HIGH-PRESSURE CRANKCASE FOR AN AIRCRAFT TURBOMACHINE

Abstract

The invention relates to a high-pressure housing (15) for an aircraft turbomachine (4) (1), the high-pressure housing (15) extending about a longitudinal axis (X) and characterized in that it comprises: - a first ferrule (150) centered on the longitudinal axis (X) and extending axially between a first front annular end and a first rear annular end, the first ferrule (150) comprising at least one first conduit (154) extending longitudinally between the first front and rear ends, - a second ferrule (152) centered on the longitudinal axis (X) and extending axially between a second front annular end located inside the first front end and a second rear annular end located around the first rear end, the second ferrule (152) comprising at least one second conduit (156) extending longitudinally between the second front and rear ends. (Shorthand figure: Figure 4)

Description

Title of the invention: HIGH-PRESSURE CRANKCASE FOR AN AIRCRAFT TURBOMACHINE Technical field of the invention

[0001] The invention relates to the field of housings for 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 flow and a secondary flow. The primary flow passes through a primary channel of the turbomachine delimited by an internal casing, while the secondary flow is directed towards a secondary channel surrounding the primary channel.

[0006] The primary flow is compressed within the compressors. Each compressor comprises at least one rotating disc about its longitudinal axis. Each compressor further comprises vanes regularly distributed around each disc, which compress the primary flow.

[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 at least one disk centered on its longitudinal axis and blades regularly distributed around each 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 more virtuous energy, in particular a combustible gas such as dihydrogen (H2). However, the current configuration of turbomachines does not 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 line or equipment of the fuel system. Furthermore, the hydrogen-air mixture with a hydrogen volume fraction 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, the hydrogen-air mixture with a hydrogen volume fraction 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] In this context, there is a need to provide a solution enabling the implementation of a combustible gas, in particular dihydrogen, in an aircraft turbomachine, limiting the risk of ignition and exposure in the turbomachine and without penalizing the mass of the turbomachine. Summary of the invention

[0013] To this end, the invention proposes a high-pressure housing for an aircraft turbomachine, the high-pressure housing extending around a longitudinal axis.

[0014] The high-pressure housing is remarkable in that it comprises:

[0015] - a first ferrule centered on the longitudinal axis and extending axially between a first front annular end and a first rear annular end, the first ferrule comprising at least one first conduit extending longitudinally between the first front and rear ends,

[0016] - a second ferrule centered on the longitudinal axis and extending axially between a a second front annular end situated inside the first front end and a second rear annular end situated around the first rear end, the second ferrule comprising at least one second conduit extending longitudinally finally between the second front and rear extremities.

[0017] The high-pressure housing according to the invention thus comprises a first conduit and a second conduit. The first conduit allows, for example, the circulation of a primary flow from front to back, while the second conduit allows, for example, the circulation of gas from back to front.

[0018] The first and second conduits thus allow the circulation of two flows which can cross in the high-pressure housing.

[0019] Thanks to such a casing, it is possible to implement a combustion chamber at the rear of the turbomachine, and therefore outside the bursting zone of the discs.

[0020] Typically, it is possible to consider a turbomachine comprising from front to back along the longitudinal axis, and in the direction of flow of the secondary airflow, a low pressure compressor, a low pressure turbine, a high pressure turbine, a high pressure casing, a high pressure compressor and a combustion chamber.

[0021] Moving the combustion chamber to the rear of the turbomachine requires a new flow circulation in the turbomachine.

[0022] The airflow entering the turbomachine splits into a primary flow and a secondary flow. The primary airflow circulates through the low-pressure compressor and then through the high-pressure crankcase, where it is compressed in the high-pressure compressor. The compressed primary airflow is mixed with the fuel gas in the combustion chamber.

[0023] The gases from the combustion flow from back to front in the high-pressure crankcase and then in the high-pressure and low-pressure turbines.

[0024] The primary airflow and the gases must therefore cross in the turbomachine. Such a crossing is made possible by the high-pressure casing according to the invention. The primary airflow circulates in the first duct while the gases circulate in the second duct.

[0025] Thus, thanks to this particular configuration of the turbomachine, the combustion chamber 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 fuel gas supply system to the fuel 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.

