CN120303467A - Three-channel aircraft turbine - Google Patents

Abstract

The application relates to a three-channel aircraft turbine (10) comprising-two coaxial annular walls defining a main annular duct (16) between them for the flow of a main air stream (18), -a rotor blade assembly (30) extending radially through the main duct (16) and forming a ducted propeller (H1), -an annular separator (24) arranged downstream of the rotor blade assembly (30) and configured to divide the main air stream (18) into two parts and form a secondary air stream (20, 22), -first variable guiding vanes (40) distributed around an axis and each comprising a leading edge (40 a) upstream of the separator (24) and a trailing edge (40 b, 40 c) in the secondary air stream (20, 22), and-fixed guiding vanes (42) distributed in the outer air stream around the axis and downstream of the first variable vanes (40), -a non-ducted propeller (H2) arranged upstream of the outer wall (14).

Description

Three-runner aircraft turbine

Technical Field

The present invention relates generally to the field of aviation. More specifically, three-channel aircraft turbines are targeted.

Background

Traditionally, aircraft turbines include a gas generator that includes at least one compressor, a combustor, and at least one turbine along a longitudinal axis.

The air stream enters the gasifier and is compressed in one or more compressors. The compressed air stream is mixed with fuel and combusted in a combustor, and the combustion gases are expanded in one or more turbines. This expansion causes one or more turbine rotors to rotate, which drives one or more compressor rotors in rotation. The combustion gases are injected through the nozzle to provide thrust, which may be in addition to the thrust provided by at least one ducted or non-ducted propeller or fan of the turbine.

The air flow flows into the turbine through the annular duct. As shown in fig. 1a, the turbine 10 thus comprises coaxial annular walls, respectively an inner annular wall 12 and an outer annular wall 14, which extend around each other and define therebetween a main annular duct 16 for the flow of a main gas stream 18.

In case the main air flow 18 is split into two secondary air flows (inner secondary air flow 20 and outer secondary air flow 22, respectively), an annular separator 24 is arranged between the two walls 12, 14 and defines with these walls 12, 14 two secondary annular ducts (inner secondary annular duct 26 and outer secondary annular duct 28, respectively) for the flow of the secondary air flows 20, 22, respectively. The splitter 24 includes an annular splitter nose 24a at an upstream end that is configured to split the primary air flow 18 into two portions and form the secondary air flows 20, 22.

Rotor blade assembly 30 may extend radially through main duct 16, passing upstream of separator 24.

As shown in FIG. 1a, the structural arms 32 may extend radially through the main duct 16 downstream of the rotor blade assemblies 30 and upstream of the separators 24.

In the present application, the arm 32 or structural arm means a stator element having a general aerodynamic cross-sectional shape as shown in FIG. 1b, but excluding the pressure side or suction side. Thus, the arm 32 cannot be compared to a fan blade or vane shaped to include a pressure side and a suction side. The arms 32 are generally symmetrical about a plane P passing through the turbine axis. The number of arms 32 is typically less than 10 and may be 4. At least one of the arms 32 may be hollow and tubular in a radial direction for passage by accessories and for passage of those accessories through a duct in the engine.

For some types of turbines, such as multi-channel or variable cycle turbines, it is useful to have the stator vane assembly 34 located directly downstream of the rotor vane assembly 30 and integrated with the splitter nose 24a for separating the flow (see FIG. 2 a), rather than being positioned between the rotor 30 and the splitter 24, in order to reduce the length of the module between the concepts shown in FIG. 1a and FIG. 2 a. The stator vane assembly 34 will include a plurality of fan blades distributed about the axis of the turbine. As described above, and as shown in fig. 2b, each of these fan blades will have an aerodynamic profile in cross section including a pressure side 34a and a suction side 34b (fig. 2 b), so that the aerodynamic profile is an asymmetric profile, which is not the case for the arm 32 shown in fig. 1 a. The stator vane assembly 34 will extend radially through the main duct 16. These will include a leading edge 36 upstream of the diverter nose 24a in the main duct 16, and trailing edges, an inner trailing edge 38a and an outer trailing edge 38b in the inner duct 26 and the outer duct 28, respectively. The stator vane assembly may be connected to the diverter nose 24a.

The stator vane assembly 34 will impart a particular direction to the airflow 16, 20, 22. However, in the case of a variable cycle turbine, it is useful to provide a variable geometry downstream of the rotor blade assembly 30 so as to be able to accommodate variations in the different modes of operation and bypass ratios of the turbine. However, adding a variable pitch blade assembly downstream of stator blade assembly 34 can be complicated for overall size reasons. In practice, such additions would require lengthening the axial dimensions of the turbine, which would result in increased mass and reduced performance of the turbine.

