PASSAGE OF EASEMENTS IN AN AIRCRAFT TURBOMACHINE EXHAUST HOUSING

Abstract

Turbomachine (10) for an aircraft, this turbomachine having a longitudinal axis (A) and comprising, from upstream to downstream, at least one compressor (14, 16), an annular combustion chamber (18) and at least one turbine (20, 22), the turbomachine further comprising an annular exhaust casing (128) located downstream of said at least one turbine (20, 22), the exhaust casing (128) comprising two coaxial annular shells (72, 74), respectively internal and external, which define between them a flow path for a gas stream and which are connected to each other by radial arms (28a) extending in this path, the internal annular shell (72) extending at least partially around equipment which is connected to utilities radially traversing said path, characterized in that said utilities are housed in at least one radial sleeve (80) located in downstream of one of said arms. Figure for the abbreviation: Figure 7

Description

Title of the invention: PASSAGE OF EASEMENTS IN A TURBOMACHINE EXHAUST HOUSING D'AERONEF Technical field of the invention

[0001] The present invention relates to an aircraft turbomachine equipped with an exhaust casing which is mounted around a piece of equipment and which is radially traversed by servicing connected to this equipment. Technical background

[0002] The prior art includes in particular document FR-A1-3 104 193.

[0003] Conventionally, an aircraft turbomachine comprises, from upstream to downstream in the direction of gas flow, a fan, a low-pressure compressor, a high-pressure compressor, an annular combustion chamber, a high-pressure turbine, and a low-pressure turbine. 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 an engine performance and fuel consumption standpoint, maximizing the rotational speed of the low-pressure turbine is advantageous because it results in better turbine efficiency. However, increasing the turbine's rotational speed increases the centrifugal forces it experiences, thus significantly complicating its design.

[0005] One suggestion for increasing the efficiency of a turbine without increasing its rotational speed is to use a counter-rotating turbine. The low-pressure turbine is then replaced by a two-rotor turbine, the first rotor of which is configured to rotate in a first direction of rotation and is connected to a first turbine shaft, and the second rotor is configured to rotate in the opposite direction of rotation and is connected to a second turbine shaft. The first rotor has turbine wheels interspersed between turbine wheels of the second rotor.

[0006] A low-pressure turbine can have a takeoff speed of approximately 4,000 rpm in a conventional design where the turbine directly drives the fan, or a takeoff speed of approximately 10,000 rpm in a design where the turbine drives the fan via a reduction gear. Replacing it with a counter-rotating turbine whose rotors rotate at takeoff speeds of approximately 3,000 and 7,000 rpm respectively allows for a relative speed of 10,000 rpm (3000+7000) while maintaining an absolute speed in a lower range of the range of aforementioned speed.

[0007] This counter-rotating turbine thus comprises a slow rotor and a fast rotor, the slow rotor driving the blower and the fast rotor meshing with a planetary type epicyclic gear mechanical reducer whose input and output are counter-rotating (rotating ring, fixed satellite carrier, rotating solar).

[0008] The gearbox couples the high-speed and low-speed rotors, thus enabling power transfer from the high-speed rotor to the low-speed rotor. The higher efficiencies of a high-speed turbine are taken advantage of while a large portion of the turbine's power is transferred to the blower directly, without passing through a gearbox but via a shaft.

[0009] This architecture is complex due to its mechanical integration: the mechanical reducer is located downstream of the turbomachine, radially inside a stator housing called the exhaust housing. This housing comprises two coaxial annular shells which define between them a flow channel for the gas stream from the counter-rotating turbine and which are connected together by radial arms.

[0010] Furthermore, since the gearbox releases considerable energy (on the order of 100 kW) during operation, it must be continuously lubricated to maintain an acceptable operating temperature. An oil circuit is therefore implemented to supply the gearbox with oil, and its consumption can be significant, for example, on the order of 60,000 L / h. Oil removal is a critical issue. Indeed, the oil must be collected and disposed of.

