Turbomachine with counter-rotating turbines for aircraft

Abstract

A turbomachine (10) with a counter-rotating turbine for an aircraft, the turbomachine comprising a counter-rotating turbine (22), a first rotor (22a) of the counter-rotating turbine being configured to rotate in a first rotational direction and being connected to a first turbine shaft (36), a second rotor (22b) of the counter-rotating turbine being configured to rotate in an opposite rotational direction and being connected to a second turbine shaft (38), the first rotor comprising turbine disks interleaved between turbine disks of the second rotor, the first shaft (36) being guided by at least two guide bearings (60, 62) mounted between the first shaft and a stator housing, the second shaft (38) being guided by at least two guide bearings (56, 58) mounted between the second shaft and another stator housing (28).

Description

Technical Field

[0001] The invention relates to a turbomachine with a counter-rotating turbine for an aircraft. Background Art

[0002] In particular, the prior art includes documents US-A1-2019 / 085701, EP-A1-3 447 243, EP-A1-2 975213 and WO-A1-2013 / 147977.

[0003] Generally, the turbomachine of an aircraft includes a fan, a low-pressure compressor, a high-pressure compressor, an annular combustion chamber, a high-pressure turbine and a low-pressure turbine from upstream to downstream in the direction of gas flow. The rotor of the low-pressure compressor is driven by the rotor of the low-pressure turbine, and the rotor of the high-pressure compressor is driven by the rotor of the high-pressure turbine.

[0004] From the point of view of engine performance and fuel consumption, it is advantageous to maximize the speed of the low-pressure turbine, as this leads to a higher efficiency of the turbine. However, increasing the speed of the turbine means increasing the centrifugal forces on the turbine, thus making the design of the turbine very complicated.

[0005] One suggestion to improve the efficiency of the turbine without increasing the rotational speed of the turbine is to use a counter-rotating turbine. The low-pressure turbine is replaced by a twin-rotor turbine having a first rotor configured to rotate in a first rotational direction and connected to a first turbine shaft and a second rotor configured to rotate in an opposite rotational direction and connected to a second turbine shaft. The first rotor has a turbine wheel portion inserted between the turbine wheel portions of the second rotor.

[0006] In a conventional architecture where the turbine directly drives the fan, the low-pressure turbine may have a takeoff speed of about 4,000 rpm, or in an architecture where the turbine drives the fan through a speed reduction device, the low-pressure turbine may have a takeoff speed of about 10,000 rpm. Replacing the low-pressure turbine with a counter-rotating turbine whose rotors rotate at takeoff speeds of about 3,000 rpm and 7,000 rpm, respectively, enables a relative speed of 10,000 (3000+7000) rpm, while having an absolute speed in the lower range of the above speed interval.

[0007] The counter-rotating turbine thus comprises a low-speed rotor that drives the fan and a high-speed rotor that meshes with a planetary epicyclic mechanical reduction gear whose input and output are counter-rotating (rotating ring gear, fixed planet carrier, rotating sun gear).

[0008] The reduction gear couples the high-speed rotor to the low-speed rotor, allowing power to be transferred from the high-speed rotor to the low-speed rotor. This takes advantage of the higher efficiency of the high-speed turbine while transferring most of the power from the turbine to the fan through the shaft rather than the reduction gear.

[0009] This architecture is complex due to its mechanical integration: the mechanical reduction device is located downstream of the turbine, radially inside the stator casing, known as the exhaust casing.

[0010] Existing integration solutions are particularly complex because there are many inter-shaft bearings that are particularly difficult to lubricate. The radial loads of the turbine are likely to pass through the gearbox that radially holds the two shafts, which is extremely detrimental to the good performance of the reduction gear. Finally, the space requirements are not optimized, especially in the case of bearings under the planetary gears or sun gears of the reduction gear, which limits the radial integration of the reduction gear. Summary of the invention

[0011] The present invention provides improvements to the above-described art which represent simple, effective and economical solutions to at least some of the problems discussed above.

[0012] The invention proposes a turbomachine with a counter-rotating turbine for an aircraft.

