TURBOMACHINE FOR AIRCRAFT

Abstract

The invention relates to a turbomachine (4) comprising: - a low-pressure compressor (10), - a high-pressure compressor (11), - a high-pressure turbine (12), - a low-pressure turbine, - an annular combustion chamber (14), - a primary flow path (v1) for a primary air stream (F1) which passes successively, from upstream to downstream following the flow of the primary air stream (F1), through the rotors of the low-pressure compressor (10) and high-pressure compressor (11), the combustion chamber (14), and the rotors of the high-pressure and low-pressure turbines (12, 13), and - a heat exchanger (28) located in the primary flow path (v1) and comprising a first air passage circuit from the high-pressure compressor (11) to supply the combustion chamber (14), and a second gas passage circuit from the combustion chamber (14). (Shorthand figure: Figure 2)

Description

Title of the invention: TURBOMACHINE FOR AIRCRAFT Technical field of the invention

[0001] The invention relates to the field of aircraft turbomachines.

[0002] The invention relates in particular to the field of twin-body turbojets comprising a combustion chamber intended to be supplied with combustible gas, in particular dihydrogen. Technical background

[0003] An aircraft turbomachine, such as a twin-spool, twin-flow turbojet, typically comprises, from upstream to downstream in the direction of gas flow along a longitudinal axis, a fan rotatable about a longitudinal axis, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine and a gas exhaust nozzle.

[0004] The turbomachine further comprises a system for supplying the combustion chamber with fuel, in particular kerosene. The supply system comprises a pipe for conducting the fuel from a fuel source to the combustion chamber.

[0005] The blower allows the suction of an air flow dividing for example into a primary air flow and a secondary air flow. The primary air flow passes through a primary vein of the turbomachine delimited by an internal casing while the secondary air flow is directed towards a secondary vein surrounding the primary vein.

[0006] The primary air flow is compressed within the compressors. Each compressor comprises a mobile disc rotating around the longitudinal axis. Each compressor further comprises blades regularly distributed around the disc and which make it possible to compress the primary air flow.

[0007] The compressed air is then mixed with the fuel and burned in the combustion chamber. The gases from the combustion pass through the turbines. Each turbine has a disc centered on the longitudinal axis and blades regularly distributed around the disc which allow a force to be exerted on the gases from the combustion chamber. The gases finally escape through the nozzle, the section of which allows the acceleration of these gases to generate propulsion.

[0008] In order to limit the environmental impact of aircraft, it has been proposed to replace the kerosene used in the combustion chamber with a potentially more virtuous energy, in particular a combustible gas such as dihydrogen (H2). However, the current configuration of turbomachines does not easily allow the implementation of such a gas.

[0009] Indeed, during operation, there is a risk of one or more disks bursting, a phenomenon known by the English acronym UERF for "Uncontained Engine Rotor Failure". Such a burst can generate debris that can damage the pipe or the equipment of the fuel system. However, the hydrogen-air mixture in a volume proportion of hydrogen of between 13% and 65% becomes detonating. Thus, given the gas flow rate of the order of several hundred g / s, the gas pressure that can be greater than 100 bars and the energy level involved in the disk, the bursting of at least one disk is likely to cause an explosion. Such an explosion represents a catastrophic event for the aircraft. Also, the hydrogen-air mixture, in a volume proportion of hydrogen of between 4% and 75%, is flammable.

[0010] In order to avoid damaging the pipe and therefore limit the risk of explosion in the turbomachine, it was considered to reinforce the internal casing with an annular reinforcement shield in order to contain the disc(s) in the event of a burst.

[0011] However, such a solution is not feasible in current turbomachine configurations. In particular, reinforcing the internal casing would severely penalize the turbomachine in terms of cost and mass. Furthermore, the integration of such a reinforcement shield is not easy given the thickness required for such a shield to fulfill its protective function.

