Damper engine mounting connecting rod

Abstract

A damper engine mounting link is disclosed. An example engine assembly includes a core comprising a front and a rear section, the front section of which is connected to a fan hub frame, and the rear section of which is connected to a turbine rear frame or turbine center frame. The engine assembly also includes a damper link connecting the fan hub frame to a front mount of the aircraft, and a damper link connecting the turbine rear frame or turbine center frame to the aircraft mount.

Description

Technical Field

[0001] The present disclosure relates generally to aircraft engines and, more particularly, to damper engine mounting links. Background Art

[0002] Aircraft engines typically include mountings to couple the engine to a structure of the aircraft, such as a pylon. The mountings provide a linkage system between the engine and the pylon at the front of the engine (e.g., front mountings) and / or at the rear of the engine (e.g., rear mountings). The linkage system maintains the position of the engine relative to the aircraft. The linkage system affects the loads encountered by the engine, the alignment of the rotor and stator flow path steps throughout the engine, and the tip clearance between the blades and the engine casing. Summary of the invention

[0003] A damper engine mounting link is disclosed.

[0004] Some examples herein provide an engine assembly. The engine assembly includes a core, the core including a front portion and a rear portion, the front portion of the core coupled to a fan hub frame, the rear portion of the core coupled to a turbine aft frame or a turbine center frame. The engine assembly also includes a front mount coupling the fan hub frame to a first aircraft mount, and at least one damper link configured to couple the turbine aft frame or the turbine center frame to a second aircraft mount.

[0005] Some examples herein provide an apparatus including a damper link including a stiffness corresponding to at least one of a position of the damper link, a length of the damper link, or a rate of change of the length of the damper link. The apparatus also includes an engine coupled to a first longitudinal end of the damper link, and a nacelle or pylon connected to a wing of an aircraft, the nacelle or pylon coupled to a second longitudinal end of the damper link.

[0006] Certain examples herein provide an apparatus comprising: a first device for mounting a front portion of an engine to an aircraft; a second device for mounting a rear portion of the engine to the aircraft, the second device for mounting comprising a first device for damping to support the engine in response to movement of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] FIG1 shows a schematic cross-sectional view of a prior art example of an aircraft engine.

[0008] FIG. 2 shows a prior art example of an aircraft engine having a front mount.

[0009] FIG. 3 shows a prior art example of the aircraft engine of FIG. 2 under extreme operating conditions.

[0010] FIG. 4 shows a prior art example of an aircraft engine with forward and rearward mounts under the same extreme operating conditions as FIG. 3 .

[0011] Figure 5A -C shows an example damper link mount implemented in an example aircraft engine.

[0012] Fig. 6A -C shows the Figure 5A A plurality of damper link mounts implemented in an example aircraft engine of FIG.

[0013] Figure 7 Shows from Figure 5A -C aircraft rear front view looking from the rear of the engine.

[0014] Figure 8 Shows from Fig. 6A -C is a rear-forward view of an example aircraft engine looking from the rear.

[0015] Fig. 9 Shows Figure 5A -C, 6A-C, 7 and / or 8 and a first example associated connection.

[0016] Fig.10 Shows Figure 5A -C, 6A-C, 7 and / or 8 and a second example associated connection of a second example longitudinal end of the damper link.

[0017] The drawings are not drawn to scale. Generally, the same reference numerals will be used throughout the drawings and the accompanying written description to refer to the same or similar parts. As used herein, unless otherwise specified, connection references (e.g., attachment, connection, connection, and engagement) may include intermediate members between elements referenced by the connection reference and / or relative movement between those elements. Therefore, connection references do not necessarily infer that two elements are directly connected and / or are in a fixed relationship to each other. As used herein, a statement that any part is "in contact" with another part is defined to mean that there is no intermediate part between the two parts.

[0018] Unless otherwise specifically stated, descriptors such as "first", "second", "third", etc. are used herein without imposing or otherwise indicating any meaning of priority, physical order, arrangement in a list, and / or ordering in any manner, but are used only as labels and / or arbitrary names to distinguish elements for ease of understanding the disclosed examples. In some examples, the descriptor "first" may be used to refer to an element in the detailed description, while the same element may be referred to in the claims with different descriptors (such as "second" or "third"). In such cases, it should be understood that such descriptors are only used to clearly identify those elements that may, for example, share the same name in other ways. As used herein, "substantially real time" means occurring in a nearly instantaneous manner, recognizing that there may be real-world delays in computing time, transmission, etc. Therefore, unless otherwise stated, "substantially real time" means real time + / - 1 second. DETAILED DESCRIPTION

[0019] Damper engine mounting links are disclosed herein. Typically, engine mounts can be used to attach an engine of an aircraft to a structure, such as a pylon or nacelle. In some examples, a front mount is implemented to couple the front of the engine and / or a fan hub frame to the aircraft. In some examples, the front of the engine is coupled to the fan hub frame. In some examples, a housing surrounding the engine is also coupled to the fan hub frame. In some examples, in addition to the front mount, a rear mount is implemented to couple the rear of the engine (e.g., a portion coupled to a turbine rear frame or a turbine center frame) to the aircraft. The front mount and the rear mount provide a rigid and / or fixed connection between the engine and the pylon or nacelle. Typically, the front mount and / or the rear mount maintain the relative position of the engine during operation of the aircraft.

[0020] In some embodiments, the cantilever engine includes a front mounting and no rear mounting. No rear mounting allows the cantilever engine and the associated housing to pivot around the fan hub frame connected with the front mounting. In addition, no rear mounting prevents the cantilever engine from encountering bending moments. Therefore, compared with the engine including the rear mounting, the cantilever engine provides improved clearance between the turbine blades and the housing of the engine (e.g., the turbine blade tip and interstage seal clearance of the entire engine, gap closure characteristics, housing deformation). However, under extreme operating conditions (e.g., initial impact, hard landing, fan blade shedding resulting in wind milling imbalance (e.g., fan blade loss or damage during flight), etc.), the cantilever engine oscillates (e.g., swings, shakes, vibrates, etc.). The oscillation of the cantilever core can generate a large load on the front mounting and the associated fan hub frame. For example, when the cantilever engine pivots, the movement of the rear of the engine applies a significant load on the front mounting and / or the fan hub frame. In some cases, the loads generated on the front mounting and / or the fan hub frame can lead to some major design challenges that are difficult to overcome. Furthermore, the front mount requires a significant amount of additional material to withstand the loads that may be encountered during flight of the aircraft.

