SUSPENSION OF A THREE-FLEET AIRCRAFT TURBOMACH ENGINE

Description

Domaine technique de l'invention

[0001] The present invention relates to the field of suspension of a triple-flow turbomachine, in particular of an aircraft. Arrière-plan technique

[0002] The state of the art includes in particular documents WO-A1-2010 / 092263, FR-A1-2 981 989, FR-A1-2 987 416, FR-A1-3 008 462, FR-A1-3 008 463, FR-A1-3041 054, FR-A1-3 078 998, US-B2-9,091,207, US-A1-2021 / 323690 and US-B1-6,209,311.

[0003] As depicted in the figure 1 , an aircraft turbomachine 10 generally comprises a gas generator 12 including at least one compressor 14, an annular combustion chamber 16 and at least one turbine 18.

[0004] The turbomachine 10 further includes at least one propulsion propeller 20 which is driven by a shaft of the turbine 18.

[0005] The gas generator 12 includes an annular nozzle 22 for separating two annular veins V1, V2, with respective flow of an internal primary flow F1 inside the gas generator 12, and an external secondary flow F2 around the gas generator 12.

[0006] When the propulsion propeller 20 is located upstream of the gas generator 12, as in the example shown, it is called a fan and its purpose is to accelerate the intake air. Part of the airflow exiting the fan flows around the gas generator 12 to form the secondary flow F2, which generates most of the thrust produced by the turbomachine 10. The remaining airflow exiting the fan enters the gas generator 12 to form the primary flow F1. This air is compressed in the compressor(s) 14, then mixed with fuel and burned in the combustion chamber 16. The combustion gases are then expanded in the turbine(s) 18 to rotate the turbine rotor and thus the propulsion propeller 20.

[0007] The propeller 20 comprises blades 24 which generally have an angular position, called pitch, that is fixed around an axis Y perpendicular to the longitudinal axis X of the turbomachine 10. Even in low engine speed phases, such as idle, this pitch is configured so that the flow rate of the generated secondary flow F2 is sufficient to supply surface or non-surface fluid / air heat exchangers 26, installed in the flow path of the secondary flow F2. These heat exchangers 26 are necessary to cool equipment of the turbomachine.

[0008] THE figures 2 And 3 show two possible configurations for the suspension of turbomachine 10 of the figure 1 A pylon 28 is used to attach and suspend the turbomachine 10 from a part of the aircraft, such as a wing 46 for example. The pylon 28 has a generally elongated shape along the longitudinal axis X of the turbomachine 10.

[0009] The turbomachine 10 comprises upstream suspension elements 48 and downstream suspension elements 50. In the suspension configuration of the figure 2 The downstream suspension members 50 are fixed to a turbine casing 52 of the turbomachine 10, and the upstream suspension members 48 are fixed to an intermediate casing 54 of the turbomachine 10. Thrust-retaining connecting rods 56 extend from the upstream suspension members 48 to the gas generator 12.

[0010] In the suspension configuration of the figure 3 The downstream suspension members 50 are fixed to the turbine casing 52 of the turbomachine 10, and the upstream suspension members 48 are fixed to a blower casing 58 of the turbomachine 10. Thrust-recovery rods 56 extend from the downstream suspension members 50 to the gas generator 12.

[0011] To improve engine performance (higher thrust and lower fuel consumption), some engine architectures aim to increase the bypass ratio, also known as BPR (which is the acronym for the Anglo-Saxon term). Bypass Ratio ), thanks to an increase in the diameter of the propeller 20. In order to limit the mass of the turbomachine 10, the fairing 28 (with the fan casing 58) located around the propulsion propeller 20 can be removed, and the blades 24 of the propeller 20 can be made steerable, and therefore variable pitch, to control the thrust level of the turbomachine.

[0012] However, it sometimes happens that in certain phases such as idle on the ground, stationary point or thrust reversal phases, the flow rate of the secondary flow F2 is no longer sufficient for the thermal management of the turbomachine which needs heat exchange between fluids and air in order to ensure the cooling of its equipment and its optimal operation.

[0013] To address this problem, it has already been proposed to provide an additional bypass flow dedicated to the thermal management of the turbomachine.

[0014] As depicted in the figure 4 , the turbomachine 10 then includes a secondary propeller 30 driven in rotation by a shaft of the gas generator 12. This propeller 30 is located in the flow vein V1 of the primary flow F1 and upstream of a second annular nozzle 32 separating two annular veins V11, V12 of respective flow of a first internal flow F11 and a second external flow F12.