[0013] The turbine includes a counter-rotating turbine, a first rotor of the counter-rotating turbine is configured to rotate in a first rotational direction and is connected to a first turbine shaft, a second rotor of the counter-rotating turbine is configured to rotate in an opposite rotational direction and is connected to a second turbine shaft, the first rotor includes turbine wheel portions that are interleaved between turbine wheel portions of the second rotor.

[0014] The turbine further includes a planetary epicyclic mechanical reduction device, which includes: a sun gear, which is driven to rotate by the second shaft; a ring gear, which is driven to rotate by the first shaft; and a planet carrier, which is attached to the first stator housing of the turbine, and the first stator housing is located upstream of the counter-rotating turbine relative to the direction of gas flow in the turbine.

[0015] The turbine comprises bearings for guiding the first shaft and the second shaft,

[0016] characterised in that the first shaft is guided by at least two guide bearings, which are mounted between the first shaft and a second stator housing located downstream of the contra-rotating turbine or between the first shaft and an element connected to the second stator housing, and the second shaft is guided by at least two guide bearings, which are mounted between the second shaft and the first stator housing or between the second shaft and an element connected to the first stator housing,

[0017] Thus, each of the two rotors of the free-style turbine is guided in rotation by two bearings, as opposed to a single bearing in some prior art architectures.

[0018] In particular, this solution is advantageous for the following reasons:

[0019] - the turbine is easier to assemble because the centering of each of the first shaft and the second shaft is provided by the bearing pair during assembly;

[0020] - since the reduction gear can be framed by bearings that carry a rigid force path, the reduction gear is better held under gravity and maneuvers;

[0021] The turbine according to the invention may include one or more of the following features, which may be employed independently of one another or in combination with one another:

[0022] - the second shaft is guided by only two guide bearings;

[0023] - The two guide bearings of the second shaft are respectively an upstream ball bearing and a downstream roller bearing;

[0024] - one of the two guide bearings of the second shaft is associated with an oil film compression damping system, for example a downstream roller bearing;

[0025] - the two guide bearings of the second shaft are radially inside the first and second rotors; this is advantageous since it makes it possible to avoid cantilevers that would generate a significant loss of clearance under maneuvers and also to better control the dynamics of the turbine (no bending modes of the shaft of the turbine within the operating range, better damping of the suspension modes of the turbine); moreover, these bearings will ensure better support of the turbine in the case where they are arranged respectively in line with the upstream and downstream stages of the turbine;

[0026] - the first shaft is guided by only two guide bearings;

[0027] - a first guide bearing of the first shaft is located upstream of the reduction gear unit, and a second guide bearing of the first shaft is located downstream of the reduction gear unit;

[0028] - the first guide bearing of the first shaft is located in a first plane perpendicular to the longitudinal axis of the turbomachine, the first plane passing downstream of the first rotor and the second rotor and upstream of the reduction gear and the second housing, and the second guide bearing of the first shaft is located in a second plane perpendicular to this axis, the second plane being located downstream of the reduction gear and the second housing;

[0029] - The two guide bearings of the first shaft are roller bearings;

[0030] - one of the two guide bearings of the first shaft is associated with an oil film compression damping system, for example the upstream bearing;

[0031] - the first shaft is connected to a downstream stage of the first rotor, and / or the second shaft is connected to a downstream stage of the second rotor; this enables the torque to pass through the stage of the last rotor, which is much cooler than the first or upstream stages,

[0032] at least one of the guide bearings of the first shaft is shaped to have a certain flexibility in radial direction; this flexibility may be provided by an annular portion of the ring of the bearing having a C-shaped cross section,

[0033] at least one of the guide bearings of the second shaft is shaped to have a certain flexibility in radial direction; this flexibility may be provided by an annular portion of the ring of the bearing having a C-shaped cross section,

[0034] - at least one of the annular gear carrier and the planet carrier is shaped to have a certain radial flexibility and a certain tilting flexibility (flexibility in rotation about an axis perpendicular to the motor axis); this flexibility can be provided by an annular portion of the annular gear carrier with a C-shaped cross section or an S-shaped cross section, or by an annular portion of the planet carrier (if the planet carrier is integrated in the annular gear carrier) with a C-shaped cross section or an S-shaped cross section, preferably this flexibility is integrated outside the force path from the downstream end of the annular gear carrier to the bearing,