[0011] In the current technique, the oil is evacuated through pipes that pass through at least two arms of the exhaust housing. These exhaust housing arms are tubular and oversized to allow the passage of these pipes. For this reason, they have a cross-section or circumferential thickness that is greater than that of the other housing arms, which generates significant aerodynamic losses in the flow.

[0012] This problem also arises when electrical equipment, such as an electric generator, is mounted inside a turbomachine exhaust casing, and this equipment has to be connected to utilities that cross the channel defined by the exhaust casing.

[0013] There is therefore a need to find a solution to optimize the passage of oil lines, and more generally of auxiliary lines, through the channel defined by an exhaust casing of a turbomachine, the turbomachine being a conventional or specific type (for example, with a counter-rotating turbine). Summary of the invention

[0014] The invention provides a turbomachine for an aircraft, this turbomachine having a longitudinal axis and comprising, from upstream to downstream, at least one compressor, a annular combustion chamber and at least one turbine, the turbomachine further comprising an annular exhaust casing located downstream of said at least one turbine, the exhaust casing comprising two coaxial annular shells, respectively internal and external, which define between them a flow path of a gas flow and which are connected to each other by radial arms extending in this path, the internal annular shell extending at least in part around equipment which is connected to servicing radially through said path, characterized in that said servicing is housed in at least one radial sleeve located downstream of one of said arms.

[0015] The service lines therefore no longer pass through the arms but through dedicated sleeves located downstream of the arms. This avoids oversizing the arms to accommodate the service lines and thus reduces the impact of the service line passage on the gas flow in the channel.

[0016] The turbomachine according to the invention may comprise one or more of the following features, taken individually or in combination with each other:

[0017] - said easements are oil pipelines; - in which said equipment is a mechanical reducer or electrical equipment such as an electric generator; - said shirt extends in line with said arm and is formed in one piece with this arm; - said shirt extends in line with said arm and is attached and fixed to this arm; - said shirt is located at an axial distance from said arm; - said shirt extends between an inner ring and an outer ring which are respectively configured to be fixed to the internal and external ferrules; - said shirt is located at 6 o'clock with reference to the face of a clock; and

[0018] - the turbomachine comprises a counter-rotating turbine located just upstream of said exhaust casing; and

[0019] - the turbomachine is of the dual-flow type, respectively primary and secondary;

[0020] - said at least one radial sleeve is located downstream of said primary flow. Brief description of the figures

[0021] The invention will be better understood and other details, features and advantages of the invention will become more apparent upon reading the following description, given by way of non-limiting example and with reference to the accompanying drawings in which:

[0022] [Fig-1] [Fig.1] is a very schematic axial cross-sectional view of a turbomachine counter-rotating turbine,

[0023] [Fig.2] [Fig.2] is another very schematic view of a turbine turbomachine counter-rotating,

[0024] [Fig.3] [Fig.3] is a schematic perspective view of an exhaust housing according to the prior art;

[0025] [Fig.4] [Fig.4] is a schematic axial cross-sectional view of the exhaust housing of the [Fig.3];

[0026] [Fig.5] [Fig.5] is a schematic cross-sectional view of an arm of the exhaust housing of [Fig.3];

[0027] [Fig.6] [Fig.6] is a schematic perspective view of an exhaust housing according to a first embodiment of the invention;

[0028] [Fig.7] [Fig.7] is a schematic axial cross-sectional view of the exhaust housing of the [Fig.6];

[0029] [Fig.8] [Fig.8] is a schematic perspective view of an exhaust housing according to a second embodiment of the invention;

[0030] [Fig.9] [Fig.9] is a schematic axial cross-sectional view of the exhaust housing of the [Fig.8];

[0031] [Fig. 10] [Fig. 10] is a schematic perspective view of an exhaust housing according to a third embodiment of the invention;

[0032] [Fig. 11] [Fig. 11] is a schematic axial cross-sectional view of the exhaust housing of [Fig. 10];

[0033] [Fig. 12] [Fig. 12] is a partial schematic perspective view of an exhaust housing, without a liner to visualize any auxiliary components; and

[0034] [Fig. 13] [Fig. 13] is a partial schematic perspective view of a cylinder liner of an exhaust housing and the components housed within this liner. Detailed description of the invention

[0035] Fig. 1 represents in a very schematic way a counter-rotating turbine turbomachine 10 for an aircraft.

