EP4103827A1

EP4103827A1 - Turbofan engine

Info

Publication number: EP4103827A1
Application number: EP21703040.2A
Authority: EP
Inventors: Claude JOUVE
Original assignee: Individual
Current assignee: Individual
Priority date: 2020-02-13
Filing date: 2021-02-08
Publication date: 2022-12-21
Also published as: WO2021160568A1; FR3107316A1; US20230108704A1; FR3107316B1; CA3164374A1

Abstract

The present invention relates to a turbofan engine (10), comprising a ducted fan (22) and an engine (16) having a nozzle (24) for expelling burnt gases (N1) by the engine in the upstream direction of the turbofan engine, the free end (26) of the nozzle being located upstream of the fan (22).

Description

TURBOJET

Technical area

The present description relates generally to turbojets.

State of the art

The turbojet is a propulsion system which transforms the potential of chemical energy contained in a fuel, associated with an oxidizer corresponding to the ambient air, into kinetic energy making it possible to generate a reaction force, the thrust, in the opposite direction to ejection. The thrust generated results from the acceleration of a certain quantity of air between the inlet (air inlet nozzle) and the outlet (ejection nozzle).

A turbojet comprises at least one compressor mechanically linked by a shaft to a turbine. A combustion chamber is provided between the compressor and the turbine. The compressor can include several compressor stages. Likewise, the turbine can include several turbine stages. When all the stages of the compressor rotate at the same speed as the turbine, the turbojet is said to be single-body. In a double-body turbojet, the turbojet comprises a first compressor, called a high-pressure compressor, driven by a first turbine, called a high-pressure turbine, and a second compressor, called a low-pressure compressor, driven by a second turbine, called a low-pressure turbine. , the rotational speeds of the first and second compressors being different.

The turbojet is said to be single-flow when all of the air entering the turbojet enters the combustion chamber. The turbojet is said to be a double-flow when the air flow entering the turbojet is divided into two flows, the primary flow and the secondary flow. The primary flow, or hot flow, passes through the entire reactor passing through the compressor (s), the combustion chamber, and the turbine (s). The secondary flow or cold flow bypasses the entire hot part of the reactor. On some turbojets, the secondary flow also drives the low pressure compressor. Other turbojets include a fan with a diameter significantly greater than the low pressure compressor and located at the front thereof. The blower is driven by the same shaft as the low pressure compressor. It makes it possible to obtain maximum thrust from the secondary flow.

The ratio of the secondary flow air flow to the primary flow air flow is called the dilution ratio or dilution ratio. In a single-flow turbojet, the air circulating in the reactor is accelerated very strongly, resulting in a high ejection speed, creating strong turbulence by mixing with the ambient air, resulting in significant noise. On the other hand, in a double-flow turbojet engine, the large quantity of air passing through the secondary flow is weakly accelerated and “sheath” the strongly accelerated primary flow, hence reducing noise.

However, for certain applications, it would be desirable to further reduce the noise emitted by the downstream turbojet engine.

Disclosure of the invention

Thus, an object of one embodiment is to at least partially overcome the drawbacks of the turbojets described above.

An object of an embodiment is for the turbojet engine to have reduced noise emission.

For this purpose, one embodiment provides a bypass turbojet comprising a ducted fan and a reactor comprising a nozzle for expelling gases burnt by the reactor upstream of the turbojet, the free end of the nozzle being located upstream. of the blower.

According to one embodiment, the turbojet engine is intended to receive an air flow dividing into a primary flow and a secondary flow, the primary flow supplying the reactor and forming the burnt gases ejected upstream of the turbojet engine, the secondary flow being ejected downstream from the turbojet.

According to one embodiment, the turbojet comprises a first fairing having a first air inlet at least for the secondary flow, the fan being contained in the first fairing, the turbojet comprising a second air inlet for the primary flow.

According to one embodiment, the second air inlet is located upstream of the first air inlet.

According to one embodiment, the rate of dilution of the turbojet is between 8 and 15.

According to one embodiment, the nozzle is divergent.

