Support device for at least one stator blade of an aircraft turbomachine and method of using same

Abstract

A support device (1) for at least one stator blade (22) of an aircraft turbomachine with a longitudinal axis (X) comprising an external wall (21) from which the stator blade (22) projects, the support device (1) comprising: - a housing (2) configured to radially receive a mounting foot (23) of the stator blade (22); - an upstream support foot (5) and a downstream support foot (6) connected respectively to an upstream end (3) and to a downstream end (4) of the housing (2) and configured to be mounted on the external wall (21) of the aircraft turbomachine; - the downstream support foot (6) comprising a protective portion (8) having a porosity greater than that of a connecting portion (7), the protective portion (8) comprising a plurality of pores (9) adapted to be compressed under the effect of an impact (A) exerted on the stator blade (22). Abstract figure: Figure 2

Description

Title of the invention: Support device for at least one stator blade of an aircraft turbomachine and method of using it Technical field

[0001] The present invention relates to an aircraft turbomachine and more specifically relates to a device for supporting at least one stator blade as well as its method of use.

[0002] Civil aviation has been mobilizing for several years now to make a contribution to the fight against climate change. Technological research efforts have already made it possible to significantly improve the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

[0003] This sustained research and development work relates in particular to new generations of aircraft turbomachines known as "unducted" in that they are free of a nacelle surrounding the aircraft turbomachine. An unducted turbomachine conventionally comprises, upstream, a fan configured to guide an air flow in an air stream defined from upstream to downstream by one or more compressors, a combustion chamber and one or more turbines. No fairing surrounds the fan, which makes it possible to provide a fan with a very large diameter greater than 4 m. The mass of the aircraft turbomachine is advantageously reduced and the thrust improved, making it possible to reduce fuel consumption by 20% compared to the latest generations of aircraft turbomachines.

[0004] In practice, a rectifier is mounted externally on the aircraft turbomachine, downstream of the fan, in order to straighten the air flow not admitted into the air stream and promote thrust. Like the fan, no fairing surrounds the rectifier. Such a rectifier comprises a plurality of stator blades comprising a radial blade and a mounting foot connected by a platform. The mounting foot of each stator blade is mounted in a housing which is fixed to the outer casing of the compressor(s). In the case of variable-pitch stator blades, the mounting foot is mounted to pivot radially in the housing in order to adapt the angle of attack of the stator blades during flight.

[0005] Because it is unducted, such a rectifier is more exposed to a risk of impact by a solid body, such as a bird, which in the long term could damage the parts of the aircraft turbomachine, in particular by deformation. To limit the risk, one solution would be to reinforce the housing and the compressor casing, which would however have an undesirable impact on the mass of the aircraft turbomachine. Another solution would be to reinforce only the housing, but this would transmit the forces undesirably to the external compressor casing.

[0006] The invention thus aims to protect an unducted turbomachine in the event of impact of a solid body, such as a bird, on the stator blades of an unducted rectifier. PRESENTATION OF THE INVENTION

[0007] The invention relates to a support device for at least one stator blade of an aircraft turbomachine, the aircraft turbomachine extending along a longitudinal axis oriented from upstream to downstream and comprising an external wall from which the stator blade extends projecting along a radial axis, the support device comprising: • At least one housing configured to radially receive a mounting foot of the stator blade, • An upstream support foot and a downstream support foot connected respectively to an upstream end and to a downstream end of the housing and configured to be mounted on the external wall of the aircraft turbomachine.

[0008] The invention is remarkable in that the downstream support foot comprises at least one connecting portion and at least one protective portion, the protective portion having a porosity greater than that of the connecting portion, the protective portion comprising a plurality of pores adapted to be compressed under the effect of an impact exerted on the stator blade.

[0009] The support device according to the invention advantageously makes it possible to absorb the forces experienced during an impact exerted from upstream to downstream by a solid body, such as a bird, on the stator blade, in particular at the blade. The protective portion of the downstream support foot is advantageously adapted to compact under the effect of the impact, which integrally tilts the housing and the stator blade downstream. This has the effect of limiting the forces experienced by the stator blade, but also those transmitted to the external wall of the aircraft turbomachine. The protective portion thus makes it possible to avoid deformation or even damage to the aircraft turbomachine, without impacting the mass and complexity thereof.

