AIRCRAFT TURBOMACHINE

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 (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 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). Abbreviated 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 turbomachinery.

[0002] The invention relates particularly 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 movable fan rotating 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 includes a pipeline for conveying the fuel from a fuel source to the combustion chamber.

[0005] The blower allows the intake of an airflow which is divided, for example, into a primary airflow and a secondary airflow. The primary airflow passes through a primary channel of the turbomachine delimited by an internal casing, while the secondary airflow is directed towards a secondary channel surrounding the primary channel.

[0006] The primary airflow is compressed within the compressors. Each compressor comprises a rotating disc about its longitudinal axis. Each compressor further comprises vanes evenly distributed around the disc, which compress the primary airflow.

[0007] Compressed air is then mixed with the fuel and burned in the combustion chamber. The combustion gases pass through the turbines. Each turbine has a disk centered on its longitudinal axis and blades regularly distributed around the disk, which exert force on the gases exiting the combustion chamber. The gases finally escape through the nozzle, the cross-section of which accelerates 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 cleaner energy source, in particular a combustible gas such as dihydrogen (H2). However, the current configuration of turbomachinery does not easily allow the implementation of such a gas.

[0009] Indeed, during operation, there is a risk of one or more discs bursting, a phenomenon known by the English acronym UERF for "Uncontained Engine Rotor Failure." Such a burst can generate debris that may damage the fuel system piping or equipment. Furthermore, a hydrogen-air mixture with a hydrogen volume between 13% and 65% becomes explosive. Thus, given the gas flow rate of several hundred g / s, the gas pressure which can exceed 100 bar, and the energy level involved in the disc, the bursting of at least one disc is likely to cause an explosion. Such an explosion represents a catastrophic event for the aircraft. Also, a dihydrogen-air mixture with a dihydrogen volume between 4% and 75% is flammable.

[0010] In order to avoid damaging the pipeline and thus limit the risk of explosion in the turbomachine, it was envisaged to reinforce the internal casing with an annular reinforcement shield in order to contain the disc(s) in the event of bursting.

[0011] However, such a solution is not feasible in current turbomachine configurations. In particular, reinforcing the internal casing would significantly increase the turbomachine's cost and weight. Furthermore, integrating such a reinforcing shield is not easy, given the thickness required for such a shield to fulfill its protective function.

[0012] Furthermore, the use of combustible gases, particularly dihydrogen, requires a significant energy input. Indeed, in order to increase the energy density while reducing the volume of the combustible gas, especially 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 at low temperatures, particularly combustion instabilities, and to increase combustion efficiency, it is necessary to heat the combustible gas before its injection into the combustion chamber. However, heating this combustible gas requires a significant energy input, which limits the turbomachine's energy capacity.

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

[0014] To this end, the invention proposes a turbomachine for an aircraft, the turbomachine having a longitudinal axis and extending from front to rear along the Ion- axis gitudinal with respect to the flow of a secondary air stream 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 high-pressure turbine rotor mechanically connected to the compressor rotor high pressure,

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

[0019] - an annular combustion chamber,

[0020] the turbomachine further comprising:

[0021] - a primary flow vein of a primary airflow which passes successively, from upstream to downstream following the flow of the primary air stream, 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, is located outside the burst zone of the turbine disk. In the event of a burst of such a disk, the risk of damaging the combustion chamber's fuel gas supply system 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-to-air type, improves the combustion efficiency of the turbomachine. Indeed, such a heat exchanger transfers heat from the combustion gases to the cold, compressed air from the high-pressure compressor. The compressed air supplying the combustion chamber is thus preheated, and the amount of fuel gas required is significantly reduced. Since the amount of fuel gas used is reduced, the amount of energy required to heat the fuel gas in the combustion chamber is also limited. Furthermore, because the amount of fuel gas is reduced, the size of the heat exchanger is also minimal, which optimizes 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 individually or in combination with each other:

[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 flow vein of the secondary airflow 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 housing arranged coaxially inside the inter-vein housing and delimiting, with the inter-vein casing, a portion of the primary vein,

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

[0034] - the high-pressure compressor includes 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 includes a primary airflow inlet connected to the axial compressor and a primary airflow outlet connected to the first circuit of the first heat exchanger,

[0036] - the axial compressor includes 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 features and advantages will become apparent from the following description of non-limiting embodiments of the invention with reference to the accompanying drawings in which:

[0041] [Fig. 1] is a schematic representation of an aircraft according to the invention,

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

[0043] An aircraft 1 is shown for example 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 attached 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 flow of a secondary airflow 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" refer to the direction of flow of a primary airflow Fl in the turbomachine 4. In particular, the primary airflow Fl flows in a primary channel vl within a secondary channel v2 surrounding this primary channel vl. A secondary airflow F2 flows in the secondary channel v2.

[0047] The terms "axial", "axially", "radial", "radially", are defined with respect 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 engine. "Twin-spool" means 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. "Twin-flow" means that two flows pass through the turbomachine, in particular the primary and secondary air flows Fl, F2.

[0050] The turbomachine 4 comprises an engine M optionally including 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 disk centered on the longitudinal axis X and rotatable about the longitudinal axis X. The disk 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 disk.

[0052] The blower 7 is driven in rotation by a blower shaft 7c. The blower shaft 7c is connected to the blower disk 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 back, 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 includes 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 compressor and low pressure turbine 10, 13 each comprise at least one disk centered on the longitudinal axis X and rotor blades regularly distributed around the disk (not shown).