According to one embodiment, the turbojet has a double body and comprises a low pressure hitch having at least one low pressure turbine stage and a high pressure hitch, in which part of the fan forms the rotor of the low turbine stage. pressure. According to one embodiment, the turbojet comprises first blades, each first blade comprising a first part and a second part, the first parts of the first blades forming the rotor of the low pressure turbine stage and the second parts of the first blades forming the blower. According to one embodiment, the low pressure coupling comprises at least one low pressure compressor stage and another part of the fan forms the rotor of the low pressure compressor stage.

According to one embodiment, each first blade further comprises a third part, the third parts of the first blades forming the rotor of the low pressure compressor stage.

According to one embodiment, the low pressure compressor stage is crossed by the primary flow while the primary flow flows from upstream to downstream of the turbojet engine and the low pressure turbine stage is crossed by the gases. burnt while the burnt gases flow from downstream to upstream of the turbojet. According to one embodiment, the low pressure hitch comprises at least first and second low pressure turbine stages and first and second low pressure compressor stages. A part of the fan forms the rotor of the first stage of the low pressure turbine, the turbojet comprising second blades, each second blade comprising a first part and a second part, the first parts of the second blades forming the rotor of the second low turbine stage pressure and the second parts of the second blades forming the rotor of the second stage of the low pressure compressor.

According to one embodiment, the turbojet further comprises movable flaps and a mechanism for actuating the movable flaps between a first position in which the movable flaps allow the expulsion of the burnt gases upstream of the turbojet and a second position in which the movable flaps release openings for expelling the burnt gases downstream of the turbojet.

Brief description of the drawings

These characteristics and advantages, as well as others, will be explained in detail in the following description of particular embodiments given without limitation in relation to the appended figures, among which: FIG. 1 represents an embodiment of a turbojet engine ; FIG. 2 represents another embodiment of a turbojet; FIG. 3 represents an embodiment of part of the reactor of the turbojet engine represented in FIG. 1; FIG. 4 is a perspective view, partial and schematic, of an embodiment of a blade; FIG. 5 represents another embodiment of a part of the reactor of the turbojet engine represented in FIG. 1; FIG. 6 represents another embodiment of a part of the reactor of the turbojet engine represented in FIG. 1; FIG. 7 represents a more detailed embodiment of the turbojet engine represented in FIG. 1; FIG. 8 represents a more detailed embodiment of the turbojet engine represented in FIG. 2; FIG. 9 is a perspective view, partial and schematic, of another embodiment of a blade; FIG. 10 represents a variant of the embodiment of the turbojet engine represented in FIG. 7 in a first mode of operation; FIG. 11 represents a variant of the embodiment of the turbojet engine represented in FIG. 7 in a second mode of operation; FIG. 12 represents a variant of the embodiment of the turbojet engine represented in FIG. 8 in a first mode of operation; and FIG. 13 represents a variant of the embodiment of the turbojet engine represented in FIG. 8 in a second mode of operation.

Detailed description of the embodiments

The same elements have been designated by the same references in the various figures. In particular, the structural and / or functional elements common to the different embodiments may have the same references and may have identical structural, dimensional and material properties. For the sake of clarity, only the steps and elements useful for understanding the embodiments described have been shown and are detailed. In particular, the structures of turbines, of the combustion chamber, and of turbojet compressors are well known to those skilled in the art and are not described in detail.

Unless otherwise specified, the terms “upstream” and “downstream” refer to the air flow which enters the turbojet. Unless otherwise specified, the expressions "approximately", "approximately", "substantially", and "on the order of" mean to within 10%, preferably within 5%.

FIG. 1 is a partial and schematic sectional view of an embodiment of a turbojet 10. The turbojet 10 comprises a nacelle 11 comprising an external fairing 12 and an internal fairing 14, the external fairing 12 delimiting a channel 15 extending in a direction D and containing the internal fairing 14. The channel 15 may be substantially symmetrical about revolution about the axis D. The turbojet 10 comprises a reactor 16 disposed at least in part in the channel 15 and in the internal fairing 14. Beams, not shown, connect the external fairing 12, the internal fairing 14 and the engine 16.

