Aircraft turbomachine including pylon attachment elements

Abstract

The application relates to an aircraft turbomachine, the turbomachine (1) having a longitudinal axis (X) extending from upstream to downstream along a flow direction of an airflow within the turbomachine (1), the turbomachine (1) comprising: a core (3) including a low-pressure compressor (4), a high-pressure casing, and a low-pressure turbine (8); an upstream mounting element (34a) to a pylon (101); and a downstream mounting element (34b) to the pylon (101); the upstream mounting element (34a) and the downstream mounting element (34b) being separated along the longitudinal axis (X) by at least 1000 millimeters. Figure for the abstract: Fig. 5

Description

Title of the invention: Aircraft turbomachine comprising pylon attachment elements technical field

[0001] This disclosure relates generally to the field of aircraft propulsion systems and more particularly to aeronautical turbomachinery comprising a shrouded or unshrouded fan with a high bypass ratio. It relates more specifically to the field of attaching a turbomachine to an aircraft pylon. STATE OF THE ART

[0002] An aircraft turbomachine comprises, for example, from upstream to downstream in the direction of gas flow, a core which includes a fan section, a compressor section which may include a low-pressure compressor and a high-pressure compressor, a combustion chamber, and a turbine section which may include a high-pressure turbine and a low-pressure turbine. The high-pressure compressor is driven in rotation by the high-pressure turbine via a high-pressure shaft. The fan and, where applicable, the low-pressure compressor are driven in rotation by the low-pressure turbine via a low-pressure shaft. The high-pressure compressor, the combustion chamber, and the high-pressure turbine form a high-pressure unit. Environmental aspects

[0003] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new types of aircraft and those already in service, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working for several years now to contribute to the fight against climate change.

[0004] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes these factors into consideration in all phases of design and development in order to obtain less energy-intensive and more environmentally friendly aeronautical components and products, the integration and use of which in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0005] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0006] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, in particular through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[0007] Thus, in order to improve the propulsive efficiency of the propulsion system and reduce its specific fuel consumption, it has been found that, to obtain the same thrust, it is advantageous to accelerate a large quantity of air less than to accelerate a smaller quantity of air more strongly. With this in mind, a new generation of propulsion systems, known as Open Fan, has been proposed, in which the fan blades are unshod and have variable pitch, as do the fan stator blades, generally called outlet guide vanes (OGVs).Open Fan propulsion systems, still with this aim of improving propulsive efficiency, offer a high BPR (bypass ratio, corresponding to the ratio between the secondary airflow and the primary airflow) by increasing the diameter of the fan and by extension the external dimensions of the propulsion system (and therefore its mass and drag), which makes the integration of the propulsion system more difficult in addition to increasing its mass. Structural aspects

[0008] The attachment of a turbomachine to an aircraft, for example to a wing of the aircraft, is done by means of a pylon, also called a mast, connected to the rest of the aircraft.

[0009] The pylon includes upstream and downstream turbomachine suspension components located upstream of the turbomachine, respectively on an inlet casing around the low-pressure compressor and on an intermediate casing surrounding an inter-compressor part and / or the high-pressure compressor.

[0010] Thus, the entire part of the core downstream of the downstream suspension component, including in particular the majority of the high-pressure body and the low-pressure turbine, is cantilevered, which induces several important problems, which are aggravated by the fact that this part of the core supports substantial additional masses (accessory relay box, fuel tank, pipes, electrical harnesses, supports, sheet metal, ...).

[0011] In operation, this configuration ensures the transmission of forces between the upstream and downstream suspension components. The high-pressure body is, for its part, held simply upstream, which results in deformations of the core, for example by a sagging of the high-pressure body under acceleration, and by changes in the relative clearances between the turbomachine in its downstream part and the pylon.

[0012] More specifically, the core is subject to maneuvering loads induced by acceleration variations when the aircraft performs maneuvers such as turns in flight, to gravitational loads induced during the aircraft's cruise phase, and, when the fan is unfaired or enclosed by a large-diameter casing, for example, at least 82 inches, to loads induced by a tilting force known as the IP mode. This IP mode more precisely describes the forces and moments experienced by the turbomachine during climb or descent. In the case of an unfaired turbomachine, these forces and moments arise from a difference in load encountered by, on the one hand, the upward-moving fan blades (i.e., approaching the pylon) and, on the other hand, the downward-moving fan blades (i.e., moving away from the pylon).