[0012] Furthermore, the use of combustible gas, in particular dihydrogen, requires a significant energy input. Indeed, in order to increase the energy density while reducing the volume of the combustible gas, in particular dihydrogen, the latter is typically stored in the aircraft at a temperature of approximately -253°C. However, in order to reduce the risks associated with the use of this combustible gas when cold, in particular the combustion instabilities associated with the use of this gas, and to increase the combustion efficiency, it is necessary to reheat the combustible gas before its injection into the combustion chamber. However, the reheating of this combustible gas requires a significant energy input which limits the energy capacities of the turbomachine.

[0013] In this context, there is a need to provide an aircraft turbomachine comprising a combustion chamber which can be supplied with combustible gas, while being lighter, economical, with a limited or even zero risk of explosion and with improved energy capacities. Summary of the invention

[0014] To this end, the invention provides a turbomachine for an aircraft, the turbomachine having a longitudinal axis and extending from front to rear along the axis Ion- horizontal with respect to the flow of a secondary airflow in the turbomachine, the turbomachine being remarkable in that it comprises from front to back:

[0015] - a rotor of a low pressure compressor,

[0016] - a rotor of a high pressure compressor,

[0017] - a rotor of a high pressure turbine mechanically connected to the rotor of the compressor high pressure,

[0018] - a rotor of a low pressure turbine mechanically connected to the rotor of the compressor low pressure,

[0019] - an annular combustion chamber,

[0020] the turbomachine further comprising:

[0021] - a primary flow vein of a primary air flow which passes successively, from upstream to downstream following the flow of the primary air flow, through the rotors of the low pressure and high pressure compressors, the combustion chamber, and the rotors of the high pressure and low pressure turbines, and

[0022] - a heat exchanger located in the primary vein and comprising a first air passage circuit from the high pressure compressor and supply to the combustion chamber, and a second gas passage circuit from the combustion chamber.

[0023] According to the invention, the combustion chamber is located at the rear of the low pressure turbine and therefore of the turbomachine.

[0024] Thus, thanks to this particular configuration of the turbomachine, the combustion chamber and therefore its potential fuel gas supply system are located outside the burst zone of the turbine disk. In the event of such a disk bursting, the risk of damaging the fuel gas supply system of the combustion chamber is limited. The risk of fuel gas leakage is therefore greatly reduced, thereby limiting the risk of explosion in the turbomachine and / or the aircraft.

[0025] Thanks to the invention, it is therefore possible to supply the turbomachine with combustible gas, in particular dihydrogen, without significant risk of explosion or fire.

[0026] The heat exchanger of the invention, which is in particular of the air / air type, makes it possible to improve the combustion efficiency of the turbomachine. Indeed, such a heat exchanger makes it possible to transfer the heat from the combustion gases to the cold, compressed air coming from the high-pressure compressor. The compressed air supplying the combustion chamber is thus heated and the quantity of combustible gas required is significantly reduced. Since the quantity of combustible gas used is reduced, the amount of energy required to heat the combustible gas in the combustion chamber is also limited. Also, since the quantity of combustible gas is reduced, the size of the heat exchanger is also minimal, which makes it possible to optimize the quantity of combustible gas consumed.

[0027] The energy performance of the turbomachine is thus significantly improved.

[0028] The invention may comprise one or more of the following features, taken in isolation from one another or in combination with one another:

[0029] - the combustion chamber is radially further from the longitudinal axis than the low pressure and high pressure compressors and the low pressure turbine,

[0030] - a secondary vein of flow of the secondary air flow located around the primary vein,

[0031] - an inter-vein casing separating the primary and secondary veins and delimiting a engine compartment in which the combustion chamber is located, and

[0032] - a gas outlet casing arranged coaxially inside the inter-vein casing and delimiting with the inter-vein casing a portion of the primary vein,

[0033] - the heat exchanger is located radially between the inter-vein casing and the casing gas outlet,

[0034] - the high pressure compressor comprises a first compression stage comprising at least one axial compressor and at least one second compression stage comprising a centrifugal compressor located at the rear of the axial compressor,

