Turbomachine comprising rows of stator vanes and a diffuser in a channel in which a third flow circulates

Abstract

The invention relates to a turbomachine (1), in particular an aircraft turbomachine, comprising: - a splitter nose (19) arranged in a primary duct (13) in which a primary flow circulates and intended to separate the primary flow into a radially inner flow (F11) and into a radially outer flow (F12); - stator equipment (26) installed in a channel (22) in which the radially outer flow (F12) circulates; and - first stator vanes (25) installed in the channel and intended to straighten the radially outer flow (F12) towards the equipment. According to the invention, the turbomachine comprises a diffuser (29) which is arranged downstream of the first stator vanes (25) and which is divergent between an inlet section and an outlet section, and second stator vanes (31) arranged downstream of the diffuser in order to deflect the radially outer flow at the outlet of the diffuser towards the equipment, the second stator vanes being situated upstream of the equipment.

Description

Description TITLE: TURBOMACHINE COMPRISING ROWS OF STATOR BLADES AND A DIFFUSER IN A CHANNEL WHERE A THIRD FLOW CIRCULATES. Technical field of the invention

[0001] The present invention relates to the field of turbomachines, and in particular multi-flow turbomachines for aircraft. Technological background

[0002] The prior art includes documents BE-A1-1029617, FR-A1-3120898, FR-A1-3109796 and US-A1-2021 / 108597.

[0003] A turbomachine generally comprises, from upstream to downstream, following the flow of gases in the turbomachine, a compressor assembly, an annular combustion chamber and a turbine assembly. A fan or at least one propeller is mounted upstream of the compressor assembly. The fan or propeller generates, on the one hand, a primary flow circulating in a primary vein and feeding the compressor assembly and, on the other hand, a secondary flow. The secondary flow circulates in a secondary vein or outside any fairing. The secondary flow generates the majority of the turbomachine's thrust.

[0004] The triple-flow turbomachine includes a tertiary flow or third flow which circulates in a tertiary vein and which is derived from the division of the primary flow. This tertiary flow also contributes to the thrust of the turbomachine and makes it possible to reduce the diameter of the fan or propeller while respecting the thrust specification. The turbomachine can then be installed under a low wing of an aircraft, such as an airplane. This tertiary flow can also make it possible to more efficiently cool the fluids used in the turbomachine for its operation thanks to heat exchangers which would be placed in the tertiary vein.

[0005] However, installing a heat exchanger can cause a pressure drop. Furthermore, the tertiary flow can pass through a row of stator blades which is placed in the tertiary vein and upstream of the heat exchanger. This row of stator blades, which is part of the compressor assembly, must in some cases achieve a very significant deviation of the tertiary flow in order to straighten the flow at its outlet. This deviation can be greater than 40° and lead to an increase in aerodynamic losses.

[0006] There is a need to address some or all of the above drawbacks. Summary of the invention

[0007] The objective of the present invention is to provide a solution for better distributing aerodynamic loads in a channel where a third flow circulates while avoiding impacting the mass and performance of the turbomachine.

[0008] We achieve this objective in accordance with the invention by means of a turbomachine, in particular for an aircraft, comprising: - a dividing nozzle arranged in a primary vein in which a primary flow circulates and intended to separate the primary flow into a radially internal flow and a radially external flow; - stator equipment installed in a channel in which the radially external flux circulates; and, - a first row of first stator blades installed in the channel and intended to straighten the radially external flow towards the equipment, the turbomachine comprising a diffuser which is arranged, downstream of the first stator blades and which is divergent between an inlet section and an outlet section, and a second row of second stator blades arranged, downstream of the diffuser, so as to deflect the radially external flow leaving the diffuser towards the stator equipment, the second stator blades being located upstream of the stator equipment.

[0009] Thus, this solution achieves the above-mentioned objective. In particular, the integration of two rows of stator blades in combination with a diffuser in the channel where the radially external flow (third flow) circulates allows a better distribution of the aerodynamic load throughout the channel. The highly deflecting row of blades is replaced by two separate first and second rows of stator blades, one of which is less deflecting than the other and produces a residual gyration at the inlet of the diffuser. The presence of a residual gyration at the inlet of the diffuser allows for a greater aerodynamic diffusion length and thus reduces the physical axial length of the diffuser as well as that of the second stator blades and the stator equipment. The second stator blades take up the remainder of the deflection to cancel any gyration. The stator equipment allows the radially external flow to be slowed down before to sweep the second stator blades. This therefore allows for a better compromise between aerodynamic performance, operability and length of the channel and the elements that equip it.