[0013] These different loads suffered by the core, its cantilevered position, and the additional loads it supports lead to its sagging under acceleration, to its deformation and that of its casings, which degrades the clearances at the top of the high-pressure compressor blade, creates wear and degrades its performance.

[0014] In addition, the positioning of the suspension elements upstream of the turbomachine and the mass of the cantilevered part of the core can cause local deformations of the pylon, particularly under IP loading.

[0015] The core must therefore be protected against the aerodynamic forces generated by the fan. The core must also be well supported to limit its deformation under maneuvering loads and IP loads.

[0016] One difficulty lies in the fact that these two objectives appear to be contradictory. Indeed, protecting the core from the forces generated by the fan would require decoupling it from the fan as much as possible to prevent the transmission of these forces, even to the point of cantilevering the core, but this would expose it to dangerous deformations under operating loads. Conversely, limiting its deformation under operating loads leads to maintaining the core as rigid as possible, which, in addition to having a significant impact on mass and size, conversely promotes the transmission of forces from the fan. The challenge, therefore, is to provide core support capable of reconciling these opposing constraints. EXPOSITION

[0017] One objective of this application is to optimize the performance of the propulsion system in terms of specific fuel consumption, mass, and drag, while ensuring the possibility of integrating the propulsion system into an aircraft. More specifically, it aims to to overcome the aforementioned disadvantages and in particular to improve the maintenance of the core and to reconcile these opposing constraints.

[0018] To this end, according to a first aspect, an aircraft turbomachine is proposed, the turbomachine having a longitudinal axis extending from upstream to downstream along a flow direction of an airflow within the turbomachine, the turbomachine comprising: - a core comprising a low-pressure compressor, a high-pressure body, and a low-pressure turbine; - an upstream fixing element to a pylon; and - a downstream fixing element to the pylon;

[0019] the upstream fixing element and the downstream fixing element being separated along the longitudinal axis by at least 1000 millimeters.

[0020] According to some embodiments, the turbomachine is devoid of an intermediate fixing element to the pylon located axially between the upstream fixing element and the downstream fixing element.

[0021] According to some embodiments, the turbomachine includes a structural hood extending axially between the upstream and downstream fastening elements, the structural hood being configured to form a predominant force path not passing through the core between the upstream and downstream fastening elements.

[0022] According to certain embodiments, the structural hood comprises: - an upstream ring; - a downstream ring; and - a first beam and a second beam, each connecting the upstream ring and the downstream ring.

[0023] According to some embodiments, the first beam and the second beam extend on either side of the longitudinal axis, the first beam extending for example above the longitudinal axis and the second beam extending for example below the longitudinal axis.

[0024] According to some embodiments, the structural hood comprises at least one lateral structural panel, extending between the upstream ring and the downstream ring and connecting the first beam and the second beam.

[0025] According to certain embodiments, the turbomachine comprises, from upstream to downstream: - an inlet casing extending around the low-pressure compressor, the upstream ring being axially adjacent to the inlet casing; - an outlet casing extending around the low-pressure turbine, the outlet casing comprising a turbine blade frame and an exhaust casing, the downstream ring extending around the turbine blade frame and / or the exhaust casing.

[0026] According to some embodiments, the inlet housing comprises an upstream section, an intermediate section and a downstream section adjacent to the upstream ring, the upstream fastening element being located on the upstream section or on the intermediate section, and the downstream fastening element being located on the downstream ring.

[0027] According to some embodiments, the upstream fixing element and the downstream fixing element are separated along the longitudinal axis by a maximum of 2400 millimeters.

[0028] According to some embodiments, the structural hood extends around the core.

[0029] According to certain embodiments, the structural hood comprises a section perpendicular to the longitudinal axis with dimensions between 1200 millimeters and 1700 millimeters.

[0030] According to some embodiments, the turbomachine includes an unshod fan.

[0031] According to some embodiments, the turbomachine includes a shrouded fan, in particular equipped with a casing with a diameter of at least 200 centimeters.