[0015] The heat exchangers 26 are housed in the V12 channel and are supplied by the second external flow F12. The first internal flow F11 supplies the compressor 14, as mentioned above.

[0016] There figure 4 allows us to see that the gas generator 12 comprises two annular compartments C1, C2 which extend around the X axis.

[0017] The first annular compartment C1 is called the hot compartment because it is relatively close to the flow stream V11 and therefore more exposed to the heat generated by compression, the combustion chamber 16, and the combustion gases. This hot compartment C1 extends axially between the second nozzle 32 and a first nozzle 34 for the ejection of the first internal flow F11 exiting the turbine(s). This hot compartment C1 is further delimited radially by first and second annular walls 36 and 38, respectively internal and external, which extend coaxially around each other. The first wall 36 externally defines the flow stream V11 of the first internal flow F11, and the second wall 38 internally defines the flow stream V12 of the second external flow F12.

[0018] The second annular compartment C2 is called the cold compartment because it is less exposed to heat, particularly because it is separated from the hot compartment by the flow channel V12. This cold compartment C2 extends axially between the first nozzle 22 and a second nozzle 40 for the ejection of the second external flow F12. This cold compartment C2 is further delimited radially by third and fourth annular walls 42 and 44, respectively internal and external, which extend coaxially around each other. The third wall 42 externally defines the flow channel V12 for the second external flow F12, and the fourth wall 44 internally defines the flow channel V2 for the secondary flow F2.

[0019] Many components of the turbomachine 10 are installed in the large hot compartment C1. The relatively small cold compartment C2 is reserved for equipment that cannot withstand the temperatures of the hot compartment C1, particularly electronic equipment.

[0020] This type of installation has several consequences: The equipment in the hot compartment C1 must be thermally protected. This thermal protection adds volume and mass to an already constrained environment, which tends to hinder accessibility for inspections and maintenance of the turbomachine. Localized ventilation must be added in the hot compartment C1 for certain (electronic) equipment that is more sensitive to heat and cannot be installed in the cold compartment C2 due to space limitations. Equipment has a reduced lifespan in the hot compartment C1 and this restrictive environment has consequences for its design: the choice of their technologies is limited and their mass is increased.

[0021] There figure 5 shows a possible configuration for the suspension of turbomachine 10 of the figure 4 The upstream suspension members 48 and downstream 50 are located at the cold compartment C2. The downstream suspension members 50 are connected to the intermediate casing 54 and the thrust linkages 56 extend from the gas generator 12 to a fixing point on the pylon 28 which is located well downstream of the suspension members 48, 50.

[0022] There figure 6 illustrates another type of triple-flow turbomachine 10, in which the main propeller 20 is located downstream of the gas generator 12. This main propeller 20 is part of a doublet of unshod and counter-rotating propellers.

[0023] This figure 6 Figure 10 shows a possible configuration for the suspension of the turbomachine. The upstream suspension members 48 are located at the cold compartment C2, and the downstream suspension members 50 are located at the hot compartment C1. The thrust take-up rods 56 extend from the gas generator 12 to a fixing point on the pylon 28 which is located between the suspension members 48 and 50.

[0024] The different suspension configurations mentioned above present constraints: The upstream and downstream suspension elements 48, 50 are located in planes P1, P2 perpendicular to the longitudinal axis X of the turbomachine 10, which are located at a relatively small axial distance (called center distance) from each other, which generates significant forces in the elements 48, 50; the distribution of forces between the upstream and downstream suspension elements 48, 50 is also problematic because the center of gravity of the turbomachine 10 is not in the middle between the two suspension planes P1, P2; and the thrust of the turbomachine 10, which is notably generated by the propeller 20, passes through casings to be transmitted to the pylon 28 by the connecting rods 56.

