Three-flow aircraft turbine engine
By adopting a combined structure of fixed guide wheel blades and variable pitch guide wheel blades in the turbine engine, the problem of reduced length and performance in the turbine engine in the variable cycle turbine engine is solved, and noise reduction and performance optimization are achieved.
Patent Information
- Application Number
- CN202280102897.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2022-12-05
- Publication Date
- 2025-08-12
AI Technical Summary
When existing turbine engines add stator blades to adapt to different operating states under variable cycles, the length of the turbine engine and the performance of the turbine engine is increased and the noise problem is difficult to solve.
The combined structure of fixed guide wheel blades and variable pitch guide wheel blades is adopted. The fixed guide wheel blades are connected to the annular nose. The variable pitch guide wheel blades are closely spaced or interlocked to optimize the gas flow path to reduce the impact of axial dimensions and mass.
In multi-flow or variable cycle turbine engines, the operation of the turbine engine is optimized, limiting the length and mass impact of the turbine engine while reducing noise and improving performance.
Smart Images
Figure CN120476244A_ABST
Abstract
Description
Technical Field
[0001] The present invention relates to the general field of aviation and more particularly to a three-flow aircraft turbine engine. Background Art
[0002] Typically, an aircraft turbine engine comprises a gas generator comprising, along a longitudinal axis, at least one compressor, a combustion chamber and at least one turbine.
[0003] Air enters the gas generator and is compressed in one or more compressors. This compressed air is mixed with fuel and combusted in a combustion chamber. The combustion gases expand in one or more turbines. This expansion causes one or more turbine rotors to rotate, which in turn drives one or more compressor rotors. The combustion gases are discharged through a nozzle to provide thrust, which may be in addition to the thrust provided by at least one propulsive ducted or non-ducted propeller or fan of the turbine engine.
[0004] The gas flow flows in the turbine engine through an annular duct. Thus, as can be seen in FIG1 a, the turbine engine 10 comprises coaxial annular walls (inner annular wall 12 and outer annular wall 14 respectively) extending around each other and defining between them a main annular flow duct 16 for the main gas flow 18.
[0005] When the primary gas flow 18 is divided into two secondary gas flows (an inner secondary gas flow 20 and an outer secondary gas flow 22), an annular separator 24 is arranged between the two walls 12, 14 and defines two secondary annular flow ducts (an inner secondary annular flow duct 26 and an outer secondary annular flow duct 28) for the secondary gas flows 20, 22 with these walls 12, 14, respectively. The separator 24 includes an annular nose 24a at its upstream end, which is configured to divide the primary gas flow 18 into two parts and form the secondary gas flows 20, 22.
[0006] The rotor blade section 30 may extend radially through the main duct 16 upstream of the separator 24 .
[0007] As shown in FIG. 1 a , a structural arm 32 may extend radially through the main duct 16 downstream of the rotor blade portion 30 and upstream of the separator 24 .
[0008] As used in this application, an arm 32 or structural arm refers to a stator element having a generally aerodynamic cross-sectional shape such as that shown in FIG1 b, but not including a pressure side or a suction side. The arm 32 is not to be compared to a 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 axis of the turbine engine. The number of arms 32 is typically less than ten and may be four. At least one of the arms 32 may be hollow and tubular in the radial direction to allow for the passage of auxiliary devices and for passing these auxiliary devices through ducts in the engine.
[0009] For certain types of turbine engines (e.g., multi-flow or variable cycle turbine engines), it is useful to position the stator blade section 34 directly downstream of the rotor blade section 30 and integrated into the flow separator nose 24a, rather than between the rotor 30 and the separator 24 (see FIG2a ), in order to reduce the module length between the concept shown in FIG1a and the concept shown in FIG2a . The stator blade section 34 will include a plurality of blades distributed around the axis of the turbine engine. As described above and shown in FIG2b , the cross-section of each of these blades will have an aerodynamic profile comprising a pressure side 34a and a suction side 34b ( FIG2b ), i.e., an asymmetric profile, which is not the case with the arms 32 visible in FIG1a . The stator blade section 34 will extend radially through the main duct 16 . In the case where the nose 24a is to be connected to the vanes of the stator blade section 34, these vanes will include a leading edge 36 and a trailing edge (inner trailing edge 38a and outer trailing edge 38b, respectively), with the leading edge 36 being located upstream of the nose 24a in the main duct 16 and the trailing edges being located in the inner duct 26 and outer duct 28.
