TURBOMACHINE FOR AIRCRAFT

Abstract

The invention relates to a turbomachine (4) comprising: - a low-pressure casing (8), - a high-pressure casing (9), - an annular combustion chamber (25), and the turbomachine (4) being characterized in that: - the high-pressure turbine (16) is located at the rear of the low-pressure casing (8), and - the combustion chamber (25) is located at the rear of the low-pressure and high-pressure casings (8, 9), and in that it further comprises: - a first heat exchanger (27) located in the primary flow (v1) between the low-pressure turbine (11) and the high-pressure turbine (16), the first heat exchanger (27) comprising a first air passage circuit from the high-pressure compressor (15) supplying the combustion chamber (25), and a second gas passage circuit from the combustion chamber (25). 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 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. 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 and more economical, the risk of explosion of which is limited or even zero and the energy capacities of which are improved. 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 axis Ion- horizontal with respect to the flow of a secondary air flow in the turbomachine, the turbomachine comprising:

[0015] - a low pressure body comprising a rotor of a low pressure compressor connected mechanically to a rotor of a low pressure turbine,

[0016] - a high pressure body comprising a rotor of a high pressure compressor connected mechanically to a rotor of a high pressure turbine,

[0017] - an annular combustion chamber, and

[0018] - 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.

[0019] The turbomachine is remarkable in that:

[0020] - the high pressure turbine is located at the rear of the low pressure body and,

[0021] - the combustion chamber is located at the rear of the low pressure and high pressure bodies pressure,

[0022] and in that it further comprises:

[0023] - a first heat exchanger located in the primary vein between the lower turbine pressure and the high pressure turbine, the first heat exchanger comprising a first circuit for passing air from the high pressure compressor and supplying the combustion chamber, and a second circuit for passing gas from the combustion chamber.

[0024] According to the invention, the turbomachine therefore comprises from front to rear along the longitudinal axis, and in the direction of flow of the secondary air flow, the low pressure body and the high pressure turbine. According to the invention, the combustion chamber is located at the rear of the low pressure and high pressure bodies.

[0025] 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 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, thus limiting the risk of explosion in the turbomachine and / or the aircraft.

[0026] Thanks to the invention, it is therefore possible to supply combustion gas, in particular dihydrogen, to the turbomachine without significant risk of explosion or fire.

[0027] In addition, the first internal heat exchanger makes it possible to improve the combustion efficiency. Indeed, such a first internal heat exchanger makes it possible to transfer the heat from the combustion gases to the cold air coming from the high-pressure compressor. The compressed air injected into the combustion chamber is thus heated and the amount of fuel gas required for combustion is significantly reduced. Such a heat exchanger thus reduces the amount of energy required to heat the fuel gas since the amount of fuel gas used is reduced. Also, since the amount of fuel gas is reduced, the size of the first internal heat exchanger is also minimal, which improves the amount of fuel gas consumed.

[0028] The energy performance of the turbomachine is thus significantly improved.

[0029] The presence of such a first heat exchanger has little impact on the configuration of the turbomachine and is on the contrary facilitated by the configuration of the turbomachine according to the invention. Indeed, the first heat exchanger is located in the primary vein in which the hot air flow and the cold air flow circulate at the outlet of the high-pressure compressor, thus making the heat exchange between the flows favorable without impacting the mass and / or the size of the turbomachine.

[0030] In addition, the high pressure turbine is located at the rear of the low pressure body, it can be directly connected to the high pressure compressor, thus reducing the need for a speed reducer.

[0031] Also, thanks to the particular configuration of the turbomachine, and in particular of the high pressure turbine located at the rear of the low pressure body, it is possible to connect the rotor of each body, such as the rotors of the low pressure compressor and the low pressure turbine by short shafts facilitating the assembly and maintenance of the turbomachine.