[0035] - the centrifugal compressor comprises an inlet of a primary air flow connected to the axial compressor and a primary air flow outlet connected to the first circuit of the first heat exchanger,

[0036] - the axial compressor comprises a rotor mechanically connected to the rotor of the com low pressure presser,

[0037] - the centrifugal compressor comprises a rotor mechanically connected to the rotor of the high pressure turbine,

[0038] - the high pressure turbine is a centripetal turbine comprising a gas inlet connected to the combustion chamber and a gas outlet connected to the low pressure turbine,

[0039] - the rotor of the low pressure compressor is connected to the rotor of the low pressure turbine by a low pressure shaft. Brief description of the figures

[0040] Other characteristics and advantages will emerge from the following description of non-limiting embodiments of the invention with reference to the appended drawings in which:

[0041] [Fig.l] is a schematic representation of an aircraft according to the invention,

[0042] [Fig.2] is a schematic representation in longitudinal section of a half- turbomachine according to the invention. Detailed description of the invention

[0043] An aircraft 1 is for example shown in [Fig. 1]. The aircraft 1 comprises a fuselage 2 and two wings 3 mounted on either side of the fuselage 2. The aircraft 1 further comprises at least two turbomachines 4 respectively secured to the two wings 3.

[0044] Each turbomachine 4 extends around and along a longitudinal axis X.

[0045] In the present application, the terms "front" and "rear" are defined with respect to the direction of circulation of a secondary air flow F2 in the turbomachine 4 along the longitudinal axis X, in particular from left to right in [Fig. 2]. In particular, the front is located at an air inlet 5 of the turbomachine 4, and the rear 6 is located at a combustion gas outlet of the turbomachine 4.

[0046] In the present application, the terms “upstream” and “downstream” are understood to refer to the direction of circulation of a primary air flow F1 in the turbomachine 4. In particular, the primary air flow F1 circulates in a primary vein v1 inside a secondary vein v2 surrounding this primary vein v1. A secondary air flow F2 circulates in the secondary vein v2.

[0047] The terms “axial”, “axially”, “radial”, “radially”, are defined relative to the longitudinal axis X of the turbomachine 4.

[0048] The terms "internal", "interior", "internally", "external", "exterior", "externally", are defined with respect to the distance from the longitudinal axis X along a radial axis extending radially with respect to the longitudinal axis X.

[0049] The turbomachine 4 is for example a twin-spool, twin-flow turbojet. By “twin-spool”, it is understood that the turbomachine 4 is equipped with a low-pressure compressor mechanically connected to a low-pressure turbine and a high-pressure compressor mechanically connected to a high-pressure turbine. By “twin-flow”, it is understood that two flows flow in the turbomachine, in particular the primary and secondary air flows F1, F2.

[0050] The turbomachine 4 comprises an engine M optionally comprising a fan 7, a low pressure body 8 and a high pressure body 9.

[0051] The fan 7 is located at the front of the low pressure and high pressure bodies 8, 9. The fan 7 comprises an air inlet cone 7a forming for example the air inlet 5 of the turbomachine 4. The fan 7 further comprises a disc centered on the longitudinal axis X and movable in rotation about the longitudinal axis X. The disc is axially connected to the air inlet cone 7a, for example by flanges. The fan 7 further comprises blades 7b regularly distributed around the longitudinal axis X and extending radially from the disc.

[0052] The fan 7 is driven in rotation by a fan shaft 7c. The fan shaft 7c is connected to the fan disc 7 and is centered on the longitudinal axis X. It extends for example inside the air inlet cone 7a.

[0053] The turbomachine 4 further comprises, from front to rear, a low-pressure compressor 10, a high-pressure compressor 11, a high-pressure turbine 12, a low-pressure turbine 13 and at least one annular combustion chamber 14.