[0026] The invention may comprise one or more of the following features, taken individually or in combination with each other:

[0027] - the first ferrule comprises a plurality of first regularly distributed around the longitudinal axis, and the second ferrule comprises a plurality of second conduits regularly distributed around the longitudinal axis, the first and second set of conduits being entangled,

[0028] - the first ferrule includes first openings between each first conduit, and the second ferrule includes second openings between each second conduit, the first conduits passing through the second openings and the second conduits passing through the first openings,

[0029] - the first front annular end has an internal diameter greater than internal diameter of the first rear annular end,

[0030] - the second front annular end has an internal diameter smaller than internal diameter of the second rear annular end,

[0031] - the first and second conduits each comprise internal and external walls and radial walls connecting the inner and outer walls,

[0032] - the first and second conduits each comprise a front section and a section rear connected by a central section, the central section extending over an angular sector smaller than the angular sectors of the downstream and rear sections.

[0033] The invention also relates to a turbomachine for an aircraft, the turbomachine comprising a high-pressure casing according to any one of the preceding characteristics.

[0034] The turbomachine may comprise one or more of the following features, taken individually or in combination with each other:

[0035] - the turbomachine extends around and along the longitudinal axis, and comprises in besides from front to back:

[0036] - a low-pressure body comprising a low-pressure compressor and a turbine low pressure,

[0037] - a high-pressure body comprising a high-pressure turbine and a compressor high pressure, and

[0038] - a combustion chamber, the high-pressure crankcase being arranged axially between the high-pressure turbine and the high-pressure compressor,

[0039] - an annular primary vein and an annular secondary vein arranged around the primary vein, the primary vein presenting:

[0040] - a first flow circuit of a primary airflow from front to rear in the low-pressure compressor, the high-pressure crankcase, and from the high-pressure compressor to the combustion chamber,

[0041] - a second gas flow circuit from the combustion chamber of from back to front in the high-pressure casing, the high-pressure turbine and the low-pressure turbine, the primary airflow flowing into the first duct and the gases flowing into the second duct. Brief description of the figures

[0042] 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:

[0043] [Fig. 1] is a schematic representation of an aircraft according to the invention,

[0044] [Fig. 2] is a schematic longitudinal sectional representation of a half- turbomachine according to the invention,

[0045] [Fig. 3] is a perspective representation of a low-pressure casing equipping the turbomachine according to the invention,

[0046] [Fig. 3a] is a front view of the low-pressure casing equipping the turbomachine according to the invention,

[0047] [Fig. 3b] is a rear view of the low-pressure casing equipping the turbomachine according to the invention,

[0048] Figure 4 is a perspective representation of a high-pressure housing according to the invention.

[0049] [Fig.4a] is a perspective representation of a first ferrule of the high-pressure housing of [Fig.4],

[0050] [Fig.4b] is a perspective representation of a second ferrule of the high-pressure housing of [Fig.4],

[0051] [Fig. 5] is a schematic longitudinal cross-sectional representation of a combustion chamber equipping the turbomachine according to the invention,

[0052] [Fig. 6] is a schematic longitudinal cross-sectional representation of half an aircraft turbomachine according to another embodiment,

[0053] [Fig. 7] is a schematic longitudinal cross-sectional representation of the aircraft turbomachine half of [Fig. 2], in which the flow circulation is illustrated. Detailed description of the invention

[0054] 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.

[0055] With reference to [Fig.2], each turbomachine 4 extends around and along a longitudinal axis X.

[0056] 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 on the [Fig.2].

[0057] The terms "axial", "axially", "radial", "radially", are defined with respect to the longitudinal axis X of the turbomachine 4.

[0058] The terms “internal”, “interior”, “external”, “outer”, “externally”, are defined with respect to the distance from the longitudinal axis X along a radial axis perpendicular to the longitudinal axis X.

[0059] The turbomachine 4 is, for example, a twin-spool, twin-flow turbojet engine. It comprises an engine M including, from front to rear, optionally a fan 5, a low-pressure casing 6 and a high-pressure casing 7. According to the invention, the low-pressure and high-pressure casings 6, 7 are mechanically independent.