[0032] The invention may comprise one or more of the following features, taken in isolation from one another or in combination with one another:

[0033] - the primary vein comprises:

[0034] - a first flow circuit of a cold primary air flow from the front to the rear in the low pressure and high pressure compressors to the combustion chamber,

[0035] - a second flow circuit for gases from the combustion chamber of back to front in the high pressure turbine, the first heat exchanger and the low pressure turbine,

[0036] - 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,

[0037] - the centrifugal compressor comprises an air inlet connected to the axial compressor and an air outlet connected to the first circuit of the first heat exchanger,

[0038] - the centrifugal compressor comprises a rotor mechanically connected to the rotor of the high pressure turbine and axial compressor includes a rotor mechanically connected to the rotor of the low pressure compressor,

[0039] - the axial compressor is located between the low pressure compressor and the low turbine pressure and the centrifugal compressor is located in front of the high pressure turbine,

[0040] - a second heat exchanger located in front of the low pressure turbine,

[0041] - a blower mechanically connected to the rotor of the low pressure compressor, the second heat exchanger being located between the blower and the low pressure turbine,

[0042] - the blower is connected to the rotor of the low pressure compressor via of a first mechanical speed reducer,

[0043] - the rotor of the high pressure turbine is directly connected to the rotor of the compressor high pressure. Brief description of the figures

[0044] 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:

[0045] [Fig.l] is a schematic representation of an aircraft according to the invention,

[0046] [Fig.2] is a schematic representation in longitudinal section of a half- turbomachine according to the invention,

[0047] [Fig. 3] is a schematic representation in longitudinal section of a half-turbomachine according to an advantageous embodiment of the invention,

[0048] [Fig.4] is a schematic representation of the first and second heat exchangers,

[0049] [Fig. 5] is a schematic representation in longitudinal section of a half-turbomachine according to another advantageous embodiment of the invention. Detailed description of the invention

[0050] An aircraft 1 is for example shown in [Fig.l]. 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. The aircraft 1 may comprise a single turbomachine 4 or more turbomachines 4.

[0051] Each turbomachine 4 extends around and along a longitudinal axis X.

[0052] In the present application, the terms “front” and “rear” are defined in relation 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 FIGS. 2 to 4. 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.

[0053] In the present application, the terms “upstream” and “downstream” are understood in relation 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 vl and in which the secondary air flow F2 circulates.

[0054] The terms “axial”, “axially”, “radial”, “radially”, are defined relative to the longitudinal axis X of the turbomachine 4.

[0055] The terms "internal", "interior", "internally", "external", "exterior", "externally", are defined with respect to the distance from the longitudinal axis X along a radial axis perpendicular to the longitudinal axis X.

[0056] The turbomachine 4 is for example a twin-spool, twin-flow turbojet. It comprises an engine M optionally comprising a fan 7, a low-pressure spool 8 and a high-pressure spool 9.

[0057] 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 5. 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 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.

[0058] The low pressure body 8 comprises a low pressure compressor 10 and a low pressure turbine 11. The low pressure compressor 10 is located in front of the low pressure turbine 11. The low pressure compressor 10 is therefore axially offset relative to the low pressure turbine 11. Preferably, the low pressure compressor 10 is radially closer to the longitudinal axis X compared to the high pressure turbine 11 which is radially further from the longitudinal axis X.

[0059] The low pressure compressor 10 comprises a rotor 10a mechanically connected to a rotor 11a of the low pressure turbine 11. The rotors 10a, 11a each comprise at least one disc centered on the longitudinal axis X and rotor blades regularly distributed around the disc (not shown).

[0060] The low-pressure compressor 10 and the low-pressure turbine 11 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 low-pressure compressor and low-pressure turbine stage 10, 11. The low-pressure compressor 10 and the low-pressure turbine 11 may comprise one or more stages.

[0061] The low pressure body 8 further comprises a first low pressure shaft 12 connecting the rotor 10 of the low pressure compressor 10 and the rotor 11a of the low pressure turbine 11.

[0062] Advantageously, the low pressure body comprises a second low pressure shaft 13 connecting the fan 7 to the rotor 10a of the low pressure compressor 10. The second low pressure shaft 13 is radially closer to the longitudinal axis X relative to the first low pressure shaft 12. The first and second low pressure shafts 12, 13 are axially offset.