[0010] The turbomachine also includes one or more of the following features, taken alone or in combination: - the second stator blades are configured so as to achieve a straightening of the radially external flow greater than the straightening of the first stator blades and so as to cancel any possible gyration of the radially external flow at the outlet of the second stator blades. - each second stator blade comprises a blade having a leading edge and a trailing edge which are connected by an intrados surface and an extrados surface, each second stator blade has a metal angle at the leading edge and a metal angle at the trailing edge the difference of which is between 0° and 65° so as to have a possible gyration of the radially external flow equal or substantially equal to zero. - the metal angle at the trailing edge of each second stator blade is zero or substantially zero. - each first stator blade comprises a blade having a leading edge and a trailing edge which are connected by an intrados surface and an extrados surface, each first stator blade having a metal angle at the leading edge and a metal angle at the trailing edge, the difference of which is between 5° and 30° so as to produce a possible non-zero gyration of the radially external flow at the outlet of the first stator blades. - the stator equipment comprises at least one heat exchanger, an annular row of structural arms or at least one pylon. - the turbomachine comprises a propeller and a separation nozzle arranged between the propeller and the division nozzle along the longitudinal axis, the separation nozzle being intended to separate an air flow into the primary flow and a secondary flow coming from the propeller, the secondary flow circulating around the primary vein. - a rectifier comprising a plurality of stator vanes is arranged downstream of the propeller and is disposed across the secondary flow. - the equipment comprises an annular row of structural arms which is integrated into the second row of second stator blades, the structural arms each having an upstream edge, the upstream edges of the structural arms and the leading edges of the second stator blades being defined in the same plane. - the second stator blades comprise identical aerodynamic profiles or groups of aerodynamic profiles different from each other. - the diffuser comprises a radially external shell and a radially internal shell between which radially external flow passages are defined, each passage extending between an inlet opening and an outlet opening opposite the longitudinal axis and each passage has a section gradually increasing from the inlet opening to the outlet opening. - the diffuser is a smooth diffuser.

[0011] The invention also relates to an aircraft equipped with such a turbomachine. Brief description of the figures

[0012] The invention will be better understood, and other aims, details, characteristics and advantages thereof will appear more clearly on reading the detailed explanatory description which follows, of embodiments of the invention given as purely illustrative and non-limiting examples, with reference to the appended schematic drawings in which: - Figure 1 illustrates an axial section of an example of a turbomachine according to the invention; - Figure 2 shows schematically and in an axial and partial section, an example of configuration of means for straightening an air flow in a turbomachine according to the invention; - Figure 3 illustrates a detail of the rectification means and stator equipment of the turbomachine according to sectional plans of figure 2; - Figure 4 schematically represents the example of circulation of an air flow in a portion of a turbomachine according to the invention; - Figure 5 illustrates a schematic embodiment of the arrangement of a row of structural arms integrated into a row of stator blades in a turbomachine according to the invention; - Figure 6 is another embodiment of the arrangement of a row of structural arms integrated into a row of stator blades in a turbomachine according to the invention. Detailed description of the invention

[0013] Figure 1 represents a multi-flow turbomachine 1, intended to be mounted on an aircraft such as an airplane. The turbomachine 1 represented comprises an unducted propeller 2. Such a turbomachine is a turboprop and is known by the English expression "open rotor" or "unducted fan" or "open fan". In this category of turbomachine, there are those which have two unducted and counter-rotating propellers (known by the English acronym UDF for "Unducted Dual Fan") or those having a single unducted propeller and a rectifier which is also unducted and which comprises several stator blades (known by the English acronym USF for "Unducted Single Fan"). Of course, the invention applies to other types of turbomachine such as turbojets.

[0014] In the present invention, and in general, the terms "upstream", "downstream", "axial" and "axially" are defined with respect to the circulation of gases in the turbomachine and with respect to the longitudinal axis X of the turbomachine 1. Similarly, the terms "radial", "radially", "internal" and "external" are defined with respect to a radial axis Z perpendicular to the longitudinal axis X and with respect to the distance from the longitudinal axis X.