The turbojet 10 is a bypass turbojet. The air flow which supplies the turbojet 10 is divided into a primary flow F1 and a secondary flow F2. The reactor 16 is supplied by the primary flow F1. The external fairing 12 comprises an air inlet 18 and a nozzle 20. The primary flow Fl and the secondary flow F2 enter the turbojet 10 through the air inlet 18. The primary flow F1 which feeds the reactor 16 is rejected in the form of an output flow of the burnt gases NI. The turbojet 10 expels an air flow N2 downstream through the nozzle 20. The turbojet 10 comprises a fan 22, driven in rotation about an axis D by the reactor 16. The fan 22 drives the secondary flow F2 for obtain the expelled flow N2 with the desired thrust.

According to one embodiment, the outlet flow of the burnt gases NI is expelled upstream from the turbojet 10. The reactor 16 comprises a nozzle 24 having one end 26 through which the outlet flow of the NI flue gases is expelled, and the The end 26 of the nozzle 24 is located upstream of the fan 22. Preferably, the end 26 of the nozzle 24 is located upstream of the air inlet 18 of the outer shroud 12.

In the embodiment shown in FIG. 1, the internal fairing 14 comprises an air inlet 28 for the primary flow NI. According to one embodiment, the air inlet 28 of the internal fairing 14 is located downstream of the air inlet 18 of the external fairing 12, for example downstream of the fan 22. In the embodiment shown in Figure 1, the air inlet 28 of the inner shroud 14 is located just downstream of the blower 22. Alternatively, the air inlet 28 of the inner shroud 14 may be located further downstream of the blower. blower 22.

According to one embodiment, the expelled flow N2 is composed substantially entirely by the secondary flow F2. According to one embodiment, at least part (arrow 30) of the burnt gas outlet flow NI can be redirected, after its expulsion from the nozzle 24, into the secondary flow F2 and / or into the primary flow F1. However, the expulsion of the burnt gas outlet flow NI being carried out upstream of the turbojet 10, the part of the gas outlet flow burnt NI which is redirected in the secondary flow F2 and / or in the primary flow F1 is stirred by the fan 22 so that it is largely diluted in the secondary flow F2.

According to one embodiment, the reactor 16 is a double-body reactor. The reactor then comprises a high pressure HP coupling and a low pressure LP coupling, shown very schematically in FIG. 1. The high pressure HP coupling comprises a high pressure compressor and a high pressure turbine, not shown, and the low coupling LP pressure comprises a low pressure turbine and possibly a low pressure compressor, not shown. The reactor 16 further comprises a combustion chamber, not shown, located between the high pressure compressor and the high pressure turbine. According to one embodiment, the primary flow Fl flows from upstream to downstream when it enters the air inlet 28 and is then diverted to flow from downstream to upstream, which is illustrated by the arrows 32. According to one embodiment, the primary flow passes through the high pressure HP hitch while it flows from downstream to upstream. In addition, in the embodiment shown in FIG. 1, the primary flow also passes through the low pressure LP hitch as it flows from downstream to upstream.

Figure 2 is a sectional view, partial and schematic, of another embodiment of a turbojet 40. The turbojet 40 comprises all the elements of the turbojet 10 shown in Figure 1 with the difference that the internal fairing 14 is not present and that the reactor 16 comprises a pipe 42 comprising an air inlet 44 into which the primary flow Fl enters. In the embodiment shown in FIG. 2, the air inlet 44 is upstream of the air inlet 18 of the outer fairing 12. In the embodiment shown in FIG. 2, the air inlet 44 is upstream of the end 26 of the nozzle 24. Advantageously, the primary flow Fl then does not include any burnt gas from the NI burnt gas outlet flow. In the present embodiment, the primary flow F1 passes through the low pressure compressor as it flows from upstream to downstream in the line 42. As described above, the primary flow is then diverted and passes through the pipe. high pressure stage as it flows from downstream to upstream.

In the embodiments shown in FIGS. 1 and 2, the output flow of the burnt gases NI is expelled upstream of the turbojet 10 or 40. The noise perceived by an observer located on the side of the nozzle 20 rejecting the flow of the expelled air N2 is therefore reduced. By way of example, in the case where the turbojet 10 or 40 is fitted to a drone, for example for transporting objects, the axis D of the turbojet 10 or 40 may, in certain operating configurations, be slightly inclined with respect to in a vertical direction with the end of the nozzle 20 oriented towards the ground. The embodiments of the turbojet 10 or 40 allow then reduce the noise perceived by an observer on the ground. In the embodiments shown in Figures 1 and 2, the nozzle 24 for expelling the flow of flue gases NI is shown schematically with a constant section. However, it may be advantageous for the nozzle 24 to have a divergent section approaching the free end 26, to further minimize the residual thrust of the flue gas flow NI. According to one embodiment, the thrust of the flow of burnt gases NI is less than 15%, preferably less than 10%, more preferably less than 5%, of the thrust of the flow N2.

