RECOVERED CYCLE TURBOENGER

Abstract

Turboshaft engine (1) for an aircraft (2) comprising: - a gas generator (3) including a compressor (4), a combustion chamber (5) and an expansion turbine (6); - a power turbine (8) driving a power take-off (9) via a reduction gear (10); - a heat exchanger (11) comprising: - a first circuit (12) having an inlet (13) connected to an outlet (14) of the compressor (4), and an outlet (15) connected to an inlet (16) of the combustion chamber (5), and - a second circuit (17) having an inlet (18) connected to an outlet (19) of the power turbine (8), characterized in that the reduction gear (10) is arranged axially at a forward end (20) of the turboshaft engine (1), such that the compressor (4) is arranged axially between the reduction gear (10) and the power turbine (8). Figure for the abstract: 1

Description

Description Title of the invention: RECOVERED CYCLE TURBOMOTOR Technical field of the invention

[0001] — The present invention relates to a recovered cycle turboshaft engine for an aircraft. Technical background

[0002] …— Application FR2962487A1 in the name of the applicant describes a first ar- Recycled cycle turboshaft engine design.

[0003] = Such a turboshaft engine architecture includes, in particular, from front to rear a air intake, a compressor, a combustion chamber, an expansion turbine (linked turbine or high-pressure turbine), a power turbine (free turbine or low-pressure turbine), and finally an exhaust nozzle.

[0004] — The compressor, the combustion chamber and the expansion turbine form a ge- gas nebulizer in which the compressor rotor is mechanically driven by the expansion turbine rotor via a transmission shaft.

[0005] — The power turbine is independent of the gas generator, and is intended to drive a power take-off (or power take-off) of the turbomotor on which are for example connected to the aircraft's propulsion system(s).

[0006] — The air entering through the air inlet is compressed by the compressor, then injected into the combustion chamber to be mixed with fuel. The air / mixture fuel is burned and expanded in the expansion turbine and then in the turbine of power before being expelled from the turbocharger through the exhaust nozzle.

[0007] — The turbocharger is said to have a "recovered cycle" because the residual thermal energy of the gases exhaust is recovered via heat exchangers placed in the nozzle exhaust, then reinjected into the turbocharger to optimize its efficiency.

[0008] — US patent application US20090277154A1 describes a second turbocharger architecture with cycle recovered.

[0009] — This second architecture has the particularity of having its combustion chamber at one rear end of the turboshaft engine, the expansion and power turbines located axially between the compressor and the combustion chamber, the gases exhaust gases are evacuated from the turbocharger via two V-shaped nozzles arranged sen- presumably in the middle of the turbocharger.

[0010] — In such an architecture, the air compressed by the compressor is routed to the rear into the combustion chamber via supply ducts arranged around the turbines, the compressed air is diverted before entering the chamber of combustion.