[0032] This disclosure further relates to an aircraft comprising:

[0033] - a turbomachine as defined above, and

[0034] - a suspension pylon comprising an upstream suspension member and a member downstream suspension respectively fixed to the upstream fixing element and the downstream fixing element of the turbomachine.

[0035] According to certain embodiments:

[0036] - the upstream suspension member is configured to allow movement of the turbomachinery along a first direction orthogonal to the longitudinal axis and passing through the suspension pylon, and to absorb forces along the longitudinal axis and along a second direction orthogonal to the longitudinal axis and to the first direction, and

[0037] - the downstream suspension element is configured to resume: • efforts in the first direction and in the second direction, and • a couple along the longitudinal axis,

[0038] the downstream suspension element being further configured to allow displacements of the turbomachine along the longitudinal axis. DESCRIPTION OF THE FIGURES

[0039] We will now present examples of implementation, by way of non-limiting example, supported by drawings on which:

[0040] The [Fig. 1] is a schematic axial cross-sectional view of a first type of motor;

[0041] The [Fig.2] is a schematic axial cross-sectional view of a second type of motor;

[0042] The [Fig.3] is an example of an aircraft comprising at least one engine;

[0043] Fig. 4 illustrates a cross-sectional view of a turbomachine;

[0044] Fig. 5 illustrates a side view of part of the turbomachine in Fig. 5;

[0045] Figure 6 illustrates a perspective view of part of the turbomachine of the [Fig. 4]; and

[0046] Fig. 7 illustrates a partial cross-sectional view of the view in Fig. 6.

[0047] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION Turboengines in Figures 1 and 2

[0048] Fig. 1 illustrates a first type of engine forming a propulsion system. The propulsion system 1 has a principal direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of gas flow in the propulsion system 1 when it is in operation, a fan section 2 and a primary body 3, often called the "gas generator" or "core," including a compressor section 4, 5, a combustion chamber 6, and a turbine section 7, 8. As will be seen, the primary body 3 forms the high-pressure body. The propulsion system 1 is here an aeronautical propulsion system configured to be mounted on an aircraft 100 by means of a pylon (or mast) 101.

[0049] The compressor section 4, 5 comprises a series of stages, each including a rotating blade wheel (rotor) 4a, 5a in front of a fixed blade wheel (stator) 4b, 5b. The turbine section 7, 8 also comprises a series of stages, each including a fixed blade wheel (stator) 7a, 8a behind which a rotating blade wheel (rotor) 7b, 8b is rotating.

[0050] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, corresponding to the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to and passing through this axis X. Furthermore, the circumferential direction corresponds to a direction perpendicular to and not passing through the longitudinal axis X. Unless otherwise specified, internal (respectively, inside) and external (respectively, outside) are used with reference to a radial direction such that the inner part or face of an element is closer to the axis X than the outer part or face of the same element.

[0051] In operation, an airflow F entering the propulsion system 1 is divided between a primary airflow Fl and a secondary airflow F2, which flow from upstream to downstream in the propulsion system 1.

[0052] The secondary airflow F2 (also called "bypass airflow") flows around the primary body 3. The secondary airflow F2 cools the periphery of the primary body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0053] The primary airflow Fl flows in a primary channel inside the primary body 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidizer, and the turbine section 7, 8. The passage of the primary airflow Fl through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes a rotation of the rotor of the turbine section 7, 8, which in turn drives the rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the blower section 2.

[0054] In a twin-body propulsion system 1, the compressor section 4, 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of the high-pressure compressor 5 is driven in rotation by the rotor of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the blower section 2 are driven in rotation by the rotor of the low-pressure turbine 8 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body including the high-pressure compressor 5, the combustion chamber 6, the high-pressure turbine 7 and the high-pressure shaft 10, and a low-pressure body including the blower section 2, the low-pressure compressor 4, the low-pressure turbine 8 and the low-pressure shaft 11.The rotational speed of the high-pressure body is greater than the rotational speed of the low-pressure body. In a three-body propulsion system 1, the turbine section 7, 8 further includes an intermediate turbine, positioned between the high-pressure turbine 7 and the low-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via an intermediate shaft. The blower rotor 9 and the rotor of the high-pressure compressor 5 remain driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0055] The low-pressure shaft 11 is generally housed, along a portion of its length, within the high-pressure shaft 10 and is coaxial with the latter. The low-pressure shaft 11 and the high-pressure shaft 10 may be co-rotating, that is, driven in the same direction around the longitudinal axis X. Alternatively, they may be counter-rotating, that is, driven in opposite directions around the longitudinal axis X. If applicable, the intermediate shaft is housed between the high-pressure shaft 10 and the low-pressure shaft 11. The intermediate shaft and the low-pressure shaft 11 may be co-rotating or counter-rotating.