The dimensions of the turbojet 10 or 40 will depend on the intended applications. By way of example, in the case where the turbojet 10 or 40 is intended to equip a drone, for example for the transport of objects, the characteristics of the turbojet 10 or 40 can be the following:

- dilution rate of between 8 and 15, for example equal to approximately 12;

- Diameter of the inlet 18 of the outer fairing 12 of between 25 cm and 30 cm, for example equal to approximately 28 cm;

- Diameter of the inlet of the internal fairing 14 of between 11 cm and 13 cm, for example about 12 cm;

- length of the outer fairing 12 measured in direction D, between 50 cm and 60 cm, for example approximately 55 cm;

- distance between the end 26 of the nozzle 24 having one end 26 through which is expelled the outlet flow of the burnt gases NI and the fan 26 between 10 cm and 15 cm;

- expulsion speed of the NI burnt gas outlet flow between 2 m / s and 5 m / s.

Figure 3 is a sectional view, partial and schematic, of a more detailed embodiment of part of the reactor 16 of the turbojet 10 shown in Figure 1. In the present embodiment, the internal fairing 14 is not not visible. In this embodiment, the high pressure HP coupling comprises a turbine stage 50. The reactor 16 further comprises a frame 52 fixed to the internal fairing 14, not visible, and to the external fairing 16, and a shaft 54, mounted so as to be able to rotate relative to the frame 52 about the axis D, by means of bearings 56, a single bearing 56 being visible in FIG. 3. The turbine stage 50 comprises a rotor 58 integral with the shaft 54 and a stator 60 integral with the armature 52. The rotor 58 is driven in rotation by the flow of gas (arrow C1) expelled from the combustion chamber, not shown in FIG. 3. The shaft 54 drives at least one in rotation. high pressure compressor stage, not shown, receiving the primary flow NI.

In the present embodiment, the LP low pressure hitch does not include a compressor and comprises a low pressure turbine stage 62. This turbine stage 62 comprises a stator 64, integral with the frame 52, and a rotor 66. In the present embodiment, the rotor 66 corresponds to the central part of the fan 22. The rotor 66 is integral with a shaft 68 mounted movably in rotation relative to the frame 52 around the axis D by means of bearings 70. The rotor 66 is driven in rotation around the axis D by the flow of gas having passed through the high pressure turbine 50. The shafts 68 and 54 can rotate at different rotational speeds and / or in opposite directions.

Connecting elements are arranged between the outer fairing 12 and the internal fairing 14 and between the internal fairing 14 and the frame 52 of the reactor 16. Some or all of the connecting elements can act as connecting beams to ensure the cohesion of the turbojet 10. Some or all of the connecting elements may correspond to diffusers, also called rectifiers or distributors, serving in particular to properly orient the gas flow. By way of example, in Figure 3, there is shown connecting elements 72 extending between the outer fairing 12 and the internal fairing 14 upstream of the fan 22, connecting elements 74 extending between the outer fairing 12 and the internal fairing 14 downstream of the fan 22, and connecting elements 76 extending between the internal fairing 14 and the frame 52 downstream of the low pressure turbine 60.

FIG. 4 is a perspective view of a blade 80 of the fan 22 of FIG. 3. The blade 80 may correspond to a single piece, in particular a foundry piece. The blade 80 comprises successively, from bottom to top in FIG. 4, a root 82, a first heel or plate 84, the blade part 86 functioning as a turbine rotor blade, a second heel or plate 88, and a part of the blade. blade 90 functioning as a fan blade. The foot 82 is intended to cooperate with an opening provided at the periphery of a disc, not shown, integral with the shaft 68, not shown, and secures the blade 80 to the shaft 68. When the blades 80 are assembled, the first beads 84 form a first ring intended to cooperate with the frame 52 to form a sliding connection, substantially sealed. When the blades 80 are assembled, the second beads 88 form a second crown intended to cooperate with the internal fairing 14 to form a sliding connection, substantially sealed. In the present embodiment, it is the central part of the fan 22, which functions as a rotor of a low pressure turbine, which drives the entire fan 22 in rotation around the axis D.