[0054] The low pressure compressor 10 comprises a rotor 10a which is mechanically connected to a rotor 13a of the low pressure turbine 13. Advantageously, the rotor 10a of the low pressure compressor 10 is mechanically connected to the rotor 13a of the low pressure turbine 13 by a low pressure shaft 15 thus forming the low pressure body 8. The low pressure shaft 15 is annular and centered on the longitudinal axis X.

[0055] The high pressure compressor 11 comprises at least one rotor 18a mechanically connected to a rotor 13a of the high pressure turbine 13. The rotor 18a of the high pressure compressor 11 is mechanically connected to the rotor 13a of the high pressure turbine 13 by a high pressure shaft 16 forming the high pressure body 9. The high pressure shaft 16 is advantageously located outside the low pressure shaft 15.

[0056] The rotors 10a, 13a of the low-pressure compressor and turbine 10, 13 each comprise at least one disc centered on the longitudinal axis X and rotor blades regularly distributed around the disc (not shown).

[0057] The low-pressure compressors and turbines 10, 13 may each further comprise at least one row of stator blades mounted around the longitudinal axis X. The stator blades are mounted upstream or downstream of the rotor blades. A pair of rows of rotor blades and stator blades respectively forms a compressor or turbine stage. The low-pressure compressors and turbines 10, 13 may comprise one or more stages.

[0058] Preferably, the high-pressure turbine 12 is a centripetal turbine. The centripetal turbine comprises a gas inlet 12b connected to the combustion chamber 14 and a gas outlet 12c connected to the low-pressure turbine 13. The centripetal turbine further comprises a rotating impeller forming the rotor 12a of the high-pressure turbine 12. The movable impeller is located between the gas inlets and outlets 12b, 12c according to the flow of the gas stream in the centripetal turbine. The movable impeller has a bladed disc. The centripetal turbine further comprises a volute connected to the gas inlet 12b and a fixed distributor connecting the volute to the movable impeller. The gas inlet 12b is oriented axially while the volute makes it possible to generate a tangential component of the gas flow.

[0059] Preferably, the high pressure compressor 11 comprises a first compression stage comprising at least one axial compressor 17 and at least one second compression stage comprising a centrifugal compressor 18. The axial compressor 17 is located in front of the centrifugal compressor 18. The axial compressor 17 is for example axially aligned with the low pressure compressor 10.

[0060] The axial compressor 17 comprises at least one rotor 17a comprising at least one disc centered on the longitudinal axis X and rotor blades regularly distributed around the disc (not shown). The axial compressor 17 may further comprise at least one row of stator blades mounted around the longitudinal axis X. The stator blades are mounted upstream or downstream of the rotor blades. A pair of rows of rotor blades and stator blades respectively forms an axial compressor stage 17. The axial compressor 17 may comprise one or more stages.

[0061] The centrifugal compressor 18 comprises a rotor 18a and an inlet of the primary air flow F1 connected to the low pressure compressor 10 and an outlet of the primary air flow F1. The inlet of the primary air flow F1 is oriented axially while the outlet of the primary air flow F1 is oriented radially.

[0062] The rotor 18a of the centrifugal compressor 18 is connected to the rotor 12a of the high pressure turbine 12, in particular to the rotor 12a of the centripetal turbine by the high pressure shaft 16 and the rotor 17a of the axial compressor 17 is connected to the rotor 10a of the low pressure compressor 10.

[0063] The turbomachine 4 may further comprise a speed reducer 19 connecting the fan 7 to the low pressure shaft 8. The speed reducer 19 in particular connects the fan shaft 7c to the low pressure shaft 15. The speed reducer 19 may be of the planetary or epicyclic type.

[0064] Advantageously, the turbomachine 4 further comprises an air inlet casing 20. The inlet casing 20 is centered on the longitudinal axis X. The inlet casing 20 is arranged axially between the fan 7 and the low-pressure compressor 10.