[0035] - the first shaft is connected to the ring gear of the reduction gear and to the drive shaft of the fan of the turbine,

[0036] - one of the guide bearings of the first shaft is located on a circumference having a diameter greater than the circumference of the other guide bearing passing through the first shaft and preferably greater than the outer diameter of the reduction gear unit, and

[0037] The guide bearing of the second shaft is located on a circumference with different diameters, the downstream bearing being located on a circumference with a smaller diameter, which may be smaller than the inner diameter of the reduction unit.

[0038] Advantageously, an oil film compression damping system is used to damp dynamic unbalance modes of the shaft or rotor.

[0039] Preferably, the turbine is of the type having a single ducted fan, the first shaft driving the single fan in rotation. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] The invention will be better understood and other details, features and advantages of the invention will become more apparent on reading the following description given by way of non-limiting example and with reference to the accompanying drawings, in which:

[0041] [ Figure 1 ] Figure 1 is a very schematic view of a turbomachine with counter-rotating turbines according to one embodiment of the invention,

[0042] [ Figure 2 ] Figure 2 yes Figure 1 A magnified view of a counter-rotating turbine.

[0043] [ Figure 3 ] Figure 3 yes Figure 1 Schematic axial cross-sectional view of a guide bearing of a turbine shaft,

[0044] [ Figure 4 ] Figure 4 yes Figure 1 A schematic axial cross-sectional view of a guide bearing of another shaft of a turbine, and

[0045] [ Figure 5 ] Figure 5 yes Figure 1 Schematic axial sectional view of another guide bearing of a shaft of a turbine. DETAILED DESCRIPTION

[0046] Figure 1 A turbomachine 10 with contra-rotating turbines for an aircraft is shown in a very schematic manner.

[0047] The turbomachine 10 includes, from upstream to downstream in the direction of gas flow, a fan 12 , a low-pressure compressor 14 , a high-pressure compressor 16 , an annular combustion chamber 18 , a high-pressure turbine 20 , and a counter-rotating turbine 22 .

[0048] Reference numeral 24 refers to an intermediate casing situated between compressors 14 and 16. Reference numeral 26 refers to a turbine casing (a turbine casing of the TVF type, TVF being the abbreviation for Turbine Vane Frame, a turbine casing of the turbine vane frame type indicating a turbine casing equipped with arms forming straightening blades) situated between turbines 20 and 22. Finally, reference numeral 28 refers to an exhaust casing (an exhaust casing of the TRF type, TRF being the abbreviation for Turbine Rear Frame, an exhaust casing of the turbine rear frame type indicating the last turbine casing). These casings form the structure of the turbomachine: they support the bearings guiding the rotating shaft and are coupled to the suspension of the turbomachine.

[0049] The rotor of the high-pressure turbine 20 drives the rotor of the high-pressure compressor 16 to rotate via a high-pressure shaft 30, which is centrally positioned and guided to rotate via bearings such as an upstream ball bearing 32 and a downstream roller bearing 34. The bearing 32 is mounted between the upstream end of the shaft 30 and the intermediate housing 24, and the bearing 34 is mounted between the downstream end of the shaft 30 and the turbine housing 26.

[0050] The counter-rotating turbine 22 includes a first rotor 22a having a wheel portion 22aa configured to rotate in a first rotational direction and connected to a first turbine shaft 36, and a second rotor 22b having a wheel portion 22ba configured to rotate in an opposite rotational direction and connected to a second turbine shaft 38, the wheel portions 22ba of the rotor 22b being interleaved between the wheel portions 22aa of the rotor 22a (see Figure 2 ).

[0051] Each turbine wheel includes an annular row of blades, each of the blades having an aerodynamic profile with a pressure side and a suction side that join to form leading and trailing edges for the gas in the turbine flow.