[0025] The invention offers a solution to resolve all or part of these problems. Résumé de l'invention

[0026] The invention relates to a triple-flow turbomachine for an aircraft, this turbomachine having a longitudinal axis and comprising: a gas generator comprising at least one compressor, one annular combustion chamber, and at least one turbine, at least one main propeller driven in rotation by a shaft of the gas generator, the gas generator comprising: a first annular nozzle for separating two annular flow paths, respectively of a primary flow internal to the gas generator and a secondary flow external around the gas generator, a second annular nozzle for separating two annular flow paths, respectively of a first internal flow and a second external flow, a first ejection nozzle for the first internal flow, a second ejection nozzle for the second external flow, a first annular compartment, called the hot compartment, extending axially between the second nozzle and the first nozzle, and between first and second annular walls, respectively internal and external, which extend, preferably coaxially, around each other, the first wall defining externally said flow path of the first internal flow, and said second wall defining internally said flow path of the second external flow,a second annular compartment, called the cold compartment, extending axially between the first nozzle and the second nozzle, and between third and fourth annular walls, respectively internal and external, which extend, preferably coaxially, around each other, the third wall defining externally said flow path of the second external flow, and said fourth wall defining internally said flow path of the secondary flow, upstream suspension members of the turbomachine, which are located in a first plane perpendicular to said axis and which are connected or fixed to the gas generator, downstream suspension members of the turbomachine, which are located in a second plane perpendicular to said axis and which are connected or fixed to the gas generator,and thrust-retaining connecting rods comprising first ends connected or fixed to the gas generator and second opposite ends situated in a third plane perpendicular to said axis, characterized in that said first, second and third planes are situated axially between the first nozzle and the second jet nozzle, and in that at least a portion of said upstream and downstream suspension members are connected or fixed to said third wall.

[0027] In this application, the terms "upstream" and "downstream" refer to the direction of gas flow in the turbomachine.

[0028] The invention proposes to axially bring the turbomachine's suspension planes closer together and position them at the level of the third wall, and therefore the cold compartment of the gas generator. The suspension components are connected or fixed to the third wall, which is thus configured to transmit the forces related to the weight and thrust of the turbomachine to the tower. This third wall is therefore at least partially structural, meaning that it has the rigidity and mechanical strength necessary to transmit the aforementioned forces.

[0029] Advantageously, the turbomachine is devoid of additional suspension elements downstream of said downstream suspension elements or thrust take-up rods so that the turbomachine is fixed in cantilever to the pylon.

[0030] The turbomachine according to the invention may comprise one or more of the following features, taken individually or in combination with each other: The downstream suspension elements are fixed or connected to said third wall; the downstream suspension elements are located in the cold compartment; the upstream suspension elements are fixed to the third wall and / or the fourth wall; the upstream suspension elements are located just downstream of said first nozzle; said thrust linkages extend from the upstream suspension elements or downstream of these elements to connection points on the gas generator, these connection points being located inside the third wall; said connection points are located upstream of said second nozzle; alternatively, they could be located downstream of said second nozzle; at least a portion of said second and third walls are connected together by tubular reinforcements or arms to form a single unit; the second and third walls include openings and removable hatches for closing these openings;said assembly forms a single-piece section; said assembly comprises a first single-piece section and a second section which can be detached from the first section and which can itself be single-piece or formed of two half-shells hinged to each other about an axis substantially parallel to said longitudinal axis; at least one heat exchanger is located in said second external flow, and / or at least one piece of equipment, such as an accessory or gear box, is located in said cold compartment; said at least one main propeller is located upstream of the gas generator; said at least one main propeller is located downstream of the gas generator; the turbomachine includes a secondary propeller driven in rotation by a shaft of the gas generator, this propeller being located in the flow path of the primary flow and upstream of said second annular nozzle;- The pylon comprises an upstream end situated between two adjacent stator blades extending in the secondary flow channel, or situated at the base of one of these blades; - the upstream suspension members are situated between two adjacent stator blades extending in the secondary flow channel, or situated at the base of one of these blades; - the upstream suspension members are situated downstream of the first nozzle, and upstream or opposite the second nozzle; - said assembly is fixed downstream of a housing situated between two compressors of the gas generator.