[0060] The low-pressure body 6 comprises, from front to back, a low-pressure compressor 8 and a low-pressure turbine 9. The low-pressure body 6 further comprises a low-pressure shaft 10 connecting the low-pressure compressor 8 and the low-pressure turbine 9. Advantageously, the low-pressure shaft 10 is connected to the blower 5, for example, via a speed reducer (not shown).

[0061] The high-pressure body 7 comprises, from front to back, a high-pressure turbine 11 and a high-pressure compressor 12. The high-pressure body 6 further comprises a high-pressure shaft 13 connecting the high-pressure turbine 11 to the high-pressure compressor 12.

[0062] The low-pressure and high-pressure shafts 10, 13 are centered on the longitudinal axis X. They are axially spaced apart. The low-pressure and high-pressure shafts 10, 13 are thus separated by an axial distance d.

[0063] Each compressor 8, 12 and each turbine 9, 11 comprises at least one disk centered on the longitudinal axis X and rotor blades regularly distributed around the disk (not shown). The disk is mounted to rotate about the longitudinal axis and, together with the rotor blades, forms a rotor. Each compressor and each turbine also comprises 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 a stage. The compressor and / or the turbine may comprise one or more stages.

[0064] According to the invention, the turbomachine 4 optionally includes a low-pressure housing 14 and includes a high-pressure housing 15. The low-pressure housing 14 and the high-pressure housing 15 are centered on the longitudinal axis X.

[0065] The low-pressure housing 14 is located axially between the low-pressure compressor 8 and the low-pressure turbine 9. With reference to Figures 3, 3a, 3b, the low-pressure housing 14 comprises an outer shell 140, an inner shell 142 and first and second intermediate shells 144, 146 arranged radially between the outer and inner shells 140, 142. The second intermediate shell 146 is arranged around the first intermediate shell 144.

[0066] The outer, inner, and intermediate ferrules 140, 142, 144, 146 are annular. They are coaxial and centered on the longitudinal axis X. The outer ferrule 140 and the second intermediate ferrule 146 delimit between themselves a portion of a vein se secondary v2 of flow of a secondary air flow F2. They are connected by first radial arms 148 extending radially between the outer ferrule 140 and the second intermediate ferrule 146.

[0067] The first arms 148 are regularly distributed around the longitudinal axis X. Each first arm 148 comprises a hollow fairing 148a and first fins 148b located within the fairing 148a. The fairing 148a has a general U-shape open at the rear. More precisely, the fairing 148a comprises radial first and second walls 148c, 148d connected at their leading edges by a nose 148e. The nose 148e has an aerodynamic surface that minimizes disturbances to the secondary airflow F2. The rear edges of the first and second walls 148c, 148d are free. The first fins 148b are radially overlapped within the fairing 148a.

[0068] The first and second intermediate ferrules 144, 146 define between them a portion of a primary flow channel vl of a primary airflow FL. The low-pressure housing 14 further comprises second arms 149a extending radially between the first and second intermediate ferrules 144, 146. The second arms 149a are radially aligned with the first arms 148. The second arms 149a have an aerodynamic surface to minimize disturbances in the flow of the primary airflow FL.

[0069] The low-pressure housing 14 further comprises third arms 149b extending radially between the inner ferrule 142 and the first intermediate ferrule 144. The third arms 149b are circumferentially offset with respect to the first and second arms 148, 149a. Advantageously, each third arm 149b is arranged circumferentially between two second arms 149a.

[0070] The low-pressure housing 14 further comprises second fins 149c arranged circumferentially between the third arms 149b. The second fins 149c are radially superimposed between the inner ferrule 142 and the first intermediate ferrule 144.

[0071] The high-pressure casing 15 is located axially between the high-pressure turbine 11 and the high-pressure compressor 12. With reference to [Fig.4], the high-pressure casing 15 according to the invention comprises a first annular ferrule 150 centered on the longitudinal axis X and a second annular ferrule 152 centered on the longitudinal axis X.

