FR3130897A1 - AIRCRAFT TURBOMACHINE - Google Patents

Abstract

Aircraft turbomachine (10), comprising: - two coaxial annular walls (12, 14), - a rotor blade (30), - an annular separator (24) disposed downstream of the rotor blade (30) and between the two walls (12, 14), and comprising upstream an annular slat (24a), - fixed stator vanes (42) connected to the slat (24a), and - variable-pitch stator vanes (44) in downstream of the fixed stator vanes (42). Figure for abstract: Figure 3a

Description

AIRCRAFT TURBOMACHINE

Technical field of the invention

The present invention relates to the general field of aeronautics. It relates more particularly to an aircraft turbomachine.

Technical background

Conventionally, an aircraft turbomachine comprises a gas generator comprising along a longitudinal axis at least one compressor, one combustion chamber, and at least one turbine.

An air stream enters the gas generator and is compressed in the compressor(s). This flow of compressed air is mixed with fuel and burned in the combustion chamber, then the combustion gases are expanded in the turbine(s). This expansion causes the rotation of the turbine rotor(s), which causes the rotation of the compressor rotor(s). The combustion gases are ejected through a nozzle to provide thrust which may be added to thrust conferred by at least one propeller or propulsion fan of the turbomachine.

The gas flows flow into the turbomachine through annular veins. As can be seen in FIG. 1a, the turbomachine 10 thus comprises coaxial annular walls, respectively internal 12 and external 14, extending one around the other and defining between them a main annular flow path 16 a main gas flow 18.

In the case where the main gas flow 18 must be divided into two secondary gas flows, respectively internal 20 and external 22, an annular separator 24 is arranged between the two walls 12, 14 and respectively defines with these walls 12, 14 two secondary annular veins, respectively internal 26 and external 28, for the flow of the secondary gas flows 20, 22. This separator 24 comprises at an upstream end an annular spout 24a configured to separate the main gas flow 18 into two and form the flows of secondary gases 20, 22.

A rotor blade 30 may extend radially through the main stream 16, therefore upstream of the separator 24.

As shown in Figure 1a, structural arms 32 may extend radially through main stream 16 downstream of rotor blade 30 and upstream of separator 24.

In the present application, arm 32 or structural arm is understood to mean a stator element which has in section a general aerodynamic shape such as that shown in FIG. 1b, but which does not include any intrados or extrados. An arm 32 is therefore not comparable to a vane or blade which is profiled so as to include an intrados and an extrados. An arm 32 generally has a symmetry with respect to a plane P passing through the axis of the turbomachine. The number of arms 32 is generally less than 10 and may be 4. At least one of the arms 32 may be hollow and of tubular shape in the radial direction to be crossed by easements and thus serve for the passage of these easements in the motor through the veins.

For certain types of turbomachine, such as those with multi-flow or variable cycle, it would be useful to arrange a stator vane 34 directly downstream of the rotor vane 30 and integrated into the flow separation nozzle 24a instead of be positioned between the rotor 30 and the separator 24 (cf. FIG. 2a), so as to reduce the length of the module between the concept illustrated in FIG. 1a and that illustrated in FIG. 2a. The stator vane 34 would comprise several vanes distributed around the axis of the turbomachine. As mentioned in the foregoing and illustrated in Figure 2b, each of these blades would have in section an aerodynamic profile comprising an underside 34a and an upper surface 34b (Figure 2b), therefore a non-symmetrical profile which is not the case of the arm 32 visible in Figure 1a. The stator blade 34 would extend radially across the main stream 16. In the event that the blade 24a is connected to the blades of the stator blade 34, these blades would include leading edges 36 located upstream of the blade 24a, in the main vein 16, and trailing edges, respectively internal 38a and external 38b, located in the internal 26 and external 28 veins.

The stator blade 34 would impose a particular direction on the gas flows 16, 20, 22. However, in the case of a variable cycle turbomachine, it would be useful to provide a variable geometry downstream of the stator blade 34 in order to be able to adapt to the various operating speeds and variations in the bypass rate of the turbomachine. However, for space reasons, the addition of a variable-pitch vane downstream of the stator vane 34 can be complex. Indeed, this addition would require extending the axial dimension of the turbomachine, which would result in an increase in the mass of the turbomachine and a reduction in its performance.