[0031] The present invention also relates to an assembly comprising a turbomachine as described above, and a pylon for attaching this turbomachine to an aircraft, the pylon having a generally elongated shape and being connected or fixed respectively to the upstream and downstream suspension elements of the turbomachine. Brève description des figures

[0032] Other features and advantages will become apparent from the following description of a non-limiting embodiment of the invention with reference to the accompanying drawings in which: [ Fig.1 ] there figure 1 is a schematic axial cross-sectional view of a turbofan engine for an aircraft, [ Fig.2 ] there figure 2 is a schematic axial cross-sectional view of the turbomachine of the figure 1 and its suspension pylon to an aircraft, and illustrates a configuration of the suspension components, [ Fig.3 ] there figure 3 is a schematic axial cross-sectional view of the turbomachine of the figure 1 and its suspension pylon to an aircraft, and illustrates another configuration of the suspension components, [ Fig.4 ] there figure 4 is a schematic half-view in axial cross-section of a triple-flow turbomachine for an aircraft, [ Fig.5 ] there figure 5 is a schematic axial cross-sectional view of the turbomachine of the figure 4 and its suspension pylon to an aircraft, and illustrates a configuration of the suspension components, [ Fig.6 ] there figure 6 is a schematic axial cross-sectional view of a downstream pusher propeller turbomachine and its aircraft suspension pylon, and illustrates a configuration of the suspension components, [ Fig.7 ] there figure 7 is a view similar to that of the figure 5 and illustrates an embodiment of a turbomachine according to the invention and its suspension pylon for an aircraft, [ Fig.8 ] there figure 8 is a schematic perspective view of an assembly for the turbomachine of the figure 7 , [ Fig.9 ] there figure 9 is a schematic perspective view of an alternative embodiment of the assembly for the turbomachine of the figure 7 , [ Fig.10 ] there figure 10 is a schematic view of one variant implementation of the entire figure 9 , front view from downstream, [ Fig.11 ] there figure 11 is a view similar to that of the figure 7 and illustrates a variant embodiment of a turbomachine according to the invention and its suspension pylon for an aircraft, [ Fig.12 ] there figure 12 is a view similar to that of the figure 7 and illustrates another embodiment of a turbomachine according to the invention and its suspension pylon for an aircraft, [ Fig.13 ] there figure 13 is a view similar to that of the figure 7 and illustrates another embodiment of a turbomachine according to the invention and its suspension pylon for an aircraft. Fig.14 ] there figure 14 is a view similar to that of the figure 7 and illustrates another embodiment of a turbomachine according to the invention and its suspension pylon for an aircraft, [ Fig.15 ] there figure 15 is a view similar to that of the figure 7 and illustrates another embodiment of a turbomachine according to the invention and its suspension pylon for an aircraft, and [ Fig.16 ] there figure 16 is a view similar to that of the figure 7 and illustrates another variant of the embodiment of a turbomachine according to the invention and its suspension pylon to an aircraft. Description détaillée de l'invention

[0033] THE figures 1 à 6 have been described above.

[0034] There figure 7 represents an aircraft turbomachine 10 according to an embodiment of the invention, as well as a pylon 28 for suspending the turbomachine 10 from a part of an aircraft, such as a wing, for example. The pylon 28 has a generally elongated shape along an axis Y that is parallel to the longitudinal axis X of the turbomachine 10. The pylon 28 comprises an external diameter denoted Dext1 and an internal diameter denoted Dint1. The turbomachine 10 is similar to that of the figure 4 It includes a gas generator 12 comprising at least one compressor 14, an annular combustion chamber 16 and at least one turbine 18.

[0035] The gas generator 12 preferably comprises two successive compressors 14a, 14b. The upstream compressor 14a is a low-pressure compressor. The downstream compressor 14b is a high-pressure compressor. The two compressors 14a, 14b are separated from each other by an intermediate or inter-compressor housing 54, which includes an internal hub 54a and a ferrule 54b extending around the hub 54a and connected to the hub by arms (not shown).

[0036] The turbomachine 10 further includes a propulsion propeller 20 which is located upstream of the gas generator 12 and which is driven by a shaft of the turbine 18. The propeller 20 has an external diameter denoted Dext2 and an internal diameter denoted Dint2.

[0037] The gas generator 12 includes an annular nozzle 22 for separating two annular veins V1, V2, with respective flow of an internal primary flow F1 inside the gas generator 12, and an external secondary flow F2 around the gas generator 12.

[0038] The turbomachine 10 includes a secondary propeller 30 driven in rotation by a shaft of the gas generator 10. This propeller 30 is located in the flow vein V1 of the primary flow F1 and upstream of a second annular nozzle 32 separating two annular veins V11, V12 of respective flow of a first internal flow F11 and a second external flow F12.

[0039] Heat exchangers 26 are housed in the V12 channel and are supplied by the second external flow F12. The first internal flow F11 supplies the compressor 14, as mentioned above.