[0010] The stator blade sections 34 impose a specific direction on the gas flows 16 , 20 , 22 . However, in the case of a variable cycle turbine engine, it is useful to provide a variable geometry downstream of the stator blade sections 34 to accommodate different operating conditions and variations in the bypass ratio of the turbine engine. However, for reasons of overall size, adding variable pitch blade sections downstream of the stator blade sections 34 can be complex. Indeed, such an addition would require extending the axial dimension of the turbine engine, which would result in an increase in the turbine engine's mass and a decrease in its performance.
[0011] Furthermore, due to noise reasons it would not be possible to move the stator blade portion 34 axially closer to the rotor blade portion 30 .
[0012] In this application, a variable cycle turbine engine is a turbine engine that can vary its specific thrust at a given engine speed by controlling the turbine engine's variable geometry. An example of a variable geometry structure is a variable pitch stator blade section. In this application, a blade section is defined as an annular row of vanes.
[0013] Therefore, the present invention proposes optimizing a turbine engine such as that shown in FIG. 2 a so that it can be used in multiple configurations and in particular in the context of a multi-flow turbine engine (at least two) and / or a variable cycle turbine engine. Summary of the Invention
[0014] The present invention proposes a three-flow aircraft turbine engine, comprising a gas generator comprising, along a longitudinal axis, at least one compressor, a combustion chamber, and at least one turbine, the turbine engine further comprising:
[0015] - 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,
[0016] - a rotor blade section which extends radially through the main duct and forms a ducted propeller,
[0017] an annular separator, the annular separator being arranged downstream of the rotor blade section and between the two walls, the separator defining two secondary flow annular ducts for secondary air flows with the inner wall and the outer wall, respectively, the secondary air flows being an inner secondary air flow and an outer secondary air flow, the two secondary flow annular ducts being an inner secondary flow annular duct and an outer secondary flow annular duct, the separator comprising an annular nose at an upstream end, the annular nose being configured to divide the primary air flow into two parts and form the secondary air flows,
[0018] - stator elements extending radially through the primary duct on the one hand and through the secondary duct on the other hand, these stator elements being connected to the annular nose,
[0019] as well as
[0020] - a non-ducted propeller, arranged upstream of the outer wall, characterized in that the stator element comprises:
[0021] - fixed guide vanes distributed around the axis, each comprising a leading edge and a trailing edge, the trailing edges being respectively an inner trailing edge and an outer trailing edge, the leading edge being located upstream of the nose and the trailing edges being respectively located in the inner secondary duct and the outer secondary duct, these fixed guide vanes being connected to the nose, and
[0022] - variable pitch guide vanes distributed around the axis, the variable pitch guide vanes extending radially through at least one of the secondary ducts, each of the variable pitch guide vanes comprising a leading edge and a trailing edge,
[0023] and
[0024] - the leading edge of the variable-pitch guide vane is located upstream of the inner and / or outer trailing edge of the fixed guide vane,
[0025] or
[0026] The leading edge of the variable-pitch guide vane is located directly downstream of the inner and / or outer trailing edge of the fixed guide vane, the leading edge of the variable-pitch guide vane being separated from these trailing edges by a predetermined axial gap.
