Aircraft rotor assembly and aircraft

Abstract

The invention relates to an aircraft rotor assembly (100) comprising a rotor head (110) configured to rotate with a rotor axis (RA); a first rotor blade mounting (120), connecting the rotor head (110) and a first rotor blade ( 150); a second rotor blade mounting (220), connecting the rotor head (110) and a second rotor blade (250); each rotor blade mounting (120,220) comprises a lead-lag hinge with a lead-lag axis (LLA) that is substantially parallel to the rotor axis (RA), a flap hinge with a flap axis (FA) that is substantially perpendicular to and intersects the lead-lag axis (LLA), and a blade pitch hinge with a blade pitch axis (BPA) that is substantially perpendicular to and intersects both the lead-lag axis (LLA) and the flap axis (FA); a damping element (180) having a first damping axis (DAI) including a first end (182) and a second end (184), the first end (182) being attached to the first rotor blade mounting (120) the second end (184) being attached to the second rotor blade mounting (220); wherein in a plan view the first end (182) of the damping element (180) is attached to a first attachment point (125) offset the blade pitch axis (BPA1) of the first rotor blade (150) and the second end (184) of the damping element (310) is attached to a second attachment point (245) on the blade pitch axis (BPA2) of the second rotor blade (250); wherein a first distance from the rotor axis (RA) to the first attachment point (125) is bigger than a second distance from the rotor axis (RA) to the second attachment point (245).

Description

[0001] Aircraft rotor assembly and Aircraft

[0002] Technical Field

[0003] The invention relates to an aircraft rotor assembly comprising a rotor head configured to rotate with a rotor axis, a first rotor blade mounting, connecting the rotor head and a first rotor blade, a second rotor blade mounting, connecting the rotor head and a second rotor blade, each rotor blade mounting comprises a lead-lag hinge with a lead-lag axis that is substantially parallel to the rotor axis, a flap hinge with a flap axis that is substantially perpendicular to and intersects the lead-lag axis, and a blade pitch hinge with a blade pitch axis that is substantially perpendicular to and intersects both the lead-lag axis and the flap axis, a first damping element having a first damping axis including a first end and a second end, the first end being attached to the first rotor blade mounting the second end being attached to the second rotor blade mounting.

[0004] Background Art

[0005] Rotorcraft rotors comprising a hub driven in rotation about an axis of rotation by a drive shaft or outlet shaft from a power transmission gearbox, and also having a plurality of blades fixed to the hub. When a helicopter is flying horizontally, or hovering in the wind, different relative wind speeds cause the rotating blades to experience different horizontal static and dynamic forces throughout each rotation. For example, during forward flight, when the blade is advancing, it is encountering a larger relative air speed than when the blade is retreating. Accordingly, each blade experiences large static moments and varying moments in the leading and lagging directions.

[0006] Aeroelasticity of a helicopter rotor is one of the most challenging in all engineering fields. Loads and vibrations are continuously introduced in the airframe, causing problems of stability and fatigue. Each component stability can be studied in isolation through different kind of analysis, making use of ground vibration test (GVT) for airframe, and of various finite elements modeling techniques and of aerodynamic models for structural elements. Even when every single studied component is determined to be stable, instabilities can arise due to the coalescence of particular modes after substructures coupling. Ground resonance is the perfect example of substructures coupling instabilities. It is a dynamic instability due to the interaction of the blade lead-lag motion with a natural frequency (e.g. roll / pitch motion) of the structures that sustain the rotor, typically the fuselage on the landing gear.

[0007] A resonance is possible when the rotating lag frequency is below the rotation speed frequency of the rotor, as for articulated (soft-in plane) rotors.

[0008] A similar dynamic instability can occur in flight as well, due to the coupling of the lead-lag cyclic modes with the low frequency airframe modes (typically due to weight saving in structure design) related to flight dynamics, in this case it is called "air resonance". Lead-lag motion of the rotor is always a function of the rotor angular speed. This typically gives a critical range of rotor speeds that should be studied to guarantee that the helicopter is free from ground and air resonance.

[0009] In addition to ground and air resonance, the lead-lag blade motion is coupled collectively with the torsional dynamics of the engine and drive train system of the helicopter, which can potentially be unstable for high gains in the fuel control system. The high gains in fuel control system are typically advantageous from rotor reactiveness point of view, especially in situations of rapid power and lift demands. Therefore, an efficient and safe main rotor design has to provide a solution for the stability problems ground resonance, air resonance and torsional stability of the dynamic system.

[0010] In the prior art two different types of architecture are known. For example there are blade holder connected to the hub by a lead-lag damper called B2H (Blade to Hub) which can be viscous- elastic or hydraulic or hydro elastic. Further there are blade holder connected together by lead- lag dampers around but not to the hub called inter-blades which can be viscous-elastic or hydraulic or hydro elastic.

