Three-runner aircraft turbine
By introducing a three-channel design of variable pitch stator blades and fixed stator blades into the turbine, the problem of increasing axial size and mass in the variable cycle mode is solved, and efficient, low-noise multi-channel and variable cycle operation is achieved.
Patent Information
- Application Number
- CN202280102305.6
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2022-12-05
- Publication Date
- 2025-07-11
AI Technical Summary
Existing turbines in variable cycle mode increase the axial size and mass of the turbine, while increasing noise pollution and complexity make it difficult to operate efficiently in multi-channel and variable cycle turbines.
The three-channel turbine design includes variable pitch stator fan blades and fixed stator fan blades. By setting variable pitch stator fan blades downstream of the rotor blade assembly, combined with fixed stator fan blades, the operation of the turbine is optimized and the axial length and mass influence is limited.
It realizes efficient operation in multi-channel and variable cycle turbines, reduces noise pollution, and optimizes the performance and structural design of the turbine.
Smart Images
Figure CN120303467A_ABST
Abstract
Description
Field of the Invention
[0001] The present invention generally relates to the field of aviation. More specifically, it is directed to a three - flowpath aircraft turbine. Background Art
[0002] Traditionally, an aircraft turbine includes a gas generator that includes along a longitudinal axis at least one compressor, a combustion chamber, and at least one turbine.
[0003] An air flow enters the gas generator and is compressed in one or more compressors. The compressed air flow is mixed with fuel and burned in the combustion chamber, and the combustion gases expand in one or more turbines. This expansion causes one or more turbine rotors to rotate, which drives one or more compressor rotors to rotate. The combustion gases are ejected through a nozzle to provide thrust, which can be thrust in addition to the thrust provided by at least one ducted or unducted propeller or fan of the turbine.
[0004] The air flow enters the turbine through an annular duct. As shown in FIG. 1a, the turbine 10 thus includes coaxial annular walls, namely 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 the main air flow 18.
[0005] In the case where the main air flow 18 is divided 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 respectively 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. The separator 24 includes at its upstream end an annular diverter nose 24a which is configured to separate the main air flow 18 into two parts and form the secondary air flows 20, 22.
[0006] The rotor blade assembly 30 may extend radially through the main duct 16 and thus through the upstream of the separator 24.
[0007] As shown in FIG. 1a, the structural arm 32 may extend radially through the main duct 16 downstream of the rotor blade assembly 30 and upstream of the separator 24.
[0008] In the present application, the arm 32 or structural arm means a stator element as described below: having a generally aerodynamic cross-sectional shape as shown in Figure 1b, but not including a pressure side or a suction side. Thus, the arm 32 cannot be compared to a blade or vane shaped to include a pressure side and a suction side. The arm 32 is substantially symmetric about a plane P passing through the axis of the turbine. 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 the radial direction so as to be passed through by accessories and for passing these accessories through ducts in the engine.
[0009] For some types of turbines, such as multi-stream or variable cycle turbines, it is very 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 flows (see Figure 2a), rather than being located between the rotor 30 and the separator 24, in order to reduce the length of the module between the concept shown in Figure 1a and the concept shown in Figure 2a. The stator vane assembly 34 will include a plurality of vanes distributed around the axis of the turbine. As described above and as shown in Figure 2b, each of these vanes will have an aerodynamic profile whose cross-section includes a pressure side 34a and a suction side 34b (Figure 2b), so that the aerodynamic profile is an asymmetric profile, which is not the case for the arm 32 shown in Figure 1a. The stator vane assembly 34 will extend radially through the main duct 16. These vanes will include a leading edge 36 located upstream of the splitter nose 24a in the main duct 16, and trailing edges, the inner trailing edge 38a and the outer trailing edge 38b being located in the inner duct 26 and the outer duct 28 respectively. The stator vane assembly may be connected to the splitter nose 24a.