[0036] This turbomachine 10 comprises, from upstream to downstream, in the direction of gas flow, a blower 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.

[0037] Reference numeral 24 designates an intermediate casing located between compressors 14 and 16, and reference numeral 26 designates a turbine casing (of the TVF type, an acronym for Turbine Vane Frame, which designates a turbine casing equipped with arms forming stator blades) located between turbines 20 and 22. Finally, reference numeral 28 designates an exhaust casing (of the TRF type, an acronym for Turbine Rear Frame, which designates the last turbine casing). These casings form the structure of the tur- turbomachine: they support the bearings that guide the rotating shafts and are linked to the turbomachine suspensions.

[0038] The high-pressure turbine rotor 20 drives the high-pressure compressor rotor 16 in rotation via a high-pressure shaft 30 which is centered and guided in rotation by bearings, such as an upstream ball bearing 32 and a downstream roller bearing 34. The bearing 32 is mounted between an upstream end of the shaft 30 and the intermediate housing 24, and the bearing 34 is mounted between a downstream end of the shaft 30 and the turbine housing 26.

[0039] The counter-rotating turbine 22 comprises a first rotor 22a whose wheels 22aa are configured to rotate in a first direction of rotation and are connected to a first turbine shaft 36, and a second rotor 22b whose wheels 22ba are configured to rotate in an opposite direction of rotation and are connected to a second turbine shaft 38 and are intercalated between the wheels 22aa of the rotor 22a (see figures 1 and 3).

[0040] Each turbine wheel comprises an annular row of blades, each of which has an aerodynamic profile comprising an intrados and an extrados which meet to form a leading edge and a trailing edge of the gases in the turbine stream.

[0041] The first shaft 36 drives the blower 12 in rotation as well as the rotor of the low pressure compressor 14. This first shaft 36 is also meshed with a ring gear 40 of a planetary type epicyclic mechanical reducer 42.

[0042] The second shaft 38 is meshed with the solar 44 or planetary of the reducer 42.

[0043] The reducer 42 further comprises satellites 41 meshed respectively with the solar 44 and the crown 40 and carried by a satellite carrier 46 which is fixed to the exhaust housing 28.

[0044] Each of the housings 26 and 28 generally comprises two coaxial annular shells, respectively internal and external, which are connected to each other by a series of arms substantially radial with respect to the longitudinal axis A of the turbomachine and which pass through the turbine runner. The internal shell of the housing 28 extends around at least a portion of the gearbox 42.

[0045] Figure 2 represents another turbomachine in which the elements described above are designated by the same reference numerals. The preceding description applies here insofar as it does not contradict what follows and is not contrary to what is illustrated in Figure 2.

[0046] The differences between the turbomachines in Figures 1 and 2 are based in particular on the bearings for guiding the shafts and rotors and on the fixing of the reducer on the turbine casing 26 or the exhaust casing 28.

[0047] In [Fig. 2], the high-pressure shaft 30 is centered and guided in rotation by two upstream bearings, respectively ball bearings 32 and roller bearings 33, and by a downstream roller bearing 34. The bearings 32, 33 are mounted between an upstream end of the shaft 30 and the intermediate casing 24, and bearing 34 is mounted between a downstream end of shaft 30 and turbine casing 26.

[0048] Unlike [Fig. 1], the planet carrier 46 of the gearbox 42 of [Fig. 2] is fixed 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.

[0049] The second shaft 38 has its downstream end meshed with the solar, as mentioned above, and is also connected to the last stage or downstream stage of the second rotor 22b, that is to say to the last wheel of this rotor.

[0050] The second shaft 38 is centered and guided in rotation on this wall 50 by means of two guide bearings, respectively upstream 56 and downstream 58.