[0021] In some examples, in addition to the front mount, the engine also includes a rear mount to prevent the engine from shifting. However, under extreme operating conditions, the engine bends between the front mount and the rear mount. In addition, the bend between the front mount and the rear mount hinders the turbine blade tip and interstage seal clearance of the entire engine (e.g., causing the clearance to deform). For example, the engine bending can cause the blade tip to contact the engine case, which can cause the blade tip material of any blade at a particular rotor station to be lost. In addition, the lost blade tip material increases the blade tip clearance, which can cause engine performance loss, unacceptable operability characteristics, and / or temperature increases at certain locations throughout the engine.

[0022] The examples disclosed herein enable the damper engine mounting link to be configured to be coupled to an aircraft engine and an aircraft mounting. In some examples, the engine assembly includes a core (e.g., an engine core) having turbine blades (e.g., rotor blades, stator blades, etc.). In some examples, the front portion of the core is coupled to a fan hub frame and the rear portion of the core is coupled to a turbine rear frame. In some examples, the engine assembly includes a casing circumferentially surrounding the core. In some examples, the engine assembly includes a front mounting and a damper link (e.g., a squeeze film damper, a shock absorber, a rubber isolator, a wire mesh damper, a buffer, a magnetorheological damper, etc.). In some such examples, the first longitudinal end of the front mounting is coupled to the fan hub frame and the second longitudinal end of the front mounting is coupled to the aircraft. In some examples, the first longitudinal end of the damper link is coupled to the turbine rear frame and the second longitudinal end of the damper link is coupled to the aircraft. Therefore, the damper link is implemented as a rear mounting supporting the core. In some examples, the front mounting and the damper link are coupled to the pylon of the aircraft. In some examples, the front mount includes a front mount damper link.

[0023] In some examples, a pylon of the aircraft is connected to a nacelle of a shell circumferentially surrounding the core. In some such examples, at least one of the pylon or the nacelle is coupled to a front mount and a damper link. For example, a plurality of damper links may be positioned circumferentially around a turbine aft frame or a turbine center frame and / or a plurality of front mounts or front mounted damper links may be positioned circumferentially around a fan hub frame to couple the core to the aircraft. In some such examples, one or more of the plurality of damper links are coupled to the nacelle of the aircraft.

[0024] In some examples, the longitudinal end of the damper link includes a U-clip. In some such examples, the bearing of the engine core, the turbine rear frame, the fan hub frame, the pylon and / or the nacelle includes an opening corresponding to the opening of the U-clip. In some examples, the U-clip includes an opening having one or more slots. In some examples, a pin is inserted through the opening of the U-clip and the opening of the bearing of the engine core, the turbine rear frame, the fan hub frame, the pylon and / or the nacelle to couple the damper link to the engine core, the turbine rear frame, the fan hub frame, the pylon and / or the nacelle. In other examples, the bearing of the engine core, the turbine rear frame, the fan hub frame and / or the nacelle includes at least one U-clip, and the longitudinal end of the damper link includes an opening corresponding to the opening of at least one U-clip.

[0025] The damper link provides resistance to core movement based on flight conditions, the position of the core, the position of the fan hub frame, the position of the damper link, and / or the length of the damper link. In some examples, the damper link provides greater resistance to core movement under a first flight condition (e.g., fan blade shedding, initial impact, hard landing, etc. that causes wind milling imbalance) than a second flight condition (e.g., cruise flight, takeoff rotation, etc.). Therefore, the damper link allows the core to move like a typical cantilever core system, but provides damping resistance in response to inefficient and / or dangerous core movement. For example, the damper link can maintain a clearance distance between the turbine blades of the core and the casing surrounding them to improve the specific fuel consumption of the aircraft. In addition, the damper link maintains the alignment of the turbine blades to reduce the air flow path through the engine and further improve the efficiency of the engine.

[0026] The damper link prevents and / or reduces the oscillation of the core, which can cause destructive loads on the front mount and the associated fan hub frame in the cantilever engine. As a result, the front mount of the engine assembly experiences a reduced load compared to the front mount in the cantilever engine, and therefore requires less support material. In addition, the reduced load prevents the front mount and / or the fan hub frame from requiring heavier weight (i.e., more material) and more complex design. Therefore, the damper link alleviates some key design challenges, which can reduce the weight of the aircraft and improve the fuel efficiency of the engine. In some examples, the damper link allows the core to move with some damping to prevent the rigid rear mount from causing the shell to bend. In addition, the damper link can include an extension limit (e.g., a minimum extension length, a maximum extension length, etc.) to provide rigid support under certain flight conditions (e.g., fan blade shedding, hard landing, initial impact, etc.). Therefore, the extension limit depends on the area of ​​the engine and the damper link implementation.

[0027] In some examples, the damper link is an active damper link that is electronically controllable (e.g., by modulating an electric current). In some examples, the stiffness of the damper link is determined based on engine operating parameters. For example, the engine operating parameters may indicate instantaneous turbine blade tips and interstage seal clearances. In addition, the stiffness of the damper link is modulated to adjust the turbine blade tips and interstage seal clearances. In some examples, the stiffness of the damper link is predetermined for certain engine operating parameters based on flight simulations, wind tunnel testing, etc. In some examples, during flight of the aircraft, an open-loop control system compares the stiffness of the damper link with a predetermined stiffness for the current engine operating parameters. Furthermore, the control system adjusts the stiffness of the damper link in response to the stiffness of the damper link being different from the predetermined stiffness for the current engine operating parameters.

[0028] In some examples, during flight of the aircraft, the closed-loop control system modulates the stiffness of the damper link to determine a stiffness that optimizes an engine operating parameter. In some examples, the control system implements various stiffnesses of the damper link during flight of the aircraft to determine a stiffness associated with optimization of a flight parameter of interest. For example, the stiffness of the damper link is modulated to determine a setting that reduces vibration and / or noise output of the engine. In some examples, the stiffness of the damper link is determined in response to a signature flight condition encountered (e.g., cruise flight).

[0029] In some examples, the closed-loop control system determines the stiffness of the damper link based on parameters determined by sensors (e.g., strain gauges, pressure sensors, thermocouples, accelerometers, etc.). For example, an internal bearing, a housing, a front mount, a fan hub frame, a pylon, a nacelle, a core engine, a turbine aft frame, a turbine center frame, a flange, an internal structure, and / or a damper link may include a sensor. In some examples, the optimal output of the sensor is predetermined for certain flight conditions. As a result, the control system modulates the stiffness of the damper link to determine a setting that implements the optimal output of the sensor.