[0063] The second low pressure shaft 13 advantageously connects the blower 7 to the rotor 10a of the low pressure compressor 10 via a first mechanical speed reducer 14. The second low pressure shaft 13 thus comprises an inlet portion 13a connected to the blower 7 and to the first mechanical speed reducer 14 and an outlet portion 13b connected to the first mechanical speed reducer 14 and to the rotor 10a of the low pressure compressor 10.

[0064] The high pressure body 9 comprises a high pressure compressor 15 and a high pressure turbine 16.

[0065] According to the invention, the high pressure turbine 16 is located at the rear of the low pressure body 8. The high pressure compressor 15 is located at the front of the high pressure turbine 16. The high pressure compressor 15 can be located axially between the low pressure turbine 11 and the high pressure turbine 16 or be located inside the low pressure turbine 11 and therefore radially aligned with the latter.

[0066] Preferably, the high-pressure compressor 15 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 and is located in front of the low-pressure turbine 11. The centrifugal compressor 18 may be located between the low-pressure and high-pressure turbines 11, 16.

[0067] The high-pressure compressor 15 comprises at least one rotor 17a, 18a. In particular, 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.

[0068] The centrifugal compressor 18 comprises a rotor 18a and an inlet for a primary air flow F1 and an outlet for the primary air flow F1. The inlet for the primary air flow F1 is oriented axially while the outlet for the primary air flow F1 is oriented radially.

[0069] The rotor 17a, 18a of the high pressure compressor 15 is mechanically connected to a rotor 16a of the high pressure turbine 16. In particular, the rotor 18a of the centrifugal compressor 18 is connected to the rotor 16a of the high pressure turbine 16.

[0070] The high pressure body 9 may further comprise a high pressure shaft 19a connecting the rotor 16a of the high pressure turbine 16 to the rotor 18a of the high pressure compressor 15.

[0071] The low pressure and high pressure shafts 13, 19 are centered on the longitudinal axis X. They are axially spaced from each other.

[0072] 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.

[0073] Advantageously, the turbomachine 4 further comprises a rectifier 21 arranged at the rear of the fan 7 and at the front of 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.

[0074] The turbomachine 4 further comprises a nacelle 22. The nacelle 22 is centered on the longitudinal axis X and surrounds the fan 7. The fan 7 is of the ducted type. According to another example, the turbomachine 4 does not comprise a nacelle 22 around the fan 7. It is of the unducted type.

[0075] 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.

[0076] The low pressure shaft 13 is guided in rotation about the longitudinal axis X by a front bearing 24a and optionally a first rear bearing 24b. The front bearing 24a is for example arranged radially between the input casing 20 and the low pressure shaft 13, in particular the input portion 13a.

[0077] Furthermore, advantageously, the axial compressor 17 is guided in rotation by the low pressure shaft 13 via a second rear bearing 24c. The second rear bearing 24c is located between the low pressure shaft 13, in particular the outlet portion 13a and the axial compressor 17. The second rear bearing 24c is located at the rear of the first rear bearing 24b.

[0078] Advantageously, the rotor 16a of the high-pressure turbine 16 is guided in rotation by a third rear bearing 24d. The third rear bearing 24d is arranged radially between the rotor 16a of the high-pressure turbine 16 and a casing 25a.

[0079] Each bearing 24a, 24b, 24c, 24d comprises a bearing such as rollers or balls arranged between an outer ring carried by a housing and an inner ring carried by a shaft.

[0080] According to the invention, the turbomachine 4 further comprises an annular combustion chamber 25. The combustion chamber 25 comprises a casing 25a defining a internal cavity 25b.

[0081] According to the invention, the combustion chamber 25 is arranged at the rear of the high-pressure turbine 16, in particular at the rear of the last rotor of the high-pressure turbine 16.

