this production being preferably 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.

In addition, the integration of the arms and the distributor makes it possible to reduce the losses mentioned above. 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 linked to the transport of the wakes in the downstream distributor.

The module according to the invention may comprise one or more of the following characteristics, taken separately from each other or in combination with each other:

- 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 this arm is associated with; the skeleton of an airfoil 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 α with respect to said axis, and each of the blades has an orientation inclined at an angle β with respect to said axis, β possibly being different from α,

- β is equal to or greater than α,

- the angles α and β are adapted to the angle of incidence of the flow; in the present application, the flow or the main flow corresponds to the flow of gas which flows from upstream to downstream; the thickness can be substantially parallel to this flow,

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

The present invention also relates to an aircraft turbine engine, 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 figures

Other characteristics and advantages of the invention will appear during the reading of the detailed description which will follow for the understanding of which reference will be made to the appended drawings in which:

FIG. 1 is a diagrammatic half view in axial section of part of an aircraft turbine engine,

FIG. 2 is a very schematic cross-sectional view from above of a distributor arm and vanes, according to a technique prior to the invention

Figure 3 is a very schematic view in axial and side section of the arm and the blades of Figure 2,

FIG. 4 is a very schematic cross-sectional view from above of a distributor arm and vanes, according to one embodiment of the module of the invention,

Figure 5 is a very schematic view in axial and side section of the arm and the blades of Figure 4.

Detailed description of the invention

FIG. 1 represents 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 at least one least 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 of 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 combustion chamber outlet 16 and includes a stator distributor 28 and a rotor wheel 26. A low pressure turbine stage 18b is located downstream of the 18a and also includes a distributor 28 and a rotor wheel 26.

A turbine nozzle includes an annular row of stationary 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 attachment to sealing ring support flanges 36, 38.

A casing 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 housing 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 the stage 18b. This crown 34 includes a flange 34a for fixing to the flanges 32a, 24b.

The wheels 26 are interconnected by a shaft 40 which is also connected to the impeller of the centrifugal compressor 14. The shaft 40 is guided in rotation around an axis A by roller 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, respectively internal 42a and external 42b, interconnected by an annular row of arms 44 extending substantially radially with respect to the axis A of rotation of the shaft 40. The arms 44 are tubular and can be used to pass services 46 such as fluid pipes or electrical cables.

The bearing support 42 is mounted inside the casing 32 and carries a bearing housing which comprises a crown 48 for supporting the outer rings 41a of the bearings 41. The bearings 41 are here two in number, an upstream roller bearing and a ball bearing, the inner rings 41b of which are directly mounted on the shaft 40

Figures 2 and 3 very schematically show the current state of the art in terms of manufacturing and assembling the bearing support 42, the arm 44 and the distributor 28. As mentioned in the foregoing, it can be seen that these parts are manufactured independently of each other.

The distributor 28 comprises an annular row of blades 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 aiding sealing can 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 vein in functioning.

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 plane P1. These blades comprise downstream trailing edges located in a plane P3 perpendicular to the axis A and located downstream of the plane P2.

It can be seen in FIG. 2 that the plane P2 is downstream and separated 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 hooks vein at the interfaces between the walls 42a, 42b and the platforms 28b, 28c, as well as the mixing of intergrid leaks at the hub and casing.

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

Unlike the prior art, the module 50 is formed in one piece. The vanes 54a here extend 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, N2 in number, namely those located directly downstream of the arms in relation to the direction of gas flow in the stream (figure 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 the normals to the gas flow in the vein, as mentioned above.

In addition, each arm 52 has a general orientation inclined by an angle α relative to the axis A. The blades 54a each have an orientation inclined by an angle β relative to the axis A. The angles α and β are adapted to the angle of incidence of the upstream flow. β is equal to or greater than α.

The annular walls 50a, 50b extend continuously from the upstream ends located upstream of the plane P1, to the downstream ends located downstream of the plane P3 (FIG. 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 to 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 wake downstream 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 from a single piece. It is thus possible to aerodynamically design these two profiles in one piece, thus reducing the overall aerodynamic losses per wake. In addition, the creation of a single and continuous vein eliminates the effects of stalling or walking when assembling the parts, which reduces the pressure drops associated both with this accidental shape for the flow of the vein, and the mixing of the leak with the seam air. All problematic assembly issues are thus resolved by additive manufacturing, which produces a one-piece part.

In a particular embodiment of the invention, the numerical 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 (about 25 to 30% in the example shown) its mass compared to the prior art.

Additive manufacturing makes it possible to achieve these manufacturing and optimization objectives.

Claims (8)

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) 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 at least some of these blades comprising upstream leading edges (54b) situated in a plane P2 perpendicular to said axis and situated downstream of plane P1, and downstream of trailing edges (54c) situated in a plane P3 perpendicular to said axis and located downstream of said plane P2, characterized in that this module is formed in one 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.

A module (50) according to claim 1, wherein each of the arms (52) comprises an airfoil in cross-section and has a thickness normal to the skeleton which varies continuously and regularly from the leading edge (52a) of the arm to 'at the trailing edge (54c) of the blade (54a) with which this arm is associated. Module (50) according to Claim 2, in which 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. Module (50) according to one of the preceding claims, in which each of the arms (52) has a general orientation inclined at an angle α with respect to said axis (A), and each of the vanes (54a) has an inclined orientation d an angle β with respect to said axis. A module (50) according to claim 4, wherein β is equal to or greater than α. Module (50) according to one of the preceding claims, in which the said annular walls (50a, 50b) extend continuously from upstream ends located upstream of the plane P1, to downstream ends located downstream of the plane P3 . Aircraft turbomachine, comprising at least one module (50) according to one of the preceding claims. Process for manufacturing a module (50) according to one of Claims 1 to 6, characterized in that it is obtained by additive manufacturing.