[0030] Referring now to the drawings, FIG1 is a schematic cross-sectional view of a prior art example of an aircraft engine 100 that may incorporate various examples disclosed herein. As shown in FIG1 , the aircraft engine 100 defines a longitudinal or axial centerline axis 102 extending therethrough for reference. Generally speaking, the aircraft engine 100 may include a core turbine or core turbine engine 104 disposed downstream of a fan section 106.

[0031] The core turbine engine 104 may generally include a substantially tubular casing 108 defining an annular inlet 110. The casing 108 may be formed of a plurality of segments. The casing 108 encloses, in serial flow relationship: a compressor section having a supercharger or low-pressure compressor 112 (“LP compressor 112”) and a high-pressure compressor 114 (“HP compressor 114”); a combustion section 116; a turbine section having a high-pressure turbine 118 (“HP turbine 118”) and a low-pressure turbine 120 (“LP turbine 120”); and an exhaust section 122. A high-pressure shaft or spool 124 (“HP shaft 124”) drivingly couples the HP turbine 118 and the HP compressor 114. A low-pressure shaft or spool 126 (“LP shaft 126”) drivingly couples the LP turbine 120 and the LP compressor 112. The LP shaft 126 may also be coupled to a fan shaft or spool 128 of the fan section 106. In some examples, LP shaft 126 may be directly coupled to fan shaft 128 (ie, a direct drive configuration). In an alternative configuration, LP shaft 126 may be coupled to fan shaft 128 via reduction gear 130 (ie, an indirect drive or gear drive configuration).

[0032] As shown in FIG1 , the fan section 106 includes a plurality of fan blades 132 (“fans” 132) coupled to and extending radially outward from the fan shaft 128. A first annular fan casing or first nacelle 134 circumferentially surrounds the fan section 106 and / or at least a portion of the core turbine engine 104. The aircraft engine 100 includes a second nacelle 135 opposite the first nacelle 134. The nacelles 134, 135 may be supported relative to the core turbine engine 104 by a front mount 136. In addition, a downstream section 138 of the nacelles 134, 135 may surround an outer portion of the core turbine engine 104 to define a bypass airflow passage 140 therebetween.

[0033] As shown in FIG. 1 , air 142 enters an intake or inlet portion 144 of the aircraft engine 100 during operation of the aircraft engine 100. A first portion 146 of the air 142 flows into the bypass flow passage 140, while a second portion 148 of the air 142 flows into the inlet 110 of the LP compressor 112. One or more successive stages of LP compressor stator vanes 150 and LP compressor rotor blades 152 (e.g., turbine blades) coupled to the LP shaft 126 progressively compress the second portion 148 of the air 142 flowing through the LP compressor 112 and directed to the HP compressor 114. Next, one or more successive stages of HP compressor stator vanes 154 and HP compressor rotor blades 156 coupled to the HP shaft 124 further compress the second portion 148 of the air 142 flowing through the HP compressor 114. This provides the compressed air 158 to the combustion section 116, where the compressed air 158 is mixed with fuel and combusted to provide combustion gases 160.

[0034] The combustion gases 160 flow through the HP turbine 118, wherein one or more successive stages of HP turbine stator blades 162 and HP turbine rotor blades 164 coupled to the HP shaft 124 extract a first portion of kinetic and / or thermal energy from the combustion gases 160. This energy extraction supports the operation of the HP compressor 114. The combustion gases 160 then flow through the LP turbine 120, wherein one or more successive stages of LP turbine stator blades 166 and LP turbine rotor blades 168 coupled to the LP shaft 126 extract a second portion of thermal and / or kinetic energy from the combustion gases 160. This energy extraction causes the LP shaft 126 to rotate, thereby supporting the operation of the LP compressor 112 and / or the rotation of the fan shaft 128. The combustion gases 160 then exit the core turbine 104 through the exhaust section 122 of the core turbine 104.

[0035] As with the aircraft engine 100, the core turbine 104 serves a similar purpose and sees similar circumstances in land-based turbines, turbojets, and non-ducted fan engines, where the ratio of the first portion 146 of air 142 to the second portion 148 of air 142 is less than that of turbofan engines, and where the fan section 106 does not have a nacelle 134, 135. In each of the turbofan engines, turbojets, and non-ducted engines, a speed reduction device (e.g., a speed reduction gearbox 130) may be included between any shaft and the spool. For example, the speed reduction gearbox 130 may be disposed between the LP shaft 126 and the fan shaft 128 of the fan section 106.

[0036] As depicted therein, the aircraft engine 100 defines an axial direction A, a radial direction R, and a circumferential direction C. Generally, the axial direction A extends generally parallel to the axial centerline axis 102 , the radial direction R extends orthogonally outward from the axial centerline axis 102 , and the circumferential direction C extends concentrically about the axial centerline axis 102 .

[0037] FIG. 2 shows a prior art example of a cantilever aircraft engine 200. In FIG. 2, the cantilever engine 200 includes the core turbine engine 104, the outer casing 108, the turbine blades 132, and the front mounting member 136 of FIG. 1. In addition, the cantilever engine assembly 200 includes a pylon 202, a fan hub frame 208, a turbine rear frame 210, and a turbine center frame. In FIG. 2, the pylon 202 includes a front attachment point 204 and a rear attachment point 206. In FIG. 2, the front mounting member 136 is coupled to the front attachment point 204 and the fan hub frame 208. In FIG. 2, the core turbine engine 104 is supported by the front mounting member 136. As a result, the front mounting member 136 and the fan hub frame 208 experience significant loads (e.g., destructive loads and / or other undesirable loads, etc.) during operation, as further discussed in conjunction with FIG. 3.

[0038] FIG. 3 shows the cantilever aircraft engine 200 of FIG. 2 in an example operation. In FIG. 3 , the core turbine engine 104 is coupled to the fan hub frame 208 and the turbine aft frame 210. In some examples, the front of the core 302 and the outer shell 108 are coupled to the fan hub frame 208. In FIG. 3 , the front mount 136 couples the fan hub frame 208 to the front attachment point of the pylon 204. In FIG. 3 , the rear of the core 304 is coupled to the turbine aft frame 210. In FIG. 3 , the turbine aft frame 210 is not coupled to the rear attachment point of the pylon 206, which allows the core turbine engine 104 to be cantilevered. In some examples, the core turbine engine 104 is fixed to the fan hub frame 208 together with the outer shell 108, thereby allowing the outer shell 108 to pivot around the fan hub frame 208 together with the core turbine engine 104. In addition, the gap between the core turbine engine 104 and the outer shell 108 is maintained because they share a similar pivot point. In some examples, maintaining a clearance between the core turbine engine 104 and the outer casing 108 maintains the efficiency of the cantilever aircraft engine 200 .