[0052] The first shaft 36 drives the fan 12 and the rotor of the low-pressure compressor 14 to rotate. The first shaft 36 is further meshed with a ring gear 40 of a planetary epicyclic mechanical reduction device 42.

[0053] The second shaft 38 meshes with a sun gear 44 or a planetary gear of the reduction gear 42 .

[0054] The reduction gear 42 further includes a sun gear 41 which meshes with a sun gear 44 and the ring gear 40 , respectively, and is carried by a planet carrier 46 attached to the turbine housing 26 .

[0055] Each of the housings 26 and 28 generally includes a central hub and an outer annular portion surrounding the hub and connected to the hub by a series of arms generally radial to the longitudinal axis of the turbine. The central hub of the housing 28 extends around at least a portion of the reduction gear 42 .

[0056] In the example shown, the high-pressure shaft 30 is centrally positioned and guided in rotation by two upstream ball bearings 32 and a roller bearing 33 and a downstream roller bearing 34. The bearings 32, 33 are mounted between the upstream end of the shaft 30 and the intermediate housing 24, and the bearing 34 is mounted between the downstream end of the shaft 30 and the turbine housing 26.

[0057] The planet carrier 46 of the reduction device 42 is attached to the turbine housing 26. The planet carrier 46 is thus connected to the turbine housing 26 by a cylindrical wall 50, which is advantageously rigid. This wall 50 passes axially through the rotors 22a, 22b and the second shaft 38.

[0058] As described above, the downstream end of the second shaft 38 meshes with the sun gear and is also connected to the final or downstream stage of the second rotor 22 b , ie, to the final wheel portion 22 ba 1 of the rotor.

[0059] The second shaft 38 is centered on this wall 50 and guided in rotation by means of two guide bearings, respectively an upstream guide bearing 56 and a downstream guide bearing 58. Examples of the design of these bearings 56, 58 will be referred to below. Figure 5 Describe in more detail.

[0060] The downstream end of the first shaft 36 is fixed to the ring gear 40 of the reduction gear, and the upstream end of the first shaft is fixed to the last stage or downstream stage of the first rotor 22a, that is, to the last wheel portion of the rotor. The ring gear 40 is also fixed to the upstream end of the ring gear carrier 40a, and the downstream end of the ring gear carrier is fixed to the downstream end of the shaft 36 or meshes with the downstream end of the shaft 36.

[0061] The planet carrier 46 here may include an annular portion with a C-shaped cross section or an S-shaped cross section on the downstream side of the reduction gear 42 to give the planet carrier a certain flexibility (flexibility of rotation around an axis perpendicular to the motor axis) through elastic deformation, especially in the radial direction and in tilting. Due to this flexibility provided by the planet carrier 46, the ring gear carrier 40a of the reduction gear 42 can be rigid. Under certain conditions, the opposite is possible. In this case, the ring gear carrier 40a will be flexible or provide flexibility, while the planet carrier 46 will be rigid. Therefore, the ring gear carrier 40a will include an annular portion with a C-shaped cross section or an S-shaped cross section to give the ring gear a certain flexibility (flexibility of rotation around an axis perpendicular to the motor axis) through elastic deformation, especially in the radial direction and in tilting. In this second configuration, the flexibility is advantageously integrated outside the force path from the downstream end of the ring gear carrier 40a to the bearing 60.

[0062] The shaft 36 is guided upstream by bearings 52, 54 mounted between the shaft and the intermediate housing 24. A first of these bearings is, for example, an upstream roller bearing 52 and a second of these bearings is, for example, a downstream ball bearing 54.

[0063] The shaft 36 is also centrally positioned and guided in rotation by two guide bearings, respectively an upstream guide bearing 60 and a downstream guide bearing 62. Advantageously, these bearings are arranged on both sides of the reduction device 42. Figure 3 and Figure 4 Examples of designs of these bearings 60 , 62 are described in more detail.

[0064] Figure 3A more detailed example of a guide bearing 60 (for example a guide bearing with rollers) is shown. Preferably, this bearing 60 is located upstream of the reduction device 42. More precisely, this bearing can be located on a plane P1 perpendicular to the axis of the turbomachine, which is located downstream of the turbine 22, upstream of the reduction device 42 and the exhaust housing 28 (see also Figure 1 ).