[0015] In Figure 1, the turbomachine 1 comprises a gas generator or gas turbine engine 3 which typically comprises, from upstream to downstream, a low pressure compressor or ("booster" in English) 5, a high pressure compressor (not shown), a combustion chamber (not shown), a high pressure turbine (not shown) and a low pressure turbine (not shown). The low pressure compressor 5 comprises an upstream part 4 which is located in a primary vein 13 as explained later.

[0016] The rotors of the low pressure compressor 5, of the upstream part 4 of the low pressure compressor 5 and of the low pressure turbine are mechanically connected by a low pressure shaft (not shown) so as to form a low pressure body. The rotors of the high pressure compressor 5, of its upstream part 4 and of the high pressure turbine are mechanically connected by a high pressure shaft (not shown) so as to form a high pressure body. The low pressure shaft extends inside the high pressure shaft and are coaxial with the longitudinal axis X.

[0017] The propeller 2 is mounted upstream of the gas generator 3 and preferably upstream of the low-pressure compressor 5 (and its upstream part). The propeller 2 comprises a plurality of blades 6 movable around the longitudinal axis X and extending radially from an internal casing 7 forming the hub of the propeller 2. The movable blades 6 may have variable pitch around a pitch axis transverse to the longitudinal axis X.

[0018] A rectifier 8 is arranged downstream of the propeller 2. The rectifier 8 comprises a plurality of stator vanes 9 (or fixed vanes) known by the English acronym “OGV” for Outlet Guide Vane. In the present invention, we understand by the term “stator vane” or “fixed vane”, a vane which is not driven in rotation about the longitudinal axis X of the turbomachine. The stator vanes 9 are distributed about the longitudinal axis X and are arranged downstream of the moving vanes 6 of the propeller 2 so as to straighten the air flow generated by them.

[0019] The stator vanes 9 of the rectifier 8 are variable pitch or fixed.

[0020] The air flow F which passes through the propeller 2 is split into a primary flow F1 and a secondary flow F2 by a separation nozzle 10. The latter is carried by an inlet casing 11 centered on the longitudinal axis X. The inlet casing 11 is fixed and is extended downstream by an external casing or inter-stream casing 12 which extends radially outside the internal casing 7.

[0021] The primary air flow F1 circulates in a primary vein 13 while the secondary flow F2 circulates radially outside the primary vein 13. In particular, the secondary flow F2 circulates radially outside the casings and sweeps the unducted rectifier 8.

[0022] The primary vein 13 is delimited radially by a radially internal wall 14 (carried by the external casing and the inlet casing) and a radially external wall 15 (carried by the internal casing 7).

[0023] Each compressor and each turbine comprise at least one stage composed, along the longitudinal axis X, of a blading of rotor blades 4a, 5a and a blading of stator blades 4b, 5b. The rotor blades 4a, 5a extend radially from the inner casing 7 (in particular from the radially outer wall 15) to which they are integral in rotation about the longitudinal axis X. The rotor blades may be of variable pitch. The stator blades 5b of the low-pressure compressor 5 extend radially between the outer casing 7 (in particular from the radially outer wall 15) and a dividing beak 19 to which they are integral. The stator blades may be of variable pitch. The dividing beak 19 comprises in particular a radially inner wall 20 and a radially outer wall 21.

[0024] In the present example, the upstream part 4 of the low pressure compressor is arranged upstream of the dividing nozzle 19. The stator vanes 4b of the upstream part 4 extend between the inner casing 7 and the outer casing 12.

[0025] Upstream of the low-pressure compressor 5 are arranged several inlet guide vanes 16 which are known by the English acronym “IGV” for Inlet Guide Vane. The inlet guide vanes 16 are arranged around the longitudinal axis X and are carried by a ferrule 17 extending the radially internal wall 14 upstream.

[0026] Casing arms 18 are advantageously arranged downstream of the inlet guide vanes 16 and allow the passage of services (not shown) for example. The casing arms 18 extend radially between the inlet casing 11 and the internal casing 7 to which they are secured. The casing arms 18 are mounted upstream of the low pressure compressor 4 (and in particular of the upstream part 4 of the low pressure compressor).

[0027] The primary flow F1 sweeps these inlet guide vanes 16 and the casing arms 18 before undergoing a first compression by the upstream part 4 of the low pressure compressor 5 then a second compression by the high pressure compressor 5. The casing arms 18 generally have no function of straightening the radially external flow.