[0010] The stator vane assembly 34 will impose a specific direction on the airflows 16, 20, 22. However, in the case of a variable cycle turbine, it is very useful to provide a variable geometry downstream of the rotor vane assembly 30 in order to be able to adapt to different operating modes of the turbine and changes in the bypass ratio. However, for reasons of overall dimensions, it would be complex to add a variable pitch vane assembly downstream of the stator vane assembly 34. In fact, such an addition would require lengthening the axial dimension of the turbine, which would result in an increase in the mass of the turbine and a reduction in performance.
[0011] In addition, due to reasons of noise pollution, the stator vane assembly 34 cannot be moved axially closer to the rotor vane assembly 30.
[0012] In the present application, a variable cycle turbine means a turbine as described below: by controlling the variable geometry of the turbine, the specific thrust of the turbine can be varied at a given engine speed. An example of a variable geometry is a variable pitch stator vane assembly. In the present application, a vane assembly means an annular array of vanes.
[0013] Accordingly, the present invention proposes to optimize a turbine as shown in Figure 2a so that it can be used in a variety of configurations and, in particular, in the context of multi-channel (at least two channels) and / or variable cycle turbines. Summary of the Invention
[0014] The present invention proposes a three-channel aircraft turbine comprising a gas generator which includes along a longitudinal axis at least one compressor, a combustion chamber and at least one turbine. The turbine further comprises:
[0015] - two coaxial annular walls, namely an inner annular wall and an outer annular wall, the inner annular wall and the outer annular wall extending around each other and defining therebetween a main annular duct for the flow of a main air stream,
[0016] - a rotor blade assembly which extends radially through the main duct and forms a ducted propeller,
[0017] - an annular separator which is provided downstream of the rotor blade assembly and is located between the two walls. The separator together with the inner wall and the outer wall respectively defines two secondary annular ducts, namely an inner secondary duct and an outer secondary duct, for the flow of secondary air streams, namely an inner secondary air stream and an outer secondary air stream. The separator includes at its upstream end an annular diverter nose which is configured to divide the main air stream into two parts and form the secondary air streams,
[0018] - stator elements which on the one hand extend radially through the main duct and on the other hand extend radially through the secondary ducts,
[0019] and
[0020] - a non-ducted propeller arranged upstream of the outer wall,
[0021] Characterized in that the stator elements include first variable pitch stator vanes which are distributed around the axis, and each of the first variable pitch stator vanes includes a leading edge and a trailing edge located upstream of the diverter nose. The trailing edge includes an inner trailing edge and an outer trailing edge respectively located in the inner secondary duct and the outer secondary duct,
[0022] And the turbine further includes fixed stator vanes which are distributed around the axis in the outer secondary duct and are located downstream of the outer trailing edge of the first variable pitch vanes.
[0023] Therefore, the present invention proposes to provide variable pitch stator blades at the splitter. In order to enable angular displacement of these blades around their pitch axis, it will be understood that the blades will be separated from the splitter nose and the separator by a small gap in order to limit gas leakage in these areas.
[0024] The fixed stator blades are associated with variable pitch blades and are located in the outer secondary duct. This configuration optimizes the operation of the turbine, allowing multi-flow or variable cycle applications, while limiting the impact on the axial length or size and mass of the turbine. In fact, the provision of variable pitch blades at the splitter nose makes it possible to reduce the flow distance between the rotor and the splitter nose in the axial direction, while also allowing to vary the airflow flowing in the inner and outer secondary ducts.
[0025] In this application, "annular" means a shape of revolution about an axis, which shape may be continuous or discontinuous.