[0011] = The thermal energy recovered by the heat exchangers is used here to re- heat the compressed air coming out of the compressor before it enters the combustion chamber. Such an architecture makes it possible to improve the performance of the turbocharger since the amount of fuel to be injected to reach operating temperatures is less than that required in the case of a conventional cycle turbocharger (i.e. a turbocharger in which the compressed air coming out of the compressor directly feeds the combustion chamber without being preheated). The power take-off is driven by the power turbine via a reducer arranged axially between the compressor and the turbines. Compared to the first architecture described above, this second architecture, described in US20090277154A1, has the advantage of crossing the airflow from the compressor, which is delivered to the combustion chamber via two supply ducts, with the hot gas flow from the two exhaust nozzles. It is therefore particularly well-suited to the integration of heat exchangers for the heat recovery cycle, while minimizing the required duct lengths and thus the mass and size of the heat recovery system. However, the second architecture has drawbacks. Indeed, firstly, the aforementioned reducer is generally associated with an accessory box which is intended to transmit mechanical power taken from the turbines to various accessories of the turbomotor such as a pump, an alternator-starter, an air / oil separator, etc. The gearbox and accessory box are bulky, which significantly limits the available space around the compressor and turbines, and therefore the possibilities for installing the heat exchanger(s) intended for the recovery of residual thermal energy from the exhaust gases. Secondly, the arrangement of the reducer almost in the middle of the turbomotor requires a significant offset of the power take-off to bypass in particular the compressor and the supply ducts, to the detriment of the mass and the balance (or mass distribution) of the turbomotor. The objective of the present invention is therefore to provide a simple, effective and economical solution to address the aforementioned drawbacks. Summary of the invention The invention thus proposes a turboshaft engine for an aircraft, the turboshaft engine comprising: - a gas generator comprising a compressor, a combustion chamber and an expansion turbine, the compressor and the expansion turbine extending along the same longitudinal axis X and being mechanically connected to each other, the chamber of combustion being arranged axially at a rear end of the turbomotor; - a power turbine arranged axially between the compressor and the expansion turbine, the power turbine driving in rotation a power take-off via a reduction gear; - a heat exchanger comprising: - a first circuit comprising an input connected to an output of the compressor, and an output connected to an inlet of the combustion chamber, and - a second circuit comprising an input connected to an output of the power turbine, characterized in that the reducer is arranged axially at a front end of the turbomotor, so that the compressor is arranged axially between the reducer and the power turbine. This gearbox arrangement frees up space around the compressor and turbines for installing the heat exchanger. It is now possible to consider installing an annular heat exchanger around the X-axis. Such a reduction gear arrangement also simplifies access to and drive of the power take-off. This type of turbocharger is called a "recovered cycle" turbocharger because the residual thermal energy of the exhaust gases is recovered via the heat exchanger to preheat the compressed air exiting the compressor before it enters the combustion chamber. This reclaimed cycle architecture improves turbocharger performance since the amount of fuel injected to reach operating temperatures is less than that required in a conventional cycle turbocharger. The turbocharger according to the invention may comprise one or more of the following characteristics, taken individually or in combination with each other: - the heat exchanger is annular around said longitudinal axis X; - the heat exchanger is arranged at least partially around the power turbine; - the heat exchanger is sectorized and comprises at least two sectors placed circumferentially end to end or circumferentially distant from each other, each sector of the heat exchanger comprising a sub-inlet of the first circuit connected to the outlet of the compressor; - the heat exchanger is a single unit; - the turboshaft engine includes at least one first bypass duct having an inlet connected to the compressor outlet and an outlet connected to the combustion chamber inlet, such that the first bypass duct supplies the combustion chamber combustion in compressed air exiting the compressor without passing through the heat exchanger; - The turboshaft engine includes at least one second bypass duct having an inlet connected to the outlet of the power turbine and an outlet connected to the nozzle, such that the second bypass duct supplies the nozzle with exhaust gas exiting the power turbine without passing through the heat exchanger; - The reduction gear is part of a transmission housing, the transmission housing comprising