[0056] The blower section 2 comprises at least the blower rotor 9 adapted to be driven in rotation relative to a blower housing 12 by the turbine section 7, 8. Each blower rotor 9 comprises a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 can be fixed relative to the hub 13 or have variable pitch. In this case, the root of the blades 14 of each rotor 9 is pivotally mounted about a pitch axis and is connected to a pitch-changing mechanism 15 mounted in the propulsion system 1, the pitch being adjusted according to the flight phases by the pitch-changing mechanism 15. This mechanism 15 is shown in dashed lines in [Fig. 1] to show that this feature is optional.

[0057] The fan section 2 may further include a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub 18 of the fan stator 16 and whose function is to rectify the secondary airflow F2 exiting the fan rotor 9. The fan stator blades 17 may be fixed relative to the hub 18 or have variable pitch. Similar to the rotor blades 14, the base of the stator blades 17 is pivotally mounted about a pitch axis X and is connected to a pitch-changing mechanism 15a, which is generally separate from that of the fan rotor 9, the pitch being adjusted according to the flight phases by the pitch-changing mechanism. This mechanism 15a is shown in dashed lines in [Fig. 1] to show that this feature is optional.

[0058] In order to improve the propulsive efficiency of propulsion system 1 and reduce its specific fuel consumption as well as the noise emitted by the fan section 2, propulsion system 1 has a high bypass ratio. A high bypass ratio is understood here to be a ratio greater than or equal to 10, for example, between 10 and 80 inclusive. To calculate the bypass ratio, the mass flow rate of the secondary airflow F2 and the mass flow rate of the primary airflow Fl are measured when propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) Manual, Doc 7488 / 3, 3rd edition) and at sea level. It should be noted that, in this application, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions.By "not installed", we understand here that the measurements are carried out when the propulsion system 1 is in a test bench (and not installed on an aircraft 100, as shown in [Fig.3]), the measurements then being simpler to carry out.

[0059] The fan rotor 9 can be provided to be decoupled from the low-pressure shaft 11 by means of a reduction mechanism 19, placed between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speeds. In this case, the propulsion system 1 further includes an additional shaft, referred to as the fan shaft 20. The low-pressure shaft 11 connects the low-pressure turbine 8 to an inlet of the reduction mechanism 19, while The blower shaft 20 connects the output of the reduction mechanism 19 to the blower rotor 9. The blower rotor 9 is therefore driven by the low pressure shaft 11 via the reduction mechanism 19 and the blower shaft 20 at a rotational speed lower than the rotational speed of the low pressure turbine 8.

[0060] This decoupling makes it possible to reduce the rotational speed and pressure ratio of the fan rotor 9 and to increase the power extracted by the low-pressure turbine 8. Indeed, the overall efficiency of the propulsion systems is conditioned to the first order by the propulsion efficiency, which is favorably influenced by minimizing the variation in kinetic energy of the air as it passes through the propulsion system 1. In a propulsion system 1 with a high dilution ratio, the bulk of the flow generating the propulsive force consists of the secondary airflow F2 of the propulsion system 1, the kinetic energy of the secondary airflow F2 being mainly affected by the compression that the secondary airflow F2 undergoes when passing through the fan section 2.The propulsive efficiency and the pressure ratio of the blower section 2 are therefore linked: the lower the pressure ratio of the blower section 2, the better the propulsive efficiency.

[0061] The blower section 2 may be shrouded or unshrouded. In the case of a shrouded blower section 2 as illustrated in [Fig. 1], the blower section 2 comprises a blower housing 12 and the blower rotor 9 is housed in the blower housing 12.