[0065] The bearing 60 comprises an inner ring 60a integral with the shaft 36 and an outer ring 60b integral with a shell 28a fixed to the casing 28. Advantageously, the shell 28a has a substantially frustoconical shape flaring towards the downstream end.

[0066] Advantageously, the bearing 60 is associated with an oil film compression damping system 70. In the example shown, this system 70 is present at the level of the outer ring of the bearing, between this ring 60 b and the radially inner end of the shell 28 a. The oil film is formed in the annular space extending between the outer cylindrical surface of the ring 60 b and the inner cylindrical surface of the ring portion 72 mounted around the ring, this space being supplied with oil by suitable means.

[0067] The outer ring 60b is further shaped to provide a certain flexibility to the bearing 60. In the example shown, one of the axial ends of the outer ring (in this case the downstream end) has a C-shaped axial cross section or a pin-shaped axial cross section, which enables the ring 60b to have radial flexibility by deformation.

[0068] Figure 4 A more detailed example of a guide bearing 62, for example a guide bearing with rollers, is shown. Preferably, this bearing 62 is located downstream of the reduction device 42. More precisely, this bearing can be located in a plane P2 perpendicular to the axis of the turbine, which plane is located downstream of the reduction device 42 and the exhaust housing 28 (see Figure 1 ).

[0069] The bearing 62 comprises an inner ring 62a integral with the downstream end of the shaft 64 and an outer ring 62b integral with a shell 28b fixed to the casing 28. Advantageously, this shell 28b has a substantially frustoconical shape flaring towards the upstream end. This shell 28b forms a support for the bearing 62.

[0070] As mentioned above, the bearing 62 participates in guiding the shaft 64 and, since the shaft 36 is connected to the shaft 64 , in particular via the ring gear carrier 40 a , it also participates in guiding the shaft 36 .

[0071] The shell portion 28 b or the downstream axial extension of the shell portion 28 b may cooperate with a dynamic seal 74 carried by the shaft 64 or an element attached to the downstream end of this shaft.

[0072] The bearing 60 is located on the circumference C1 , and the diameter of the circumference C1 is greater than the diameter of the circumference C2 where the bearing 62 is located.

[0073] Figure 5 A more detailed example of guide bearings 56, 58 is shown. The upstream bearing 56 is a ball bearing and the downstream bearing 58 is a roller bearing. Preferably, these bearings are located upstream of the reduction device 42 and radially inside the turbine 22. The upstream bearing 56 can be located at the upstream end or upstream stage of the turbine 22, and the downstream bearing 58 can be located at the downstream end or downstream stage of the turbine 22.

[0074] Each of the bearings 56, 58 comprises an inner ring 56a, 58a integral with the annular wall 76 fixed to the shaft 38, and an outer ring 56b, 58b integral with the shell 26a fixed to the housing 26. Advantageously, the shell 26a has a generally frustoconical shape flaring upstream. The shell 26a forms a support for the bearings 56, 58.

[0075] Advantageously, the bearing 58 is associated with a damping system 78 with oil film compression. In the example shown, this system 78 is present at the level of the outer ring 58b of the bearing 58, between this ring 58b and the radially inner end of the shell 26a. The oil film is formed in an annular space extending between the outer cylindrical surface of the ring 58b and the inner cylindrical surface of the ring 80 mounted around the ring 58b, to which space the oil is supplied by suitable means.

[0076] The outer ring 58b is further shaped to provide a certain flexibility to the bearing 58. In the example shown, one of the axial ends of the outer ring (in this case the upstream end) has a C-shaped axial cross section or a pin-shaped axial cross section, which enables the ring 58b to have radial flexibility by deformation.

[0077] The bearing 56 is located on a circle C3 , and the diameter of the circle C3 is greater than the diameter of the circle C4 on which the bearing 58 is located.

[0078] The invention thus proposes a turbomachine with a counter-rotating turbine and a reduction gear with bearing positioning, which makes it possible to solve the problems of dynamics, clearance consumption under maneuvers and misalignment, while proposing a better mechanical integration of the various components.