[0026] In addition, in the present application, a "variable pitch" element refers to an element in which at least a portion of the element has an adjustable position around an axis referred to as a pitch axis. The entire element or only a portion thereof may be variable pitch. In the case of a blade, for example, the blade may be a single piece and have an adjustable position around a wedge-shaped axis. Alternatively, the position of only one portion thereof (e.g., a leading edge or a trailing edge) may be adjusted relative to the rest of the blade around the pitch axis. In the case where a blade assembly includes a plurality of blades, each blade has an adjustable position around its own pitch axis. Therefore, for the same blade assembly, there are as many pitch axes as blades with variable pitch. Each of these axes may have a radial orientation or an inclined orientation relative to the longitudinal axis of the turbine.
[0027] The turbine may include one or more of the following features, which may be independent of each other or in combination:
[0028] - the stator element further comprises a second variable pitch stator blade located in the inner secondary duct;
[0029] - The second variable pitch blades include a leading edge and a trailing edge. The leading edges of these second variable pitch blades are directly 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; thus, the first variable pitch blades and the second variable pitch blades are very close axially, such that within the meaning of the present invention they are considered to be components forming a stator element; the aforementioned axial gap between these blades is preferably as small as possible. Minimizing these axial gaps enables restricting 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; thus it should be understood that the gas flowing through the pressure side of the first variable pitch blade must subsequently flow through the pressure side of the second variable pitch blade, and the gas flowing through the pressure side of the first variable pitch blade must subsequently flow through the pressure side of the second variable pitch blade; when the gap is large, a portion of the air flow flowing through the suction side of the first variable pitch blade subsequently flows to the pressure side of the second blade and will cause energy to be added to the fluid flowing through the pressure side of the first blade;
[0030] - The fixed stator blades include a leading edge, and the leading edge of the fixed stator blades is spaced apart from the trailing edge of the first variable pitch blades by a predetermined axial gap;
[0031] - Preferably, the gap is less than 10 mm, and more preferably, less than or equal to 5 mm;
[0032] - The number of the second variable pitch blades is equal to the number of the first variable pitch blades;
[0033] - The number of the second variable pitch blades is a multiple of the number of the first variable pitch blades;
[0034] - The number of the fixed blades is equal to the number of the first variable pitch blades;
[0035] - The number of the fixed blades is a multiple of the number of the first variable pitch blades;
[0036] - The turbine further includes at least one system for controlling the pitch angle of the variable pitch blades;
[0037] - The control system is installed in the separator or is located radially outside the outer wall;
[0038] - At least some of the fixed blades have a different profile from other fixed blades and thus form a multi-profile blade grid;
[0039] - The rotor blade assembly is a propulsion fan or a compressor rotor blade assembly; and
[0040] - The fixed stator blade includes a pressure side and a suction side, and the variable pitch stator blade includes a pressure side and a suction side.
[0041] The invention also relates to an aircraft (in particular a transport aircraft) comprising a turbine as described above. Description of the Drawings
[0042] Other features and advantages of the present invention will become apparent from the following detailed description, and the description is facilitated by reference to the accompanying drawings, in which:
[0043] [Figure 1a] Figure 1a is a very schematic half view of an axial section of an aircraft turbine according to the prior art before the present invention;
[0044] [Figure 1b] Figure 1b is a very schematic cross-sectional view of the arm of the turbine of Figure 1a;
[0045] [Figure 2a] Figure 2a is a very schematic half view of an axial section of an aircraft turbine;
[0046] [Figure 2b] Figure 2b is a very schematic cross-sectional view of the stator blade of the turbine of Figure 2a;
[0047] [Figure 3a] Figure 3a is a very schematic half view of an axial section of an aircraft turbine according to a first embodiment of the present invention;
[0048] [Figure 3b] Figure 3b is a very schematic cross-sectional view of a variable pitch stator blade, the fixed stator blade of the turbine of Figure 3a following immediately behind the variable pitch stator blade, and two different positions of the pitch of the variable pitch stator blade being shown on the left and right sides of the figure respectively;
[0049] [Figure 3c] Figure 3c is a view similar to the left side of Figure 3b and shows an alternative embodiment of the present invention;
[0050] [Figure 4a] Figure 4a is a very schematic half view of an axial section of an aircraft turbine according to a second embodiment of the present invention;
[0051] [Figure 4b] Figure 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 Figure 4a following immediately behind the first variable pitch stator blade, and two different pitch positions of these blades being shown on the left and right sides of the figure respectively;
[0052] [Figure 4c] Figure 4c is a view similar to the left side of Figure 4b and shows an alternative embodiment of the present invention;
[0053] Figure 5 Figure 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 grille consists of at least two different blade profiles; and
[0054] Figure 6 Figure 6 is a schematic view of a three-flow turbine according to the invention. Detailed implementation mode
[0055] FIGS. 1a, 1b, 2a and 2b have been described above.