[0062] A shrouded fan section 2 comprises a fan rotor 9 extending upstream of a fan stator. The fan stator blades are then generally called outlet guide vanes (OGVs). They may have a fixed or variable pitch relative to the fan stator hub. Furthermore, the bypass ratio of the propulsion system 1 is preferably greater than or equal to 10, for example, between 10 and 35 inclusive, preferably between 10 and 18 inclusive.

[0063] In an unshod fan section 2, which corresponds to the second type of engine illustrated in [Fig. 2], the fan section 2 is not enclosed by a fan housing. Since the fan section 2 is unshod, the blades 14 of the fan rotor 9 have variable pitch. Propulsion systems comprising at least one unshod fan rotor 9 are known as "open rotor" or "unducted fan." The propulsion system 1 may comprise two unshod, counter-rotating fan rotors 9. Such a propulsion system 1 is known by the English acronym CROR for "Contra-Rotating Open Rotor" or UDF for "Unducted Double Fan." The blower rotor(s) 9 can be placed at the rear of the primary body 3 so as to be of the pusher type or at the front of the primary body 3 so as to be of the tractor type. Alternatively, the propulsion system 1 may comprise a single unshod blower rotor 9 and an unshod blower stator 16 (rectifier). Such a propulsion system 1 is known by the English acronym USF for "Unducted Single Fan". In the case of a USF-type propulsion system 1, the blades 17 of the rectifier 16 are fixed in rotation with respect to the X-axis of rotation of the upstream blower rotor 9 and consequently do not experience centrifugal force. The blades 17 of the rectifier 16 also have variable pitch.

[0064] Removing the fairing around the fan section 2 allows the dilution ratio to be increased very significantly without the propulsion system 1 being penalized by the mass of the housings or the nacelle intended to surround the fan section 2. The dilution ratio is thus greater than or equal to 40, for example between 40 and 80 inclusive. Attaching the turbomachine to an aircraft

[0065] We will present an example of an implementation applied to an aircraft turbomachine illustrated in [Fig.4] and which may be of the type of those in Figures 1 and 2. [Fig.4] is a cross-sectional view of the internal structure of the turbomachine 1.

[0066] This turbomachine 1 is here of the type of that of [Fig.2]. It has a main longitudinal axis X and includes in particular a fan 2 and a core 3 or gas generator, which comprises, from upstream to downstream in the direction of the gas flow in the turbomachine 1 when it is in operation, a low-pressure compressor 4, a high-pressure compressor 5, a combustion chamber 6, a high-pressure turbine 7 and a low-pressure turbine 8. The high-pressure compressor 5, the combustion chamber 6 and the high-pressure turbine 7 together form a high-pressure body.

[0067] In an alternative not shown, the fan 2 can be shrouded. The proposed turbomachine offers a particular advantage in this case when the fan 2 is shrouded with a large-diameter casing, i.e., for example, with a dimension greater than or equal to 82 inches in a plane perpendicular to the longitudinal axis X, since it is above a certain fan casing diameter that the effects of the IP mode become detrimental. One inch corresponds in the International System of Units (SI) to 25.40 millimeters, so a diameter of at least 82 inches corresponds to a diameter of at least 208.28 centimeters. Possibly, the fan casing 2 could be at least 200 centimeters in diameter.

[0068] Fig. 5 illustrates a profile view of the core 3 of the turbomachine 1 of Fig. 1 on which structural housings can be observed around the core 3.

[0069] The turbomachine 1 thus comprises an inlet housing 30 extending circumferentially around the low-pressure compressor 4. The inlet housing 30 includes an upstream section 30a, an intermediate section 30b and a downstream section 30c adjacent to each other along the longitudinal axis X.

[0070] Similarly, the turbomachine 1 includes an outlet casing 31 extending around the low-pressure turbine 8. The outlet casing 31 includes a turbine blade frame 41 extending around the blades of the low-pressure turbine, and an exhaust casing 31b attached directly to the turbine frame 41, the exhaust casing 31b being located axially downstream of the turbine blade frame 41. The exhaust casing 31b forms a downstream end of the outlet casing 31. The turbine blade frame 41 is not shown in [Fig. 5], as it is covered by a structural cowling which will be described below. It is, however, visible in [Fig. 6].