[0027] Therefore, the present invention proposes using fixed guide vanes and variable-pitch guide vanes to replace the arms in FIG. 1a or the stator blade sections in FIG. 1b . The fixed guide vanes and variable-pitch guide vanes are axially closely spaced or axially interlocked, such that the fixed guide vanes and variable-pitch guide vanes can be considered to form a component of a stator element within the meaning of the present invention. In practice, the leading edge of the variable-pitch guide vane is located upstream of the trailing edge of the fixed guide vane, or the variable-pitch guide vane is separated from the trailing edge of the fixed guide vane by a predetermined axial gap, preferably as small as possible. By minimizing this axial gap, the passage of gas between the trailing edge of the fixed guide vane and the leading edge of the variable-pitch guide vane is restricted or prevented during operation. Therefore, it should be understood that gas flowing on the pressure side of the fixed guide vane must flow on the pressure side of the variable-pitch guide vane, and gas flowing on the suction side of the fixed guide vane must flow on the suction side of the variable-pitch guide vane.
[0028] This configuration is particularly advantageous because it allows the operation of the turbine engine to be optimized for multi-flow or variable cycle applications while limiting the impact on the length, axial dimensions, and mass of the turbine engine. In fact, reducing the axial gap between the guide vanes and positioning them in the nose makes it possible to limit the impact of these vanes on the axial dimensions of the turbine engine.
[0029] Upstream of the rotor blade section located in the first duct there may be any configuration for a turbine engine.
[0030] In the present application, "annular" refers to a shape of rotation around an axis, which can be continuous or discontinuous. Furthermore, in the present application, a "variable pitch" element is defined as an element having a portion having a position that can be adjusted around an axis called the pitch axis. The entire element or only a portion of the element can have a variable pitch. For example, in the case of a blade, the blade can be a single piece and have an adjustable position around the pitch axis. Alternatively, the blade can include only a portion, such as a leading edge or a trailing edge, the position of which can be adjusted around the pitch axis relative to the rest of the blade. In the case where a blade section includes multiple blades, each of the blades has an adjustable position around its own pitch axis. For the same blade section, the number of pitch axes is the same as the number of variable pitch blades. Each of these axes can have a radial or inclined orientation relative to the longitudinal axis of the turbine engine.
[0031] The turbine engine may include one or more of the following features, taken alone or in combination with one another:
[0032] - the gap is less than 10 mm, preferably less than or equal to 5 mm;
[0033] - the clearance is less than 10% of the chord of the fixed blade or the variable pitch blade, preferably less than or equal to 5% of the chord;
[0034] --The fixed guide vane includes a pressure side and a suction side, and the variable pitch guide vane includes a pressure side and a suction side;
[0035] - the number of the variable-pitch guide blades is greater than or equal to the number of the fixed guide blades;
[0036] - the variable pitch guide vanes are located in the inner secondary duct;
[0037] - the trailing edge of the variable-pitch guide vane is located downstream of the outer trailing edge of the fixed guide vane;
[0038] - the turbine engine further comprises a system for controlling the angular pitch of the variable-pitch guide vanes, the system being installed in the separator;
[0039] - the turbine engine further comprises a system for controlling the angular pitch of the variable-pitch guide vanes, the system being mounted radially outside said outer wall;
[0040] - the variable pitch guide vanes are located in the outer secondary duct;
[0041] - a first variable pitch guide vane is located in the inner secondary duct, and a second variable pitch guide vane is located in the outer secondary duct;
[0042] - the turbine engine further comprises a common system for controlling the angular pitch of the first and second variable-pitch guide vanes, or independent systems for controlling the angular pitch of the first and second variable-pitch guide vanes, respectively;
[0043] - the turbine engine further comprises structural arms distributed around said axis in said outer secondary duct;
[0044] - the number of structural arms is less than the number of fixed guide vanes;
[0045] - a structural arm connected to some of said fixed guide vanes;
[0046] - the rotor blade section is a fan or compressor rotor blade section;
[0047] the leading edge of the variable-pitch guide vane is located at a distance from the inner and / or outer trailing edge of the fixed guide vane that is greater than 10% of the chord of one of these vanes, more preferably greater than or equal to 20% of this chord;
[0048] - some of the fixed guide vanes have a different profile or curvature than other fixed guide vanes;
[0049] - At least some of the arms have a different profile than the other arms.