an accessory gearbox intended to transmit mechanical power taken from the compressor and / or the expansion turbine and / or the power turbine to various accessories of the turboshaft engine: - the exchanger includes a retaining shield for the moving blades of the power turbine configured to contain said moving blades in the event of overspeed of the power turbine; - the heat exchanger includes a sound attenuation layer; - the compressor includes a first shaft driven in rotation by a second shaft of the expansion turbine via a transmission mechanism, said transmission mechanism and said reduction gear forming part of a transmission housing which is axially disposed at a front end of the turboshaft engine, so that the compressor is axially disposed between the transmission housing and the power turbine; - the power turbine includes a third shaft, the first, second and third shafts being coaxial with said longitudinal axis X; - the third shaft of the power turbine is arranged radially between the first and second shafts; - said transmission mechanism has a transmission ratio greater than 1, so that the speed of the first shaft of the compressor is greater than the speed of the second shaft of the expansion turbine; - said transmission mechanism is a gear mechanism; - said transmission mechanism includes a first toothed wheel fixed to the first shaft of the compressor and a second toothed wheel fixed to the second shaft of the expansion turbine, said reducer including a third toothed wheel fixed to a third shaft of the power turbine, the third toothed wheel of the reducer being arranged axially between said first and second toothed wheels of the transmission mechanism; - the transmission case has, from front to back, the second gear, the third gear and the first gear; - said power take-off is coaxial or vertically aligned with said longitudinal axis X; - The first shaft of the compressor is guided in rotation via a first bearing and a second bearing arranged in the transmission housing; - said second shaft of the expansion turbine is guided in rotation via a third bearing disposed in the transmission housing, and a fourth bearing disposed either between the expansion and power turbines or at a rear end of said second shaft located on the side opposite the power turbine; - said third shaft of the power turbine is guided in rotation via a fifth bearing disposed in the transmission housing, and a sixth bearing disposed axially between the compressor and the power turbine. The present invention also relates to an aircraft, preferably a single-engine helicopter, comprising a turboengine as described above. Brief description of the figures 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: [Fig.1] [Fig.1] is a schematic view of a turbomotor according to the invention; [Fig.2] [Fig.2] is a detailed view of a first variant of the embodiment; [Fig.3] [Fig.3] is a detailed view of a second variant embodiment; [Fig.4] [Fig.4] is a detailed view of a third variant embodiment; [Fig. 5] [Fig. 5] is a detailed view of a fourth embodiment. Detailed description of the invention Figure 1 schematically represents a turboshaft engine for aircraft 2. Aircraft 2 is preferably a single-engine helicopter, and in particular a light single-engine helicopter whose maximum take-off weight (known by the English acronym MTOW for "Maximum Take-Off Weight") does not exceed 3175 kg (or 7000 lbs). Turboshaft 1 comprises: - a gas generator 3 comprising a compressor 4, a combustion chamber 5 and an expansion turbine 6, the compressor 4 and the expansion turbine 6 being mechanically linked to each other, the combustion chamber 5 being arranged axially at a rear end 7 of the turbomotor 1; - a power turbine & arranged axially between the compressor 4 and the expansion turbine 6, the power turbine 8 driving in rotation a power take-off 9 via a reducer 10; - a heat exchanger 11 comprising: - a first circuit 12 comprising an input 13 connected to an output 14 of the compressor 4, and an output 15 connected to an input 16 of the combustion chamber 5, and - a second circuit 17 comprising an input 18 connected to an output 19 of the turbine power of 8. According to the invention, the reducer 10 is arranged axially at a front end 20 of the turbomotor 1, so that the compressor 4 is axially arranged between the reducer 10 and the power turbine 8. Such an arrangement of the reducer 10 makes it possible in particular to free up space around the compressor 4 and the turbines 6, 8 to install the heat exchanger 11, but also to simplify access to and drive of the power take-off 9. Such a turboshaft engine architecture 1 is called a "recovered cycle" because the residual thermal energy of the exhaust gases is recovered via the heat exchanger 11 to preheat the compressed air exiting the compressor 4 before it enters the combustion chamber 5. Such a recovered cycle architecture makes it possible to improve the performance of the turboshaft engine | since the amount of fuel to be injected to reach operating temperatures is less than that required in the case of a conventional cycle turboshaft engine. The turbomotor | is defined along a longitudinal axis X which corresponds to the axis of rotation of the shafts 22, 23, 26 respectively of the compressor 4 and