[0057] The low-pressure compressors and turbine 10, 13 may further each 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 forms, respectively, one compressor or turbine stage. The low-pressure compressors and turbine 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 impeller is located between the gas inlet and outlet 12b, 12c, depending on the gas flow direction in the centripetal turbine. The 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 impeller. The gas inlet 12b is axially oriented, while the volute generates a tangential component of the gas flow.

[0059] Preferably, the high-pressure compressor 11 comprises a first compression stage including at least one axial compressor 17 and at least one second compression stage including 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 The axial compressor 17 consists of a disk centered on the longitudinal axis X and rotor blades evenly distributed around the disk (not shown). It may further include 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 each form one stage of the axial compressor 17. The axial compressor 17 may comprise one or more stages.

[0061] The centrifugal compressor 18 includes a rotor 18a and a primary airflow inlet Fl connected to the low pressure compressor 10 and a primary airflow outlet FL. The primary airflow inlet Fl is axially oriented while the primary airflow outlet Fl is radially oriented.

[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 include 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 housing 20. The inlet housing 20 is centered on the longitudinal axis X. The inlet housing 20 is arranged axially between the fan 7 and the low-pressure compressor 10.

[0065] Advantageously, the turbomachine 4 further comprises a straightener 21 arranged axially between the fan 7 and the low-pressure compressor 10. The straightener 21 comprises fixed blades 21a rotating about the longitudinal axis X. The straightener 21 allows the secondary airflow F2 to be straightened at the outlet of the fan 7.

[0066] The turbomachine 4 may further include 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 enclosed type. According to another example, the turbomachine 4 is without a nacelle. The fan 7 is therefore of the unenclosed type.

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

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

[0069] The blower shaft 7c is guided in rotation around the longitudinal axis X by a First bearing 24a. The first bearing 24a is arranged radially between the inlet housing 20 and the blower shaft 7c. Advantageously, the first bearing 24a is supported by a first bearing support 25a extending radially inwards from the inlet housing 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 housing 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 housing 20. The third bearing 24c is arranged radially between the low-pressure shaft 15 and the gas outlet housing 26. Advantageously, the third bearing 24c is supported by a third bearing support 25c extending radially inward from the gas outlet housing 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 inwards 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 exhaust housing 26. The combustion chamber 14 is, for example, connected to the inter-flow housing 23 by radial arms 23a. The radial arms 23a extend between the combustion chamber 14 and the inter-flow housing 23.

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

[0077] The turbomachine 4 further comprises a chamber feeding system 27 combustion 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 includes a pipeline 27a for the passage of combustible gas and optionally a source 27b of combustible gas connected to the pipeline 27a.

[0079] The pipe 27a is connected to the combustion chamber 14, and can 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 aircraft 1 or in the fuselage 2 of 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 vl, v2 are annular. The secondary vein v2 surrounds the primary vein vl. The primary and secondary veins vl, v2 are separated, for example, by the inter-vein housing 23. The secondary vein v2 is, for example, located between the blower housing 22 and the inter-vein housing 23.

[0082] In operation, the blower 7 allows the aspiration of a main airflow F which divides into a primary airflow Fl and a secondary airflow F2.

[0083] The primary airflow Fl passes from front to back through the low-pressure compressor 10 and the high-pressure compressor 11. The compressed primary airflow Fl is then directed into the combustion chamber 14. The compressed primary airflow Fl is mixed with the fuel 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 flow vl. Preferably, the heat exchanger 28 is located at the rear of the low-pressure turbine 13 and is advantageously arranged radially between the interflow 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. Moreover, since the heat exchanger 28 is located between these casings 23, 26, heat exchange is enhanced because the gas flow exiting 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 the passage of 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 allows the primary airflow Fl compressed to be heated by heat exchange thanks to the passage of the gas flow from the combustion in the second circuit of the heat exchanger 28. The first circuit allows the heated primary airflow Fl to be conveyed to the combustion chamber 14.

[0087] Thus, according to the invention, the primary airflow Fl 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 exchange between the cold, compressed air flow exiting the high-pressure compressor 11 and the hot gas flow from combustion. Since these two flows are in close proximity in the primary stream vl, they promote the heating of the compressed primary air flow Fl before its introduction into the combustion chamber 14.

[0089] The secondary airflow 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 bursting zone of the turbine disc 12, 13.

[0091] In the event of a burst of a turbine disc 12, 13, the risks of damaging the supply system 27 are greatly limited and the risk of fuel gas leakage 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] In addition, 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 fuel gas, in particular of dihydrogen, by promoting heat exchanges between the primary air flow Fl compressed at the outlet of the high-pressure compressor 12 and the gases from combustion, reducing the fuel gas requirements.

[0094] Finally, and particularly preferably, the combination of the centripetal turbine 12 and the centrifugal compressor 18 further promotes these heat exchanges and further reduces the need to heat the fuel gas, in particular dihydrogen.

Claims

Demands

1. 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 stream (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 duct (vl) of flow of a primary airflow (Fl) which passes successively, from upstream to downstream following the flow of the primary airflow (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 stream (vl) and comprising a first air passage circuit from the high-pressure compressor (11) and supply to the combustion chamber (14), and a second gas passage circuit from the combustion chamber (14).

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

3. Turbomachine according to any one of the preceding claims, characterized in that it comprises: - a secondary airflow channel (v2) located around the primary airflow channel (vl), - an inter-channel housing (23) separating the primary and secondary airflow channels (vl, v2) and defining an engine compartment (C) in which the combustion chamber (14) is located, and - a gas outlet housing (26) arranged coaxially inside the inter-channel housing (23) and defining with the inter-channel housing (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 housing (23) and the gas outlet housing (26).

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

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

7. Turbomachine according to any 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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