[0072] The first ferrule 150 extends axially between a first front annular end 150a and a first rear annular end 150b. The first front annular end 150a advantageously has an internal diameter greater than the internal diameter of the first rear annular end 150b. Thus, the first ferrule 150 has a diameter that decreases towards the rear. The first ferrule 150 has, for example, a substantially frustoconical shape.

[0073] As more clearly seen in [Fig.4a], the first ferrule 150 comprises a first conduit 154 and advantageously a plurality of first conduits 154 regularly distributed around the longitudinal axis X. Each first conduit 154 extends longitudinally between the first front and rear ends 150a, 150b.

[0074] Each first conduit 154 comprises an inner wall 154a, an outer wall 154b, and a passage 154c defined between the inner and outer walls 154a, 154b. Each first conduit 154 further comprises side walls 1541 connecting the inner and outer walls 156a, 156b. The side walls 1541 thus separate the first conduits 154 from each other. Each first conduit 154 extends longitudinally between a front section 154d and a rear section 154e connected by a central section 154f. The front sections 154d advantageously each extend over an angular sector greater than the angular sector of each rear section 154e. Each central section 154f extends over an angular sector less than the angular sectors of the front and rear sections 154d, 154e. The central sections 154f flare out towards the front and towards the rear. Thus, each first conduit 154 has essentially an hourglass shape.

[0075] Advantageously, the first ferrule 150 further comprises first openings 155 between each first conduit 154. The first openings 155 are through-holes. They thus open both inside and outside the first ferrule 150. The first openings 155 are regularly distributed around the longitudinal axis X.

[0076] The second ferrule 152 extends axially between a second front annular end 152a and a second rear annular end 152b. The second front annular end 152a advantageously has a smaller internal diameter than the second rear annular end 152b. Thus, the second ferrule 150 has a progressive diameter, that is, one that increases towards the rear. The second ferrule 150, for example, has a substantially frustoconical shape.

[0077] Preferably, the first front annular end 150a of the first ferrule 150 is arranged around the second front annular end 152a of the second ferrule 152 and the second rear annular end 152b of the second ferrule 152 is arranged around the first rear annular end 150b. The first and second ferrules 150, 152 are thus nested one inside the other.

[0078] As more clearly seen in [Fig.4b], the second ferrule 152 comprises at least one second conduit 156 and advantageously a plurality of second conduits 156 regularly distributed around the longitudinal axis X. Each second conduit 156 extends longitudinally between the second front and rear ends 152a, 152b.

[0079] Each second conduit 156 comprises an inner wall 156a, an outer wall 156b, and a passage 156c defined between the inner and outer walls 156a, 156b. Each second conduit 156 further comprises side walls 1561 connecting the walls internal and external 156a, 156b. The lateral walls 1561 thus separate the second conduits 156 from each other.

[0080] Each second conduit 156 extends longitudinally between a forward section 156d and a rear section 156e connected by a central section 156f. The forward sections 156d advantageously each extend over an angular sector smaller than the angular sector of each rear section 156e. Each central section 156f extends over an angular sector smaller than the angular sectors of the forward and rear sections 156d, 156e. The central section 156f flares out upstream and downstream. Thus, each arm 156 has a substantially hourglass shape.

[0081] The second ferrule 152 further includes second openings 157 between each second conduit 156. The second openings 157 are through-holes. They thus open both inside and outside the second ferrule 152. The second openings 157 are regularly distributed around the longitudinal axis X.

[0082] With reference to [Fig.4], the first and second conduits 154, 156 are intertwined with each other. The first conduits 154 pass through the second openings 157 respectively, and the second conduits 156 pass through the first openings 155 respectively. Thus, the front sections 156d of the second conduits 156 are arranged radially inside the front sections 154d of the first conduits 154, and the rear sections 156e of the second conduits 156 are arranged radially around the rear sections 154e of the second conduits 154. In other words, along a median plane perpendicular to the longitudinal axis X, the high-pressure housing 15 comprises an alternation of central sections 154f of the first conduits 154 and central sections 156f of the second conduits 156 around the longitudinal axis X. Similarly, a first conduit 154 and an adjacent second conduit 156 substantially form an X in a radial plane passing through the longitudinal axis X.