[0051] The first shaft 36 has its downstream end fixed to the ring gear 40 of the reduction gear and its upstream end fixed to the last stage or downstream stage of the first rotor 22a, that is to say, to the last wheel of this rotor. The ring gear 40 is also fixed to the upstream end of a ring carrier 40a, the downstream end of which is fixed or meshed with the downstream end of the shaft 36.

[0052] The planet carrier 46 may include, here on the downstream side of the gearbox 42, an annular portion having a C- or S-shaped cross-section so as to impart a certain degree of flexibility to the planet carrier through elastic deformation, particularly in the radial direction as well as in tilting (rotational flexibility about axes perpendicular to the drive axis). Due to this flexibility provided by the planet carrier 46, the ring gear 40a of the gearbox 42 can itself be rigid. The reverse is also possible, under certain conditions. In this case, the ring gear 40a would be flexible or would provide flexibility, and the planet carrier 46 would be rigid. The crown carrier 40a would then include an annular portion having a C or S shaped cross-section so as to give the crown a certain flexibility by elastic deformation, in particular radially as well as in tilting (flexibility in rotation around axes perpendicular to the drive axis).In this second configuration, the flexibility is advantageously integrated outside the load path going from the downstream end of the crown carrier 40a to the bearing 60.

[0053] The shaft 36 is guided upstream by bearings 52, 54 mounted between this shaft 36 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.

[0054] The shaft 36 is further centered and guided in rotation downstream by two guide bearings, respectively upstream 60 and downstream 62, both supported by the exhaust housing 28. These bearings are advantageously placed on either side of the reducer 42.

[0055] During operation, the reducer 42 is lubricated and the oil that lubricated the reducer The oil must be collected and removed for recycling to prevent it from accumulating in the gearbox and turning into coke due to the relatively high temperatures in that environment. The gearbox lubricating oil also serves to cool it by absorbing heat. The heated oil must be removed and cooled before being reintroduced into the system to dissipate heat from the gearbox, which must not exceed 200°C.

[0056] In operation, the oil is centrifuged and ends up in the ring 40 of the reducer 42. Passages may be provided at the external periphery of the ring 40 and / or the ring carrier 40a so as to project the oil by centrifugation onto an annular oil recovery device which is not shown in the drawings for clarity.

[0057] This oil recovery device is connected to one or more oil lines 70 which pass through arms 28a of the exhaust housing 28 and which therefore result in a transverse oversizing of these arms 28a in the prior art.

[0058] Figures 3 and 4 show an exhaust casing 28 according to the prior art.

[0059] The exhaust casing 28 comprises two coaxial annular ferrules, respectively internal 72 and external 74, which define between them the flow path of the gas flow exiting the counter-rotating turbine 22 and which are connected to each other by arms 28a. There are 16 arms 28a in the example shown.

[0060] The internal ferrule 72 delimits an internal annular space for housing the reducer 42 and therefore extends at least partly around the reducer 42.

[0061] As can be seen in Figures 4 and 5, each of the arms 28a comprises an aerodynamic profile and includes a leading edge 28al and a trailing edge 28a2. The leading edges 28al of the arms 28a are located substantially in the same plane PI perpendicular to the longitudinal axis A. The trailing edges 28a2 of the arms 28a are located substantially in the same plane P2 perpendicular to the longitudinal axis A.

[0062] In the prior art, at least two of the arms 28a, such as the arms 28a', are tubular and oversized in the transverse direction to allow the passage of oil pipes 70 connected to the aforementioned evacuation device.

[0063] The transverse oversizing of the arms 28a' of the housing 28 complicates the design and manufacture of the latter and generates aerodynamic losses in the flow of the gas stream in the vein.

[0064] The invention proposes a solution to this problem by housing the pipes 70, and more generally the servicing, in at least one radial sleeve which is located downstream of at least one of the arms 28a. In other words, the pipes 70 are no longer located in arms 28a but outside of these arms 28a and therefore do not risk impacting their transverse dimensions.