[0040] The gas generator 12 comprises two annular compartments C1, C2 which extend around the X axis.

[0041] The first annular compartment C1, or hot compartment, extends axially between the second nozzle 32 and a first ejection nozzle 34 for the first internal flow F11 exiting the turbine(s). This hot compartment C1 is further delimited radially by first and second annular walls 36 and 38, respectively internal and external, which extend coaxially around each other. The first wall 36 externally defines the flow path V11 of the first internal flow F11, and the second wall 38 internally defines the flow path V12 of the second external flow F12.

[0042] The second annular compartment C2, called the cold compartment, extends axially between the first nozzle 22 and a second nozzle 40 for the ejection of the second external flow F12. This cold compartment C2 is further delimited radially by third and fourth annular walls 42 and 44, respectively internal and external, which extend coaxially around each other. The third wall 42 externally defines the flow channel V12 of the second external flow F12, and the fourth wall 44 internally defines the flow channel V2 of the secondary flow F2.

[0043] The turbomachine 10 includes, downstream of the propeller 20 and the nozzle 22, stator blades 21 which extend radially outwards from the wall 44 and through the V2 channel.

[0044] The pylon 28 includes a free end 28a at its upstream end. This free end 28a is located here between two adjacent stator blades 21 or at the foot of one of these blades 21. The free end 28a of the pylon 28 is located between the leading and trailing edges of the stator blades 21.

[0045] The C2 cold compartment is designed to house equipment from the turbomachine 10, as will be described in more detail below.

[0046] The turbomachine 10 is suspended from the pylon 28 by means of upstream suspension members 48 and downstream suspension members 50.

[0047] The upstream suspension members 48 are located in a first plane P1 perpendicular to the axis X and are connected or fixed to the gas generator 12. The downstream suspension members 50 are located in a second plane P2 perpendicular to the axis and are connected or fixed to the gas generator 12.

[0048] According to the invention, these planes P1, P2 are located axially between the first nozzle 22 and the second nozzle 40, i.e. at the level of the cold compartment C2. Furthermore, at least a part of these suspension elements 48, 50 are connected or fixed to the wall 42.

[0049] In the example shown, Dext1 is less than Dext2, and Dint1 is less than the maximum diameter of wall 44 so that a lower part of pylon 28 is housed in the cold compartment C2, as shown in the figure.

[0050] The organ(s) 48 are located at the upstream end 28a of the pylon 28 and they may be located just downstream of the beak 22. They may be fixed or connected to the wall 42 and / or to the wall 44.

[0051] The organs 48 can be located substantially opposite the spout 32. In other words, the plane P1 can pass substantially at the level of the spout 32 or just upstream of it, as illustrated in the drawing.

[0052] Thrust-recovery connecting rods 56 extend from the components 48 to the gas generator 12. The connecting rods 56 extend from upstream to downstream radially outwards and include first radially internal ends 56a connected, for example by clevises, to the wall 42 or to points W of the gas generator 12 located radially inside this wall 42. These points W may be located upstream of the nozzle 32. However, this is not always the case, as illustrated by variants 14 to 16 described below.

[0053] The connecting rods 56 include radially external second ends 56b which are located in a third plane P3 perpendicular to the X-axis. The plane P3 is located between the planes P1 and P2. These second ends 56b are, for example, connected by a rocker arm to the pylon 28, just downstream of the components 48.

[0054] The components 50 are located downstream of the components 48 and are fixed or connected to the wall 42. They are located inside the cold compartment C2.

[0055] Wall 42 is connected to wall 38 by tubular arms 60 that extend radially within channel V12 and are used for the passage of utilities through this channel. Utilities include, for example, electrical harnesses and cables, oil lines, ventilation ducts, etc. Walls 38, 42, and arms 60 form a single unit called assembly 62.

[0056] Alternatively, the 60 arms could be replaced by (non-tubular) reinforcements.

[0057] The assembly 62 is intended to be fixed to the downstream end of the housing 54. The wall 42 has its upstream end which is fixed to the downstream end of the ferrule 56b, and the wall 38 has its upstream end which is fixed to the hub 56a.

[0058] THE figures 8 à 10 illustrate examples of the realization of this set 62. In the case of the figure 8 The assembly 62 is a single unit. The wall 42 includes openings 42a that lead into the internal cavities of the arms 60 for the passage of the aforementioned utilities. The wall 42 also includes openings 42b configured to allow for maintenance and inspection operations. These openings 42b are designed to be sealed by removable hatches, which are not shown. These hatches can be disassembled and removed through similar openings provided on the wall 44, for example.