[0065] Advantageously, the turbomachine 4 further comprises a rectifier 21 arranged axially between the fan 7 and the low-pressure compressor 10. The rectifier 21 comprises fixed blades 21a rotating around the longitudinal axis X. The rectifier 21 makes it possible to straighten the secondary air flow F2 at the outlet of the fan 7.

[0066] The turbomachine 4 may further comprise a nacelle 22. The nacelle 22 is centered on the longitudinal axis X and surrounds the fan 7. The fan 7 is therefore of the ducted type. According to another example, the turbomachine 4 is without a nacelle. The fan 7 is therefore of the unducted type.

[0067] The turbomachine 4 further comprises an inter-vein casing 23 centered on the longitudinal axis X. The inter-vein casing 23 is arranged inside the fan casing 22. The inter-vein casing 23 delimits an engine compartment C.

[0068] The turbomachine 4 may further comprise a gas outlet casing 26 located coaxially inside the inter-vein casing 23. The gas outlet casing 26 delimits with the inter-vein casing 23 a portion of the primary vein vl. It is arranged at the rear of the low-pressure turbine 13.

[0069] The fan shaft 7c is guided in rotation about the longitudinal axis X by a first bearing 24a. The first bearing 24a is arranged radially between the inlet casing 20 and the fan shaft 7c. Advantageously, the first bearing 24a is supported by a first bearing support 25a extending radially inward from the inlet casing 20.

[0070] The low pressure shaft 15 is guided in rotation about the longitudinal axis X by a second bearing 24b and a third bearing 24c. The second bearing 24b is arranged axially between the first bearing 24a and the third bearing 24c. The second bearing 24b is located radially between the inlet casing 20 and the low pressure shaft 15. Advantageously, the second bearing 24b is supported by a second bearing support 25b extending radially inward from the inlet casing 20. The third bearing 24c is arranged radially between the low pressure shaft 15 and the gas outlet casing 26. Advantageously, the third bearing 24c is supported by a third bearing support 25c extending radially inward from the gas outlet casing 26.

[0071] The axial compressor 17 is guided in rotation by the low pressure shaft 13 via a fourth bearing 24d. The fourth bearing 24d is located radially between the axial compressor 17 and the low pressure shaft 15. Advantageously, the fourth bearing 24d is supported by a fourth bearing support 25d extending radially inward from the axial compressor 17.

[0072] The high pressure shaft 16 is guided in rotation by a fifth bearing 24e. The fifth bearing 24e is located axially between the fourth bearing 24d and the third bearing 24c. The fifth bearing 24e is located radially between the high pressure shaft 16 and the low pressure shaft 15. Advantageously, the fifth bearing 24e is supported by a fifth bearing support 25e extending radially inwards.

[0073] Each bearing 24a, 24b, 24c, 24d, 24e comprises a bearing such as rollers or balls arranged between an outer ring and an inner ring.

[0074] According to the invention, the combustion chamber 14 is arranged at the rear of the low pressure turbine 13, in particular at the rear of the last rotor of the low pressure turbine 13.

[0075] Preferably, the combustion chamber 14 is radially further from the longitudinal axis X of the turbomachine 4 than the low-pressure turbine 13 and the low- and high-pressure compressors 10, 11. The combustion chamber 14 is preferably located in the engine compartment C and therefore outside the gas outlet casing 26. The combustion chamber 14 is for example connected to the inter-vein casing 23 by radial arms 23a. The radial arms 23a extend between the combustion chamber 14 and the inter-vein casing 23.

[0076] The combustion chamber 14 comprises an annular wall 14a delimiting an internal combustion space 14b.

[0077] The turbomachine 4 further comprises a supply system 27 for the chamber combustion system 14, in particular in combustible gas. The feed system 27 is located at the rear of the low pressure turbine 13.

[0078] The supply system 27 comprises a pipe 27a for the passage of the combustible gas and optionally a source 27b of combustible gas connected to the pipe 27a.