[0050] The invention also relates to an aircraft, in particular a transport aircraft, comprising a turbine engine as described above. BRIEF DESCRIPTION OF THE DRAWINGS
[0051] Other features and advantages of the present invention will become apparent from the following detailed description and reference is made for the understanding of the description to the accompanying drawings, in which:
[0052] [ FIG. 1 a ] FIG. 1 a is a very schematic axial cross-sectional half view of an aircraft turbine engine according to the prior art;
[0053] [ FIG. 1 b ] FIG. 1 b is a very schematic cross-sectional view of an arm of the turbine engine of FIG. 1 a ;
[0054] [ FIG2 a ] FIG2 a is a very schematic axial sectional half view of a portion of an aircraft turbine engine;
[0055] [ FIG. 2 b ] FIG. 2 b is a very schematic cross-sectional view of a stator vane of the turbine engine of FIG. 2 a ;
[0056] FIG3 a is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a first embodiment of the invention;
[0057] FIG3b is a very schematic cross-sectional view of two fixed guide vanes preceding the two variable-pitch guide vanes of the turbine engine of FIG3a , and shows two different positions for arranging the variable-pitch guide vanes on the left and right sides of the figure, respectively;
[0058] FIG4a is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a second embodiment of the invention;
[0059] FIG4b is a highly schematic cross-sectional view of two fixed guide vanes preceding the three variable-pitch guide vanes of the turbine engine of FIG4a , and shows two different pitch positions of the variable-pitch guide vanes on the left and right sides of the figure, respectively;
[0060] FIG5 a is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a third embodiment of the invention;
[0061] FIG5b is a very schematic cross-sectional view of two fixed guide vanes of the turbine engine of FIG5a with two inserted variable-pitch guide vanes, and shows two different pitch positions of the variable-pitch guide vanes on the left and right sides of the figure, respectively;
[0062] [ Figure 6 ] Figure 6 is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a fourth embodiment of the invention;
[0063] [ Figure 7 ] Figure 7 is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a fifth embodiment of the invention;
[0064] [ Figure 8 ] Figure 8 is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a sixth embodiment of the invention;
[0065] [ Figure 9 ] Figure 9 is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a seventh embodiment of the invention;
[0066] [ Figure 10 ] Figure 10 is a highly schematic axial cross-sectional half view of an aircraft turbine engine according to an eighth embodiment of the invention;
[0067] [ Figure 11 ] Figure 11is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a ninth embodiment of the invention;
[0068] [ Figure 12 ] Figure 12 is a very schematic axial cross-sectional half view of an aircraft turbine engine according to a tenth embodiment of the invention;
[0069] [ Figure 13 ] Figure 13 is a highly schematic axial cross-sectional half view of an aircraft turbine engine according to an eleventh embodiment of the invention; and
[0070] [ Figure 14 ] Figure 14 is a schematic diagram of a three-flow turbine engine within the scope of the present invention. DETAILED DESCRIPTION
[0071] FIG. 1 a , FIG. 1 b , FIG. 2 a and FIG. 2 b have been described above.
[0072] refer to Figure 14 The turbine engine 10 is of the three-flow 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 engine further comprises a ducted propeller or fan, designated H1, and a non-ducted propeller or fan, designated H2. The propeller H1 is surrounded by a nacelle 4, which extends about the axis X downstream of the propeller H2. The air flow passing through the propeller H2 is divided 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 divided into two further flows F1 and F2, as described below. Within the scope of the present invention shown in Figures 3a and 3b, the turbine engine 10 comprises two coaxial annular walls (an inner annular wall 12 and an outer annular wall 14, respectively), which extend around each other and define between the two annular walls a main annular flow duct 16 for the main gas flow 18.