[0079] The present invention enables solving or avoiding a number of problems of the prior art, including:

[0080] The solution consisting in guiding each of the shafts by a single bearing is unsatisfactory for the following reasons:

[0081] The reduction gear does not or only slightly misaligns under gravity, steering or power. This would create overloads in the meshing teeth, which is detrimental to the design of the reduction gear (mechanics, quality);

[0082] The reduction gear can be installed in a relatively flexible environment (flexibility built into the planet carrier and input shaft to enable the reduction gear to accommodate displacement of the planet carrier and input shaft, thereby avoiding overloading of the teeth). This requires space requirements that may be detrimental to the length of the motor;

[0083] The high-speed rotor is not held by a single bearing, which may lead to high-speed rotor centering problems during assembly; this may also lead to significant consumption of inter-turbine clearance under maneuvers, which will again degrade the performance of the architecture;

[0084] ·wait.

Claims

1. A turbomachine (10) having a counter-rotating turbine for an aircraft, The turbine comprises a counter-rotating turbine (22), a first rotor (22a) of the counter-rotating turbine being configured to rotate in a first rotational direction and being connected to a first turbine shaft (36), a second rotor (22b) of the counter-rotating turbine being configured to rotate in an opposite rotational direction and being connected to a second turbine shaft (38), the first rotor comprising turbine wheel portions interleaved between turbine wheel portions of the second rotor, The turbine further comprises a first stator housing (26) located upstream of the counter-rotating turbine relative to the direction of gas flow in the turbine, The turbine further comprises a second stator housing (28) located downstream of the counter-rotating turbine with respect to the direction of the gas flow, The turbine further comprises a planetary epicyclic mechanical reduction device (42), the planetary epicyclic mechanical reduction device comprising: a sun gear (44), the sun gear being driven to rotate by the second turbine shaft (38); a ring gear (40) driven for rotation by the first turbine shaft (36); and a planet carrier (46) attached to the first stator housing (26), The turbine comprises bearings (56-62) for guiding the first turbine shaft and the second turbine shaft, Characterized in that the first turbine shaft (36) is guided by at least two guide bearings (60, 62), the at least two guide bearings of the first turbine shaft are carried by the second stator housing (28), the at least two guide bearings are located downstream of the counter-rotating turbine, the first guide bearing of the at least two guide bearings is located radially outside the planetary epicyclic mechanical reduction device, the second guide bearing of the at least two guide bearings is located radially inside the planetary epicyclic mechanical reduction device, the second turbine shaft (38) is guided by at least two guide bearings (56, 58), the at least two guide bearings of the second turbine shaft are installed between the second turbine shaft and the first stator housing, or installed between the second turbine shaft and an element connected to the first stator housing.

2. The turbomachine (10) with counter-rotating turbines for an aircraft according to claim 1, wherein: The second turbine shaft (38) is guided only by two guide bearings (56, 58).

3. The turbomachine (10) with counter-rotating turbines for an aircraft according to claim 2, wherein: The two guide bearings (56, 58) of the second turbine shaft (38) are respectively an upstream ball bearing and a downstream roller bearing.

4. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 2 or 3, wherein: One of the two guide bearings (56, 58) of the second turbine shaft (38) is associated with an oil film compression damping system.

5. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 2 or 3, wherein: The two guide bearings (56, 58) of the second turbine shaft (38) are located radially inside the first rotor (22a) and the second rotor (22b).

6. A turbomachine (10) with a counter-rotating turbine for an aircraft according to any one of claims 1 to 3, wherein: The first turbine shaft (36) is guided only by two guide bearings (60, 62), which are mounted between the first turbine shaft and the second stator housing (28) or between the first turbine shaft and an element connected to the second stator housing.

7. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 6, wherein: The first guide bearing of the first turbine shaft (36) is located downstream of the counter-rotating turbine and upstream of the planetary epicyclic mechanical reduction device (42), and the second guide bearing of the first turbine shaft is located downstream of the counter-rotating turbine and downstream of the planetary epicyclic mechanical reduction device.

8. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 7, wherein: The first guide bearing of the first turbine shaft (36) is located in a first plane (P1) perpendicular to the longitudinal axis of the turbine, the first plane passing through the downstream of the first rotor (22a) and the downstream of the second rotor (22b), as well as the upstream of the planetary epicyclic mechanical reduction device (42) and the upstream of the second stator housing (28), and the second guide bearing of the first turbine shaft is located in a second plane (P2) perpendicular to the longitudinal axis, the second plane being located downstream of the planetary epicyclic mechanical reduction device and downstream of the second stator housing.

9. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 6, wherein: The two guide bearings (60, 62) of the first turbine shaft (36) are roller bearings.

10. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 9, wherein: One of the two guide bearings of the first turbine shaft (36) is associated with an oil film compression damping system.

11. A turbomachine (10) with a counter-rotating turbine for an aircraft according to any one of claims 1 to 3, wherein: The first turbine shaft (36) is connected to a downstream stage of the first rotor (22a), and / or the second turbine shaft (38) is connected to a downstream stage of the second rotor (22b).

12. A turbomachine (10) with counter-rotating turbines for an aircraft according to any one of claims 1 to 3, wherein: At least one of the guide bearings (60, 62) of the first turbine shaft (36) is shaped to have a certain flexibility in the radial direction.

13. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 12, wherein: The flexibility of at least one of the guide bearings (60, 62) of the first turbine shaft (36) is provided by an annular portion of the ring of the guide bearing having a C-shaped cross section.

14. A turbomachine (10) with counter-rotating turbines for an aircraft according to any one of claims 1 to 3, wherein: At least one of the guide bearings (56, 58) of the second turbine shaft (38) is shaped to have a certain flexibility in the radial direction.

15. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 14, wherein: The flexibility of at least one of the guide bearings (56, 58) of the second turbine shaft (38) is provided by an annular portion of the ring of the guide bearing having a C-shaped cross section.

16. A turbomachine (10) with counter-rotating turbines for an aircraft according to any one of claims 1 to 3, wherein: At least one of the planet carrier (46) and the ring gear carrier (40a) supporting the ring gear (40) is formed to have certain radial flexibility and tilt flexibility.

17. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 16, wherein: The flexibility of the at least one of the planet carrier and the ring gear carrier is provided by the annular portion thereof having a C-shaped cross section or an S-shaped cross section.

18. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 17, wherein: One of the guide bearings of the first turbine shaft (36) is located on a first circumference (C1) having a diameter greater than a diameter of a second circumference (C2) passing through the other guide bearing of the first turbine shaft (36).

19. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 18, wherein: One of the guide bearings (56, 58) of the second turbine shaft (38) is a downstream roller bearing, the guide bearings (56, 58) of the second turbine shaft (38) are located on a third circumference (C3) and a fourth circumference (C4) having different diameters, and the downstream roller bearing is located on a circumference having a smaller diameter, and the smaller diameter is smaller than the inner diameter of the planetary epicyclic mechanical reduction device (42).

20. A turbomachine (10) with counter-rotating turbines for an aircraft according to any one of claims 1 to 3, wherein: The first turbine shaft (36) is connected to the ring gear (40) of the planetary epicyclic mechanical reduction device and to the drive shaft of the fan (12) of the turbine.

21. A turbomachine (10) with a counter-rotating turbine for an aircraft according to claim 20, said turbomachine comprising a single ducted fan, said first turbine shaft (36) driving said single ducted fan in rotation.

22. The turbomachine (10) with counter-rotating turbines for an aircraft according to claim 4, wherein: The one guide bearing of the two guide bearings (56, 58) of the second turbine shaft (38) is a downstream roller bearing.

23. The turbomachine (10) with counter-rotating turbines for aircraft according to claim 10, wherein: The one guide bearing of the two guide bearings of the first turbine shaft (36) is an upstream bearing.

24. A turbomachine (10) with counter-rotating turbines for an aircraft according to claim 18, wherein: The diameter of the first circle is greater than the outer diameter of the planetary epicyclic mechanical speed reducer (42).

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