[0039] In example operations (e.g., fan blade shedding, wind milling loads, initial impact, etc.), the core turbine engine 104 oscillates (e.g., rocks, sways, etc.) while pivoting about the fan hub frame 208. In other words, as the rear portion of the core 304 oscillates within the nacelle (e.g., lower nacelle 134, upper nacelle 135, etc.) 306, the front portion of the core 302 maintains its relative position while pivoting with the fan hub frame 208. As a result, as the rear portion of the core 304 swings, the front mount 136 and / or the fan hub frame 208 encounter significant loads. In some examples, additional support materials and / or structures are implemented within the front mount 136 and / or the fan hub frame 208 to help ensure that they are able to support the significant loads encountered when the rear portion of the core 304 moves. In some examples, outer casing 108 contacts an inner wall of nacelle 306 , which causes structural damage at nacelle 306 , outer casing 108 , core turbine engine 104 , fan hub frame 208 , and / or forward mount 136 .

[0040] FIG. 4 shows a prior art example of an aircraft engine 400 that is supported by a rigid aft mount 402 in addition to being supported by the forward mount 136 of FIGS. 1 , 2 and 3 . In FIG. 4 , the core turbine engine 104 is coupled to the fan hub frame 208 and the turbine aft frame 210 . Specifically, the forward portion of the core 302 is coupled to the fan hub frame 208 and the aft portion of the core 304 is coupled to the turbine aft frame 210 . In FIG. 4 , the casing 108 is coupled to the fan hub frame 208 . In FIG. 4 , the forward mount 136 is coupled to the fan hub frame 208 and the forward attachment point of the pylon 204 . In FIG. 4 , the rigid aft mount 402 is coupled to the turbine aft frame and the aft attachment point of the pylon 206 .

[0041] In FIG. 4 , during extreme operations (e.g., fan blade shedding, hard landing, initial impact, etc. resulting in wind milling imbalance), the outer casing 108 oscillates about the fan hub frame 208. However, the rigid rear mount 402 and the front mount 136 maintain the relative position of the core turbine engine 104. As a result, the inner wall of the outer casing 108 may collide with the core turbine engine 104, which may damage the outer casing and / or the core turbine engine 104 and reduce the efficiency of the aircraft engine 400. In addition, the core turbine engine 104 experiences a bending moment between the front mount 136 and the rigid rear mount 402. Specifically, the rigid rear mount maintains a fixed position of the rear portion of the core 304, while the front portion of the core 302 pivots with the fan hub frame 208. In some examples, the fixed position of the rear portion of the core 304 hinders the outer casing 108 when the outer casing 108 pivots about the fan hub frame 208. In some examples, the gap between the core turbine engines 104 deforms as the outer casing 108 pivots, and the rigid aft mounts 402 maintain the relative positions of the core turbine engines 104 .

[0042] Figure 5A An example engine assembly 500 is shown that includes a rear mounted damper link (eg, rubber isolator, wire mesh damper, snubber, magnetorheological damper, squeeze film damper, etc.) 502. Figure 5A In the embodiment, the engine assembly includes an engine core 504 and a pylon 515 connected to the wing of the aircraft. In some examples, the pylon 515 is connected to a nacelle circumferentially surrounding the engine core 504, such as in combination with Fig. 6A -B is discussed further in Figure 5A In FIG. 5 , the engine core 504 includes a front portion 506 coupled to a fan hub frame 510 and a rear portion 508 coupled to a turbine center frame 511 and / or a turbine rear frame 512. Figure 5A In FIG. 5 , the fan hub frame 510 is coupled to a first end of the front mount 514 and the turbine aft frame 512 is coupled to a first end of the damper link 502. Figure 5A In FIG. 5 , the pylon 515 includes a front attachment point 516 and a rear attachment point 518. Figure 5A 5 , a front attachment point of the pylon 516 is coupled to the second end of the front mount 514 and a rear attachment point of the pylon 518 is coupled to the second end of the damper link 502 .

[0043] exist Figure 5AIn some examples, the damper link 502 provides resistance (e.g., damping) to the movement of the engine core 504 based on the flight conditions of the aircraft. In some examples, the damping is proportional to the rate of movement of the engine core 504. In addition, the damping provided by the damper link 502 can be based on the position of the damper link 502 relative to the engine core 504. Therefore, in some examples, the damper link 502 includes a stiffness proportional to the rate of movement of the engine core 504 and / or the direction of these movements to reduce, limit and / or prevent displacement of the engine core 504. Specifically, the damper link 502 can be active (e.g., a controllable damper link), which allows an open-loop or closed-loop control system to adjust the stiffness of the damper link 502 in response to the movement of the engine core 504 based on the rate of movement and / or position of the engine core 504. For example, the damper link 502 provides greater resistance to the movement of the engine core 504 under a first flight condition (e.g., fan blade shedding resulting in wind milling imbalance, initial impact, hard landing, maneuvering, etc.) than under a second flight condition (e.g., cruise flight). In some examples, the damper link 502 is passive (e.g., a squeeze film damper, a wire mesh damper, a shock absorber, etc.) and includes a stiffness based on the movement rate of the engine core 504. In some such examples, the stiffness of the damper link 502 is based on the area of ​​implementation and the movement rate typically encountered therein to provide a sufficient damping ratio. In some examples, the damper link 502 maintains a gap between the engine core 504 and the casing 108 to improve the specific fuel consumption, operability, and / or degradation characteristics of the engine assembly 500. In addition, the damper link 502 maintains the alignment of turbine blades (e.g., rotor blades, stator vanes, etc.) associated with the engine core 504.

[0044] exist Figure 5A In the embodiment of the present invention, the damper link 502 supports the engine core 504 to reduce the load on the fan hub frame 510 and / or the front mount 514. For example, the damper link 502 can maintain a fixed length at a maximum or minimum extension limit to provide rigid support to the engine core 504 during extreme flight conditions (e.g., wind milling imbalance, maneuvers, hard landings, initial impacts, etc.). As a result, the damper link 502 enables the front mount 514 and / or the fan hub frame 510 to be constructed with less support material (e.g., lighter weight) due to the reduced loads encountered by the front mount 514 and the fan hub frame 510.