[0073] The primary gas flow 18 is divided into two secondary gas flows, an inner secondary gas flow 20 and an outer secondary gas flow 22, respectively, by an annular separator 24, which is arranged between the two walls 12, 14. The separator 24 comprises, at its upstream end, an annular nose 24a configured to divide the primary gas flow 18 into two parts and form the secondary gas flows 20, 22.
[0074] The rotor blade section 30 extends radially across the main conduit 16 upstream of the separator 24. Figure 14 Within the scope of the turbine engine, the rotor blade section 30 forms a ducted propeller H1.
[0075] The stator element 40 is located downstream of the rotor blade portion 30 and at the splitter nose 24a.
[0076] According to the invention, these stator elements 40 include fixed guide vanes 42 and variable-pitch guide vanes 44 .
[0077] The stationary vanes 42 are distributed about the axis, each including a leading edge 42a and a trailing edge (an inner trailing edge 42b and an outer trailing edge 42c, respectively). The leading edge 42a is located upstream of the nose 24a, while the trailing edges are located in the inner secondary duct 26 and the outer secondary duct 28, respectively. It will be appreciated that the stationary vanes 42 are connected to the nose 24a, as can be seen in the accompanying drawings. It will also be appreciated that the leading edge 42a may be inclined and extend outward from upstream to downstream. This inclination is determined, for example, based on a compromise between engine size and optimization of the noise generated by the engine. To minimize noise, it is desirable to increase the height at the top of the vanes, which results in a higher inclination of the vanes.
[0078] FIG3 b shows that each of the stationary vanes 42 has an aerodynamic profile and includes a pressure side 46 (concavely curved) and a suction side 48 (convexly curved). Furthermore, each of the stationary vanes 42 has a certain curvature along its chord. C denotes the region of the stationary vane 42 having the greatest curvature. Preferably, this region is located upstream of the nose 24 a.
[0079] Preferably, all the fixed blades 42 are identical. Preferably, the leading edges 42a of the fixed blades 42 are passed through by the same transverse plane. For example, the number of fixed blades 42 ranges from 10 to 200. The variable pitch blades 44 are distributed around the axis only in the inner secondary duct 26.
[0080] The variable pitch blades 44 each include a leading edge 44 a located downstream of the nose 24 a and a trailing edge 44 b located in the inner secondary duct 26 .
[0081] FIG3 b shows that each of the variable-pitch blades 44 has an aerodynamic profile and includes a pressure side 46 (concavely curved) and a suction side 48 (convexly curved). Furthermore, each of the variable-pitch blades 44 has a curvature along its chord. In this embodiment, the number of variable-pitch blades 44 is equal to the number of fixed blades 42, with the variable-pitch blades 44 located directly downstream of and axially extending from the fixed blades 42. The leading edges 44 a of the variable-pitch blades 44 are separated from the trailing edges 42 b of the fixed blades 42 by a predetermined axial gap J. Preferably, this gap J is less than 10 mm, more preferably less than or equal to 5 mm. Preferably, this gap J is less than 10% of the chord of the blade 42 or blade 44, more preferably less than or equal to 5% of the chord. Preferably, each of these gaps J is constant across the entire radial extent of the associated edge 42 b, 44 a, and therefore across the entire radial extent of the inner duct 26. Of course, these clearances J may vary during operation depending on the pitch position of blades 44 relative to blades 42 .
[0082] Preferably, the variable pitch blades 44 are all identical. Preferably, the leading edges 44a of the variable pitch blades 44 lie in or are passed through the same transverse plane.
[0083] For example, the number of variable-pitch blades 44 ranges from 10 to 200. Each of the variable-pitch blades 44 is rotatable about a pitch axis Y having a generally radial orientation. The rotation of each of the variable-pitch blades 44 is achieved by a control system 50 located in the separator 24.