[0011] However, these known architectures are creating several couplings between the lag and the flap and the pitch of the blade, and both these effects generate high static and dynamic damper displacements. Therefore, some flight regimes require very high static and dynamic damper displacements which can quickly damage the damper and compromise its life. Summary of the invention

[0012] It is the object of the invention to define an efficient and safe main rotor design that provides a solution to stability problems of ground resonance, air resonance and torsional stability of the dynamic system.

[0013] The solution of the invention is specified by the features of claim 1. According to the invention an aircraft rotor assembly comprises a rotor head configured to rotate with a rotor axis, a first rotor blade mounting, connecting the rotor head and a first rotor blade, a second rotor blade mounting, connecting the rotor head and a second rotor blade, each rotor blade mounting comprises a lead-lag hinge with a lead-lag axis that is substantially parallel to the rotor axis, a flap hinge with a flap axis that is substantially perpendicular to and intersects the lead-lag axis, and a blade pitch hinge with a blade pitch axis that is substantially perpendicular to and intersects both the lead-lag axis and the flap axis, a first damping element having a first damping axis including a first end and a second end, the first end being attached to the first rotor blade mounting the second end being attached to the second rotor blade mounting, wherein in a plan view the first end of the first damping element is attached to a first attachment point offset the blade pitch axis of the first rotor blade and the second end of the first damping element is attached to a second attachment point on the blade pitch axis of the second rotor blade, wherein a first distance from the rotor axis to the first attachment point is bigger than a second distance from the rotor axis to the second attachment point.

[0014] A damping element in the understanding of the invention can also be understood as a lead-lag damper.

[0015] Therewith the lead-lag damper is attached at each end to two adjacent blades. Since the first attachment point is offset the blade pitch axis of the first rotor blade and the second attachment point is on the blade pitch axis of the second rotor blade there is an assymetrical chord offset arrangement of the lead-lag damper. Additionally the first distance from the rotor axis to the first attachment point is bigger than a second distance from the rotor axis to the second attachment point. Thereby there is an assymetrical radial offset arrangement of the lead-lag damper. As a consequence the invention combines an assymetrical radial offset and assymetrical chord offset.

[0016] This results in the following technical advantages. The interblade lead-lag damper is connecting two adjacent blades with an asymmetrical radial offset. This provides stiffness and damping in the collective drag mode (oscillations in phase) of all blades which are linked in torsion with the entire power chain. This includes the rotor mast, the main gearbox, the tail drive shafts, the tail gearbox and tail rotor and engine. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system.

[0017] The assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades makes it possible to define a sufficient first lever arm and second lever arm (LAI and LA2) to adjust the cyclic lead-lag frequency in order to avoid once projected into the fixed frame of the fuselage from coupling with the natural modes of the fuselage. Therewith fuselage modes on ground for ground resonance and fuselage modes in flight for air resonance can be avoided. According to an advantageous embodiment of the invention in a plan view the ratio of a first lever (LAI) to a chord offset (CO) is less or equal to 2. In other words, the ratio of a distance between the lead lag axis LLA1 and the first damping axis (DAI) to a chord offset (CO) is less or equal to 2.

[0018] According to the invention a third rotor blade mounting, connecting the rotor head and a third rotor blade, wherein a second damping element having a second damping axis including a first end and a second end, the first end being attached to the third rotor blade mounting the second end being attached to the first rotor blade mounting.

[0019] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0020] According to an advantageous embodiment of the invention in a plan view the first end of the second damping element is attached to a first attachment point offset the blade pitch axis of the third rotor blade and the second end of the second damping element is attached to a second attachment point on the blade pitch axis of the first rotor blade.

[0021] The technical advantages are comparable to those of the previous embodiment. In principle, the advantages exist for each interblade lead-lag damper between two adjacent rotor blades, regardless of the total number of rotor blades. The invention combines an assymetrical radial offset and assymetrical chord offset at each interblade lead-lag damper. This results in stiffness and damping in the collective drag mode (oscillations in phase) of all blades which are linked in torsion with the entire power chain. This includes the rotor mast, the main gearbox, the tail drive shafts, the tail gearbox and tail rotor and engine. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system. The assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades makes it possible to define a sufficient first lever arm and second lever arm (LAI and LA2) to adjust the cyclic lead-lag frequency in order to avoid once projected into the fixed frame of the fuselage from coupling with the natural modes of the fuselage.

[0022] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0023] According to the invention a first lever between the lead lag axis LLA1 and the first damping axis is bigger than a second lever between the lead lag axis and the second damping axis. In other words, the distance between the lead lag axis LLA 1 and the first damping axis is greater than the distance between the lead lag axis and the second damping axis.

[0024] The technical advantages are similar to those of the previous embodiments. The interblade lead- lag damper is connecting two adjacent blades with an asymmetrical radial offset. This provides stiffness and damping in the collective drag mode (oscillations in phase) of all blades which are linked in torsion with the entire power chain. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system.