In addition, for reasons of noise pollution, it would also not be possible to bring the stator blade 34 axially closer to the rotor blade 30.

In the present application, a variable-cycle turbomachine means a turbomachine whose specific thrust can be modified at a given engine speed, by controlling variable geometries of the turbomachine. An example of variable geometry is a variable-pitch stator vane. In the present application, blading is understood to mean an annular row of blades.

The invention thus proposes to optimize a turbomachine as illustrated in FIG. 2a so as to be able to use it in several configurations and in particular within the framework of a turbomachine with several streams (at least two) and/or a variable cycle turbomachine.

The present invention proposes an aircraft turbomachine, comprising a gas generator comprising along a longitudinal axis at least one compressor, one combustion chamber and at least one turbine, the turbomachine further comprising:

- two coaxial annular walls, respectively internal and external, extending one around the other and defining between them a main annular vein for the flow of a main air flow,

- a rotor blade extending radially through said main stream,

- an annular separator arranged downstream of the rotor blade and between the two walls, the separator respectively defining with the internal and external walls two secondary annular veins, respectively internal and external, for the flow of secondary air flows, respectively internal and external, the separator comprising at an upstream end an annular nozzle configured to separate the main air flow into two and form the secondary air flows,

- stator elements extending radially on the one hand through said main vein and on the other hand through said secondary veins, these stator elements being connected to said annular nozzle,

characterized in that said stator elements comprise:

- fixed stator vanes which are distributed around said axis and which each comprise a leading edge located upstream of said nozzle, and trailing edges, respectively internal and external, located respectively in the internal and external secondary veins, these blades fixed rectifiers being connected to said nozzle, and

- variable-pitch stator vanes which are distributed around said axis and which extend radially through at least one of said secondary streams, each of the variable-pitch stator vanes comprising a leading edge and a trailing edge,

and in that :

- the leading edges of the variable-pitch stator vanes are located upstream of the inner and/or outer trailing edges of the fixed stator vanes, or

- the leading edges of the variable-pitch stator vanes are located directly downstream of the inner and/or outer trailing edges of the fixed stator vanes, and are separated by predetermined axial clearances from these trailing edges.

The present invention thus proposes to put both fixed stator vanes and variable-pitch stator vanes in place of the arms of FIG. 1a or the stator vane of FIG. 1b. The vanes of fixed and variable-pitch stators are very close together axially or are embedded axially in each other so that they are considered as an assembly forming the stator elements within the meaning of the invention. Indeed, either the variable-pitch stator vanes have their leading edges located upstream of the trailing edges of the fixed stator vanes, or the variable-pitch stator vanes are separated by predetermined axial clearances, the smallest possible preferably trailing edges of the stationary stator vanes. The minimization of these axial clearances makes it possible to limit or even prevent the passage of gas in operation between the trailing edges of the fixed stator vanes and the leading edges of the variable-pitch stator vanes. It is thus understood that the gases which flow over the lower surfaces of the fixed stator vanes must then flow over the lower surfaces of the variable-pitch stator vanes, and that the gases which flow over the upper surfaces of the fixed stator vanes must then flow over the upper surfaces of the variable-pitch stator vanes.

This configuration is particularly advantageous because it makes it possible to optimize the operation of the turbomachine, by authorizing multi-flow or variable-cycle applications, while limiting the impact on the length or axial dimension as well as the mass of the turbomachine. In fact, the fact of reducing the axial play between the stator vanes and positioning them at the level of the nozzle makes it possible to limit the impact of these vanes on the axial dimension of the turbomachine.

Upstream of the rotor blading located in the first vein, there can be any configuration for the turbomachine.

In the present application, the term “annular” means a shape of revolution around an axis, this shape possibly being continuous or interrupted.

In addition, in the present application, the term “variable pitch” element is understood to mean an element of which at least one part has a position which can be adjusted around an axis, which is called the pitch axis. All of this element or only part of this element can be variably pitched. In the case of a blade for example, it can be one-piece and have an adjustable position around a setting axis. As a variant, it could comprise only one part, comprising for example a leading edge or a trailing edge, the position of which would be adjustable around a setting axis relative to the rest of the blade. In the case of a blading comprising several blades, each of the blades has an adjustable position around a setting axis which is specific to it. For the same blading, there are therefore as many pitch axes as there are variable-pitch blades. Each of these axes can have a radial or inclined orientation with respect to the longitudinal axis of the turbomachine.