[0084] 3 b shows a first angular or pitch position of the variable pitch blades 44 on the left and a second angular or pitch position of these blades on the right. For example, the variable pitch blades 44 can be displaced around their axis Y over an angular range of approximately 60°.
[0085] Figures 4a and 4b illustrate a second embodiment of the present invention. This second embodiment differs from the previous embodiment primarily in that the number of variable-pitch blades 44 is different from, and in this case, greater than, the number of fixed blades 42. In this embodiment, the number of variable-pitch blades 44 is twice the number of fixed blades 42. Therefore, it should be understood that the circumferential pitch between fixed blades 42 is twice the circumferential pitch between variable-pitch blades 44. Alternatively, the number of fixed blades 42 is equal to a multiple of the number of variable-pitch blades 44, where this multiple is not equal to 2, for example, 3, 4, or the like.
[0086] 3a and 3b, half of the variable pitch blade 44 extends downstream and is axially aligned with the fixed blade 42. The other half of the variable pitch blade 44 is inserted between the fixed blades 42 and therefore does not extend into the extension of the fixed blade 42.
[0087] Preferably, the variable pitch blades 44 are all identical. Preferably, as in the case of the fixed blades 42, the leading edges 44a of the variable pitch blades 44 lie in or are crossed by the same transverse plane.
[0088] 5 a and 5 b show a third embodiment of the present invention, which differs from the first embodiment primarily in the positioning of the variable-pitch blades 44 relative to the fixed blades 42 .
[0089] The variable-pitch blades 44 are axially inserted between the fixed blades 42 and arranged between these blades 42. The variable-pitch blades 44 are not located in the axial extension of the fixed blades 42, but are angularly offset by half the pitch relative to the axis of the turbine engine, so that each variable-pitch blade 44 is located midway between two fixed blades 42. The leading edge 44a of the variable-pitch blade 44 is located upstream of the trailing edge 42b of the fixed blade 42. The trailing edge 44b of the variable-pitch blade 44 is located downstream of the trailing edge 42b of the fixed blade 42.
[0090] The interlocking distance of the variable pitch blades 44 between the fixed blades 42 is denoted as W, which can be estimated as a percentage of the chord of one of the blades 42 or one of the blades 44. Preferably, the distance W is greater than 10% of the chord of the blade 42 or the blade 44, and more preferably greater than or equal to 20% of the chord.
[0091] Figure 6 A fourth embodiment of the invention is shown, which differs from the first embodiment primarily in that the control system 50 is 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 occur in the gas generator. Furthermore, this environment is less constrained and contains sufficient free space to accommodate a system of this type.
[0092] The system 50 is connected to the variable pitch blades 44 and passes through the fixed blades 42. Thus, the blades 42 may extend in the axial direction and include an inner passage extending in the radial direction through the outer duct 28 to enable the system 50 to be installed and connected to the variable pitch blades 44. Thus, it should be understood that the trailing edges 42c of the fixed blades 42 may be located downstream of the trailing edges 42b of the blades.
[0093] Figure 7A fifth embodiment of the invention is shown, which differs from the first embodiment in the position of the variable pitch vanes 44. The variable pitch vanes 44 are distributed around the axis only in the outer secondary duct 28.
[0094] The variable pitch blades 44 each include a leading edge 44 a located downstream of the nose 24 a and a trailing edge 44 b located in the outer secondary duct 28 .
[0095] Each of the variable pitch blades 44 has an aerodynamic profile and includes a pressure side and a suction side. In addition, each of the variable pitch blades 44 has a certain curvature along its chord. The number of variable pitch blades 44 can be equal to or greater than the number of fixed blades 42, as discussed above with respect to Figures 3a to 4b.