of the turbines 6, 8 of the turbomotor 1. By convention, in the present application, the terms "front" and "rear" define the axial positions of the elements of the turbomotor 1 relative to each other, knowing that the reducer 10 is arranged axially at a front end 20 of the turbomotor 1, and the combustion chamber 5 is arranged axially at a rear end 7 of the turbomotor. The term "axial" or "axially" means any direction parallel to the X axis of the turbomotor |, and "radial" or "radially" means any direction perpendicular to the X axis of the turbomotor 1. Similarly, by convention in this application, the terms "internal" and "external" associated with turboshaft 1 are defined radially with respect to the X-axis of turboshaft 1. As illustrated in [Fig. 1], the compressor 4 is supplied with air via an air inlet 21 and comprises a first shaft 22 that rotates about the X-axis. The compressor 4 may comprise one or more compression stages, each stage being either axial or centrifugal. The rotors of each stage (impeller or wheel) are rotationally fixed to the first shaft 22. Advantageously, the last stage of compressor 4 is a centrifugal stage. The combustion chamber 5 is axially arranged at a rear end 7 of the turbocharger 1. The chamber 5 is supplied with compressed and heated air (via the heat exchanger 11), and with fuel via one or more injectors depending on the combustion chamber technology used. The air / fuel mixture is burned by the action of a or several ignition devices. The combustion chamber 5 can be with separate pots or with direct flow or reverse flow. A combustion chamber with separate pots offers the advantage of low production costs, provided the number of injectors is reduced (it can be limited to a single injector), and of a compact design. This allows, for example, the integration of devices aimed at minimizing pollutant emissions at the rear of the chamber. Such a combustion chamber also has a reduced-sized casing, which minimizes the required cooling air and simplifies the implementation of Lean Premix Prevaporized (LPP) technology. A direct-flow combustion chamber also has the advantage of being compact, which minimizes the required cooling air. The amount of cooling air needed for a direct-flow combustion chamber is less than that required for a reverse-flow combustion chamber. A reverse flow combustion chamber (or return combustion chamber) has the advantage of minimizing the axial footprint of the turbomachine and freeing up a central space in which it is possible to integrate one or more elements such as a shaft guide bearing 23 of the expansion turbine 6. The expansion turbine 6 (also called the high-pressure turbine) includes a second shaft 23 that rotates about the X-axis. The expansion turbine 6 may have one or more expansion stages, each stage being either axial or centripetal. The rotors of each stage (wheel or impeller) are rotationally fixed to the second shaft 23. The exhaust gases from the combustion chamber 5 are expanded in the expansion turbine 6 and then in the power turbine 8. Advantageously, as illustrated in [Fig.!], the first shaft 22 of the compressor 4 is driven in rotation by the second shaft 23 of the expansion turbine 6 via a transmission mechanism 24. Such a transmission mechanism 24 makes it possible to transmit the rotational movement initiated by the expansion turbine 6 to the compressor 4 while having the possibility of modifying its speed, so as to operate the compressor 4 in the desired operating range. Advantageously, as illustrated in [Fig.1], the transmission mechanism 24 and the reducer 10 are part of a transmission housing 25 which is axially arranged at a front end 20 of the turbomotor 1, so that the compressor 4 is axially arranged between the transmission housing 25 and the power turbine 8. The power turbine 8 (also called a free turbine or low-pressure turbine) is independent of the gas generator 3 and includes a third shaft 26 movable around the X-axis. The power turbine 8 may include one or more expansion stages, each stage being either axial or centripetal. The rotors of each stage (wheel or wheel) are rotationally fixed to the third shaft 26. The exhaust gases exiting the power turbine 8 first pass through the heat exchanger 11 (and more specifically the second circuit) before being evacuated into the external environment via an exhaust nozzle 27. The power take-off 9 (also called the power take-off) is located at the output of the reduction gear 10 and drives, for example, one or more propellers of the aircraft 2, or an alternator-generator for electrical generation. When the aircraft 2 is a helicopter, the power take-off 9 can drive a main rotor via a main gearbox 28 (known by the acronym BTP) and a tail rotor (also known by the acronym RAC for anti-torque rotor) via a rear gearbox (known by the acronym BTA) (not shown). As illustrated in [Fig.1], the turboshaft engine 1 comprises, from front to back, the transmission housing 25 (reducer 10 and transmission mechanism 24), the compressor 4, the power turbine 8, the expansion turbine 6 and the combustion chamber 5. The gas generator 3 forms the high-pressure body of the turbomotor 1, and the power turbine 8 and the reducer 10 form the low-pressure body of the turbomotor 1. As indicated