[0071] The turbomachine 1 is here an aircraft turbomachine configured to be fixed on an aircraft by means of a pylon (or mast) 101 schematically illustrated in [Fig.5].

[0072] Pylon 101 has a general elongated shape in a direction parallel to the longitudinal axis X, above turbomachine 1.

[0073] The pylon 101 includes suspension members 33 for attaching the turbomachine 1 to the aircraft. Thus, the pylon 101 includes an upstream suspension member 33a and a downstream suspension member 33b.

[0074] To fix the turbomachine 1 to the pylon 101, the turbomachine 1 includes fastening elements 34 to which the suspension members 33 of the pylon can be attached. In particular, the turbomachine 1 includes an upstream fastening element 34a to a pylon and a downstream fastening element 34b to a pylon.

[0075] The attachment of the upstream suspension member 33a to the upstream fixing element 34a is made in an upstream suspension plane which is perpendicular to the longitudinal axis X. Similarly, the attachment of the downstream suspension member 33b to the downstream fixing element 34b is made in a downstream suspension plane which is perpendicular to the longitudinal axis X.

[0076] The upstream suspension member 33a is, for example, configured to absorb forces along the longitudinal axis X and in a lateral direction relative to the longitudinal axis X in the upstream suspension plane. The upstream suspension member 33a is further, for example, configured to allow movement of the turbomachine 1 in a vertical direction in the upstream suspension plane. A vertical direction in a plane perpendicular to the longitudinal axis X is understood to be an upward direction, with the pylon 101 positioned above the turbomachine 1 in the vertical direction. A lateral direction in a plane perpendicular to the longitudinal axis X is understood to be a left-to-right direction, i.e., a direction perpendicular to the vertical direction. The absorption of forces along the vertical direction relative to the longitudinal axis X at the level of the upstream suspension is ensured by the upstream fixing element 34a.

[0077] The second suspension member 33b is, for example, configured to absorb forces in the vertical and lateral directions in the downstream suspension plane, and to allow movement of the turbomachine 1 along the longitudinal axis X. This prevents the assembly formed by the turbomachine 1 and the pylon 101 from becoming over-static. The second suspension member 33b also absorbs the torque defined along the longitudinal axis X.

[0078] Preferably, the turbomachine 1 is free of any other attachment element to the pylon 101, and in particular of an intermediate attachment element located axially between the upstream attachment element 34a and the downstream attachment element 34b. In other words, the upstream attachment element 34a and the downstream attachment element 34b are preferably the only attachment points of the turbomachine 1 to the pylon 101.

[0079] The upstream fastening element 34a can be provided for on the inlet housing 30, for example on the upstream section 30a or the intermediate section 30b, and / or the downstream fastening element 34b is axially at the outlet housing 31, for example at the turbine blade frame or the exhaust housing 31b. The downstream fastening element 34b is located, for example, on the structural cover, which will be described below, surrounding the turbine blade frame and / or the exhaust housing 31b.

[0080] In all cases, the wheelbase, that is to say the axial distance along the longitudinal axis X separating the upstream fixing element 34a from the downstream fixing element 34b, is at least 1000 millimeters, and at most 2400 millimeters.

[0081] Thus, the axial distance between the upstream fixing element 34a and the downstream fixing element 34b is longer than in the prior art, in particular because the downstream fixing element 34b is located further downstream.

[0082] Consequently, the downstream portion of the turbomachine 1 in the cantilevered position is reduced, which in turn reduces the relative clearances between the pylon 1 and the turbomachine 1 in this downstream portion. The influence of mechanical loads induced by the IP mode and operating loads is thus reduced, as are the deformations of the core 3. Wear is therefore reduced and the performance of the high-pressure compressor is improved. Structural cover around the core 3

[0083] Furthermore, in order to protect the core 3, the turbomachine 1 may be provided to include a structural cowling 35 extending along the longitudinal axis X between the upstream mounting element 34a and the downstream mounting element 34b. The structural cowling 35 circumferentially surrounds an axial portion of the core 3 defined between the upstream mounting element 34a and the downstream mounting element 34b, which in this example comprises a portion of the low-pressure compressor 4, the high-pressure body, and part of the low-pressure turbine 8. [Fig.6] illustrates a perspective view of the turbomachine 1 and the structural hood 35, and [Fig.7] illustrates the same view in partial section so as to partially illustrate the internal structure of the turbomachine 1.