[0056] Reference Figure 6 , the turbine 10 is of the three-flow type and generally includes a gas generator 2 which includes at least one compressor, a combustion chamber and at least one turbine along a longitudinal axis X. The turbine includes a ducted propeller H1 and an unducted propeller H2. The propeller H1 is surrounded by a nacelle 4 which extends downstream of the propeller H2 about the axis X. The air flow passing 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. Then the main flow F2 is divided into two other flows F1, F2 as described below.
[0057] In the present application, as shown in FIGS. 3a and 3b, the turbine 10 thus includes two coaxial annular walls which are an inner annular wall 12 and an outer annular wall 14 respectively. The inner annular wall and the outer annular wall extend around each other and define a main annular duct 16 for a main air flow 18 therebetween.
[0058] The main air flow 18 is divided into two secondary air flows by an annular separator 24 disposed between the two walls 12, 14. The two secondary air flows are an inner secondary air flow 20 and an outer secondary air flow 22 respectively. The separator 24 includes an annular diverter nose 24a at an upstream end, and the diverter nose is configured to separate the main air flow 18 into two parts and form the secondary air flows 20, 22.
[0059] The rotor blade assembly 30 extends radially through the main duct 16 and thus through the upstream of the separator 24. In the Figure 6 turbine, the rotor blade assembly 30 forms the ducted propeller H1.
[0060] The stator elements are located downstream of the rotor blade assembly 30 and at the diverter 24a.
[0061] According to the invention, these stator elements include a first variable pitch stator vane 40.
[0062] In addition, a fixed stator vane 42 is located in the outer secondary duct 28 at a position downstream of the first variable pitch stator vane 40.
[0063] The first variable pitch fan blades 40 are distributed around an axis, and each first variable pitch fan blade includes a leading edge 40a upstream of the splitter nose 24a and a trailing edge, the trailing edge being an inner trailing edge 40b and an outer trailing edge 40c located in the inner secondary duct 26 and the outer secondary duct 28 respectively. Thus, it can be understood that the first variable pitch fan blades 40 are located at the position of the splitter nose 24a, as can be seen in the figure. There are gaps (not shown) between the splitter nose 24a and the first variable pitch fan blades 40 to allow them to move. These gaps are preferably as small as possible to limit or prevent gas from passing between these fan blades 40 and the nose 24a. It can also be seen that the leading edge 42a can be inclined and extend outward from upstream to downstream. The inclination is determined, for example, based on a compromise between the size of the engine and the optimization of the noise generated by it. 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.
[0064] Figure 3b shows that each of the first variable pitch fan blades 40 has an aerodynamic profile and includes 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 certain curvature along its chord. The region of the maximum curvature of the variable pitch fan blade 40 is referred to as C. This region is preferably upstream of the splitter nose 24a.
[0065] Preferably, the first variable pitch fan blades 40 are all identical. Preferably, their leading edges 40a are intersected by the same transverse plane.
[0066] The number of the first variable pitch fan blades 40 is, for example, between 10 and 200.