[0065] Advantageously, the exhaust casing 128 comprises a single liner 80 of pipe passage 70, this sleeve 80 being located at 6 o'clock (6h) by reference with the face of a clock.

[0066] Figures 6 and 7 illustrate a first embodiment of the invention in which the sleeve 80 extends in the downstream extension of the arm 28a located at 6h and is formed in one piece with this arm 28a.

[0067] There are 16 arms 28a in the example shown.

[0068] Each of the arms 28a comprises an aerodynamic profile and includes a leading edge 28al and a trailing edge 28a2. The leading edges 28al of the arms 28a are located substantially in the same plane PI perpendicular to the longitudinal axis A. The trailing edges 28a2 of the arms 28a are located substantially in the same plane P2 perpendicular to the longitudinal axis A.

[0069] The arm 28a at 6h does not, however, include a trailing edge because its downstream end is directly connected to the upstream end of the sleeve 80.

[0070] The sleeve 80 has an aerodynamic profile and includes a trailing edge 80b located downstream of the plane P2. The sleeve 80 is tubular and defines an internal space for the radial passage of the pipes 70, and more generally of the services.

[0071] By comparing figures 4 and 7, it can be seen that the ferrules 72, 74 can be lengthened to allow the integration of the sleeve 80. The exhaust casing 128 then has an axial dimension greater than that of the prior art and this extra length can for example be compensated by a reduction in length of a nozzle located just downstream of the casing 128, in order to avoid impacting the total length of the turbomachine 10.

[0072] Figures 8 and 9 illustrate a second embodiment of the invention in which the sleeve 80 extends in the downstream extension of the arm 28a located at 6h and is attached and fixed to this arm 28a.

[0073] There are 16 arms 28a in the example shown.

[0074] Each of the arms 28a comprises an aerodynamic profile and includes a leading edge 28al and a trailing edge 28a2. The leading edges 28al of the arms 28a are located substantially in the same plane PI perpendicular to the longitudinal axis A. The trailing edges 28a2 of the arms 28a are located substantially in the same plane P2 perpendicular to the longitudinal axis A.

[0075] The arm 28a at 6h does not, however, include a trailing edge because its downstream end is directly connected to the upstream end of the sleeve 80.

[0076] The sleeve 80 has an aerodynamic profile and includes a trailing edge 80b located downstream of the plane P2. The sleeve 80 is tubular and defines an internal space for the radial passage of the pipes 70, and more generally of the services.

[0077] By comparing Figures 4 and 9, it can be seen that the ferrules 72, 74 can be lengthened to allow the integration of the sleeve 80. Alternatively, the sleeve 80 could extend between two coaxial rings 82, 84, respectively internal and external, which would be mounted downstream of the ferrules 72, 74 and would be fixed to these ferrules by clamping for example.

[0078] The exhaust casing 128 has an axial dimension greater than that of the prior art and this extra length can for example be compensated by a reduction in length of a nozzle located just downstream of the casing 128, in order to avoid impacting the total length of the turbomachine.

[0079] Figures 10 and 11 illustrate a third embodiment of the invention in which the sleeve 80 extends at an axial distance from the arm 28a located at 6h.

[0080] There are 16 arms 28a in the example shown.

[0081] Each of the arms 28a comprises an aerodynamic profile and includes a leading edge 28al and a trailing edge 28a2. The leading edges 28al of the arms 28a are located substantially in the same plane PI perpendicular to the longitudinal axis A. The trailing edges 28a2 of the arms 28a are located substantially in the same plane P2 perpendicular to the longitudinal axis A.

[0082] The sleeve 80 has an aerodynamic profile and comprises a leading edge 80a located substantially in a plane P3 perpendicular to the longitudinal axis A and located downstream of P2, and a trailing edge 80b located in a plane P4 perpendicular to the longitudinal axis A and located downstream of P3. The sleeve 80 is tubular and defines an internal space for the radial passage of the pipes 70, and more generally of the utilities.