[0084] The structural hood 35 can successively comprise from upstream to downstream an upstream ring 36, a hood body 37 and a downstream ring 38.

[0085] The upstream ring 36 is axially adjacent to the inlet housing 30. The upstream ring 36 is, for example, fixed downstream of the inlet housing 30.

[0086] The downstream ring 38 surrounds the outlet housing 31 and extends, for example, around the turbine blade frame and / or the exhaust housing 31b. The downstream fastening element 34b is thus located on the downstream ring 38.

[0087] The hood body 37 may include a first beam 37a and a second beam 37b. Each of the beams 37a, 37b extends axially so as to connect the upstream ring 36 to the downstream ring 37. The first beam 37a and the second beam 37b are opposite each other with respect to the longitudinal axis X. For example, the first beam 37a extends above the longitudinal axis X and the second beam 37b extends below the longitudinal axis X, in the vertical direction.

[0088] The hood body 37 may include at least one lateral structural panel 37c axially connecting the upstream ring 36 to the downstream ring 38, and circumferentially connecting the first beam 37a to the second beam 37b. Thus, preferably, the hood body 37 includes a first lateral structural panel 37c and a second lateral structural panel 37d connecting the first beam 37a to the second beam 37b in a circumferential direction around the core 3.

[0089] Preferably, the first beam 37a, the second beam 37b, and the lateral structural panels 37c and 37d are all concave with respect to the longitudinal axis X, in order to improve the transfer of forces.

[0090] Thus, in any plane perpendicular to the longitudinal axis X, the structural hood 35 has a substantially circular or rounded cross-section. Preferably, the first beam 37a is located, in any plane perpendicular to the longitudinal axis X, at a higher point of the cross-section of the structural hood 35 in the vertical direction, for example at 12 o'clock when this cross-section is circular, and the second beam 37b is located, in any plane perpendicular to the longitudinal axis X, at a lower point of the cross-section of the structural hood 35 in the vertical direction, for example at 6 o'clock when this cross-section is circular. In this way, the first beam 37a forms the shortest path on the structural hood 35 to connect the upstream suspension member 34a to the downstream suspension member 34b, and therefore a preferred path through which the forces pass.In any case, the first 37a and the second beam 17b are symmetrical to each other with respect to a plane including the longitudinal axis X.

[0091] It can be foreseen that in any plane perpendicular to the longitudinal axis X between the upstream fixing element 34a and the fixing element 34b, the structural hood 35 has a cross-section with dimensions between 1200 and 1700 millimeters.

[0092] For example, in a plane perpendicular to the longitudinal axis X, the upstream ring 36 has a dimension between 1600 and 1700 millimeters, the downstream ring 38 has a dimension between 1200 and 1300 millimeters.

[0093] Such dimensions of the structural hood 35 contribute to enabling satisfactory rigidity.

[0094] This low flexibility reflects the high rigidity of the structural cover 35. The structural cover 35 is thus configured to form a predominant load path C between the upstream fixing element 34a and the downstream fixing element 34b. This load path C is represented by dashed lines in [Fig. 5]. The load path C does not pass through the core 3. It bypasses it, passing mainly through the first and second beams 37a, 37b, and secondarily through the lateral structural panels 37c, 37d. The high-pressure body is therefore no longer located on the predominant load path C, and is thus protected. Preferably, the first beam 37a, which extends close to the fixing elements 34a, 34b, is larger than the second beam 37b, because it is required to bear more loads due to its positioning.

[0095] Furthermore, the presence of the structural cover 35 allows for an increase in the distance between the upstream and downstream suspension planes. By protecting the high-pressure body, the structural cover 35 reduces the mass required for the housings and the pylon. Thus, despite the mass of the structural cover 35 itself, which is added to the overall mass of the turbomachine, the presence of the structural cover 35 results in an overall reduction in mass.