Figure 5 is a sectional view, partial and schematic, of another more detailed embodiment of a part of the reactor 16 of the turbojet 10 shown in Figure 1. The embodiment of the reactor 16 shown in Figure 5 comprises the 'All of the elements of the reactor 16 shown in FIG. 3, the low pressure LP coupling comprising two additional low pressure turbine stages 92, 94 in addition to the low pressure turbine stage 62. Each low pressure turbine stage 92, 94 comprises a rotor 96, 98 driving the shaft 68 and a stator 100, 102 connected to the frame 52.

Figure 6 is a sectional view, partial and schematic, of another more detailed embodiment of part of the reactor 16 of the turbojet 10 shown in Figure 1. The embodiment of the reactor 16 shown in Figure 6 comprises the 'All the elements of the reactor 16 shown in Figure 3 with the difference that the rotor 66 of the turbine stage 62 is connected to the shaft via a speed reducer 110. In addition, the low hitch LP pressure comprises an additional low pressure turbine stage 92 in addition to the turbine stage 62.

The speed reducer 110 may correspond to a gear system. By way of example, FIG. 6 shows a gear system comprising three gears 112, 114, 116, including a first gear 112 rotating integrally with the shaft 68, a second gear 114 meshing with the first gear 112 and integral with of a shaft 118 mounted freely in rotation on the frame 52 by bearings 120. The third gear 116 is integral with the shaft 118 and cooperates with a toothed ring 122 integrated into the fan 22. The fan 22 is mounted freely in rotation on the shaft 68 via bearings 124. The use of the speed reducer 110 allows the rotor 96 of the turbine stage 62 and the fan 22 to rotate at different rotational speeds.

Figure 7 is a sectional view, partial and schematic, of a more detailed embodiment of the turbojet 10 shown in Figure 1. In the present embodiment, the air inlet 28 of the internal fairing 14 is located in the downstream half of the internal fairing 14, or even in the quarter at the downstream end of the internal fairing 14. The low pressure LP coupling comprises the turbine stage 62, including the rotor 22 formed by the central part of the fan 22 and the stator 64, and three other turbine stages 124, 126, 128. Each low pressure turbine stage 124, 126, 128 comprises a rotor 130, 132, 134, driving the shaft 68 in rotation, preceded by a stator 136 , 138, 140 integral with the frame 52.

The high pressure HP hitch comprises an axial compressor 142, a radial compressor 144 and the high pressure turbine stage 50. The axial compressor 142 comprises a rotor 146 followed by a stator 148 and the radial compressor 144 comprises a rotor 150, also called impeller, followed by a radial diffuser 152 and an axial rectifier 154. The rotors 146, 150 are driven in rotation by the shaft 54. In FIG. 7, the fan 22 and the rotors 58, 130, 132, 134, 146, 150 are shown in perspective. The reactor 16 comprises an annular combustion chamber 156 between the radial compressor 144 and the high pressure turbine stage 50. The combustion chamber 154 may be of the annular type or of the cellular type. FIG. 8 represents a more detailed embodiment of the turbojet 40 represented in FIG. 2. In the embodiment illustrated in FIG. 8, the free end 26 of the nozzle 24 for expelling the outlet flow of the burnt gases NI is located upstream of the inlet opening 18 of the outer shroud 12. As a variant, the free end 26 of the nozzle 24 can be located upstream of the air inlet 18 of the outer shroud 12 as shown in FIG. 2. The low pressure LP coupling comprises the turbine stage 62 and the other three stages of the low pressure turbine 124, 126, 128 previously described in relation to FIG. 7. However, it also comprises four compressor stages low pressure 160, 162, 164, 166. The rotor of the first stage of the low pressure compressor 160 receiving the primary flow F1 corresponds to the blower 22. The other three stages of compressor 162, 164, and 166 use the rotors 134, 132, and 130 respectively of the turbine stages e low pressure 124, 126, and 128.