[0045] In some examples, the damper link 502 is a controllable damper (e.g., a magnetorheological damper) with adjustable fluid stiffness. In some such examples, the controllable damper modulates the adjustable fluid stiffness based on flight conditions, the position of the fan hub frame 510, and / or the position of the engine core 504. In some examples, the controllable damper modulates its stiffness substantially in real time to control the clearance between the turbine blade tip and the interstage seal of the engine core 504. In some examples, the damper link 502 controls the vibration of the engine core 504. In some such examples, the damper link 502 controls the noise output of the engine assembly 500.

[0046] Figure 5B An example engine assembly 550 is shown, which includes a motor coupled to Figure 5A The damper link 520 of the turbine center frame 511 relative to the turbine rear frame 512 shown in FIG. Figure 5B In FIG. 5 , engine assembly 550 includes engine core 504 including front portion 506 coupled to fan hub frame 510 and rear portion 508 coupled to turbine center frame 511. Engine assembly 550 also includes pylon 515 including front attachment point 516 and rear attachment point 522. Figure 5B , a first end of the damper link 520 is coupled to the turbine center frame 511, and a second end of the damper link 520 is coupled to a rear attachment point of the pylon 522. In some examples, the damper link 520 spans the outer casing 108 to couple the engine core 504 to the pylon 515. For example, a strut may cover the damper link 520 and span the outer casing 108. Figure 5B , the damper link 520 provides stiffness to support the engine core 504 based on the length of the damper link 520 , the rate of change of the length of the damper link 520 , and / or determined operating conditions of the engine assembly 550 .

[0047] Figure 5C An example engine assembly 575 is shown that includes a front mounted damper link 524. Figure 5C 510, a first end of the front mounted damper link 524 is coupled to the front attachment point of the pylon 516 and a second end of the front mounted damper link 524 is coupled to the fan hub frame 510. In some examples, the front mounted damper link 524 provides stiffness to support the engine core 504 based on the position and / or movement of the fan hub frame 510. In some examples, the front mounted damper link 524 is smaller than the rear mounted damper links 502, 520. In some examples, the front mounted damper link 524 experiences lower temperatures than the rear mounted damper links 502, 520 and therefore may have reduced temperature capabilities compared to the rear mounted damper links 502, 520.

[0048] Fig. 6A An example engine assembly 600 is shown that includes a first damper link (e.g., upper damper link rear mount, damper link 502) 602 and a second damper link (e.g., lower damper link rear mount, damper link 502) 604 to couple an engine core 504 to an aircraft. Fig. 6A In the embodiment, the engine assembly 600 further includes a first front mounting member (eg, an upper front mounting member, the front mounting member 514, etc.) 610 and a second front mounting member (eg, a lower front mounting member, the front mounting member 514, etc.) 612. Fig. 6A In the embodiment, engine assembly 600 also includes an upper portion of nacelle 606 and a lower portion of nacelle 608. In some examples, the upper portion of nacelle 606 is connected to and / or includes a pylon 515 of the aircraft.

[0049] exist Fig. 6A , the front portion of the core 506 is mounted to the fan hub frame 510 by an upper front mount 610 and a lower front mount 612 coupled to the fan hub frame 510. Fig. 6A , upper front mount 610 is coupled to an upper portion of fan hub frame 510 and a front attachment point of pylon 515 and / or an upper portion of nacelle 606. Fig. 6A , lower front mount 612 is coupled to a lower portion of fan hub frame 510 and a lower portion of nacelle 608 .

[0050] exist Fig. 6A , the rear portion of the core 508 is mounted by an upper damper link rear mount 602 and a lower damper link rear mount 604. Specifically, the upper damper link rear mount 602 is coupled to the rear attachment point of the pylon 518 and / or the upper portion of the nacelle 606 and the upper portion of the turbine rear frame 512. In addition, the lower damper link rear mount 604 is coupled to the lower portion of the turbine rear frame 512 and the lower portion of the nacelle 608. In some examples, the upper damper link rear mount 602 and the lower damper link rear mount 604 are positioned relative to the engine core 504 in a radial direction, a lateral tilt direction, and / or an axial tilt direction. In some examples, additional damper links in the damper link 502 are coupled to a lateral portion of the turbine rear frame 512 and a lateral portion of the nacelle, such as in combination with the damper link. Figure 8 Further discussion.

[0051] exist Fig. 6A, the upper damper link aft mount 602 and the lower damper link aft mount 604 provide resistance to movement of the engine core 504 based on flight conditions, the position of the fan hub frame 510 (e.g., the angular orientation of the fan hub frame 510), the position of the turbine aft frame 512, the position of the engine core 504, the position of the upper damper link 602 or the lower damper link 604, and / or the length of the upper damper link 602 or the lower damper link 604. In some examples, the upper damper link aft mount 602 provides a first resistance to movement of the engine core 504, and the lower damper link aft mount 604 provides a second resistance to movement of the engine core 504 that is different from the first resistance. Fig. 6A , the upper damper link aft mount 602 and the lower damper link aft mount 604 maintain clearance between the turbine blades of the engine core 504 and the casing 108. In some examples, the upper damper link aft mount 602 and the lower damper link aft mount 604 maintain alignment of the turbine blades of the engine core 104 to maintain or improve the efficiency (e.g., specific fuel consumption, power output, etc.) of the engine core 104, which would otherwise be affected by flight conditions.

[0052] Figure 6B An example engine assembly 650 is shown, which includes an upper damper link 614 and a lower damper link 616 to couple to the Fig. 6A The turbine rear frame 512 shown in FIG. is opposite to the turbine center frame 511. Figure 6B In FIG. 6 , engine assembly 650 includes engine core 504 including front portion 506 coupled to fan hub frame 510 and rear portion 508 coupled to turbine center frame 511. Figure 6B 608. In some examples, a first end of a lower damper link 616 is coupled to the turbine center frame 511 and a second end of the lower damper link 616 is coupled to a lower portion of the nacelle 608. Additionally, a first end of an upper damper link 614 is coupled to the turbine center frame 511 and a second end of the upper damper link 614 is coupled to an upper portion of the nacelle 606. In some examples, the stiffness of the upper damper link 614 and / or the lower damper link 616 is based on the position of the turbine center frame 511. In some examples, the stiffness of the upper damper link 614 and the lower damper link 616 is modulated during flight of the aircraft to optimize engine operating parameters, such as turbine blade tips and interstage seal clearances of the engine core 504.