[0096] With this configuration, the core 3, and more specifically the high-pressure casing and the low-pressure turbine 8, is protected, as the impact of IP and gravitational loads is significantly reduced. The spacing of the mounting elements on the turbomachine 1 and the addition of the large-radius, high-stiffness structural cowling 35, which forms a predominant load path bypassing the core 3, allow for better management of IP loads and minimize problems related to the cantilevered position of the downstream part of the turbomachine 1 by increasing the lever arm between the upstream 33a and downstream 33b suspension members. Indeed, the relative displacements between the turbomachine 1 and the pylon 101 are significantly reduced, and the performance of the high-pressure casing is not affected. Similarly, the IP loads are not amplified. The proposed turbomachine thus allows for better load distribution.

[0097] A non-structural hood 40, partially illustrated in [Fig.6], can be arranged around the inlet housing, the structural hood and / or the outlet housing.

Claims

Demands

1. Aircraft turbomachine (1), the turbomachine (1) having a longitudinal axis (X) extending upstream to downstream in a flow direction of an airflow in the turbomachine (1), the turbomachine (1) comprising: - a core (3) including a low-pressure compressor (4), a high-pressure body, and a low-pressure turbine (8); - an upstream attachment element (34a) to a pylon (101); and - a downstream attachment element (34b) to the pylon (101); the upstream attachment element (34a) and the downstream attachment element (34b) being separated along the longitudinal axis (X) by at least 1000 millimeters.

2. Turbomachine (1) according to claim 1, wherein the turbomachine (1) is devoid of an intermediate fixing element to the pylon located axially between the upstream fixing element (34a) and the downstream fixing element (34b).

3. Turbomachine (1) according to any one of claims 1 and 2, comprising a structural hood (35) extending axially between the upstream fixing element (34a) and the downstream fixing element (34b), the structural hood (35) being configured to form a predominant force path (C) not passing through the core (3) between the upstream fixing element (34a) and the downstream fixing element (34b).

4. Turbomachine (1) according to claim 3, wherein the structural hood (35) comprises: - an upstream ring (36); - a downstream ring (38); and - a first beam (37a) and a second beam (37b) each connecting the upstream ring (36) and the downstream ring (38).

5. Turbomachine (1) according to claim 4, wherein the first beam (37a) and the second beam (37b) extend on either side of the longitudinal axis (X), the first beam (37a) extending for example above the longitudinal axis (X) and the second beam (37b) extending for example below the longitudinal axis (X).

6. Turbomachine (1) according to any one of claims 4 or 5, wherein the structural cowling (35) comprises at least one lateral structural panel (37c, 37d), extending between the upstream ring (36) and the downstream ring (38) and connecting the first beam (37a) and the second beam (37b).

7. Turbomachine (1) according to any one of claims 4 to 6, comprising from upstream to downstream: - an inlet casing (30) extending around the low-pressure compressor (4), the upstream ring (36) being axially adjacent to the inlet casing (30); - an outlet casing (31) extending around the low-pressure turbine (8), the outlet casing (31) comprising a turbine blade frame and an exhaust casing (31b), the downstream ring (38) extending around the turbine blade frame and / or the exhaust casing (31b).

8. Turbomachine (1) according to claim 7, wherein the inlet casing (30) has an upstream section (30a), an intermediate section (30b) and a downstream section (30c) adjacent to the upstream ring (36), the upstream fastening element (37a) being located on the upstream section (30a) or on the intermediate section (30b), and the downstream fastening element (37b) being located on the downstream ring (38).

9. Aircraft comprising: - a turbomachine (1) according to any one of the preceding claims, and - a suspension pylon (101) comprising an upstream suspension member (33a) and a downstream suspension member (33b) respectively fixed to the upstream attachment element (34a) and the downstream attachment element (34b) of the turbomachine (1).

10. Aircraft according to claim 9, wherein: - the upstream suspension member (33a) is configured to allow displacement of the turbomachine (1) in a first direction orthogonal to the longitudinal axis (X) and passing through the suspension pylon (101) and to resist forces in the longitudinal axis (X) and in a second direction orthogonal to the longitudinal axis (X) and to the first direction, and - the downstream suspension member (33b) is configured to resist: • forces in the first direction and in the second direction, and • a couple along the longitudinal axis (X), the downstream suspension element (33b) being further configured to allow displacements of the turbomachine (1) along the longitudinal axis (X).

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