FIG. 9 is a perspective view of the blade 80 of the fan for the embodiment illustrated in FIG. 8. With respect to FIG. 4, each blade 80 of the fan 22 for the embodiment illustrated in FIG. 8 comprises in besides a blade portion 168 shaped to function as a compressor blade and which is interposed between the heel 84 and the root 82, an additional heel 170 being provided between the blade portion 168 and the root 82. The embodiment of the blade 80 shown in FIG. 9 therefore comprises, from the axis of rotation towards the periphery of the fan, the root 82, the heel 170, the blade part 168 functioning as a LP compressor blade, the heel 84, the part of the blade. blade 86 functioning as a LP turbine blade, the heel 88, and the blade portion 90 functioning as a fan blade.

Each rotor blade 130, 132, or 134 may have a shape similar to the embodiment of the blade 80 shown in Figure 4 and include a blade portion shaped to function as a turbine rotor blade and a blade portion shaped to operate. like a compressor rotor blade. In the present embodiment, for each rotor blade 130, 132 or 134, the blade portion shaped to function as a compressor rotor blade is closer to the D axis than the blade portion shaped to function as a compressor rotor blade. turbine rotor blade.

The high pressure HP coupling comprises the axial compressor 142, the radial compressor 144 and the high pressure turbine stage 50 as described previously in relation to FIG. 7, with the difference that the rotor 58 of the high turbine stage pressure 50 also corresponds to the rotor of the axial compressor 142. Further, each blade of the rotor 58 thus comprises a blade portion shaped to function as a turbine rotor blade and a blade portion shaped to function as a compressor rotor blade. . In the present embodiment, for each rotor blade 58, the blade portion shaped to function as a compressor rotor blade is closer to the D axis than the blade portion shaped to function as a compressor rotor blade. turbine.

In the embodiment illustrated in Figure 8, the primary flow flows from upstream to downstream with respect to the turbojet 40 as it passes through the low pressure compressor stages 160, 162, 164 and 166 and the axial compressor 142 of the engine. high pressure hitch. The primary flow flows from downstream to upstream with respect to the turbojet 40 when it passes through the radial compressor 144, the combustion chamber 156, the high pressure turbine stage 50 and the low pressure turbine stages 124, 126, 128 and 62.

An example of application of the turbojets described above relates to a drone comprising a body equipped with turbojets as described above. The turbojets can be mounted so as to pivot relative to the body of the drone. Thus, in a vertical take-off or landing phase, the D axis of each turbojet can be substantially vertical, and in a horizontal displacement phase, the D axis of at least one turbojet can be inclined relative to the vertical. The turbojet embodiments described above are particularly suitable when the speed of movement of the drone relative to the ground is not too high, for example in the take-off and landing phases. When the speed of movement of the drone relative to the ground increases, it may be desirable for the exhaust flow of the burnt gases NI to be directed at least in part towards the downstream of the turbojet, like the expelled flow N2. For this purpose, the turbojet may include a device for reversing the outlet flow of the burnt gases NI.

FIGS. 10 and 11 represent an embodiment of a turbojet 50 in two operating configurations and FIGS. 12 and 13 represent an embodiment of a turbojet 60 in two operating configurations. The turbojet 50 comprises all the elements of the turbojet 10 shown in Figure 7 and the turbojet 60 comprises all the elements of the turbojet 40 shown in Figure 8. Each of the turbojets 50 and 60 further comprises a device 52 of NI flue gas outlet flow reversal. The device 52 comprises movable flaps 54, which make up the end part of the nozzle 24, and a mechanism 56 for actuating the movable flaps 54. The actuating mechanism is shown in FIG. 10 schematically by control pins connecting the movable flaps 54 to the outer fairing 12 and is not shown in Figures 11, 12, and 13.