[0083] The pressure shaft 10 is guided in rotation around the longitudinal axis X by a first front bearing 16 and optionally a second front bearing 17. The first and second front bearings 16, 17 are arranged radially between the low pressure housing 14 and the low pressure shaft 10.

[0084] The high-pressure shaft 13 is guided in rotation around the longitudinal axis X by a first rear bearing 18 and optionally a second rear bearing 19. The first and second rear bearings 18, 19 are arranged radially between the high-pressure housing 15 and the high-pressure shaft 13.

[0085] Each front bearing 16, 17 comprises a bearing such as rollers or balls arranged between an outer ring and an inner ring. The inner ring is connected to the low-pressure shaft 10 and the outer ring is connected to the low-pressure housing 14 via a bearing support 16a, 17a.

[0086] Each rear bearing 18, 19 comprises a bearing such as rollers or balls arranged between an outer ring and an inner ring. The inner ring is connected to the high-pressure shaft 13 and the outer ring is connected to the high-pressure housing 15 via a bearing support 18a, 19a.

[0087] Advantageously, the turbomachine 4 is connected to the wing 3 by a mast 20. The mast 20 includes a frame 20a connecting the low-pressure housing 14 to the high-pressure housing 15. In particular, the frame 20a is fixed to the outer ferrule 140 of the low-pressure housing 14 and to the first and / or second ferrules 150, 152 of the high-pressure housing 15. The mast 20 is also connected to a nacelle 20' surrounding the engine M of the turbomachine 4. The nacelle 20' is centered on the longitudinal axis X.

[0088] The turbomachine 4 further comprises an exhaust nozzle E. The nozzle E is annular and centered on the longitudinal axis X. It is mounted around the combustion chamber 23. It is axially connected to the nacelle 20'.

[0089] The turbomachine 4 further comprises a fan housing 21. The fan housing 21 is centered on the longitudinal axis X and surrounds the fan 5. It is arranged inside the nacelle 20' and connected to the latter. It is also axially connected to the outer shell 140 of the low-pressure housing 14.

[0090] The turbomachine 4 further comprises an inter-flow casing 22 centered on the longitudinal axis X. The inter-flow casing 22 is arranged inside the fan casing 21. It is axially connected to the intermediate ferrule 144 of the low-pressure casing 140. The inter-flow casing 22 has a cylindrical portion 22a and a frustoconical portion 22b. In the present example, the frustoconical portion 22b is arranged behind the cylindrical portion 22a. The nozzle E is mounted around the cylindrical portion 22a. The frustoconical portion 22b is connected to the mast 20.

[0091] According to the invention, the turbomachine 4 further comprises a combustion chamber 23 and a fuel gas supply system 24 for the combustion chamber 23.

[0092] With reference to [Fig. 5], the combustion chamber 23 comprises an outer wall 230 and an inner wall 232. The outer and inner walls 230, 232 are annular and centered on the longitudinal axis X. The outer and inner walls 230, 232 define between each other an annular enclosure 234. They are connected at their rear ends by a chamber bottom 236. The chamber bottom 236 is thus located at a rear end of the combustion chamber 22. The chamber bottom 236 has a fuel gas injection port 238.

[0093] The combustion chamber 23 further includes an injector 239 for injecting combustible gas into the enclosure 234. The injector 239 passes through the injection port 238 and opens into the annular enclosure 234.

[0094] According to the invention, the combustion chamber 23 is arranged at the rear of the high-pressure turbine 11, in particular at the rear of the last rotor of the high-pressure turbine 11. According to a particularly advantageous embodiment, the combustion chamber 23 is arranged at the rear of the high-pressure compressor 12, in particular at the rear of the last rotor of the high-pressure compressor 12. It is, for example, connected to the casing of the high-pressure compressor 12.

[0095] According to a first embodiment illustrated in [Fig.6], the annular enclosure 234 is located radially outside the high-pressure compressor 12 and the chamber bottom 236 is located at the rear of the high-pressure compressor 12.

[0096] According to a second embodiment illustrated in figures 2 and 7, the combustion chamber 23 is located at the rear of the high-pressure compressor 12. According to this embodiment, the annular enclosure 234 is located at the rear of the high-pressure compressor 12, in particular at the rear of the last rotor of the high-pressure compressor 12.