The turbomachine may include one or more of the following features, taken alone or in combination with each other:
-- Said clearances are less than 10mm, and preferably less than or equal to 5mm;

-- said clearances are less than 10% of the chord of one of the fixed or variable-pitch blades, and preferably less than or equal to 5% of this chord;

-- Said fixed stator vanes comprise an intrados and an extrados, and said variable-pitch stator vanes comprise an intrados and an extrados;

- the number of said variable-pitch stator vanes is greater than or equal to the number of said fixed stator vanes;

- Said variable-pitch stator vanes are located in said internal secondary stream;

- the trailing edges of said variable-pitch stator vanes are located downstream of the outer trailing edges of the fixed stator vanes;

the turbomachine further comprises a system for controlling the angular pitch of the variable-pitch stator vanes, this system being mounted in said separator;

- the turbomachine further comprises a system for controlling the angular pitch of the variable-pitch stator vanes, this system being mounted radially outside of said outer wall;

- Said variable-pitch stator vanes are located in said outer secondary stream;

- First variable-pitch stator vanes are located in said inner secondary stream, and second variable-pitch stator vanes are located in said outer secondary stream;

the turbomachine further comprises a common system for controlling the angular setting of the first and second variable-pitch stator vanes, or independent systems for controlling the angular setting respectively of the first and second variable-pitch stator vanes;

- the turbomachine further comprises structural arms distributed around said axis in said external secondary stream;

- the number of structural arms is lower than the number of fixed stator vanes;

- the structural arms are connected to some of said fixed stator vanes;

- the rotor blading is a fan or a compressor rotor blading;

- the leading edges of the variable-pitch stator vanes are located at a distance from the inner and/or outer trailing edges of the fixed stator vanes, which is greater than 10% of the chord of one of these vanes, and more preferably greater than or equal to 20% of this chord;

- at least some of the fixed stator vanes have different profiles or cambers from the other fixed stator vanes;

-- at least some of the arms have different profiles from the other arms.

The present invention also relates to an aircraft, in particular a transport aircraft, comprising a turbine engine as described above.

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. 1a is a very schematic half-view in axial section of an aircraft turbine engine, according to the technique prior to the invention; FIG. 1b is a very schematic cross-sectional view of an arm of the turbomachine of FIG. 1a;

FIG. 2a is a very schematic half-view in axial section of part of an aircraft turbomachine; FIG. 2b is a very schematic cross-sectional view of a stator vane of the turbomachine of FIG. 2a;

FIG. 3a is a very schematic half-view in axial section of an aircraft turbine engine, according to a first embodiment of the invention; FIG. 3b is a very schematic cross-sectional view of two fixed stator vanes followed by two variable-pitch stator vanes of the turbomachine of FIG. 3a, and illustrate, respectively on the left and on the right of the figure, two positions separate variable-pitch stator vane pitch;

FIG. 4a is a very schematic half-view in axial section of an aircraft turbine engine, according to a second embodiment of the invention; FIG. 4b is a very schematic cross-sectional view of two fixed stator vanes followed by three variable-pitch stator vanes of the turbomachine of FIG. 4a, and illustrate, respectively on the left and on the right of the figure, two positions separate variable-pitch stator vane pitch;

FIG. 5a is a very schematic half-view in axial section of an aircraft turbine engine, according to a third embodiment of the invention; FIG. 5b is a very schematic cross-sectional view of two fixed stator vanes interposed with two variable-pitch stator vanes of the turbomachine of FIG. 5a, and illustrate, respectively on the left and on the right of the figure, two positions separate variable-pitch stator vane pitch;

there is a very schematic half-view in axial section of an aircraft turbine engine, according to a fourth embodiment of the invention;

there is a very schematic half-view in axial section of an aircraft turbine engine, according to a fifth embodiment of the invention;

there is a very schematic half view in axial section of an aircraft turbine engine, according to a sixth embodiment of the invention;

there is a very schematic half-view in axial section of an aircraft turbomachine, according to a seventh embodiment of the invention;

there is a very schematic half view in axial section of an aircraft turbomachine, according to an eighth embodiment of the invention;

there is a very schematic half-view in axial section of an aircraft turbine engine, according to a ninth embodiment of the invention;

there is a very schematic half-view in axial section of an aircraft turbomachine, according to a tenth embodiment of the invention;

there is a very schematic half view in axial section of an aircraft turbine engine, according to an eleventh embodiment of the invention.