[0067] Each of the first variable pitch fan blades 40 can rotate about a pitch axis Y (the pitch axis has a generally radial orientation). The rotation of each first variable pitch fan blade 40 is achieved by means of a control system 50, which is located radially outside the outer wall 14. This is advantageous because it places the system in a relatively cool environment (compared to the high temperatures that can generally prevail in the gas generator). In addition, this environment is not very restrictive and contains free space to accommodate this type of system.
[0068] The fixed fan blades 42 are distributed around an axis in the outer secondary duct 28. They each include a leading edge 42a downstream of the splitter nose 24a and a trailing edge 42b located in the outer secondary duct 28.
[0069] Figure 3b shows that each of the fixed fan blades 42 has an aerodynamic profile and includes a pressure side 46 (concave curved shape) and a suction side 48 (convex curved shape). In addition, each fixed fan blade 42 has a certain curvature along its chord.
[0070] The number of fixed blades 42 is equal to the number of first variable pitch blades 40, or is a multiple of the number of first variable pitch blades 40, and the fixed blades 42 are directly downstream of the first variable pitch blades 40 and are axially successive with 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 10 mm, and more preferably, less than or equal to 5 mm. Preferably, these gaps I are less than 10% of the chord of the blade 40 or the 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 relevant edges 40c, 42a and over the entire radial extent of the outer duct 28. Of course, these gaps I may vary during operation depending on the pitch position of the blade 40 relative to the blade 42.
[0071] Preferably, all of the fixed blades 42 are identical. Preferably, their leading edges 42a are in the same transverse plane and are traversed by the same transverse plane.
[0072] The number of fixed blades 42 is, for example, between 10 and 200.
[0073] The left side of Figure 3b shows the first angular position or first pitch position of the first variable pitch blades 40, and the right side shows the second angular position or second pitch position of these blades. For example, the first variable pitch blades 40 can move around their axis Y within an angular range of approximately 60°.
[0074] Figure 4c shows an alternative embodiment, in which the number of fixed blades 424 is a multiple of the number of first variable pitch blades 40. The multiple is, for example, 2, 3, 4, etc.
[0075] Figures 4a and 4b show a second embodiment of the present invention, which is essentially different from the previous embodiment in that the turbine further includes a plurality of second variable pitch stator blades 44 located downstream of the trailing edge 40b of the first variable pitch blades 40 in the inner secondary duct 26.
[0076] Each of the second variable pitch stator blades 44 includes a leading edge 44a located downstream of the splitter nose 24a and a trailing edge 44b located in the inner secondary duct 26.
[0077] Each second variable pitch blade 44 has an aerodynamic profile and includes a pressure side and a suction side. In addition, each variable pitch blade 44 has a certain curvature along its chord.
[0078] The number of second variable pitch blades 44 may be equal to the number of first variable pitch blades 40.
[0079] The second variable pitch blade 44 is located directly downstream of the fixed blade 42 and is axially successive with the fixed blade 42. The leading edge 44a of the second variable pitch blade 44 is spaced apart from the trailing edge 42c of the fixed blade 42 by a predetermined axial clearance J. Preferably, these clearances J are less than 10 mm, and more preferably, less than or equal to 5 mm. Preferably, these clearances J are less than 10% of the chord of the blade 40 or 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 relevant edges 40b, 44a and over the entire radial extent of the inner duct 26. Of course, these clearances J may vary during operation depending on the pitch positions of the blades 40, 44.
[0080] Preferably, the second variable pitch blades 44 are all identical. Preferably, their leading edges 44a lie in the same transverse plane and are traversed by the same transverse plane.
[0081] 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). The rotation of each second variable pitch blade 44 is achieved by a control system 50' located in the separator 24.
[0082] Figure 4c shows an alternative embodiment in which the number of variable pitch stator blades 44 is a multiple of the number of first variable pitch blades 40. The multiple is, for example, 2, 3, 4, etc.
[0083] Figure 5 A third embodiment of the invention is shown, which differs essentially from the previous embodiments in that the fixed blades 42 are not all identical. The fixed blades 42 have at least two types, which differ from each other in terms of dimensions and / or geometry and / or camber, etc. The different types of fixed blades 42 are evenly distributed around the axis, so as to obtain a cyclic distribution of these fixed blades 42 around the axis.