[0082] The turbomachine 4 further comprises a supply system 26 for the combustion chamber 25, in particular for combustible gas. The supply system 26 is located at the rear of the high-pressure turbine 16.

[0083] The supply system 26 comprises a pipe 26a for the passage of the combustible gas and optionally a source 26b of combustible gas connected to the pipe 26a.

[0084] The pipe 26a is connected to the combustion chamber 25, and may be located in a compartment extending from the combustion chamber 25 to the source 26b. Such a configuration makes it possible to separate the supply system 26 from the other elements of the turbomachine 4.

[0085] The source 26b 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 advantageously dihydrogen (H2).

[0086] 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 defined between the fan casing 22 and the inter-vein casing 23.

[0087] The low pressure compressor 10, the low pressure turbine 11, the high pressure compressor 15 and the high pressure turbine 16 are located in the primary vein vl.

[0088] 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.

[0089] The primary vein vl comprises a first circuit cl configured to conduct the primary air flow Fl from front to back. The primary air flow Fl passes from front to back, the low pressure compressor 10, and the high pressure compressor 15 in the first circuit cl.

[0090] The compressed primary air flow Fl is then directed into the combustion chamber 23. The compressed primary air flow Fl is mixed with the combustible gas in the combustion chamber 23. The combustion reaction in the combustion chamber 23 forms gases G.

[0091] The primary vein v1 further comprises a second circuit c2 configured to conduct the flow of gas G resulting from the combustion from rear to front. The gas G passes from rear to front through the high pressure turbine 16 and the low pressure turbine 11 to the secondary vein v2.

[0092] The first circuit cl of the primary vein vl is advantageously located radially closer to the longitudinal axis X than the second circuit c2 of the primary vein vl.

[0093] According to the invention, the turbomachine 4 further comprises at least one first heat exchanger 27. The first heat exchanger 27 is located in the primary flow path vl and arranged axially between the low-pressure turbine 11 and the high-pressure turbine 15. The first heat exchanger 27 comprises a first circuit for passing air from the high-pressure compressor 15 and supplying the combustion chamber 25 and a second circuit for passing gas from the combustion chamber 25. The first circuit is connected to the combustion chamber 25 and to the axial compressor 17a.

[0094] Thus, the heat exchanger 27 makes it possible to heat the compressed primary air flow F1 by heat exchange thanks to the passage of the flow of gases resulting from the combustion in the second circuit of the heat exchanger 27. The first circuit makes it possible to convey the heated primary air flow F1 to the combustion chamber 25 in order to limit the quantity of combustible gas in the combustion chamber 25 and thus to reduce the need for heating the combustible gas, in particular di-hydrogen. The first heat exchanger 27 is of the air / air or gas / gas type.

[0095] Thus, according to the invention, the primary air flow F1 at the outlet of the high-pressure compressor 15 is heated in the first heat exchanger 27 before being transferred into the combustion chamber 25.

[0096] The secondary air flow F2 flows into the secondary vein v2.

[0097] The gases G mix with the secondary flow F2 in the secondary vein v2. This Mixture Me of the secondary air flow F2 and the gases G allows on the one hand to reduce the noise in the turbomachine 4 and on the other hand to cool the gases G to limit the formation of condensation trails. The mixture Me finally escapes through the turbomachine 4 to generate propulsion.

[0098] According to the invention, the combustion chamber 25 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 11, 16.

[0099] In the event of a turbine disk 11, 16 bursting, the risks of damaging the supply system 26 are limited and the risk of fuel gas leaking is therefore also limited.

[0100] 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.

[0101] According to an embodiment illustrated in [Fig. 3], the turbomachine 4 further comprises a second heat exchanger 28 located at the front of the low-pressure turbine 11. The second heat exchanger 28 is preferably located in the primary stream vl, in particular in the second circuit c2 of the primary vein vl. The low pressure turbine 11 is thus located axially between the first and second heat exchangers 27, 28. With reference to [Fig.4], the second heat exchanger 28 is for example fluidically connected to the first heat exchanger 27 and to the source 26b of combustible gas. It thus makes it possible to heat the combustible gas at the outlet of the source 26b by heat exchange, thereby reducing the need to heat the combustible gas before its introduction into the combustion chamber 25.