[0079] The pipe 27a is connected to the combustion chamber 14, and may be located in a compartment extending from the combustion chamber 14 to the source 27b. Such a configuration makes it possible to separate the supply system 27 from the other elements of the turbomachine 4.

[0080] The source 27b of combustible gas is located for example in the wing 3 of the aircraft 1 or in the fuselage 2 of the aircraft 1. The combustible gas is for example natural gas, methane (CH4), or ammonia (NH3). The combustible gas is particularly preferably dihydrogen (H2).

[0081] The primary and secondary veins v1, v2 are annular. The secondary vein v2 surrounds the primary vein v1. The primary and secondary veins v1, v2 are for example separated by the inter-vein casing 23. The secondary vein v2 is for example located between the fan casing 22 and the inter-vein casing 23.

[0082] In operation, the blower 7 allows the suction of a main air flow F which is divided into a primary air flow F1 and a secondary air flow F2.

[0083] The primary air flow F1 passes from front to back through the low pressure compressor 10 and the high pressure compressor 11. The compressed primary air flow F1 is then directed into the combustion chamber 14. The compressed primary air flow F1 is mixed with the combustible gas in the combustion chamber 14. The combustion reaction in the combustion chamber 14 forms gases. The gases pass from back to front through the high pressure turbine 12 which redirects the gases from front to back into the low pressure turbine 13.

[0084] According to the invention, the turbomachine 4 further comprises at least one heat exchanger 28. The heat exchanger 28 is located in the primary vein v1. In a particularly preferred manner, the heat exchanger 28 is located at the rear of the low-pressure turbine 13, and is advantageously arranged radially between the inter-vein casing 23 and the gas outlet casing 26. In this way, the heat exchanger 28 has little impact on the configuration of the turbomachine 4. Also, the pressure losses in the high-pressure and low-pressure turbines 12, 13 are minimized. Furthermore, since the heat exchanger 28 is located between these casings 23, 26, the heat exchange is favored since the gas flow at the outlet of the low-pressure turbine 13 directly feeds the heat exchanger 28.

[0085] The heat exchanger 28 is of the air / air or gas / gas type. The heat exchanger 28 comprises a first air passage circuit from the high-pressure compressor 11 and supply of the combustion chamber 14 and a second circuit for passing gases from the combustion chamber 14. The first circuit is connected to the combustion chamber 14 and to the centrifugal compressor 18, in particular to the outlet of the primary air flow Fl of the centrifugal compressor 18.

[0086] Thus, the heat exchanger 28 makes it possible to heat the compressed primary air flow Fl by heat exchange thanks to the passage of the flow of gases resulting from the combustion in the second circuit of the heat exchanger 28. The first circuit makes it possible to convey the heated primary air flow Fl to the combustion chamber 14.

[0087] Thus, according to the invention, the primary air flow F1 at the outlet of the high-pressure compressor 11 is heated in the heat exchanger 28 before being transferred into the combustion chamber 14.

[0088] Furthermore, the combination of the centripetal turbine 12 and the centrifugal compressor 18 also promotes heat exchanges between the flow of compressed and cold air at the outlet of the high-pressure compressor 11 and the flow of hot gases resulting from the combustion. These two flows being close in the primary vein vl promote the heating of the flow of compressed primary air Fl before its introduction into the combustion chamber 14.

[0089] The secondary air flow F2 flows into the secondary vein v2.

[0090] According to the invention, the combustion chamber 14 is located at the rear of the low pressure 8 and high pressure 9 bodies, and therefore outside the burst zone of the turbine disc 12, 13.

[0091] In the event of a turbine disk 12, 13 bursting, the risks of damaging the supply system 27 are greatly limited and the risk of fuel gas leaking is consequently also limited.

[0092] Thanks to the particular configuration of the turbomachine 4 according to the invention, the risks of explosion in the turbomachine 4 and / or the aircraft 1 are therefore greatly reduced.