[0028] With reference to Figure 2, the primary flow F1 is divided into a radially internal flow F11 and a radially external flow F12. The division is achieved by means of the dividing nozzle 19 which is annular. The latter is arranged in the primary stream 13. Advantageously, but not limitingly, the dividing nozzle 19 is located downstream of the separating nozzle 10 splitting the air flow F in two. The radially internal flow F11 circulates inside the dividing nozzle 19 while the radially external flow F12 circulates radially outside the dividing nozzle 19 in a channel 22. The channel 22 is advantageously delimited by the radially internal wall 14 (carried by the external casing) and the radially external wall 21 (carried by the dividing nozzle 19). The radially internal flow F11 continues its flow towards the combustion chamber and the turbines.

[0029] In Figure 2, the turbomachine 1 comprises a first row of first stator blades 25 and stator equipment 26 which are installed in the channel 22 where the radially external flow F12 circulates.

[0030] The stator equipment 26 is selected from a group comprising at least one heat exchanger, an annular row of structural arms or at least one pylon.

[0031] The first stator blades 25 are arranged upstream of the stator equipment 26. These first stator blades 25 are intended to straighten the flow radially external flow F12 towards the stator equipment 26. The first stator blades 25 are distributed, preferably regularly, around the longitudinal axis X. Each first stator blade 25 comprises a root (not shown) and a blade 27 which extends radially from the root. Each blade 27 is intended to extend through the radially external flow F12.

[0032] With reference to Figure 3 which is a detail of the sections of Figure 2, in the section planes AA and BB, each blade 27 of the first stator blades 25 comprises a leading edge 27a, a trailing edge 27b, a pressure surface 27i and an extrados surface 27e. The pressure and extrados surfaces 27i, 27e are opposite each other and are each connected by the leading edge 27a and the trailing edge 27b. The root of the blades 27 is secured to the radially external wall 21 of the dividing beak 19 and the head of the blade 27 is secured to the radially internal wall 14 of the external casing 12. Alternatively, the blades could be fixed only by their root or head to the internal casing 7 or external casing 12.

[0033] With reference to Figures 2 and 4, the turbomachine 1 comprises a diffuser 29 which is configured so as to slow down the radially external flow F12. The diffuser 29 is therefore installed in the channel 22. The diffuser 29 is arranged more precisely between the first row of the first stator blades 25 and the stator equipment 26. The diffuser 29 makes it possible to reduce the Mach number upstream of the equipment 26 and thus the aerodynamic pressure losses of the equipment 26 which is installed across the radially external flow F12.

[0034] In this embodiment, the diffuser 29 is annular and extends along the longitudinal axis X between an inlet section 29a and an outlet section 29b. The diffuser 29 is advantageously separate from the first stator vanes 25 and the equipment 26. Alternatively, the diffuser 29 is secured upstream to the first stator vanes 25 and downstream to the equipment 26.

[0035] The diffuser 29 advantageously comprises a radially outer shell and a radially inner shell which are centered around the longitudinal axis X. The radially outer shell extends outside the radially inner shell. In the present example, the radially outer shell defines a portion of the radially inner wall 14 of the outer casing 12 and the radially inner shell defines a portion of the radially outer wall 21 of the dividing nozzle. The shells define several passages for the radially outer flow which are distributed around the longitudinal axis. Each passage comprises an inlet opening, located at level of the inlet section 29a, and opposite the first stator blades 25. Each passage comprises an outlet opening, located at the level of the outlet section 29b, and opposite the equipment 26.

[0036] Advantageously, but not limitatively, the diffuser 29 is divergent between the inlet section 29a and the outlet section 29b. In other words, the section of the diffuser 29 increases from upstream to downstream. More precisely, each passage of the diffuser 29 has an increasing section from the inlet opening to the outlet opening, i.e. along the longitudinal axis. As shown in FIG. 2, the radially internal wall 14 widens from upstream to downstream (and in particular between the inlet section 29a and the outlet section 29b) generating this variation in section. This can be represented mathematically with the following formula: with R29ai4 the inner radius at the inlet section, R29bu the outer radius at the inlet section, R29t>i4 the inner radius at the outlet section, and R29b2i the outer radius at the outlet section. The radius being measured from the longitudinal axis and the radially inner or outer shell of the diffuser 29.