[0053] Figure 6C An example engine assembly 675 is shown, which includes an upper front mounted damper link 618 and a lower front mounted damper link 620. Figure 6CIn FIG. 5 , a first end of the upper forward mounted damper link 618 is coupled to an upper portion of the nacelle 606 , and a second end of the upper forward mounted damper link 618 is coupled to the fan hub frame 510 . Figure 6C 6 , a first end of lower front mount damper link 620 is coupled to a lower portion of nacelle 608, and a second end of lower front mount damper link 620 is coupled to fan hub frame 510. In some examples, in response to the stiffness of upper front mount damper link 618 and / or lower front mount damper link 620 being different than a predetermined stiffness for current engine operating parameters, the stiffness of upper front mount damper link 618 and / or lower front mount damper link 620 is adjusted.

[0054] Figure 7 Shows Figure 5A A first example cross-section AA of an example engine assembly 500 is shown. Figure 7 In FIG. 5 , the damper link 502 is coupled to the pylon 515 and the turbine rear frame 512. Figure 7 In FIG. 5 , the pylon 515 is connected to a nacelle (e.g., an upper portion of the nacelle 606, a lower portion of the nacelle 608, etc.) 702 that circumferentially surrounds the turbine center frame 511 and the turbine aft frame 512, and thus circumferentially surrounds the engine core 504. Figure 7 , the enlarged view 704 further illustrates a first longitudinal end of a damper link 706 coupled to the pylon 515 and a second longitudinal end of a damper link 708 coupled to the engine core 504 via the turbine aft frame 512 .

[0055] exist Figure 7 , the damper link 708 provides resistance to the movement of the rear portion of the engine core 508. In some examples, the rear portion of the engine core 508 moves in the radial direction R and / or the circumferential direction C. As such, the damper link 708 resists movement of the rear portion of the engine core 508 in the radial direction R and the circumferential direction C to maintain the clearance between the engine core 504 and the casing 108 and the alignment of the turbine blades of the engine core 104. In addition, the damper link 708 adjusts the resistance to the movement of the rear portion of the engine core 508 to allow the engine core 504 to pivot about the fan hub frame 510 with the casing 108 while preventing collisions between the engine core 504, the casing 108, the nacelle 702, and / or the pylon 515.

[0056] In some examples, the pylon 515, the nacelle 702, the engine core 504, the turbine center frame 511, and / or the turbine aft frame 512 include a U-shaped clip (not shown). In some such examples, the first and second longitudinal ends of the damper links 706, 708 include bearings having slots corresponding to the openings of the U-shaped clips, such as in combination with Fig. 9Further discussed. In some examples, the first and second longitudinal ends of the damper links 706, 708 include U-shaped clips. In some such examples, the pylon 515, the nacelle 702, the engine core 504, and / or the turbine aft frame 512 include a bearing having a slot corresponding to the opening of the U-shaped clip, such as in combination Fig.10 Further discussed. In some examples, the opening of the U-shaped clip includes one or more slots that align with the slots of the bearing. In addition, the engine assembly 500 includes a pin to be inserted through the opening of the U-shaped clip and the associated slot to couple the damper link 502 to the engine core 504 (via the turbine aft frame 512) and the pylon 515 or nacelle 702.

[0057] Figure 8 Shows Fig. 6A An example cross section BB of an example engine assembly 600 is shown. Figure 8 In FIG. 7 , the engine assembly 600 includes an upper damper link rear mount 602 and a lower damper link rear mount 604 to couple to the turbine rear frame 512 and the pylon 515 or nacelle 702. Figure 8 , the engine assembly 600 also includes a first transverse damper link rear mount 802 and a second transverse damper link rear mount 804 to couple to the nacelle 702 and the turbine rear frame 512. Thus, a plurality of damper links 602, 604, 802, 804 are circumferentially positioned around the turbine rear frame 512 to support the rear portion of the engine core 508.

[0058] exist Figure 8, the plurality of damper links 602, 604, 802, 804 each include a stiffness corresponding to a position of the damper link 602, 604, 802, 804, a length of the damper link 602, 604, 802, 804, and / or a rate of change (e.g., velocity) of the length of the damper link 602, 604, 802, 804. In some examples, movement of the engine core 504 results in a change in the position and / or length of the damper link 602, 604, 802, 804, which changes the stiffness of the damper link 602, 604, 802, 804. In some examples, the upper damper link rear mount 602 and the lower damper link rear mount 604 and the first and second lateral damper link rear mounts 802, 804 include first, second, third, and fourth stiffnesses, respectively, corresponding to the length and / or position of the damper links 602, 604, 802, 804. For example, under extreme operating conditions (e.g., wind milling imbalance, initial impact, hard landing, maneuvering, etc.) that cause the engine core 504 to oscillate, the stiffness of the damper links 602, 604, 802, 804 corresponds to the position of the engine core 504. Therefore, the stiffness of the damper links 602, 604, 802, 804 is modulated based on the position of the engine core 504, the movement of the engine core 504, and / or the flight conditions of the aircraft when the engine core 504 oscillates to reduce the load on the front mounts 610, 612 and the fan hub frame 510.

[0059] Fig. 9 Shows Figure 5A -C, 6A-C, 7, and / or 8 and an associated connection 900. In some examples, the associated connection 900 is between the damper link 902 and the turbine center frame 511, the turbine aft frame 512, the fan hub frame 510, the pylon 515, and / or the nacelle 702. For example, when the damper link 902 is implemented as an aft mount attachment of the engine core 504, the first longitudinal end of the damper link 906 is coupled to the turbine aft frame 512, and the second longitudinal end of the damper link 907 is coupled to the pylon 515 or the nacelle 702.

[0060] Fig. 9 An enlarged view 904 of a first longitudinal end of a damper link 906 is included. Fig. 9 , the first longitudinal end of the damper link 906 includes a slot (e.g., an opening) 908. In some examples, the first longitudinal end of the damper link 906 is positioned within a U-shaped clip 910. In some examples, the turbine aft frame 512, the fan hub frame 510, the pylon 515, and / or the nacelle 702 include a U-shaped clip 910. Fig. 9In the embodiment, the U-shaped clamp 910 includes an opening having one or more slots 912. In some examples, the one or more slots 912 of the U-shaped clamp 910 are aligned with the slot 908 of the first longitudinal end of the damper link 906. Fig. 9 , a pin 914 is inserted through one or more slots 912 of the U-clip 910 and a slot 908 of the first longitudinal end of the damper link 906 to couple the damper link 902 to the fan hub frame 510 , the turbine aft frame 512 , the pylon 515 and / or the nacelle 702 .