[0093] Furthermore, the heat exchanger 28 according to the invention improves the energy capacities of the turbomachine 4 in that it reduces the heating requirements of the combustible gas, in particular of the dihydrogen, by promoting heat exchanges between the primary air flow F1 compressed at the outlet of the high-pressure compressor 12 and the gases resulting from the combustion, reducing the requirements for combustible gas.

[0094] Finally, in a particularly preferred manner, the combination of the centripetal turbine 12 and the centrifugal compressor 18 further promotes these heat exchanges and further reduces the need for heating the combustible gas, in particular the dihydrogen.

Claims

Claims

1. A turbomachine (4) for an aircraft (1), the turbomachine (4) having a longitudinal axis (X) and extending from front to rear along the longitudinal axis (X) relative to the flow of a secondary air flow (F2) in the turbomachine (4), the turbomachine (4) being characterized in that it comprises from front to rear: - a rotor (10a) of a low-pressure compressor (10), - a rotor (17a, 18a) of a high-pressure compressor (11), - a rotor (12a) of a high-pressure turbine (12) mechanically connected to the rotor (18a) of the high-pressure compressor (11), - a rotor (13a) of a low-pressure turbine (13) mechanically connected to the rotor (10a) of the low-pressure compressor (10), - an annular combustion chamber (14), the turbomachine (4) further comprising: - a primary vein (vl) flow of a primary air flow (Fl) which passes successively, from upstream to downstream following the flow of the primary air flow (Fl),through the rotors of the low pressure (10) and high pressure (11) compressors, the combustion chamber (14), and the rotors of the high pressure and low pressure turbines (12, 13), and - a heat exchanger (28) located in the primary vein (vl) and comprising a first circuit for the passage of air from the high pressure compressor (11) and supply to the combustion chamber (14), and a second circuit for the passage of gases from the combustion chamber (14).,

2. Turbomachine according to the preceding claim, characterized in that the combustion chamber (14) is radially further from the longitudinal axis (X) than the low pressure and high pressure compressors (10, 11) and than the low pressure turbine (13).

3. Turbomachine according to any one of the preceding claims, characterized in that it comprises: - a secondary flow stream (v2) of the secondary air flow (F2) located around the primary flow stream (vl), - an inter-flow casing (23) separating the primary and secondary flow streams (vl, v2) and delimiting an engine compartment (C) in which the combustion chamber (14) is located, and - a gas outlet casing (26) arranged coaxially inside the inter-flow casing (23) and delimiting with the inter-flow casing (23) a portion of the primary vein (vl).

4. Turbomachine according to the preceding claim, characterized in that the heat exchanger (28) is located radially between the inter-vein casing (23) and the gas outlet casing (26).

5. Turbomachine according to any one of the preceding claims, characterized in that the high pressure compressor (11) comprises a first compression stage comprising at least one axial compressor (17) and at least one second compression stage comprising a centrifugal compressor (18) located at the rear of the axial compressor (17).

6. Turbomachine according to the preceding claim, characterized in that the centrifugal compressor (18) comprises an inlet of a primary air flow (Fl) connected to the axial compressor (17) and an outlet of the primary air flow (Fl) connected to the first circuit of the first heat exchanger (28).

7. Turbomachine according to one of claims 5 or 6, characterized in that the axial compressor (17) comprises a rotor (17a) mechanically connected to the rotor (10a) of the low pressure compressor (10).

8. Turbomachine according to the preceding claim, characterized in that the centrifugal compressor (18) comprises a rotor (18a) mechanically connected to the rotor (12a) of the high pressure turbine (12).

9. Turbomachine according to any one of the preceding claims, characterized in that the high pressure turbine (12) is a centripetal turbine comprising a gas inlet (12b) connected to the combustion chamber (14) and a gas outlet (12c) connected to the low pressure turbine (13).

10. Turbomachine according to any one of the preceding claims, characterized in that the rotor (10a) of the low pressure compressor (10) is connected to the rotor (13a) of the low pressure turbine (13) by a low pressure shaft (15).

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