[0096] The variable-pitch blades 44 are located directly downstream of the fixed blades 42 and in the axial extension of the fixed blades 42. The leading edges 44a of the variable-pitch blades 44 are separated from the trailing edges 42c of the fixed blades 42 by a predetermined axial gap J. Preferably, this gap J is less than 10 mm, more preferably less than or equal to 5 mm. Preferably, this gap J is less than 10% of the chord of the blade 42 or blade 44, more preferably less than or equal to 5% of this chord. Preferably, each of these gaps J is constant over the entire radial extent of the associated edge 42c, 44a, and therefore over the entire radial extent of the outer duct 28. Of course, these gaps J may vary during operation depending on the pitch position of the blade 44 relative to the blade 42.
[0097] Preferably, the variable pitch blades 44 are all identical. Preferably, the leading edges 44a of the variable pitch blades 44 lie in or are passed through the same transverse plane.
[0098] For example, the number of the variable-pitch blades 44 ranges from 10 to 200.
[0099] Each of the variable pitch blades 44 is rotatable about a pitch axis Y having a substantially radial orientation. The rotation of each of the variable pitch blades 44 is achieved by a control system 50 which is situated here radially outside the outer wall 14 .
[0100] Figure 8 3a and 3b or 4a and 4b, and the variable pitch blades 44 of the outer duct 28 may be similar to those described above with respect to FIG. Figure 7 a and Figure 7b are those variable pitch blades 44 described.
[0101] The angular pitch of the variable pitch blades 44 located in the two ducts is controlled by independent systems 50. A first control system 50 is located in the separator 24 and controls the pitch of the variable pitch blades 44 in the inner duct 26, and a second control system 50 is located radially outside the wall 14 and controls the pitch of the variable pitch blades 44 in the outer duct 28. Figure 9 In the seventh embodiment shown, a single control system 50 is used to control the angular pitch of the variable pitch blades 44 located in the two ducts 26, 28. This control system 50 is located radially outside the wall 14.
[0102] Figure 10 FIG. 8 shows an eighth embodiment of the present invention. The eighth embodiment differs from the first embodiment primarily in that not all stationary vanes 42 are identical. The stationary vanes 42 comprise at least two different types, each differing in size, geometry, and / or curvature. The different types of stationary vanes 42 are evenly distributed around the axis, resulting in a cyclic distribution of the vanes 42 around the axis.
[0103] exist Figure 11 In the ninth embodiment of the invention shown, the structural arms 32 are located in the outer duct 14 downstream of the trailing edges 42c of the stationary vanes 42. The number of arms 32 is smaller than the number of stationary vanes 42, and the arms 32 may extend in the axial extension of some of the stationary vanes 42. The arms 32 may all be identical. Figure 12 In the tenth embodiment of the invention shown, the structural arms 32 are shifted axially upstream toward each other and are connected to certain stationary vanes 42. Thus, the arms 32 are integral with the stationary vanes 42. The trailing edges 42 c of the stationary vanes 42 that are not connected to the arms 32 are located upstream of the trailing edges 32 a of the arms.
[0104] Arm 32 Figure 10 are all the same, arm 32 is Figure 11 In the eleventh embodiment, the distribution is different and cyclic. Figures 11 to 13 In an embodiment, for example, some arms 32 may be solid while other arms may be tubular for passing auxiliary devices from the outer wall 14 to the separator 24.
[0105] In general, the invention is applicable to any turbine engine in which a primary flow is divided into two secondary flows downstream of a ducted rotor blade section.