[0010] According to one aspect of the invention, the pores are adapted to be irreversibly compressed. The protective portion is an inexpensive fusible or sacrificial element capable of withstanding the forces inflicted by the impact in place of the stator blade. and the external wall which are expensive parts. The protective portion therefore protects the aircraft turbomachine, and is easily replaceable after impact.

[0011] According to one aspect of the invention, the pores extend parallel to each other, preferably in a direction parallel to the radial axis. This makes it possible to give the protective portion a preferential crushing direction, to promote the inclination of the stator blade downstream during an impact.

[0012] According to one aspect of the invention, the protective portion is in the form of a honeycomb structure, preferably in the form of a honeycomb. Such a structure is advantageously light, inexpensive and ensures sufficient support of the stator blade during normal operation.

[0013] According to one aspect of the invention, the protective portion is made of steel. Such a material makes it possible to ensure sufficient support for the stator blade during normal operation while being compactable during an impact.

[0014] According to one aspect of the invention, the connecting portion extends between the housing and the protective portion, preferably radially outwardly relative to the protective portion. The connecting portion thus exerts a force on the protective portion and promotes its compression.

[0015] According to one aspect of the invention, the support device defines a cavity delimited upstream by the upstream support foot and downstream by the downstream support foot, the cavity being adapted to house a device for controlling the pivot angle along the radial axis of the stator blade in the housing. The variable setting device is thus integrated in a simple and practical manner into the support device.

[0016] The invention also relates to an assembly of a support device as described above and an external wall of an aircraft turbomachine, the aircraft turbomachine extending along a longitudinal axis oriented from upstream to downstream, the stator blade being configured to extend projecting from the external wall along a radial axis, the upstream support foot of the support device being pivotally mounted relative to the external wall of the aircraft turbomachine, preferably along a pivot axis orthogonal to the radial axis extending in a plane transverse to the longitudinal axis.

[0017] Such a pivoting assembly maximizes the forces transmitted into the protective portion during an impact, and reduces the stress on the upstream support foot.

[0018] The invention also relates to an aircraft turbomachine extending along a longitudinal axis oriented from upstream to downstream and comprising: • An external wall, preferably in the form of an outer compressor casing, • An unshrouded rectifier comprising at least one stator vane extending projecting from the external wall along a radial axis, • At least one support device as described above, in which the upstream support foot and the downstream support foot are mounted on the external wall, the stator blade comprising a mounting foot mounted radially in the housing of the support device.

[0019] Preferably, the aircraft turbomachine is of the unducted type. The stator vanes of the rectifier are not protected by a fairing and are more exposed to the risk of impact from foreign external objects such as birds.

[0020] The invention also relates to a method of using an aircraft turbomachine as described above, in which the pores of the protective portion are compressed under the effect of an impact exerted on the stator blade of the aircraft turbomachine. Such a method is implemented passively, that is to say that it advantageously does not require any manual or electronic control. PRESENTATION OF THE FIGURES

[0021] The invention will be better understood on reading the description which follows, given by way of example, and referring to the following figures, given by way of non-limiting examples, in which identical references are given to similar objects.

[0022] [Fig.l] is a schematic perspective representation of an aircraft turbomachine according to one embodiment of the invention.

[0023] [Fig. 2] is a schematic representation in longitudinal half-section of the support device before an impact according to a first embodiment of the invention.

[0024] [Fig. 3] is a close-up perspective schematic representation of the protective portion of the support device of [Fig. 2].

[0025] [Fig.4] is a schematic representation in longitudinal half-section of the support device before an impact according to a second embodiment of the invention.

[0026] [Fig.5] is a schematic representation of a method of using the support device of [Fig.4] during an impact according to one embodiment of the invention.