[0102] The second heat exchanger 28 comprises a passage for the combustible gas in the liquid state and a passage for a second fluid. The second heat exchanger 28 may be of the two-phase type. According to this example, the second fluid is, for example, the gases resulting from combustion passing through the first heat exchanger 27. According to another example, the second fluid is an inert fluid such as nitrogen.

[0103] According to another advantageous embodiment of the invention illustrated in [Fig.5], the turbomachine 4 may further comprise a second mechanical speed reducer 29 connecting the rotor of the high pressure turbine 16 to the rotor of the high pressure compressor 18.

[0104] The second speed reducer 29 makes it possible to reduce the speed of the rotor 16a of the high pressure turbine 16 while preserving the speed of the rotor 17a, 18a of the high pressure compressor 15.

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) comprising: - a low pressure body (8) comprising a rotor (10a) of a low pressure compressor (10) mechanically connected to a rotor (11a) of a low pressure turbine (11), - a high pressure body (9) comprising a rotor (17a, 18a) of a high pressure compressor (15) mechanically connected to a rotor (16a) of a high pressure turbine (16), - an annular combustion chamber (25), and - a primary vein (vl) for the 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 (10a, 17a, 18a) of the low pressure and high pressure compressors (10, 15), the combustion chamber (25), and the rotors (11a, 16a) of the high pressure and low pressure turbines (16, 11), the turbomachine (4) being characterized in that: - the high pressure turbine (16) is located at the rear of the low pressure body (8), and, - the combustion chamber (25) is located at the rear of the low pressure and high pressure bodies (8, 9), and in that it further comprises: - a first heat exchanger (27) located in the primary vein (vl) between the low pressure turbine (11) and the high pressure turbine (16), the first heat exchanger (27) comprising a first circuit for passing air from the high pressure compressor (15) and supplying the combustion chamber (25), and a second circuit for passing gas from the combustion chamber (25).

2. Turbomachine according to the preceding claim, characterized in that the primary vein (vl) comprises: - a first circuit (cl) for flowing a cold primary air flow from front to rear in the low pressure and high pressure compressors (10, 15) to the combustion chamber (25), - a second circuit (c2) for the flow of gases from the chamber combustion (25) from back to front in the high pressure turbine (16), the first heat exchanger (27) and the low pressure turbine (H).

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

4. Turbomachine according to the preceding claim, characterized in that the centrifugal compressor (18) comprises an air inlet connected to the axial compressor (17) and an air outlet connected to the first circuit of the first heat exchanger (27).

5. Turbomachine according to one of claims 3 or 4, characterized in that the centrifugal compressor (18) comprises a rotor (18a) mechanically connected to the rotor (16a) of the high pressure turbine (16) and in that the axial compressor (17) comprises a rotor (17a) mechanically connected to the rotor (10a) of the low pressure compressor (10).

6. Turbomachine according to any one of claims 3 to 5, characterized in that the axial compressor (17) is located between the low pressure compressor (10) and the low pressure turbine (11) and in that the centrifugal compressor (18) is located in front of the high pressure turbine (16).

7. Turbomachine according to any one of the preceding claims, characterized in that it comprises a second heat exchanger (28) located in front of the low pressure turbine (11).

8. Turbomachine according to the preceding claim, characterized in that it comprises a fan (7) mechanically connected to the rotor (10a) of the low pressure compressor (10), the second heat exchanger (28) being located between the fan (7) and the low pressure turbine (H).

9. Turbomachine according to the preceding claim, characterized in that the fan (7) is connected to the rotor (10a) of the low pressure compressor (10) via a first mechanical speed reducer (14).

10. Turbomachine according to any one of the preceding claims, characterized in that the rotor (16a) of the high pressure turbine (16) is directly connected to the rotor (17a, 18a) of the high pressure compressor (15).

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