[0037] The variation of the section is progressive so as not to cause aerodynamic separations of the radially external flow F12.

[0038] According to Figure 4, the diffuser 29 has a predetermined length L1. The length L1 is measured between the inlet section 29a and the outlet section 29b. Advantageously, the length of the diffuser 29 is less than the length L2 measured between the trailing edge 27b of the first blades 25 and an upstream edge 30 of the equipment 26. Indeed, the diffuser 29 is configured so that the radially external flow moves along a helicoid H. This helical displacement of the radially external flow F12 is obtained thanks to an angle a2 in Figure 3 which is not zero. Such a displacement of the radially external flow F12 implies that the distance traveled by the radially external flow in the diffuser 29 is greater than the predetermined length L1 of the diffuser 29. Such a configuration makes it possible to have a diffuser 29 of reduced axial length L1, while maintaining a length traveled by the fluid equivalent to the case where a2 is greater than 0.Because the shorter the diffusion length, the greater the risk of detachment of the radially external flow.

[0039] Advantageously, the length of the diffuser 29 is defined in such a way that the ratio between the inlet 29a and outlet 29b sections and the length of the diffuser is between 1 and 5.

[0040] Advantageously, but not limited to, the diffuser 29 is smooth. In other words, the diffuser 29 does not include any physical element other than the ferrules delimiting the channel 22.

[0041] With reference to figures 2, 3 and 4, the turbomachine 1 comprises a second row of second stator vanes 31 which is intended to achieve a straightening of the radially external flow F12 at the outlet of the diffuser 29. The radially external flow is slowed at the outlet of the diffuser 29. The second stator vanes 31 are installed in the channel 22. Advantageously, the second stator vanes 31 are arranged downstream of the diffuser 29. In particular, the second row of second stator vanes 31 is arranged between the diffuser 29 and the equipment 26 along the longitudinal axis X. As can be seen in figure 2, the second stator vanes 31 are arranged in a divergent portion of the radially internal wall 14. This divergent portion corresponds to a portion of the radially external shell which extends downstream and at least partly envelops the equipment 26.In other words and with reference to Figure 4, the helical gyration or displacement extends beyond the outlet section 29b of the diffuser 29.

[0042] The second stator vanes 31 also have a blade 32 having a leading edge 32a, a trailing edge 32b which are connected by an internal intrados surface 32i and an extrados surface 32e.

[0043] The second stator blades 31 are configured so as to achieve a straightening of the radially external flow F12 that is greater than the straightening of the first stator blades 25 and so as to cancel any possible gyration of the radially external flow in the second stator blades 31. The straightening can be defined by the difference between the angles at the leading edges of the blades of the first and second blades, and the angles at the trailing edges of the blades of the first and second blades, for example [32 - [31 > a2- a1. The incidence at the outlet of the second stator blades is considered almost constant.

[0044] For this purpose, a predetermined metal angle o1 is fixed at the leading edge 27a of the blades 27 of the first stator vanes 25 and a predetermined metal angle a2 is fixed at the trailing edge 27b of the blades of the first stator vanes 25. The metal angle o1 at the leading edge 27a, respectively trailing edge 27b, is the angle formed between the general direction (parallel to the longitudinal axis X) of the radially external flow F12 and a straight line tangent to the skeleton line of the blade 27 at the leading edge, respectively trailing edge. The deflection or straightening of the radially external air flow F12 is slightly deflected by the first stator blades 25. The difference between the metal angle at the trailing edge and the metal angle at the leading edge, (a2 - a1), defines the deflection angle of each first stator blade 25 and is between 5° and 30°. With this low deflection, the first stator blades 25 are aerodynamically less loaded and are designed to be able to tolerate variations in incidence, in other words variations in the fluid angle defined as the angle between the direction of the flow and the longitudinal axis X. The residual gyration (angle between the flow velocity vector and the longitudinal axis) is not zero.

[0045] A predetermined metal angle 01 is also fixed to the leading edge 32a of each blade 32 of the second blades 31 and a predetermined metal angle 02 is fixed to the trailing edge 32b of each blade 32 of the second blades 31. The difference (or deviation) between the metal angle at the trailing edge 32b and the metal angle at the leading edge 32a, (02 - 01) is between 0° and 65° so as to present a possible gyration of the radially external flow equal or substantially equal to zero. Preferably, the deflection angle of each second stator blade 31 is between 15° and 45°.