[0059] Similarly, the wall 38 may include similar openings 38a to permit maintenance and inspection operations into the assembly 62. The openings 42b, 38a of the walls 38, 42 are preferably aligned radially with each other.

[0060] Set 62 of the figure 8 is structural and configured to transmit the forces of the turbomachine 10 to the pylon 28 in operation.

[0061] In the case of the figure 9 The assembly 62 comprises two sections, upstream 62a and downstream 62b respectively. The upstream section 62a is similar to the assembly 62 of the figure 8 The downstream section 62b is intended to be detachably fixed to the downstream end of the upstream section 62a. More specifically, the downstream section 62b comprises a portion of the wall 42 which is intended to be detachably fixed to the portion of the wall 42 of section 62a, and a portion of the wall 38 which is intended to be detachably fixed to the portion of the wall 38 of section 62a.

[0062] The upstream section 62a and the downstream section 62b are each of the type O - Duct, that is to say that they each form an uninterrupted annular conduit, as described in document FR-A1-2 981 989.

[0063] There figure 10 illustrates a variant embodiment of assembly 62 in which the upstream section 62a of the type O - Duct is associated with a downstream section 62b of the type D-Duct. Section 62b is formed of two half-shells 62b1, 62b2 articulated to each other around an axis Z substantially parallel to the axis X. Opening section 62b, by pivoting the half-shells 62b1, 62b2 around the axis X, allows access to the interior of section 62b during a maintenance or inspection operation.

[0064] There figure 11 shows 26 exchangers housed in the V12 vein.

[0065] There figure 12 shows equipment 64 housed in the cold compartment C2. The center distance between the planes P1, P2 and the positioning of the connecting rods 56 at the level of the components 48 allows space to be freed up in the compartment C2 for the installation of the equipment 64. In particular, this equipment 64 may include electronic equipment, an accessory or gear box, etc.

[0066] In yet another variant not shown, the turbomachine 10 according to the invention could be of the type shown in the figure 6 and include at least one main propeller 20, or a propeller pair, downstream of the gas generator 12.

[0067] This is particularly the case with the variant implementation of the figure 13 in which the upstream suspension members 48 are located at the cold compartment C2, the thrust take-up rods 56 extend from the gas generator 12 to a fixing point on the pylon 28, and the downstream suspension members 50 are located between the suspension members 48 and the thrust take-up rods 56.

[0068] The variants illustrated in figures 14 à 16 show that the thrust take-up rods 56 can be located upstream of the separation nozzle 32, between the gas generator 12 and the pylon 28 ( figure 14 ), downstream of the separation nozzle 32, between the internal casing 54 and the pylon 28 ( figure 15 - the P3 plane is here set back), or downstream of the separation beak 32, between the internal casing 54 and the pylon 28 ( figure 16 - Plan P3 is here set back and almost indistinguishable from plan P2).

[0069] The turbomachine 10 according to the invention has many advantages, including: The cold compartment C2 is relatively large, which is advantageous for housing equipment 64. Another advantage is having a cold compartment C2 protected from fire and thermal radiation by the flow vein V12 of the flow F12, which allows increasing the life of the equipment 64 installed in this compartment C2 while optimizing their design without very high temperature resistance constraints. Furthermore, from an aerodynamic standpoint, since the diameter of the V12 duct is relatively small compared to previous techniques, the radial dimension of the duct is larger for the same cross-sectional area compared to current techniques, which facilitates the integration of heat exchangers. Moreover, controlling the Mach number in the V12 duct is easier (the Mach number in the duct being directly related to the cross-sectional area). Indeed, with a larger diameter duct, even a very small variation in the radial dimension of the duct has a significant impact on the cross-sectional area.