[0061] Fig.10 Shows Figure 5A -C, 6A-C, 7 and / or 8 and an associated connection 1000. In some examples, the associated connection 1000 is between the damper link 1002 and the turbine center frame 511, the turbine aft frame 512, the fan hub frame 510, the pylon 515 and / or the nacelle 702. For example, when the damper link 1002 is implemented as a forward mounting attachment of the engine core 504, the first longitudinal end of the damper link 1006 is coupled to the fan hub frame 510, and the second longitudinal end of the damper link 1007 is coupled to the pylon 515 or the nacelle 702.

[0062] Fig.10 An enlarged view 1004 of a first longitudinal end of a damper link 1006 is included. Fig.10 In the embodiment, the first longitudinal end of the damper link 1006 includes a U-shaped clip 1008. Fig.10 In the embodiment, the U-clip 1008 includes an opening having one or more slots 1010. Fig.10 In some examples, the fan hub frame 510, the turbine aft frame 512, the pylon 515, and / or the nacelle 702 include the bearing 1012. In some examples, the U-shaped clip 1008 is positioned around the bearing 1012 and the one or more grooves 1010 of the U-shaped clip 1008 are aligned with the groove 1014 of the bearing 1012. Fig.10 , pin 1016 is inserted through one or more slots 1010 of clevis 1008 and slot 1014 of bearing 1012 to couple damper link 1002 to fan hub frame 510 , turbine aft frame 512 , pylon 515 , and / or nacelle 702 .

[0063] "Include" and "comprising" (and all their forms and tenses) are used herein as open-ended terms. Thus, whenever a claim employs any form of "include" or "comprising" (e.g., includes, contains, has, etc.) as a preamble or in any type of claim recitation, it should be understood that additional elements, terms, etc. may be present without exceeding the scope of the corresponding claim or recitation. As used herein, when the phrase "at least" is used as a transitional term, such as in the preamble of a claim, it is open-ended in the same manner as the terms "include" and "comprising" are open-ended. The term "and / or" when used in the form of, for example, A, B, and / or C, refers to any combination or subset of A, B, C, such as (1) A alone, (2) B alone, (3) C alone, (4) A and B, (5) A and C, (6) B and C, and (7) A with B and C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A and B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A or B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. As used herein in the context of describing the performance or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A and B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing the performance or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A or B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, and (3) at least one A and at least one B.

[0064] As used herein, singular references (e.g., "one," "a," "first," "second," etc.) do not exclude the plural. As used herein, the term "one" or "an" entity refers to one or more of the entity. The terms "one" (or "a"), "one or more," and "at least one" may be used interchangeably herein. In addition, although listed separately, multiple modes, elements, or method actions may be implemented by, for example, a single unit or processor. In addition, although individual features may be included in different examples or claims, these may be combined, and inclusion in different examples or claims does not mean that the combination of features is not feasible and / or advantageous.

[0065] As can be appreciated from the foregoing, an example damper engine mounting link has been disclosed that provides resistance to the rate of movement of the engine core based on flight conditions, the position of the engine core, the movement of the engine core, the position of the fan hub frame, the position of the damper link, and / or the length of the damper link. In addition, the example damper engine mounting link maintains clearance between turbine blades of the engine core and an associated casing to increase the efficiency of the engine by, for example, reducing clearance variations associated with flow streams that generate thrust. In addition, the example damper engine mounting link maintains alignment of the turbine blades of the core to increase the efficiency of the engine.

[0066] Although certain example methods, apparatus, and articles of manufacture are disclosed herein, the scope of coverage of this patent is not limited thereto. On the contrary, this patent covers all methods, apparatus, and articles of manufacture that fully fall within the scope of the claims of this patent.

[0067] The following claims are incorporated into the Detailed Description by reference, with each claim standing on its own as a separate embodiment of the present disclosure.

[0068] Further aspects of the invention are provided by the subject matter of the following clauses:

[0069] 1. An engine assembly comprising: a core, the core comprising a front portion and a rear portion, the front portion of the core being coupled to a fan hub frame, the rear portion of the core being coupled to at least one of a turbine aft frame or a turbine center frame; a front mount coupling the fan hub frame to a first aircraft mount; and a damper link coupling the turbine aft frame or the turbine center frame to a second aircraft mount.

[0070] 2. An engine assembly according to any preceding clause, further comprising an active control system to provide stiffness to the damper link based on at least one of flight conditions or sensor input.

[0071] 3. An engine assembly according to any preceding clause, wherein the active control system modulates the stiffness of the damper linkage.

[0072] 4. An engine assembly according to any preceding clause, wherein the damper link is at least one of a squeeze film damper, a wire mesh damper, a shock absorber, or a controllable damper including an adjustable fluid stiffness.

[0073] 5. An engine assembly according to any preceding clause, wherein the controllable damper modulates the adjustable fluid stiffness based on at least one of flight conditions, a position of the fan hub frame, a position of the turbine aft frame, a position of the turbine center frame, or a position of the core.

[0074] 6. The engine assembly according to any preceding clause, further comprising a pylon connected to a nacelle, the nacelle circumferentially surrounding the core, the front mount and the damper link, the damper link being configured to couple to at least one of the pylon or the nacelle.

[0075] 7. An engine assembly according to any preceding clause, further comprising a pin inserted through an opening at a longitudinal end of the damper link and an opening of the fan hub frame, the turbine aft frame, the turbine center frame, the pylon or the nacelle to couple the damper link to the fan hub frame, the turbine aft frame, the pylon or the nacelle.

[0076] 8. An engine assembly according to any preceding clause, wherein the damper link is positioned in a radial direction, a laterally inclined direction or an axially inclined direction relative to the core.

[0077] 9. An engine assembly according to any preceding clause, wherein the damper link comprises a plurality of damper links positioned circumferentially around the turbine aft frame.

[0078] 10. The engine assembly of any preceding clause, wherein the damper link is a first damper link, and wherein the front mount includes a second damper link to couple the front mount to the aircraft.

[0079] 11. An engine assembly according to any preceding clause, wherein the damper link maintains clearance between turbine blades of the core and a casing surrounding the core.

[0080] 12. An engine assembly according to any preceding clause, wherein the damper link maintains alignment of turbine blades of the core.

[0081] 13. An engine assembly according to any preceding clause, wherein the damper link controls vibrations of the core.

[0082] 14. An engine assembly according to any preceding clause, wherein the damper linkage controls the noise output of the engine assembly.