above, the heat exchanger 11 recovers the residual thermal energy from the exhaust gases to heat the compressed air exiting the compressor 4 before it enters the combustion chamber 5. The heat exchanger 11 can be, for example, a tube heat exchanger, a plate heat exchanger, or a finned heat exchanger. The heat exchanger 11 can notably be manufactured using additive manufacturing. The inlet 13 of the first circuit 12 of the heat exchanger 11 is connected to an outlet 14 of the compressor 4 via a front supply 29 comprising one or more conduits. The front supply 29 may include a radial diffuser and an axial diffuser (also called a rectifier), the outlet 14 of the compressor 4 being connected to an inlet of the radial diffuser and the inlet 13 of the first circuit 12 being connected to an outlet of the axial diffuser. The outlet 15 of the first circuit 12 of the heat exchanger 11 is connected to an inlet 16 of the combustion chamber 5 via a rear feed 30 comprising one or more conduits. This rear feed 30 may include diffusion grids to control the Mach number and vortex of the flow supplying the combustion chamber 5 via its inlet 16. The inlet 18 of the second circuit 17 of the heat exchanger 11 is connected to the outlet 19 of the power turbine 8 via one or more internal conduits of the exhaust nozzle 27. The output 33 of the second circuit 17 of the heat exchanger 11 is connected to the external environment via one or more external ducts of the exhaust nozzle 27. Advantageously, as illustrated in [Fig.1], the heat exchanger 11 is annular around the X axis. Such a configuration is possible due to the positioning of the reducer 10 at the front end 20 of the turbomotor 1. Advantageously, the heat exchanger 11 is axisymmetric with respect to the X axis, so as to obtain a balanced turboshaft engine. Advantageously, the heat exchanger 11 is disposed at least partly around the compressor 4 and / or the power turbine 8 and / or the expansion turbine 6, and preferably at least partly around the power turbine 8. According to the embodiment illustrated in [Fig. 1], the heat exchanger 11 is annular and extends continuously around the X-axis. In such a configuration, the heat exchanger 11 can be a single piece (or a single unit), and preferably manufactured using additive manufacturing (e.g., selective powder bed fusion). Additive manufacturing has the advantage of being able to produce complex shapes. The heat exchanger 11 can be divided into sectors and comprise at least two sectors 31 placed circumferentially end-to-end or circumferentially separated from each other, each sector 31 of the heat exchanger 11 comprising a sub-inlet 32 ​​of the first circuit 12 connected to the outlet 14 of the compressor 4. The first circuit 12 can be subdivided into a plurality of sub-first circuits, each sector 31 potentially comprising one sub-first circuit. When the upstream supply 29 comprises a radial diffuser and an axial diffuser, each sub-inlet 32 ​​of the first circuit 12 can be connected to a sub-outlet of the axial diffuser. A multi-outlet axial diffuser (also called a "pipe diffuser") can be used in this configuration. Each sector 31 can be a single piece (or a single unit), and preferably manufactured using additive manufacturing (e.g., selective powder bed fusion).The advantage of having a sectorized heat exchanger 11 is to simplify maintenance and allow the replacement of a defective sector 31 independently of the others. According to the first embodiment illustrated in [Fig.2], the heat exchanger 11 is sectorized and comprises an annular row of twelve sectors 31 placed circumferentially end to end, each sector 31 of the heat exchanger 11 comprising a sub-inlet 32 ​​of the first circuit 12 connected to the outlet 14 of the compressor 4. According to the second embodiment illustrated in [Fig.3], the heat exchanger 11 is sectorized and comprises six sectors 31 distributed regularly around the X axis, two successive sectors 31 being circumferentially separated from each other on the other hand, each sector 31 of the heat exchanger 11 includes a sub-input 32 of the first circuit 12 connected to the output 14 of the compressor 4. Advantageously, the turboshaft engine | includes at least a first bypass duct 34 having an inlet connected to the outlet 14 of the compressor 4 and an outlet connected to the inlet 16 of the combustion chamber 5, so that the first bypass duct 34 directly supplies the combustion chamber 5 with compressed air coming out of the compressor 4 without going through the heat exchanger 11 (or by bypassing the exchanger). The turboshaft engine | can obviously include several first conduits 34 distributed or not regularly around the X axis, the first conduits 34 being able to have identical or distinct dimensional and geometric characteristics. The first duct(s) 34 are used, in particular, to supply the combustion chamber 5 during deceleration phases, when it is necessary to send the coolest possible air to the high-pressure turbine and thus avoid being negatively impacted by the heat stored in the heat exchanger. The need for deceleration is crucial for the turboshaft engines of a helicopter to maintain its maneuverability; otherwise, there is a risk of over-revving the main rotor. The first duct(s) 34 can be