In the operating configuration shown in FIG. 10 and in FIG. 12, the movable flaps 54 are arranged so as to form the free end 26 of the nozzle 24 for that the output flow of the burnt gases NI flows upstream from the turbojet 50 and 60, in a manner analogous to what has been described previously for the turbojet 10 and 40. In the operating configuration shown in FIG. 11 and in FIG. 13, the movable flaps 54 are arranged so as to free openings 58 allowing a part, preferably the majority, or even all, of the outlet flow of the burnt gases NI to escape from the nozzle 24 directly towards downstream of the turbojet 50 and 60 and joins the expelled flow N2. The passage between the two operating configurations is obtained by moving the movable shutters 54 by means of the actuating mechanism 56.

Various embodiments and variants have been described. Those skilled in the art will understand that certain features of these various embodiments and variations could be combined, and other variations will be apparent to those skilled in the art. In particular, in the embodiment illustrated in FIG. 4, there is shown a blade 80 with a so-called fir base 82. However, other forms of foot can be provided, for example dovetail feet. In particular, the person skilled in the art can provide a system for cooling the blades of the rotors, a system for modifying the pitch angle of the blades of the stators, etc. Finally, the practical implementation of the embodiments and variants described is within the abilities of those skilled in the art based on the functional indications given above.

Claims

1. Turbojet (10; 40) bypass comprising a fan (22) shrouded and a reactor (16) comprising a nozzle (24) for expelling flue gas (NI) by the reactor upstream of the turbojet, the free end (26) of the nozzle being located upstream of the fan (22).

2. A turbojet according to claim 1, intended to receive an air flow dividing into a primary flow (Fl) and a secondary flow (F2), the primary flow supplying the reactor (16) and forming the burnt gases (NI). ejected upstream of the turbojet (10), the secondary flow being ejected downstream of the turbojet.

3. A turbojet according to claim 2, comprising a first fairing (12) having a first air inlet (18) at least for the secondary flow (F2), the fan (22) being contained in the first fairing, the turbojet comprising a second air inlet (28) for the primary flow (Fl).

4. A turbojet according to claim 3, wherein the second air inlet (28) is located upstream of the first air inlet (18).

5. A turbojet according to any one of claims 1 to 4, wherein the rate of dilution of the turbojet (10) is between 8 and 15.

6. A turbojet according to any one of claims 1 to 5, wherein the nozzle (24) is divergent.

7. A turbojet according to any one of claims 1 to 6, wherein the turbojet has a double body and comprises a low pressure hitch (LP) having at least one low pressure turbine stage (62) and a high pressure hitch (HP. ), in which part of the fan (22) forms the rotor (66) of the low pressure turbine stage.

8. A turbojet according to claim 7, comprising first blades (80), each first blade comprising a first part (86) and a second part (90), the first parts (86) of the first blades forming the rotor (66) of the low pressure turbine stage (62) and the second parts (90) of the first blades forming the fan (22).

9. A turbojet according to claim 7 or 8, wherein the low pressure hitch (LP) comprises at least one low pressure compressor stage (160) and wherein another part of the fan (22) forms the rotor of the. low pressure compressor stage.

10. A turbojet according to claim 9 in its attachment to claim 8, wherein each first blade (80) further comprises a third part (168), the third parts of the first blades forming the rotor of the compressor stage. low pressure (160).

11. A turbojet according to claim 9 or 10, wherein the low pressure compressor stage (160) is traversed by the primary flow (Fl) while the primary flow flows from upstream to downstream of the turbojet ( 10) and in which the low pressure turbine stage (62) is crossed by the burnt gases (NI) while the burnt gases flow from downstream to upstream of the turbojet.

12. A turbojet according to any one of claims 7 to 11, wherein the low pressure coupling (LP) comprises at least first and second stages of low pressure turbine (62, 124) and first and second stages of low compressor. pressure (160, 166), in which a part of the fan (22) forms the rotor (66) of the first stage of the low pressure turbine, the turbojet comprising second blades, each second blade comprising a first part and a second part, the first parts of the second blades forming the rotor of the second stage of the low pressure turbine (124) and the second parts of the second blades forming the rotor (166) of the second stage of the low pressure compressor.

13. A turbojet according to any one of claims 1 to 12, further comprising movable flaps (54) and a mechanism (56) for actuating the movable flaps between a first position in which the movable flaps allow the expulsion of gases. burnt gas (NI) upstream of the turbojet and a second position in which the movable flaps release openings (58) for expelling the burnt gases (NI) downstream of the turbojet.