Furthermore, the stator blade assembly 34 will not be able to be moved axially closer to the rotor blade assembly 30 due to noise pollution.

In the present application, a variable cycle turbine means a turbine whose specific thrust can be varied at a given engine speed by controlling the variable geometry of the turbine. An example of a variable geometry is a variable pitch stator vane assembly. In the present application, the vane assembly refers to an annular array of vanes.

The present invention therefore proposes to optimise the turbine as shown in fig. 2a so that it can be used in a variety of configurations, and in particular in the context of multi-runner (at least two-runner) and/or variable cycle turbines.

Disclosure of Invention

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

two coaxial annular walls, respectively an inner annular wall and an outer annular wall, which extend around each other and define between them a main annular duct for the main air flow,

A rotor blade assembly extending radially through the main duct and forming a ducted propeller,

An annular separator disposed downstream of the rotor blade assembly and between the two walls, the separator defining, with the inner and outer walls, respectively, two secondary annular ducts, an inner secondary duct and an outer secondary duct, respectively, for the flow of secondary air streams, respectively, the secondary air streams being inner and outer secondary air streams, respectively, the separator comprising an annular diverter nose at an upstream end configured to divide the primary air stream into two parts and form the secondary air streams,

A stator element extending radially through the main conduit on the one hand and through the secondary conduit on the other hand,

And

A non-ducted propeller arranged upstream of the outer wall,

Characterized in that said stator element comprises first variable pitch stator blades distributed about said axis and each comprising a leading edge and a trailing edge upstream of said diverter nose, said trailing edges being inner and outer trailing edges in said inner and outer secondary ducts respectively,

And, the turbine further comprises stationary stator blades distributed in the outer secondary duct around the axis and downstream of the outer trailing edge of the first variable pitch blade.

The present invention therefore proposes to provide variable pitch stator blades at the splitter. To enable angular displacement of the vanes about their pitch axis, it will be appreciated that the vanes will be separated from the diverter nose and separator by a small gap to limit gas leakage in these regions.

A fixed stator vane is associated with the variable pitch vane and is located in the outer secondary duct. This configuration optimizes the operation of the turbine, allowing for multi-flow or variable cycle applications while limiting the impact on the axial length or size and mass of the turbine. In fact, providing variable pitch blades at the diverter nose enables the flow distance between the rotor and the diverter nose to be reduced in the axial direction, while also allowing the airflow flowing in the inner and outer secondary ducts to be varied.

In the present application, "annular" means a revolution shape around an axis, which may be continuous or intermittent.

Furthermore, in the present application, a "variable pitch" element refers to an element, at least a portion of which has a position that is adjustable about an axis called the pitch axis. The whole element or only a part thereof may be pitchable. In the case of a fan blade, for example, the fan blade may be one-piece and have an adjustable position about a wedge axis. Or the position of only one portion thereof (e.g., the leading or trailing edge) may be adjusted about the pitch axis relative to the rest of the blade. Where the blade assembly includes a plurality of blades, each blade has an adjustable position about its own pitch axis. Thus, for the same blade assembly, there are as many pitch axes as there are variable pitch blades. Each of these axes may have a radial orientation or an oblique orientation relative to the longitudinal axis of the turbine.

The turbine may include one or more of the following features, which may be independent of each other or in combination with each other:

-said stator element further comprises a second variable pitch stator blade in said inner secondary duct;

The second variable-pitch blades comprise a leading edge and a trailing edge, the leading edges of which are located immediately downstream of the inner trailing edges of the first variable-pitch blades and are spaced apart from these trailing edges by a predetermined axial gap, whereby the first and second variable-pitch blades are axially very close to each other such that they are considered within the meaning of the invention an assembly forming a stator element, the above-mentioned axial gap between these blades preferably being as small as possible. Minimizing these axial clearances enables limiting or even preventing the passage of gas between the trailing edge of the first variable pitch blade and the leading edge of the second variable pitch blade during operation, it should therefore be appreciated that gas flowing through the pressure side of the first variable pitch blade must then flow through the pressure side of the second variable pitch blade and gas flowing through the pressure side of the first variable pitch blade must then flow through the pressure side of the second variable pitch blade;

-the stationary stator blades comprise a leading edge, which is spaced apart from the trailing edge of the first variable pitch blade by a predetermined axial gap;

preferably, the gap is less than 10mm, and more preferably, less than or equal to 5mm;

-the number of second variable pitch blades is equal to the number of first variable pitch blades;

-the number of second variable pitch blades is equal to a multiple of the number of first variable pitch blades;

-the number of fixed blades is equal to the number of first variable pitch blades;

-the number of fixed blades is equal to a multiple of the number of first variable pitch blades;

-the turbine further comprises at least one system for controlling the pitch angle of the variable pitch blades;

-the control system is mounted in the separator or radially outside the outer wall;

-at least some of the stationary blades have a different profile than the other stationary blades and thus form a multi-profile blade grid;

the rotor blade assembly is a propulsive fan or compressor rotor blade assembly, and

-The fixed stator blades comprise a pressure side and a suction side, and the variable pitch stator blades comprise a pressure side and a suction side.