[0027] It should be noted that the figures set out the invention in detail to implement the invention, said figures can of course be used to better define the invention where appropriate. DETAILED DESCRIPTION OF THE INVENTION

[0028] With reference to [Fig.l], the invention relates to an aircraft turbomachine 20 extending along a longitudinal axis X oriented from upstream to downstream. The aircraft turbomachine 20 conventionally comprises, upstream, a fan 27 configured to guide an air flow in an air stream 28 defined from upstream to downstream by one or more compressors, a combustion chamber and one or more turbines. The fan 27 is of the unducted type and preferably has a diameter greater than 4 m. The aircraft turbomachine 20 is preferably in the form of a turbo-propeller in which the fan 27 is driven by the turbine(s) via a reducer.

[0029] With reference to Figures 1 and 2, the aircraft turbomachine 20 also comprises a non-ducted rectifier 29 extending radially outwardly relative to the air stream 28, downstream of the fan 27. The non-ducted rectifier 29 makes it possible to straighten the air flow not admitted into the air stream 28 and promote thrust. The rectifier 29 comprises a plurality of stator blades 22 extending outwardly from an external wall 21 of the aircraft turbomachine 20, preferably along a radial axis Y. Each stator blade 22 comprises a radial blade 25 and a mounting foot 23 connected by a platform 24.

[0030] The aircraft turbomachine 20 is preferably of the unducted type, that is to say that it is free of a nacelle surrounding the aircraft turbomachine 20.

[0031] With reference to [Fig.2], the invention relates to a support device 1 for one or more stator blades 22 of the unducted rectifier 29 of the aircraft turbomachine 20, comprising: • One or more housings 2 each configured to radially receive a mounting foot 23 of a stator blade 22, • An upstream support foot 5 and a downstream support foot 6 connected respectively to an upstream end 3 and to a downstream end 4 of the housing(s) 2 and configured to be mounted on the external wall 21 of the aircraft turbomachine 20, • The downstream support foot 6 comprising a connecting portion 7 and a protective portion 8, the protective portion 8 having a porosity greater than that of the connecting portion 7, the protective portion 8 comprising pores 9 adapted to be compressed under the effect of an impact A (see [Fig.5]) exerted on the stator blade 22.

[0032] Thanks to the protective portion 8, the support device 1 according to the invention advantageously makes it possible to reduce the forces transmitted to the aircraft turbomachine 20 during an impact of a solid body, such as a bird, exerted on one or more stator blades 22 of the rectifier 29. The protective portion 8 in fact forms a fusible or sacrificial element adapted to be crushed during an impact which makes it possible to incline the stator blades 22 downstream and thus to absorb the forces transmitted to the stator blades 22 as well as to the external wall 21. The risk of deformation or even damage is advantageously reduced. The protective portion 8 does not further penalize not the mass of the aircraft turbomachine 20.

[0033] According to a preferred aspect, the support device 1 is in the form of a peripheral part of longitudinal axis X comprising a housing 2 for each stator blade 22 of the rectifier 29. According to another preferred aspect, the support device 1 is in the form of one or more angular sectors each comprising one or more housings 2 and extending in a partially peripheral or peripheral manner around the longitudinal axis X.

[0034] According to a preferred aspect, the external wall 21 on which the support device 1 is fixed is in the form of the external wall of the air stream 28, and preferably of the external casing of one or more compressors mounted in the air stream 28. In the example of [Fig. 2], the upstream support foot 5 is fixed to the external casing 21A of an upstream compressor and the downstream support foot 6 is fixed to the external casing 21B of a downstream compressor of the aircraft turbomachine 20. One or more fixing elements ensure the fixing of the support feet 5, 6 on the external wall 21, such as screws, rivets or bolts.

[0035] According to a preferred aspect illustrated in [Fig. 2], the mounting foot 23 of the stator blade 22 is pivotally mounted along the radial axis Y in the housing 2. In other words, the stator blades 22 of the rectifier 29 are of the variable pitch type, which makes it possible to modify their angle of attack depending on the flight conditions. It is specified that the stator blades 22, as opposed to rotor blades, are not mounted so as to be able to rotate along the longitudinal axis X. The only degree of freedom of the rotor blade 22 is rotation along the radial axis Y. Alternatively, the mounting foot 23 of the stator blade 22 is fixedly mounted in the housing 2.