[0097] The feed system 24 is located at the rear of the high-pressure turbine 11. Preferably, the feed system 24 is located at the rear of the high-pressure compressor 12.

[0098] The supply system 24 includes a conduit 240 for the passage of combustible gas and optionally a source 242 of combustible gas.

[0099] The pipe 240 extends radially between the combustion chamber 23 and the source 242. The pipe 240 has a fuel gas inlet section 240a connected to the source 242 and a fuel gas outlet section 240b connected to the combustion chamber 23, specifically to the injector 239. The pipe 240 can be located in a compartment extending from the combustion chamber 23 to the source 242. Such a configuration allows the fuel supply system 23 to be separated from the other elements of the turbomachine 4.

[0100] The source 242 of combustible gas is located, for example, in the wing 3 of the aircraft 1. The combustible gas is, for example, natural gas, methane (CH4), or ammonia (NH3). Advantageously, the combustible gas is dihydrogen (H2).

[0101] The primary and secondary veins vl, v2 are annular. The secondary vein v2 surrounds the primary vein vl. A portion of the secondary vein v2 is delimited by the outer ferrule 140 around the second intermediate ferrule 146 of the low-pressure housing 14 and a portion of the primary vein vl is delimited by the second intermediate ferrule 146 around the first intermediate ferrule 144.

[0102] In operation, the blower 5 allows the aspiration of a main airflow F which divides into a primary airflow Fl and a secondary airflow F2.

[0103] With reference to [Fig. 6] and [Fig. 7], the primary vein vl comprises a first circuit cl configured to conduct the primary airflow Fl from front to back. The primary airflow Fl passes successively through the low-pressure compressor 8, the low-pressure crankcase pressure 14, the high-pressure casing 15 and the high-pressure compressor 12.

[0104] The primary airflow Fl flows into the low pressure compressor 8 and then into the low pressure casing 14, between the first and second intermediate ferrules 144, 146. The primary airflow Fl then flows into the high pressure casing 15, into the first conduits 154, and then into the high pressure compressor 12.

[0105] The compressed primary airflow Fl is then directed into the combustion chamber 23. The compressed primary airflow Fl is mixed with the fuel gas in the combustion chamber 23. The combustion reaction in the combustion chamber 23 forms gas G.

[0106] The primary vein vl further comprises a second circuit c2 configured to conduct the flow of gas G from combustion from back to front. The gas G passes successively through the high-pressure casing 15, the high-pressure turbine 11, the low-pressure turbine 9 and the low-pressure casing 14.

[0107] The gases G flow into the high-pressure casing 15, into the second conduits 156, then into the high-pressure turbine 11 and the low-pressure turbine 9. The gases G then flow into the low-pressure casing 14. In the low-pressure casing 14, the gases G are directed by the second and third arms 149a, 149b towards the first vanes 148a. The vanes 148a then redirect the gases G into the secondary flow v2.

[0108] Thanks to the first hollow arms 148 open towards the rear, the gases G flowing from rear to front can undergo a 180° turn by the fins 148a and flow into the secondary vein v2.

[0109] Furthermore, it follows that the primary airflow Fl and the gases G flow in two different directions. The gases G and the primary flow Fl cross in each of the low-pressure and high-pressure housings 14, 15.

[0110] The secondary airflow F2 flows into the secondary vein v2.

[0111] The G gases mix with the secondary flow F2 in the secondary vein v2. This The mixture Me of secondary airflow F2 and gases G reduces noise in turbomachine 4 and cools the gases G to limit the formation of condensation trails. The mixture Me then exits through nozzle E to generate propulsion.

[0112] According to the invention, the low-pressure 6 and high-pressure 7 modules are arranged in series, from front to back, and the combustion chamber 23 is located behind the high-pressure compressor 12. Thus, the supply system 24 of the combustion chamber 23 is located behind the high-pressure compressor 12 and therefore outside the burst zone Z2 of the compressor disc 8, 12. In the event of a burst of a compressor disc 8, 12, the risk of damaging the supply system 24 is limited and the risk of fuel gas leakage is consequently also limited.