fitted with valves to control the flow of air and the supply rate. According to the embodiment illustrated in [Fig.1], the turboshaft engine | comprises a single first bypass conduit 34 which is arranged around the heat exchanger 11. When the heat exchanger 11 is sectorized, one or more first conduits 34 can pass through the exchanger 11 via the circumferential space defined between two successive sectors 31. Advantageously, the turboshaft engine 1] includes at least one second bypass duct 35 having an inlet connected to the outlet 19 of the power turbine 8 and an outlet connected to the nozzle 27, so that the second bypass duct 35 directly supplies the nozzle 27 with exhaust gas exiting the power turbine 8 without passing through the heat exchanger 11 (or by bypassing the exchanger). The turboshaft engine 1 can obviously include several secondary conduits 35 distributed regularly or not around the X axis, the secondary conduits 35 being able to have identical or distinct dimensional and geometric characteristics. The second duct(s) 35 are used in particular to evacuate exhaust gases more quickly while minimizing pressure losses, thereby increasing the turbocharger's power output. This extra power is especially useful when the turbocharger is operating under high acceleration conditions, for example, the engine speed of Takeoff, where the turboshaft engine must quickly reach the Maximum Takeoff Power (MTP). The second duct(s) 35 can be fitted with valves to control the passage of exhaust gases and their flow rate. According to the embodiment illustrated in [Fig.1], the turboshaft engine 1 comprises a single second bypass conduit 35 which extends radially outwards from the outlet 19 of the power turbine 8. When the heat exchanger 11 is sectorized, one or more second conduits 35 can pass through the exchanger 11 via the circumferential space defined between two successive sectors 31. Advantageously, the transmission housing 25 includes an accessory gearbox 36 for transmitting mechanical power from the compressor 4 and / or the expansion turbine 6 and / or the power turbine 8 to various accessories of the turboshaft engine 1. These accessories include, for example, a pump, a starter-alternator, an air / oil separator, etc. Positioning the accessory gearbox 36 at the front end 20 of the turboshaft engine 1 optimizes the arrangement of the accessories relative to one another and maximizes their number. Advantageously, the heat exchanger 11 includes a retaining shield 37 for the moving blades of the power turbine 8 configured to contain the moving blades in the event of overspeed of the power turbine 8. The blades of the power turbine 8 each include a frangible section which is configured to break when the power turbine 8 is overspeeding; these frangible sections form a protective device (called in English "blade shedding") allowing the rotor of the power turbine 8 to be stopped in the event of overspeed. The retention shield 37 could be configured to also contain the movable blades of the expansion turbine 6 in the event of overspeed of the expansion turbine 6. The presence of the heat exchanger 11 forms an obstacle to the noise produced by the turbomotor, the exchanger 11 thus acting as an acoustic attenuator. Advantageously, as illustrated in [Fig.!], the heat exchanger 11 can further comprise a sound attenuation layer 38. This sound attenuation layer 38 can, for example, comprise a honeycomb structure (e.g., a honeycomb structure). Advantageously, as illustrated in [Fig.1], the first, second and third trees 22, 23, 26 are coaxial with the X axis. As illustrated in [Fig.1], the third shaft 26 of the power turbine 8 is arranged radially between the first and second shafts 22, 23. As indicated above, the first shaft 22 of the compressor 4 is driven in rotation by the second shaft 23 of the expansion turbine 6 via a transmission mechanism 24. The transmission mechanism 24 can have a fixed or variable transmission ratio, this transmission ratio being different from 1, namely either less than 1 or greater than 1. When the transmission ratio is less than 1, the transmission mechanism 24 is a reducer (or multiplier), so that the speed of the first shaft 22 of the compressor 4 is less than that of the second shaft 23 of the expansion turbine 6. When the transmission ratio is greater than 1, the transmission mechanism 24 is a multiplier, so that the speed of the first shaft 22 of the compressor 4 is greater than that of the second shaft 23 of the expansion turbine 6. Advantageously, as illustrated in [Fig.1], the transmission mechanism 24 has a fixed transmission ratio which is greater than 1, so that the speed of the first shaft 22 of the compressor 4 is greater than the speed of the second shaft 23 of the expansion turbine 6. Such a configuration makes it possible to find the best compromise of speed for the speed of the compressor 4 and the speed of the expansion turbine 6, so as to maximize the efficiencies of these two components. Advantageously, as illustrated in [Fig. 1], the transmission mechanism 24 is a gear mechanism (or gear train). As indicated above, this gear mechanism 24 has a fixed transmission ratio greater than 1, so that the speed of the first shaft 22 of the compressor 4 is greater than the