[0036] As illustrated in [Fig. 2], the support device 1 preferably defines a cavity C delimited upstream by the upstream support foot 5 and downstream by the downstream support foot 6. The cavity C extends radially between the external wall 21 and the housing 2 and makes it possible to house a device (not shown) for controlling the pitch angle of one or more stator blades 22, namely their pivot angle along the radial axis Y in the housing 2.

[0037] In the example of [Fig.2], the downstream support foot 6 comprises a single connecting portion 7 and a single protective portion 8. Preferably, the downstream support foot 6 is made up of the connecting portion 7 and the protective portion 8. In this example, the connecting portion 7 extends between the housing 2 and the protective portion 8, preferably radially outwardly relative to the longitudinal axis X to promote the crushing of the protective portion 8.

[0038] With reference to [Fig. 3], the protective portion 8 is in the form of a porous structure comprising cavities, also called “pores” or “cells”, formed in the protective portion 8. The porosity of the protective portion 8 is defined as the ratio of the volume of the pores of the protective portion 8 to its total volume and is preferably between 70% and 95% in an intact state. The pores 9 of the protective portion 8 are preferably macroscopic, i.e. visible to the naked eye. The pores 9 of the protective portion 8 preferably have a cross-section greater than 0.5mm2, so as to allow sufficient compression during an impact, and less than 2mm2, so as to have sufficient mechanical strength to maintain the stator blade 22.

[0039] Still with reference to [Fig. 3], the protective portion 8 preferably has a regular geometric structure, in this example alveolar in the form of a honeycomb. The pores 9 of the protective portion 8 have substantially a section of the same size and the same shape, in this example hexagonal. As illustrated in [Fig. 3], the pores 9 extend parallel to each other, preferably in a direction parallel to the radial axis Y. This promotes radial rigidity to withstand the nominal operating forces and allows sudden and definitive crushing of the protective portion 8, namely along the radial axis Y or in a direction with a radial component, and allows the stator blade 22 to be inclined downstream.

[0040] Preferably, the protective portion 8 is made of a solid material, such as steel. The protective portion 8 thus forms a fusible or sacrificial element suitable for being crushed irreversibly upon impact of a solid body, such as a bird, on the stator blade 22, in particular at the blade 25. The material also allows the protective portion 8 to have sufficient mechanical strength to withstand the aerodynamic forces exerted on the stator blade 22 in normal operation by the air flow.

[0041] Preferably, the connecting portion 7 of the downstream support foot 6 comprises a solid structure as opposed to the porous structure of the protective portion 8. The connecting portion 7 is preferably made of titanium. The same applies to the upstream support foot 5. The connecting portion 7 of the downstream support foot 6 and the upstream support foot 5 are advantageously adapted to mechanically resist an impact from a solid body such as a bird.

[0042] The embodiment of [Fig.4] differs from that of [Fig.2] in that the upstream support foot 5 is pivotally mounted relative to the external wall 21 of the aircraft turbomachine 21, preferably along a pivot axis Z orthogonal to the radial axis Y extending in a transverse plane relative to the longitudinal axis X. This advantageously makes it possible to reduce the forces transmitted in the upstream support foot 5 during an impact. The pivot 10 advantageously makes it possible to facilitate the inclination of the stator blade 22 downstream during the crushing of the protective portion 8.

[0043] Figures 2 and 4 illustrate the support device 1 when the protective portion 8 is intact, before an impact. The protective portion 8 has a predetermined initial radial length V0 and makes it possible to maintain the stator blade(s) 22 along the radial axis Y. The protective portion 8 is dimensioned to mechanically resist in the presence of aerodynamic forces exerted on the stator blade 22, for example meteorological disturbances.