[0113] Thanks to this embodiment of the turbomachine 4, the risks of explosion in the turbomachine 4 and / or the aircraft 1 are therefore greatly reduced.

[0114] The high-pressure housing 15 according to the invention allows the circulation and crossing of two flows resulting from the configuration of the turbomachine 4 according to the invention.

[0115] Thanks to such a high-pressure housing 15, it is possible to implement the combustion chamber 23 at the rear of the turbomachine 4, and therefore outside the bursting zone of the discs.

[0116] Furthermore, the modular configuration of the low pressure and high pressure bodies 6, 7 makes it possible to reduce the axial length of the low pressure and high pressure shafts 10, 13. The total mass of the turbomachine 4 is therefore lower and the mechanical integration of the low pressure and high pressure shafts 10, 13 is simplified.

Claims

Claims

1. High pressure casing (15) for a turbomachine (4) of an aircraft (1), the high pressure casing (15) extending around a longitudinal axis (X) and being characterized in that it comprises: - a first shroud (150) centered on the longitudinal axis (X) and extending axially between a first front annular end (150a) and a first rear annular end (150b), the first shroud (150) comprising a plurality of first ducts (154) extending longitudinally between the first front and rear ends (150a, 150b) and regularly distributed around the longitudinal axis (X), - a second shroud (152) centered on the longitudinal axis (X) and extending axially between a second front annular end (152a) located inside the first front annular end (150a) and a second rear annular end (152b) located around the first rear annular end (150b),the second ferrule (152) comprising a plurality of second conduits (156) extending longitudinally between the second front and rear ends (152a, 152b) and regularly distributed around the longitudinal axis (X), the first and second conduits (154, 156) being intertwined, the first ferrule (150) further comprising first lumens (155) between each first conduit (154), and the second ferrule (152) further comprising second lumens (157) between each second conduit (156), the first conduits (154) passing through the second lumens (157) and the second conduits (156) passing through the first lumens (155).,

2. High pressure casing (15) according to the preceding claim, characterized in that the first front annular end (150a) has an internal diameter greater than the internal diameter of the first rear annular end (150b).

3. High pressure casing (15) according to any one of the preceding claims, characterized in that the second front annular end (152a) has an internal diameter smaller than the internal diameter of the second rear annular end (152b).

4. A high pressure housing (15) according to any preceding claim, characterized in that the first and second conduits (154, 156) each comprise inner and outer walls (154a, 156a, 154b, 156b) and radial walls (1541, 1561) connecting the inner and outer walls (154a, 156a, 154b, 156b).

5. High pressure casing (15) according to any one of the preceding claims, characterized in that the first and second conduits (154, 156) each comprise a front section (154d, 156d) and a rear section (154e, 156e) connected by a central section (154f, 156f), the central section (154f, 1561) extending over an angular sector smaller than the angular sectors of the downstream and rear sections (154d, 156d, 154e, 156e).

6. Turbomachine (4) for an aircraft (1), characterized in that it comprises a high pressure casing (15) according to any one of the preceding claims.

7. Turbomachine according to the preceding claim, the turbomachine (4) extending around and along the longitudinal axis (X), characterized in that it further comprises from front to rear: - a low pressure body (6) comprising a low pressure compressor (8) and a low pressure turbine (9), - a high pressure body (7) comprising a high pressure turbine (11) and a high pressure compressor (12), and - a combustion chamber (23), the high pressure casing (15) being arranged axially between the high pressure turbine (11) and the high pressure compressor (12).

8. Turbomachine according to the preceding claim, characterized in that it comprises an annular primary vein (vl) and an annular secondary vein (v2) arranged around the primary vein (vl), the primary vein (vl) having: - a first circuit (cl) for the flow of a primary air flow (Fl) from front to rear in the low pressure compressor (8), the high pressure casing (15), and the high pressure compressor (12) to the combustion chamber (23), - a second circuit (c2) for the flow of gases (G) from the combustion chamber (23) from rear to front in the high pressure casing (15), the high pressure turbine (11) and the low pressure turbine (9), the primary air flow (Fl) flowing in the first duct (154) and the gases (G) flowing in the second duct (156).

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