[0046] Advantageously, but not limitatively, the metal angle at the trailing edge of each second stator blade 31 is zero or substantially zero (between 0° and 5°). In particular, a straight line passing through the trailing edge and tangent to the extrados surface of each blade of the second stator blades 31 is parallel to the longitudinal axis. In this way, the gyration of the radially external flow passing through the second stator blades 31 is cancelled. When the mean radius of the smooth diffuser 29 is increasing, the gyration produced by the first stator blades 25 will naturally attenuate by conservation of the kinetic moment.

[0047] According to an embodiment as shown in Figure 4, the length of the chord of each second stator blade 31 is less than the length of the chord of each first stator blade. This difference in length is permitted on the one hand, because the straightening of the radially external flow F12 is shared between the first stator blades 25 and the second blades 31, and on the other hand because the variation in fluid angle seen by the second stator blades 31 is smaller. This favors the reduction of the axial length of the set of first stator blades 25, diffuser 29, second stator blades 31, and equipment 26.

[0048] According to an embodiment illustrated in Figure 5, the equipment 26 comprises an annular row of structural arms 35. These structural arms 35 can slightly deflect the flow in a range between 0° and 5° but have no overall function of straightening the radially external flow. The arms 35 are distributed regularly around the longitudinal axis X. Each structural arm 35 extends radially between internal and external walls. Each arm 35 comprises an upstream edge 35a and a downstream edge 35b which are opposite along the longitudinal axis X. Intrados and extrados surfaces connect the upstream edge and the downstream edge.

[0049] In this exemplary embodiment, the row of structural arms 35 is integrated into the second row of second stator vanes 31. The structural arms 35 are independent and distinct from the second stator vanes. Alternatively, the upstream portion of the structural arms 35 may be formed by a second stator vane. In this way, the length of the second stator vanes 31 and of the structural arms 35 is reduced, which allows a saving in mass. As illustrated, the upstream edges 35a of the structural arms 35 and the leading edges 32a of the second vanes 31 are defined in the same plane P perpendicular to the longitudinal axis X. The trailing edges 35b of the arms 35 are arranged downstream of the trailing edges 32b of the second vanes 31.

[0050] The number of structural arms 35 is different from the number of second stator vanes 31. In particular, the second stator vanes 31 are more numerous than the structural arms 35. This is due to the fact that the structural arms 35 have only a mechanical structural function and make it possible to reduce the pressure losses.

[0051] The second blades 31 in this embodiment have an identical aerodynamic profile. In the present description, an aerodynamic profile is defined by geometric parameters which are the chord, the camber line, the skeleton line, the thickness of the blade along the chord, etc.

[0052] However, the number of second stator vanes 31 between two structural arms 35 in a circumferential direction (around the longitudinal axis X) may be different. Here, there are six second stator vanes 31 between two adjacent structural arms 35 and eleven second stator vanes 31 between two adjacent structural arms 35. The profiles of the structural arms 35 are identical but may be different.

[0053] Figure 6 illustrates another embodiment of the arrangement of the second stator vanes 31 integrated with the structural arms 35. This embodiment differs from that of Figure 5 in that the second stator vanes 31 are arranged in groups of aerodynamic profiles that are different from each other. A single group or different groups of aerodynamic profiles may be arranged between two adjacent structural arms 35 in the circumferential direction. In this example illustrated with twenty-six second stator vanes 31 and three structural arms 35, there are four groups of different aerodynamic profiles P1, P2, P3, P4. More precisely, between two adjacent structural arms 35 are arranged two second stator vanes of profile P1, two second stator vanes of profile P2, and two second stator vanes of profile P3.Between two other adjacent structural arms 35 are arranged three second stator blades of profile P1, four second stator blades of profile P2, two second stator blades of profile P4, and two second stator blades of profile P3. The arrangement of the aerodynamic profiles is specific to each configuration and is defined according to the arrangement of the equipment present in the radially external flow. The installation of second stator blades of different profiles makes it possible to minimize the impact of circumferential distortion coming from downstream.