[0044] With reference to [Fig. 5], in the presence of an impact A of a solid body oriented from upstream to downstream on the stator blade 22, the protective portion 8 is compressed until it reaches a reduced radial length VL. The solid body is typically in the form of a bird exerting a force on the stator blade 22. Under the action of the impact A, the pores 9 of the protective portion 8 are crushed radially, which has the effect of reducing the radial length of the protective portion 8 irreversibly. After the impact, the protective portion 8 is in a compacted state. The protective portion 8 thus acts as a fuse and must then be replaced once the aircraft has landed on the ground.

[0045] As illustrated in [Fig. 5], the protective portion 8 advantageously makes it possible to tilt the stator blade 22 downstream during the impact A, which makes it possible to absorb the shock and the forces transmitted to the stator blade 22, but also to the external wall 21. The protective portion 8 thus makes it possible to protect the aircraft turbomachine 20 in a simple, practical and inexpensive manner. The method is implemented passively, no manual or electrical control being necessary.

Claims

Claims

1. Support device (1) for at least one stator blade (22) of an aircraft turbomachine (20), the aircraft turbomachine (20) extending along a longitudinal axis (X) oriented from upstream to downstream and comprising an external wall (21) from which the stator blade (22) projects along a radial axis (Y), the support device (1) comprising: • At least one housing (2) configured to radially receive a mounting foot (23) of the stator blade (22), • An upstream support foot (5) and a downstream support foot (6) connected respectively to an upstream end (3) and to a downstream end (4) of the housing (2) and configured to be mounted on the external wall (21) of the aircraft turbomachine (20), • The support device (1) being characterized in that the downstream support foot (6) comprises at least one connecting portion (7) and at least one protective portion (8),the protective portion (8) having a porosity greater than that of the connecting portion (7), the protective portion (8) comprising a plurality of pores (9) adapted to be compressed under the effect of an impact (A) exerted on the stator blade (22).,

2. Support device (1) according to claim 1, wherein the pores (9) are adapted to be irreversibly compressed.

3. Support device (1) according to one of claims 1 and 2, wherein the pores (9) extend parallel to each other, preferably in a direction parallel to the radial axis (Y).

4. Support device (1) according to one of claims 1 to 3, in which the protective portion (8) is in the form of a honeycomb structure.

5. Support device (1) according to one of claims 1 to 4, wherein the protective portion (8) is made of steel.

6. Support device (1) according to one of claims 1 to 5, wherein the connecting portion (7) extends between the housing (2) and the protective portion (8), preferably radially outwardly relative to the protective portion (8).

7. Support device (1) according to one of claims 1 to 6, de- ending a cavity (C) delimited upstream by the upstream support foot (5) and downstream by the downstream support foot (6), the cavity (C) being adapted to house a device for controlling the pivot angle along the radial axis (Y) of the stator blade (22) in the housing (2).

8. Assembly of a support device (1) according to one of claims 1 to 7 and of an external wall (21) of an aircraft turbomachine (20), the aircraft turbomachine (20) extending along a longitudinal axis (X) oriented from upstream to downstream, the stator blade (22) being configured to extend projecting from the external wall (21) along a radial axis (Y), the upstream support foot (5) of the support device (1) being pivotally mounted relative to the external wall (21) of the aircraft turbomachine (20), preferably along a pivot axis (Z) orthogonal to the radial axis (Y) extending in a plane transverse to the longitudinal axis (X).

9. Aircraft turbomachine (20) extending along a longitudinal axis (X) oriented from upstream to downstream and comprising: • An external wall (21), preferably in the form of an external compressor casing, • An unducted rectifier (29) comprising at least one stator blade (22) extending projecting from the external wall (21) along a radial axis (Y), • At least one support device (1) according to one of claims 1 to 7, in which the upstream support foot (5) and the downstream support foot (6) are mounted on the external wall (21), the stator blade (22) comprising a mounting foot (23) mounted radially in the housing (2) of the support device (1).

10. Method of using an aircraft turbomachine (20) according to claim 9, in which the pores (9) of the protective portion (8) are compressed under the effect of an impact (A) exerted on the stator blade (22) of the aircraft turbomachine (20).

Images