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

Drawings

Other features and advantages of the present invention will become apparent from the following detailed description, and it is convenient to understand the description with reference to the accompanying drawings, in which:

Fig. 1a is a very schematic half view of an axial section of an aircraft turbine according to the prior art;

FIG. 1b is a very schematic cross-sectional view of an arm of the turbine of FIG. 1 a;

FIG. 2a is a very schematic semi-view of an axial section of an aircraft turbine;

FIG. 2b is a very schematic cross-sectional view of the stator vanes of the turbine of FIG. 2 a;

Fig. 3a is a very schematic half view of an axial section of an aircraft turbine according to a first embodiment of the invention;

FIG. 3b is a very schematic cross-sectional view of the variable pitch stator blades, the fixed stator blades of the turbine of FIG. 3a immediately behind the variable pitch stator blades, and showing two different positions of the pitch of the variable pitch stator blades on the left and right sides of the figure, respectively;

FIG. 3c is a view similar to the left side of FIG. 3b and showing an alternative embodiment of the present invention;

Fig. 4a is a very schematic half view of an axial section of an aircraft turbine according to a second embodiment of the invention;

FIG. 4b is a very schematic cross-sectional view of a first variable pitch stator blade, the second variable pitch stator blade of the turbine of FIG. 4a immediately behind the first variable pitch stator blade, and two different pitch positions of these blades are shown on the left and right sides of the figure, respectively;

FIG. 4c is a view similar to the left side of FIG. 4b and showing an alternative embodiment of the present invention;

FIG. 5 is a very schematic half view of an axial section of an aircraft turbine according to a third embodiment of the invention, wherein the stator grid consists of at least two different blade profiles, and

FIG. 6 is a schematic view of a three-flow path turbine according to the present invention.

Detailed Description

Figures 1a, 1b, 2a and 2b have been described above.

Referring to fig. 6, the turbine 10 is of the three-channel type and generally comprises a gas generator 2 comprising, along a longitudinal axis X, at least one compressor, a combustion chamber and at least one turbine. The turbine includes a ducted propeller H1 and a non-ducted propeller H2. The propeller H1 is surrounded by a nacelle 4, which extends downstream of the propeller H2 about an axis X. The air flow through the propeller H2 is separated by the nacelle 4 into a main flow F2 entering the nacelle 4 and another flow F3 flowing around the nacelle 4. The main flow F2 is then split into two further flows F1, F2, as described below.

In the present application, as shown in fig. 3a and 3b, the turbine 10 thus comprises two coaxial annular walls, respectively an inner annular wall 12 and an outer annular wall 14, which extend around each other and define between them a main annular duct 16 of a main air flow 18.

The main air flow 18 is divided into two secondary air flows, an inner secondary air flow 20 and an outer secondary air flow 22, respectively, by an annular separator 24 arranged between the two walls 12, 14. The splitter 24 includes an annular splitter nose 24a at an upstream end that is configured to split the primary air flow 18 into two portions and form the secondary air flows 20, 22.

Rotor blade assembly 30 extends radially through main duct 16 and thus upstream of separator 24. In the turbine of fig. 6, the rotor blade assembly 30 forms a ducted propeller H1.

The stator elements are located downstream of the rotor blade assembly 30 and at the splitter 24 a.

According to the invention, these stator elements comprise first variable pitch stator blades 40.

In addition, stationary stator blades 42 are located in outer secondary duct 28 downstream of first variable pitch stator blades 40.

The first variable pitch blades 40 are distributed about the axis and each include a leading edge 40a and a trailing edge upstream of the diverter nose 24a, the trailing edges being inner trailing edge 40b and outer trailing edge 40c in the inner secondary duct 26 and outer secondary duct 28, respectively. Thus, it will be appreciated that the first variable pitch fan blade 40 is located at the position of the diverter nose 24a, as can be seen in the figures. There is a gap, not shown, between the diverter nose 24a and the first pitchable fan blade 40 to allow them to move. The gaps are preferably as small as possible to limit or prevent the passage of gas between the vanes 40 and nose 24 a. It can also be seen that the leading edge 42a may be beveled and extend from upstream to downstream toward the outside. The inclination is determined, for example, according to a compromise between the size of the engine and the optimization of the noise it generates. In order to minimize noise, it is preferable to increase the height of the top of the blade, which results in a higher pitch of the blade.