Claims (13)

Aircraft turbomachine (10), comprising a gas generator comprising along a longitudinal axis at least one compressor, one combustion chamber and at least one turbine, the turbomachine further comprising:
- two coaxial annular walls, respectively internal (12) and external (14), extending one around the other and defining between them a main annular vein (16) for the flow of a main air flow (18),
- a rotor blade (30) extending radially through said main stream (16),
- an annular separator (24) arranged downstream of the rotor blade (30) and between the two walls (12, 14), the separator (24) defining respectively with the internal and external walls (12, 14) two veins secondary annular rings, respectively internal (26) and external (28), for the flow of secondary air flows, respectively internal (20) and external (22), the separator (24) comprising at an upstream end an annular spout (24a ) configured to separate the main airflow (18) in two and form the secondary airflows (20, 22),
- stator elements extending radially on the one hand through said main vein (16) and on the other hand through said secondary veins (26, 28), these stator elements being connected to said annular nose (24a),
characterized in that said stator elements comprise:
- fixed stator vanes (42) which are distributed around said axis and which each have a leading edge (42a) located upstream of said nozzle (24a), and trailing edges, respectively internal (42b) and external ( 42c), located respectively in the inner (26) and outer (28) secondary streams, these fixed stator vanes (42) being connected to said nozzle (24a), and
- variable-pitch stator vanes (44) which are distributed around said axis and which extend radially through at least one of said secondary streams (26, 28), each of the variable-pitch stator vanes (44) comprising a leading edge (44a) and a trailing edge (44b),
and in that :
- the leading edges (44a) of the variable-pitch stator vanes (44) are located upstream of the inner (42b) and/or outer (42c) trailing edges of the fixed stator vanes (42), or
- the leading edges (44a) of the variable-pitch stator vanes (44) are located directly downstream of the inner (42b) and/or outer (42c) trailing edges of the fixed stator vanes (42), and are separated by predetermined axial clearances (J) of these trailing edges (42b, 42c). A turbomachine (10) according to claim 1, wherein the number of said variable-pitch stator vanes (44) is greater than or equal to the number of said fixed stator vanes (42). A turbomachine (10) according to one of the preceding claims, wherein said variable-pitch stator vanes (44) are located in said internal secondary stream (26). A turbomachine (10) according to one of the preceding claims, wherein the trailing edges (44b) of said variable-pitch stator vanes (44) are located downstream of the outer trailing edges (42c) of the fixed stator vanes (42 ). Turbomachine (10) according to claim 4, in which it further comprises a system (50) for controlling the angular pitch of the variable-pitch stator vanes (42), this system being mounted radially outside of said outer wall ( 14). A turbomachine (10) according to one of claims 1 to 3, wherein said variable-pitch stator vanes (44) are located in said outer secondary stream (28). A turbomachine (10) according to one of claims 1 to 3, wherein first variable-pitch stator vanes (44) are located in said inner secondary stream (26), and second variable-pitch stator vanes (44) are located in said external secondary vein (28). A turbomachine (10) according to claim 7, wherein it further comprises a common system (50) for controlling the angular timing of the first and second variable-pitch stator vanes (44), or independent systems (50) for controlling the angular timing of the first and second variable-pitch stator vanes (44), respectively. Turbomachine (10) according to one of the preceding claims, in which it further comprises structural arms (32) distributed around said axis in said external secondary stream (28). A turbomachine (10) according to claim 9, wherein the number of structural arms (32) is less than the number of stationary stator vanes (42). A turbomachine (10) according to claim 9 or 10, wherein the structural arms (32) are connected to some of said stationary stator vanes (42). A turbomachine (10) according to one of the preceding claims, wherein the rotor blade (30) is a fan or a compressor rotor blade. Turbomachine (10) according to one of the preceding claims, in which the leading edges (44a) of the variable-pitch stator vanes (44) are located at a distance (W) from the internal trailing edges (42b) and/ or external (42c) of the fixed stator vanes (42), which is greater than 10% of the chord of one of these vanes (42, 44), and more preferably greater than or equal to 20% of this chord.