[0084] In summary, the present invention is applicable to any turbine in which the main flow is separated into two secondary flows downstream of the ducted rotor blade assembly.
Claims
1. Three - flow - path aircraft turbine (10), comprising a gas generator which includes along a longitudinal axis at least one compressor, a combustion chamber and at least one turbine, and the turbine further includes: - Two coaxial annular walls, namely an inner annular wall (12) and an outer annular wall (14), The inner annular wall and the outer annular wall extend around each other and define therebetween a main annular duct (16) for the flow of a main air flow (18), - A rotor blade assembly (30) which extends radially through the main duct (16) and forms a ducted propeller (H1), - An annular separator (24) which is arranged downstream of the rotor blade assembly (30) and is located between the two walls (12, 14), and the separator (24) Together with the inner wall and the outer wall (12, 14) respectively defines two secondary annular ducts, That is, an inner secondary duct (26) and an outer secondary duct (28), and the two secondary annular ducts are respectively for the flow of secondary air flows, and the secondary air flows are an inner secondary air flow (20) And an outer secondary air flow (22), and the separator (24) includes an annular diverter nose (24a) at its upstream end, and the annular diverter nose is configured to divide the main air flow (18) into two parts and form the secondary air flows (20, 22), - A stator element which on the one hand extends radially through the main duct (16), And on the other hand extends radially through the secondary ducts (26, 28), and - A non - ducted propeller (H2) arranged upstream of the outer wall (14), Characterized in that the stator element includes a first variable - pitch stator vane (40) which is distributed around the axis, and each of the first variable - pitch stator vanes includes a leading edge (40a) located upstream of the diverter nose (24a) and a trailing edge, and the trailing edge is an inner trailing edge (40b) and an outer trailing edge (40c) respectively located in the inner secondary duct (26) and the outer secondary duct (28), And, the turbine (10) further includes fixed stator vanes (42) which are distributed around the axis in the outer secondary duct (28) and are located downstream of the outer trailing edge (40c) of the first variable - pitch vanes (40).
2. The turbine (10) according to claim 1, wherein, The stator element further includes a second variable - pitch stator vane (44) located in the inner secondary duct (26).
3. The turbine (10) according to claim 2, wherein, The second variable - pitch vanes (44) include a leading edge (44a) and a trailing edge (44b), and the leading edges (44a) of these second variable - pitch vanes (44) are directly located downstream of the inner trailing edge (40b) of the first variable - pitch vanes (40) and are spaced apart from these trailing edges (40b) by a predetermined axial gap (J).
4. The turbine (10) according to any one of the preceding claims, wherein, The fixed stator vanes (42) include a leading edge (42a), and the leading edge of the fixed stator vanes is spaced apart from the trailing edge (40c) of the first variable - pitch vanes (40) by a predetermined axial gap (I).
5. The turbine (10) according to any one of claims 2 to 4, wherein, The number of the second variable pitch fan blades (44) is equal to the number of the first variable pitch fan blades (40), or is a multiple of the number of the first variable pitch fan blades (40).
6. The turbine (10) according to any one of the preceding claims, wherein, The number of the fixed fan blades (42) is equal to the number of the first variable pitch fan blades (40), or is a multiple of the number of the first variable pitch fan blades (40).
7. Turbine (10) according to any one of the preceding claims, wherein, The turbine further includes at least one system (50, 50') for controlling the pitch angle of the variable pitch fan blades (42).
8. The turbine (10) according to claim 7, wherein, The control system (50, 50') is installed in the separator (24) or is located radially outside the outer wall (14).
9. The turbine (10) according to any one of the preceding claims, wherein, At least some of the fixed fan blades (42) have a profile different from that of the other fixed fan 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.