Fig. 3b shows that each of the first variable pitch fan blades 40 has an aerodynamic profile and comprises a pressure side 46 (concave curved shape) and a suction side 48 (convex curved shape). In addition, each first variable pitch fan blade has a curvature along its chord. The area of maximum curvature of the variable pitch fan blade 40 is referred to as C. This region is preferably located upstream of the diverter nose 24 a.

Preferably, the first variable pitch blades 40 are all identical. Preferably, their leading edges 40a are traversed by the same transverse plane.

The number of first variable pitch blades 40 is for example between 10 and 200.

Each of the first variable pitch blades 40 is rotatable about a pitch axis Y (the pitch axis having a generally radial orientation). The rotation of each first variable pitch fan blade 40 is achieved by means of a control system 50 located radially outside the outer wall 14. This is advantageous because it allows the system to be located in a relatively cool environment (compared to the high temperatures that can be prevalent in gas generators). Furthermore, this environment is not very limited and contains free space to accommodate this type of system.

The stationary vanes 42 are distributed about an axis in the outer secondary duct 28. Each of which includes a leading edge 42a downstream of the diverter nose 24a and a trailing edge 42b in the outer secondary duct 28.

Fig. 3b shows that each of the stationary blades 42 has an aerodynamic profile and comprises a pressure side 46 (concave curved shape) and a suction side 48 (convex curved shape). In addition, each of the fixed blades 42 has a curvature along its chord.

The number of fixed blades 42 is equal to the number of first variable pitch blades 40, or a multiple of the number of first variable pitch blades 40, and the fixed blades 42 are located immediately downstream of the first variable pitch blades 40 and axially successive to the first variable pitch blades 40. The leading edge 42a of the fixed blade 42 is spaced apart from the trailing edge 40c of the first variable pitch blade 40 by a predetermined axial gap I. Preferably, these gaps I are less than 10mm, and more preferably, less than or equal to 5mm. Preferably, these gaps I are less than 10% of the chord of blade 40 or blade 42, and more preferably less than or equal to 5% of the chord. Preferably, each of these gaps I is constant over the entire radial extent of the associated edge 40c, 42a and over the entire radial extent of the outer duct 28. Of course, these clearances I may vary during operation depending on the pitch position of blade 40 relative to blade 42.

Preferably, the fixed blades 42 are all identical. Preferably, their leading edges 42a lie in and are traversed by the same transverse plane.

The number of stationary blades 42 is, for example, between 10 and 200.

The left side of fig. 3b shows a first angular position or first pitch position of the first variable pitch fan blades 40 and the right side shows a second angular position or second pitch position of these fan blades. For example, the first variable pitch blades 40 may be movable about their axis Y through an angular range of about 60 °.

Fig. 4c shows an alternative embodiment, wherein the number of fixed blades 424 is a multiple of the number of first variable pitch blades 40. The multiples are for example 2,3, 4, etc.

Fig. 4a and 4b show a second embodiment of the invention, which differs essentially from the previous embodiments in that the turbine further comprises a plurality of second variable pitch stator blades 44 downstream of the trailing edge 40b of the first variable pitch blades 40 in the inner secondary duct 26.

The second variable pitch stator blades 44 each include a leading edge 44a downstream of the diverter nose 24a and a trailing edge 44b in the inner secondary duct 26.

Each second variable pitch fan blade 44 has an aerodynamic profile and includes a pressure side and a suction side. In addition, each variable pitch fan blade 44 has a curvature along its chord.

The number of second variable pitch blades 44 may be equal to the number of first variable pitch blades 40.

The second variable pitch blade 44 is located directly downstream of the fixed blade 42 and axially successive to the fixed blade 42. The leading edge 44a of the second variable pitch fan blade 44 is spaced apart from the trailing edge 42c of the fixed fan blade 42 by a predetermined axial gap J. Preferably, these clearances J are less than 10mm, and more preferably, less than or equal to 5mm. Preferably, these gaps J are less than 10% of the chord of the fan blade 40 or the fan blade 44, and more preferably less than or equal to 5% of the chord. Preferably, each of these clearances J is constant over the entire radial extent of the associated edges 40b, 44a and over the entire radial extent of the inner conduit 26. Of course, these clearances J may vary during operation depending on the pitch position of the blades 40, 44.