speed of the second shaft 23 of the expansion turbine 6. The gear mechanism 24 may include one or more reduction stages. The gear mechanism 24 may also include one or more epicyclic gear trains. Epicyclic gear trains have the advantage of being able to achieve high reduction or multiplication ratios while remaining compact. More specifically, as illustrated in [Fig.1], the gear mechanism 24 comprises a first gear 39 fixed to the first shaft 22 of the compressor 4 and a second gear 40 fixed to the second shaft 23 of the expansion turbine 6. The reducer 10 comprises a third gear 41 fixed to the third shaft 26 of the power turbine 8. The third gear 41 of the reducer 10 is arranged axially between the first and second gears 39, 40 of the transmission mechanism 24. As illustrated in [Fig.1], the transmission housing 25 has, from front to back, the second gear 40, the third gear 41 and the first gear 39. More specifically, according to the embodiment illustrated in [Fig. 1], the second gear 40 is centered on the X-axis and is the driving gear. The second gear 40 is rotationally connected to a front end of the second shaft 23. The second gear 40 is meshed with a first intermediate gear 42. The first intermediate gear 42 is driven and fixed to an intermediate shaft 43 that is free to rotate about an axis A, which is radially offset from the X-axis and parallel to the X-axis. The transmission mechanism 24 further includes a second intermediate gear 44 that meshes with the first gear 39. The second intermediate gear 44 is driven and fixed to the intermediate shaft 43. The first gear 39 is centered on the X-axis and is driven. The gears of the transmission mechanism 24 have external contact. The first gear 39 has fewer teeth than the second intermediate gear 44. The second gear 40 has more teeth than the first intermediate gear 42. The first and second gears 39, 40 rotate in the same direction of rotation, the intermediate shaft 43 rotates in the opposite direction of rotation.The various gear wheels 39, 40, 42, 44 of the gear mechanism 24 make it possible to obtain the desired transmission ratio. The first gear wheel 39 can either be made from the material with the first shaft 22, or be attached to the first shaft 22 and rotationally connected to the first shaft 22 via coupling means such as splines or by shrink fitting. According to the embodiment illustrated in [Fig. 1], the third gear 41 is centered on the X-axis and is the driving gear. The third gear 41 is rotationally connected to a front end of the third shaft 26. The third gear 41 meshes with an intermediate third gear 45. The intermediate third gear 45 is the driven gear and is fixed to the power take-off 9, which is movable about an axis B that is radially offset from the X-axis and parallel to the X-axis. According to the third embodiment illustrated in [Fig. 4], the third intermediate gear 45 is fixed to a second intermediate shaft 46, which is movable about axis B and is radially offset from axis X and parallel to axis X. The reducer 10 further includes a fourth intermediate gear 47, which meshes with an output gear 48. The fourth intermediate gear 47 is a driving gear and fixed to the second intermediate shaft 46. The output gear 48 is a driven gear and fixed to the power take-off 9. The power take-off 9 is coaxial or vertically aligned with axis X. According to the embodiment illustrated in [Fig.1], the first shaft 22 of the compressor 4 is guided in rotation via a first bearing 49 and a second bearing 50 arranged in the transmission housing 25. According to the embodiment illustrated in [Fig. 1], the second shaft 23 of the expansion turbine 6 is guided in rotation via a third bearing 51 disposed in the transmission housing 25, and a fourth bearing 52 disposed at a rear end of the second shaft 23 located on the opposite side to the power turbine 8. Such an arrangement is notably used when the combustion chamber 5 has reverse flow, which avoids the need for servicing between the turbines 6, 8, to the benefit of the axial footprint of the turbomotor. According to a fourth embodiment illustrated in [Fig.5], the fourth bearing 52 for guiding the second shaft 23 is arranged between the expansion and power turbines 6, 8. Such an arrangement of the fourth bearing 52 is used in particular when the combustion chamber 5 does not have a free central space for the installation of a bearing. According to the embodiment illustrated in [Fig.1], the third shaft 26 of the power turbine 8 is guided in rotation via a fifth bearing 53 disposed in the transmission housing 25, and a sixth bearing 54 disposed axially between the compressor 4 and the power turbine 8. Advantageously, the bearings are rolling bearings. The gears and bearings are lubricated with a liquid lubricant such as oil. Generally, each gear 39-48 can either be machined from the same material as the corresponding shaft, or mounted onto the corresponding shaft and rotationally connected to it by means of couplings such as splines or shrink fittings. The advantage of using couplings is that the gears can be easily replaced individually when they are worn and need replacing. By convention, in this application, the different variants and configurations presented are obviously combinable with each other.