Claims
1. A three-flow aircraft turbine engine (10), comprising a gas generator comprising, along a longitudinal axis, at least one compressor, a combustion chamber, and at least one turbine, the turbine engine further comprising: - two coaxial annular walls, namely 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) for the main air flow (18), - a rotor blade section (30) which extends radially through the main duct (16) and forms a ducted propeller (H1), an annular separator (24) arranged downstream of the rotor blade section (30) and between the two walls (12, 14), the separator (24) defining with the inner wall and the outer wall (12, 14) two secondary flow annular ducts for secondary air flows, the secondary air flows being an inner secondary air flow (20) and an outer secondary air flow (22), the two secondary flow annular ducts being an inner secondary flow annular duct (26) and an outer secondary flow annular duct (28), the separator (24) comprising an annular nose (24a) at its upstream end, the annular nose being configured to divide the primary air flow (18) into two parts and form the secondary air flows (20, 22), - stator elements extending radially through the primary duct (16) on the one hand and through the secondary ducts (26, 28) on the other hand, these stator elements being connected to the annular nose (24a), and - a non-ducted propeller (H2) arranged upstream of the outer wall (14), characterized in that the stator element comprises: - fixed guide vanes (42) distributed around the axis, each comprising a leading edge (42a) and a trailing edge, the trailing edges being respectively an inner trailing edge (42b) and an outer trailing edge (42c), the leading edge being located upstream of the nose (24a) and the trailing edges being respectively located in the inner secondary duct (26) and the outer secondary duct (28), these fixed guide vanes (42) being connected to the nose (24a), and - variable pitch guide vanes (44) distributed around the axis, extending radially through at least one of the secondary ducts (26, 28), each of the variable pitch guide vanes (44) comprising a leading edge (44a) and a trailing edge (44b), and - the leading edge (44a) of the variable-pitch guide vane (44) is located upstream of the inner trailing edge (42b) and / or the outer trailing edge (42c) of the fixed guide vane (42), or - the leading edge (44a) of the variable-pitch guide vane (44) is located directly downstream of the inner trailing edge (42b) and / or outer trailing edge (42c) of the fixed guide vane (42), the leading edge of the variable-pitch guide vane being separated from these trailing edges (42b, 42c) by a predetermined axial gap (J).
2. The turbine engine (10) according to claim 1, wherein: The number of the variable-pitch guide blades (44) is greater than or equal to the number of the fixed guide blades (42).
3. A turbine engine (10) according to any one of the preceding claims, wherein: The variable pitch guide vanes (44) are located in the inner secondary duct (26).
4. A turbine engine (10) according to any one of the preceding claims, wherein: The trailing edge (44b) of the variable-pitch guide vane (44) is located downstream of the outer trailing edge (42c) of the fixed guide vane (42).
5. The turbine engine (10) according to claim 4, wherein: The turbine engine further comprises a system (50) for controlling the angular pitch of the variable-pitch guide vanes (42), which is mounted radially outside the outer wall (14).
6. The turbine engine (10) according to any one of claims 1 to 3, wherein: The variable pitch guide vanes (44) are located in the outer secondary duct (28).
7. The turbine engine (10) according to any one of claims 1 to 3, wherein: A first variable-pitch guide vane (44) is located in the inner secondary duct (26), and a second variable-pitch guide vane (44) is located in the outer secondary duct (28).
8. The turbine engine (10) according to claim 7, wherein: The turbine engine further includes a common system (50) for controlling the angular pitch of the first and second variable-pitch guide vanes (44), or independent systems (50) for controlling the angular pitch of the first and second variable-pitch guide vanes (44), respectively.
9. A turbine engine (10) according to any one of the preceding claims, wherein: The turbine engine further comprises structural arms (32) distributed around the axis in the outer secondary duct (28).
10. The turbine engine (10) according to claim 9, wherein: The number of the structural arms (32) is smaller than the number of the fixed guide vanes (42).
11. The turbine engine (10) according to claim 9 or 10, wherein: The structural arm (32) is connected to some of the fixed guide vanes (42).
12. The turbine engine (10) according to any one of the preceding claims, wherein: The rotor blade section (30) is a fan or compressor rotor blade section.
13. A turbine engine (10) according to any one of the preceding claims, wherein: The leading edge (44a) of the variable-pitch guide vane (44) is located at a distance (W) from the inner trailing edge (42b) and / or the outer trailing edge (42c) of the fixed guide vane (42), the distance being greater than 10% of the chord of one of these vanes (42, 44), more preferably greater than or equal to 20% of the chord.