Claims

1. A triple-flow turbine engine (10) for an aircraft, this turbine engine having a longitudinal axis (X) and comprising: - a gas generator (12) comprising at least one compressor (14), an annular combustion chamber (16), and at least one turbine (18), - at least one main propeller (20) rotated by a shaft of the gas generator (12), the gas generator (12) comprising: - a first annular splitter nose (22) for separating two annular ducts (V1, V2) for the respective flow of an internal primary flow (F1) inside the gas generator (12) and an external secondary flow (F2) around the gas generator (12), - a second annular splitter nose (32) for separating two annular ducts (V11, V12) for the respective flow of a first internal flow (F11) and of a second external flow (F12), - a first nozzle (34) for ejecting the first internal flow (F11), - a second nozzle (40) for ejecting the second external flow (F12), - a first annular compartment (C1), referred to as hot compartment, extending axially between the second splitter nose (32) and the first nozzle (34), and between first and second annular walls (36, 38), respectively internal and external, which extend around each other, the first wall (36) externally defining said duct (V11) for the flow of the first internal flow (F11), and said second wall (38) internally defining said duct (V12) for the flow of the second external flow (F12), - a second annular compartment (C2), referred to as cold compartment, extending axially between the first splitter nose (22) and the second nozzle (40), and between third and fourth annular walls (42, 44), respectively internal and external, which extend around each other, the third wall (42) externally defining said duct (V12) for the flow of the second external flow (F12), and said fourth wall (44) internally defining said duct (V2) for the flow of the secondary flow (F2), - upstream suspension members (48) of the turbine engine (10), which are located in a first plane (P1) perpendicular to said axis (X) and which are connected or attached to the gas generator (12), - downstream suspension members (50) of the turbine engine (10), which are located in a second plane (P2) perpendicular to said axis and which are connected or attached to the gas generator (12), and - thrust-absorbing rods (56) which comprise first ends (56a) connected or attached to the gas generator (12) and opposite second ends (56b) located in a third plane (P3) perpendicular to said axis (X), characterized in that said first, second and third planes (P1, P2, P3) are located axially between the first splitter nose (22) and the second nozzle (40), and in that at least one part of said upstream and downstream suspension members (48, 50) are connected or attached to said third wall (42).

2. The turbine engine (10) according to claim 1, wherein the downstream suspension members (50) are attached or connected to said third wall (42).

3. The turbine engine (10) according to claim 1 or 2, wherein the downstream suspension members (50) are located in the cold compartment (C2).

4. The turbine engine (10) according to any of the preceding claims, wherein the upstream suspension members (48) are attached to the third wall (42) and / or to the fourth wall (44).

5. The turbine engine (10) according to claim 4, wherein the upstream suspension members (48) are located just downstream of said first splitter nose (22).

6. The turbine engine (10) according to any of the preceding claims, wherein said thrust-absorbing rods (56) extend from the upstream suspension members (48) or downstream of these members to connection points (W) on the gas generator (12), these connection points (W) being located inside the third wall (42).

7. The turbine engine (10) according to any of the preceding claims, wherein at least one part of said second and third walls (38, 42) are connected together by tubular reinforcements or arms, to form an assembly (62).

8. The turbine engine (10) according to claim 7, wherein the second and third walls (38, 42) comprise openings (38a, 42b) and removable hatches for closing these openings.

9. The turbine engine (10) according to claim 7 or 8, wherein said assembly (62) forms a segment in one piece.

10. The turbine engine (10) according to claim 7 or 8, wherein said assembly (62) comprises a first segment (62a) in one piece and a second segment (62b) which can be detached from the first segment and which can itself be in one piece or formed of two half-shells (62b1, 62b2) articulated with respect to each other about an axis (Z) substantially parallel to said longitudinal axis (X).

11. The turbine engine (10) according to any of the preceding claims, wherein at least one heat exchanger (26) is located in said second external flow duct (V12), and / or at least one equipment (64), such as an accessory box or gearbox, is located in said cold compartment (C2).

12. The turbine engine (10) according to any of the preceding claims, wherein said at least one main propeller (20) is located upstream of the gas generator (12).

13. The turbine engine (10) according to any of claims 1 to 11, wherein said at least one main propeller (20) is located downstream of the gas generator (12).

14. The turbine engine (10) according to any of the preceding claims, wherein it comprises a secondary propeller (30) driven in rotation by a shaft of the gas generator (12), this propeller (30) being located in the duct (V1) for the flow of the primary flow (F1) and upstream of said second annular splitter nose (32).

15. An assembly comprising a turbine engine (10) according to any of the preceding claims, and a pylon (28) for hanging this turbine engine to an aircraft, the pylon (28) having a generally elongate shape and being connected or attached respectively to the upstream and downstream suspension members (48, 50) of the turbine engine (10).