Claims

Demands

1. Turboshaft engine (1) for aircraft (2), the turboshaft engine (1) comprising: - a gas generator (3) comprising a compressor (4), a chamber combustion (5) and an expansion turbine (6), the compressor (4) and the expansion turbine (6) extending along the same longitudinal axis (X) and being mechanically linked to each other, the combustion chamber (5) being arranged axially at a rear end (7) of the turboshaft engine {1}; - a power turbine (8) arranged axially between the compressor (4) and the expansion turbine (6), the power turbine (8) driving in rotation a power take-off (9) via a reducer (10); - a heat exchanger (11) comprising: - a first circuit (12) comprising an input (13) connected to an output (14) of the compressor (4), and an outlet (15) connected to an inlet (16) of the combustion chamber (5), and - a second circuit (17) comprising an input (18) connected to an output (19) of the power turbine (8), characterized in that the reducer (10) is arranged axially at a front end (20) of the turboshaft engine (1), so that the compressor (4) either axially arranged between the reducer (10) and the power turbine (8).

2. Turbomotor (1) according to any one of the preceding claims, characterized in that the heat exchanger (11) is annular around said longitudinal axis gitudinal (X).

3. Turbomotor (1) according to any one of the preceding claims, characterized in that the heat exchanger (11) is disposed at least partly around the power turbine (8).

4. Turbomotor (1) according to any one of the preceding claims, characterized in that the heat exchanger (11) is sectorized and includes at minus two sectors (31) placed circumferentially end to end or circumferentially distant from each other, each sector (31) of the heat exchanger (11) comprising a sub-inlet (32) of the first circuit (12) connected to the output (14) of the compressor (4).

5. Turbomotor (1) according to any one of claims 1 to 3, characterized in that that the heat exchanger (11) is a single unit.

6. Turbomotor (1) according to any one of the preceding claims, characterized in that the turboshaft engine (1) comprises at least one first duct of branch (34) having an input connected to the output (14) of the com- pressor (4) and an outlet connected to the inlet (16) of the chamber combustion (5), so that the first bypass conduit (34) supplies the combustion chamber (5) with compressed air exiting the com- press (4) without going through the heat exchanger (11).

7. Turbomotor (1) according to any one of the preceding claims, characterized in that the turboshaft engine (1) comprises at least one second duct of bypass (35) having an inlet connected to the outlet (19) of the turbine power (8) and an outlet connected to the nozzle (27), so that the The second bypass duct (35) supplies the nozzle (27) with gas exhaust exiting the power turbine (8) without passing through the heat exchanger (11).

8. Turbomotor (1) according to any one of the preceding claims, characterized in that the reducer (10) is part of a transmission housing (25), the transmission housing (25) including an accessory box (36) intended to transmit mechanical power taken from the com- pressor (4) and / or the expansion turbine (6) and / or the power turbine {8) to various accessories of the turbomotor (1).

9. Turbomotor (1) according to any one of the preceding claims, characterized in that the exchanger (11) includes a retention shield (37) of movable blades of the power turbine (8) configured to contain said movable blades in case of overspeed of the power turbine (8).

10. Turboshaft engine (1) according to any one of the preceding claims, characterized in that the heat exchanger (11) comprises a layer acoustic attenuation (38).

11. Aircraft (2), preferably a single-engine helicopter, comprising a turbomotor (1) according to any one of the preceding claims.