[0054] In this way, the aerodynamic loads in the channel 22 are better distributed thanks to first stator blades 25 which will fix the fluid angle defined as the angle between the direction of the flow and the longitudinal axis X upstream of the second stator blades 31, a diffuser slowing down the radially external flow F12, and the second blades 31 which take up the rest of the deviation to cancel the gyration and which operate with an almost constant incidence.

Claims

Claims [1] Turbomachine (1) in particular for an aircraft, comprising: - a dividing nozzle (19) arranged in a primary vein (13) in which a primary flow (F1) circulates and intended to separate the primary flow (F1) into a radially internal flow (F11) and a radially external flow (F12); - stator equipment (26) installed in a channel (22) in which the radially external flux (F12) circulates; and, - a first row of first stator vanes (25) installed in the channel (22) and intended to straighten the radially external flow (F12) towards the equipment (26), characterized in that the turbomachine comprises a diffuser (29) which is arranged, downstream of the first stator vanes (25) and which is divergent between an inlet section (29a) and an outlet section (29b), and a second row of second stator vanes arranged (31), downstream of the diffuser (29) so as to deflect the radially external flow (F12) leaving the diffuser (29) towards the equipment, the second stator vanes (31) being located upstream of the equipment (26). [2] Turbomachine (1) according to claim 1, characterized in that the second stator blades (31) are configured so as to achieve a straightening of the radially external flow (F12) greater than the straightening of the first stator blades (25) and so as to cancel any possible gyration of the radially external flow at the outlet of the second stator blades (31). [3] Turbomachine (1) according to one of the preceding claims, characterized in that each second stator blade (31) comprises a blade (32) having a leading edge (32a) and a trailing edge (32b) which are connected by a lower surface (32i) and an upper surface (32e), each second stator blade (31) has a metal angle (01) at the leading edge (32a) and a metal angle (02) at the trailing edge (32b) the difference of which is between 0° and 65° so as to have a possible gyration of the radially external flow (F12) equal or substantially equal to zero. [4] Turbomachine (1) according to the preceding claim, characterized in that the metal angle (02) at the trailing edge (32a) of each second stator blade (31) is zero or substantially zero. [5] Turbomachine (1) according to one of the preceding claims, characterized in that each first stator blade (25) comprises a blade (27) having a leading edge (27a) and a trailing edge (27b) which are connected by a pressure surface (27i) and an extrados surface (27e), each first stator blade (25) having a metal angle (a1) at the leading edge (27a) and a metal angle (a2) at the trailing edge (27b) the difference of which is between 5° and 30° so as to produce a possible non-zero gyration of the radially external flow (F12) at the outlet of the first stator blades (25). [6] Turbomachine (1) according to any one of the preceding claims, characterized in that the stator equipment (26) comprises at least one heat exchanger, an annular row of structural arms (35) or at least one pylon. [7] Turbomachine (1) according to one of the preceding claims, characterized in that it comprises a propeller (2) and a separation nozzle (10) arranged between the propeller (2) and the division nozzle (19) along the longitudinal axis, the separation nozzle (10) being intended to separate an air flow (F) into the primary flow (F1) and a secondary flow (F2) coming from the propeller, the secondary flow (F2) circulating around the primary vein (13). [8] Turbomachine (1) according to the preceding claim, characterized in that a rectifier (8) comprising a plurality of stator blades (9) is arranged downstream of the propeller (2) and is arranged across the secondary flow (F2). [9] Turbomachine (1) according to one of claims 3 to 8, characterized in that the equipment (26) comprises an annular row of structural arms (35) which is integrated into the second row of second stator blades (31), the structural arms (35) each having an upstream edge (35a), the upstream edges (35a) of the structural arms (35) and the leading edges (31a) of the second stator blades (31) being defined in the same plane (P). [10] Turbomachine (1) according to the preceding claim, characterized in that the second stator blades (31) comprise identical aerodynamic profiles or groups of aerodynamic profiles different from each other. [11] Turbomachine (1) according to any one of the preceding claims, characterized in that the diffuser (29) comprises a radially external shell and a radially internal shell between which are defined passages for the radially external flow (F12), each passage extending between an inlet opening and an outlet opening opposite along the longitudinal axis and each passage has a section gradually increasing from the inlet opening to the outlet opening. [12] Turbomachine (1) according to any one of the preceding claims, characterized in that the diffuser (29) is a smooth diffuser.

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