Preferably, the second variable pitch blades 44 are all identical. Preferably, their leading edges 44a lie in and are traversed by the same transverse plane.

The number of second variable pitch blades 44 is for example between 10 and 200. Each of the second variable pitch blades 44 is rotatable about a pitch axis Z (which has a generally radial orientation). Rotation of each second variable pitch fan blade 44 is achieved by a control system 50' located in the separator 24.

Fig. 4c shows an alternative embodiment, wherein the number of variable pitch stator blades 44 is a multiple of the number of first variable pitch blades 40. The multiples are for example 2, 3, 4, etc.

Fig. 5 shows a third embodiment of the invention, which differs essentially from the previous embodiments in that the fixed blades 42 are not all identical. The stationary blades 42 are of at least two types that differ from each other in terms of size and/or geometry and/or camber. The different types of stationary blades 42 are evenly distributed about the axis, thereby obtaining a circular distribution of these stationary blades 42 about the axis.

In summary, the present invention is applicable to any turbine in which a primary stream is split into two secondary streams downstream of a ducted rotor blade assembly.

Claims (10)

1. A three-channel aircraft turbine (10) comprising a gas generator including at least one compressor, a combustion chamber and at least one turbine along a longitudinal axis, the turbine further comprising:

two coaxial annular walls, an inner annular wall (12) and an outer annular wall (14), respectively,

The inner and outer annular walls extending around each other and defining therebetween a main annular duct (16) for the flow of a main air stream (18),

A rotor blade assembly (30) extending radially through the main duct (16) and forming a ducted propeller (H1),

-An annular separator (24) arranged downstream of the rotor blade assembly (30) and between the two walls (12, 14), the separator (24)

Defining two secondary annular ducts with said inner and outer walls (12, 14), respectively,

Namely an inner secondary duct (26) and an outer secondary duct (28) for the flow of secondary air streams, respectively, the secondary air streams being inner secondary air streams (20), respectively

And an outer secondary air stream (22), the separator (24) including an annular diverter nose (24 a) at an upstream end, the annular diverter nose configured to divide the primary air stream (18) into two portions and form the secondary air stream (20, 22),

A stator element which extends radially through the main conduit (16) on the one hand,

And on the other hand extends radially through the secondary duct (26, 28), and

-A non-ducted propeller (H2) arranged upstream of said outer wall (14),

Characterized in that the stator element comprises first variable pitch stator blades (40) distributed around the axis and each comprising a leading edge (40 a) and a trailing edge upstream of the diverter nose (24 a), the trailing edges being an inner trailing edge (40 b) and an outer trailing edge (40 c) in the inner secondary duct (26) and the outer secondary duct (28), respectively,

And, the turbine (10) further comprises stationary stator blades (42) distributed in the outer secondary duct (28) around the axis and downstream of an outer trailing edge (40 c) of the first variable pitch blade (40).

2. The turbine (10) of claim 1, wherein the stator element further comprises a second variable pitch stator vane (44) located in the inner secondary duct (26).

3. The turbine (10) of claim 2, wherein the second variable pitch blades (44) include leading edges (44 a) and trailing edges (44 b), the leading edges (44 a) of the second variable pitch blades (44) being located directly downstream of the inner trailing edges (40 b) of the first variable pitch blades (40) and spaced apart from the trailing edges (40 b) by a predetermined axial gap (J).

4. The turbine (10) of any one of the preceding claims, wherein the stationary stator blades (42) include a leading edge (42 a) that is spaced apart from the trailing edge (40 c) of the first variable pitch blade (40) by a predetermined axial gap (I).

5. The turbine (10) of any of claims 2 to 4, wherein the number of second variable pitch blades (44) is equal to the number of first variable pitch blades (40) or a multiple of the number of first variable pitch blades (40).

6. The turbine (10) of any of the preceding claims, wherein the number of fixed blades (42) is equal to the number of first variable pitch blades (40) or a multiple of the number of first variable pitch blades (40).

7. The turbine (10) according to any one of the preceding claims, wherein the turbine further comprises at least one system (50, 50') for controlling the pitch angle of the variable pitch fan blades (42).

8. The turbine (10) of claim 7, wherein the control system (50, 50') is mounted in the separator (24) or radially outward of the outer wall (14).

9. The turbine (10) of any of the preceding claims, wherein at least some of the stationary blades (42) have a different profile than other stationary blades (42).

10. The turbine (10) according to any one of the preceding claims, wherein the rotor blade assembly (30) is a fan or compressor rotor blade assembly.