[0083] 15. An apparatus comprising: a damper link, the damper link including a stiffness corresponding to at least one of a position of the damper link, a length of the damper link, or a rate of change of the length of the damper link; an engine, the engine coupled to a first longitudinal end of the damper link; and a nacelle or a pylon, the nacelle or the pylon being connected to a wing of an aircraft, the nacelle or the pylon being coupled to a second longitudinal end of the damper link.

[0084] 16. The apparatus of any preceding clause, wherein at least one of the first longitudinal end of the damper link, the second longitudinal end of the damper link, the engine, the nacelle, or the pylon comprises a U-clip.

[0085] 17. An apparatus according to any preceding clause, wherein at least one of the first longitudinal end of the damper link, the second longitudinal end of the damper link, the engine, the nacelle or the pylon includes a slot corresponding to an opening of the U-clip.

[0086] 18. Apparatus according to any preceding clause, wherein the opening of the U-shaped clamp comprises one or more slots.

[0087] 19. Apparatus according to any preceding clause, further comprising a pin to be inserted through said opening of said clevis and said slot to couple said damper link to said engine and said nacelle or said pylon.

[0088] 20. An apparatus comprising: a first device for mounting a front portion of an engine to an aircraft; and a second device for mounting a rear portion of the engine to the aircraft, the second device for mounting including a first device for damping to support the engine in response to a position of the engine or movement of the engine.

[0089] 21. Apparatus according to any preceding clause, wherein said first means for mounting comprises second means for damping.

[0090] 22. Apparatus according to any preceding clause, wherein the stiffness of the first means for damping is modulated based on the position of the engine, the movement of the engine or flight conditions of the aircraft.

Claims

1. An engine component, characterized in that: include: a core, the core comprising a front portion and a rear portion, the front portion of the core coupled to a fan hub frame, the rear portion of the core coupled to at least one of a turbine aft frame or a turbine center frame; a front mounting member coupling the fan hub frame to a pylon of an aircraft; as well as at least one damper link, the at least one damper link including a first longitudinal end and a second longitudinal end, the first longitudinal end being configured to be directly coupled to the turbine aft frame or the turbine center frame, the second longitudinal end being configured to be directly coupled to the pylon, Wherein a stiffness of the at least one damper link is determined based on flight conditions of the aircraft.

2. The engine assembly according to claim 1, characterized in that Further included is a control system to modulate the stiffness of the at least one damper link based on at least one of flight conditions or sensor input.

3. The engine assembly according to claim 2, characterized in that Wherein the control system modulates the stiffness of the at least one damper link based on at least one of a rate of movement of the core or a position of the core.

4. The engine assembly according to claim 1, characterized in that Wherein the at least one damper link is at least one of a squeeze film damper, a wire mesh damper, a shock absorber, or a controllable damper including an adjustable fluid stiffness.

5. The engine assembly according to claim 4, characterized in that Wherein the controllable damper modulates the adjustable fluid stiffness based on at least one of flight conditions, a position of the fan hub frame, a position of the turbine aft frame, a position of the turbine center frame, or a position of the core.

6. The engine assembly according to claim 1, characterized in that Wherein the at least one damper link is a first damper link, and wherein the engine assembly further comprises a second damper link configured to couple the turbine aft frame or the turbine center frame to a nacelle, the nacelle circumferentially surrounding the core.

7. The engine assembly according to claim 6, characterized in that Further comprising a pin inserted through an opening at a longitudinal end of the second damper link and an opening of the fan hub frame, the turbine aft frame, the turbine center frame, the pylon or the nacelle to couple the second damper link to the fan hub frame, the turbine aft frame, the pylon or the nacelle.

8. The engine assembly according to claim 1, characterized in that Wherein the at least one damper link is positioned in a radial direction, a transversely inclined direction or an axially inclined direction relative to the core.

9. The engine assembly according to claim 8, characterized in that Wherein the at least one damper link comprises a plurality of damper links positioned circumferentially around the turbine rear frame.

10. The engine assembly according to claim 1, characterized in that Wherein the at least one damper link is a first damper link, and wherein the front mount includes a second damper link to couple the front mount to the pylon.

11. The engine assembly according to claim 1, characterized in that The at least one damper link maintains a gap between turbine blades of the core and a casing surrounding the core.

12. The engine assembly according to claim 1, characterized in that Wherein the at least one damper link maintains alignment of turbine blades of the core.

13. The engine assembly according to claim 1, characterized in that Wherein the at least one damper link controls vibrations of the core.

14. The engine assembly according to claim 13, characterized in that Wherein the at least one damper link controls noise output of the engine assembly.

15. An engine device, characterized in that: include: a damper link including a fluid to provide a stiffness corresponding to at least one of a position of the damper link, a length of the damper link, or a rate of change of the length of the damper link; an engine coupled to a first longitudinal end of the damper link; as well as a nacelle or a pylon, the nacelle or the pylon being connected to a wing of the aircraft, the nacelle or the pylon being directly coupled to the second longitudinal end of the damper link, The stiffness of the damper link is determined based on the flight conditions of the aircraft.

16. The device according to claim 15, characterized in that Wherein at least one of the first longitudinal end of the damper link, the second longitudinal end of the damper link, the engine, the nacelle, or the pylon comprises a U-clip.

17. The device according to claim 16, characterized in that Wherein at least one of the first longitudinal end of the damper link, the second longitudinal end of the damper link, the engine, the nacelle, or the pylon comprises a slot corresponding to an opening of the U-clip.

18. The device according to claim 17, characterized in that The opening of the U-shaped clip includes one or more slots.

19. The device according to claim 17, characterized in that Further included is a pin to be inserted through the opening of the clevis and the slot to couple the damper link to the engine and the nacelle or the pylon.

20. An engine device, characterized in that: include: a first device for mounting a front portion of the engine to an aircraft; as well as second means for mounting a rear portion of the engine to the aircraft, the second means for mounting comprising means for damping to adjust the support provided to the engine in response to movement of the engine, the means for damping comprising means for coupling directly to the rear portion of the engine and to a pylon of the aircraft, Wherein the stiffness of the means for damping is determined based on flight conditions of the aircraft.

21. The device according to claim 20, characterized in that wherein said means for damping is second means for damping, and wherein said first means for mounting comprises first means for damping.

22. The device according to claim 20, characterized in that Wherein the stiffness of the means for damping is modulated based on the position of the engine, the movement of the engine or the flight conditions of the aircraft.

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