[0083] By comparing figures 4 and 10, it can be seen that the ferrules 72, 74 can be lengthened to allow the integration of the sleeve 80. Alternatively, the sleeve 80 could extend between two coaxial rings 82, 84, respectively internal and external, which would be mounted downstream of the ferrules 72, 74 and would be fixed to these ferrules by clamping for example.

[0084] The exhaust housing 128 has an axial dimension greater than that of the prior art and this extra length can for example be compensated by a reduction in length of a nozzle located just downstream of the housing, in order to avoid impacting the total length of the turbomachine.

[0085] Figures 12 and 13 illustrate a series of pipes 70 or service lines that can pass through a sleeve 80. The pipes 70 are generally elongated in the radial direction and may have different diameters. They are arranged one behind the other in the same plane passing through the axis A, which may be a median plane of symmetry of the sleeve 80. The pipes 70 are held in the sleeve 80 by one or more fastening systems 86. The pipes 70 may include fluid connection means 88 at their internal and external radial ends.

[0086] As mentioned above, the pipes 70 can be replaced by other types of easements and for example by electrical cables or harnesses, gas pipelines, etc.

[0087] Although the embodiments described above illustrate the invention in the context of a turbomachine 10 with a counter-rotating turbine 22 and a reduction gear 42, the invention is not limited to this type of turbomachine and can, for example, also be applied to a turbomachine equipped with electrical equipment surrounded by the exhaust casing. This equipment is, for example, an electric generator.

[0088] The invention therefore applies generally to all turbomachines requiring the passage of utilities through the exhaust casing 28.

[0089] In the case of a turbomachine or turbofan engine, the primary flow enters the engine (also called the gas generator) and successively the compressor, the combustion chamber, and the turbine. The secondary flow circulates around the engine. The cylinder liner is advantageously located downstream of the primary flow so that the auxiliary systems, and in particular the piping 70, extend downstream of the primary flow.

Claims

Claims

1. Turbomachine (10) for an aircraft, this turbomachine having a longitudinal axis (A) and comprising from upstream to downstream at least one compressor (14, 16), an annular combustion chamber (18) and at least one turbine (20, 22), the turbomachine further comprising an annular exhaust casing (128) located downstream of said at least one turbine (20, 22), a nozzle being located downstream of this exhaust casing (128), the exhaust casing (128) comprising two coaxial annular shells (72, 74), respectively internal and external, which define between them a flow path for a gas flow and which are connected to each other by radial arms (28a) extending in this path, the internal annular shell (72) extending at least in part around equipment which is connected to services passing through radially the said vein,characterized in that said services are housed in at least one radial sleeve (80) located downstream of one of said arms (28a), in the extension or at an axial distance from this arm (28a).,

2. Turbomachine (10) according to claim 1, in which said services are oil pipes (70).

3. A turbomachine (10) according to claim 1 or 2, wherein said equipment is a mechanical reducer (42) or electrical equipment such as an electric generator.

4. Turbomachine (10) according to one of claims 1 to 3, in which said sleeve (80) extends in the extension of said arm (28a) and is formed in a single piece with this arm.

5. Turbomachine (10) according to one of claims 1 to 3, in which said sleeve (80) extends in the extension of said arm (28a) and is attached and fixed to this arm.

6. Turbomachine according to one of claims 1 to 3, in which said sleeve (80) is located at an axial distance from said arm (28a).

7. A turbomachine (10) according to claim 5 or 6, wherein said liner (80) extends between an inner ring (82) and an outer ring (84) which are respectively configured to be fixed to the inner and outer shrouds (72, 74).

8. Turbomachine (10) according to one of the preceding claims, in which said jacket (80) is located at 6 o'clock with reference to the dial of a clock.

9. Turbomachine (10) according to one of the preceding claims, in which it comprises a counter-rotating turbine (22) located just upstream of said exhaust casing (128).

10. Turbomachine (10) according to one of the preceding claims, in which it is of the double flow type, respectively primary and secondary, and in which said at least one radial sleeve (80) is located downstream of said primary flow.

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