[0025] This effect is additionally increased by the assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades. The greater the chord offset at the first attachment point, the greater the difference between the first lever and the second lever.

[0026] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0027] According to an advantageous embodiment of the invention a ratio between LAI to LA2 is bigger than 1.5. This means that the ratio between the distance between the lead lag axis LLA1 and the first damping axis and the distance between the lead lag axis LLA1 and the second damping axis is bigger than 1.5. This ratio has proven to be particularly efficient in terms of the optimal effect of the asymmetrical radial offset and the asymmetrical chord offset.

[0028] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0029] According to the invention a distance between the lead lag axis and the second attachment point is smaller than a distance between the lead lag axis and an intersection point between the first damping axis and the blade pitch axis.

[0030] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0031] According to a further embodiment of the invention a rotational movement of the first damping element is blocked. This means that at least one degree of freedom of the lead lag damper is blocked at the first attachment point and / or at the second attachment point. For example, the lead lag damper may only have two degrees of freedom at the first attachment point, while there are three degrees of freedom at the second attachment point. According to another example, the lead lag damper may have three degrees of freedom at the first attachment point, while there are only two degrees of freedom at the second attachment point. According to a further example, the lead lag damper may have two degrees of freedom at the first attachment point and may have two degrees of freedom at the second attachment point. This reduces wear on the lead bearing damper and increases its service life.

[0032] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0033] According to the invention the first attachment point comprises a spherical function bearing for mounting the first end of the first damping element and the second attachment point comprises a spherical function bearing for mounting the second end of the first damping element. On the one hand, this has the technical advantage of reducing wear on the lead lag steamers. On the other hand the adaptability of the lead lag damper is sure to changes of the pitch angle of the rotor blade.

[0034] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0035] According to a further an advantageous embodiment of the invention the ratio of a first distance oc2, from the first attachment point 325 of the third rotor blade mounting 320 to the flap axis FA1 of the first rotor blade mounting 120, to a second distance / ?2, fromthe intersection point between the blade pitch axis BPA1 of the first rotor blade mounting 120 and the second damping axis DA2 to the flap axis FA1 of the first rotor blade mounting 120, is between 0 and 0,3.

[0036] Due to the assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades just the first attachment point is placed offset the blade pitch axis. This means that the intersection point between the blade pitch axis of the first rotor blade mounting and the second damping axis corresponds to the second attachment point. This achieves the technical advantage that the first rotor blade is made free of flap / lead-lag coupling. A further advantage results from minimizing the susceptibility to the air resonance phenomenon but also helps for ground resonance. According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0037] According to a further an advantageous embodiment of the invention the ratio of a first distance o from the second attachment point of the second rotor blade mounting to the flap axis of the first rotor blade mounting, to a second distance from the flap axis the first rotor blade mounting to the intersection point between the blade pitch axis of the first rotor blade mounting and the first damping axis, is between 0 and 0,3. The technical advantages are similar to the advantages of the previous embodiment. On the one hand the first rotor blade is made free of flap / lead-lag coupling. On the other hand an advantage results from minimizing the susceptibility to the air resonance phenomenon but also helps for ground resonance.

[0038] According the invention these features can be realized in a figurative sense for all rotor blade mountings of the aircraft rotor assembly. For the first rotor blade mounting, the second rotor blade mounting and the third rotor blade mounting. This can also be realized with a fourth rotor blade mounting and a fifth rotor blade mounting, depending on how many rotor blade mountings are realized.

[0039] According to the invention each damping element has a nominal stiffness between 0,5 kN / mm and 10 kN / mm. In particular according to the invention each damping element has a nominal stiffness between 1 kN / mm and 4 kN / mm.

[0040] According to a further an advantageous embodiment of the invention each damping element is fully elastomeric.

[0041] Advantageously the lead-lag hinge with a lead-lag axis, the flap hinge with a flap axis and the blade pitch hinge with a blade pitch axis are arranged by means of a single laminated spherical abutment.

[0042] According to an advantageous embodiment of the invention each rotor blade is attached to each corresponding rotor blade mounting by at least two bolts. Advantageously according to the invention the aircraft rotor assembly comprises at least three rotor blade mountings. Especially preferred the aircraft rotor assembly comprises five rotor blade mountings. As a result, there is a lead lag damper between each of the rotor blade mountings, which is why the entire aircraft rotor assembly has a total of five lead lag dampers.

[0043] Another solution of the invention is specified by the features of claim 16. According to the invention an aircraft comprises an aircraft rotor assembly according to one of the preceding embodiments. There are similar advantages to the preceding solution of independent claim 1. The lead-lag damper is attached at each end to two adjacent blades of an aircraft having an aircraft rotor assembly according to one of the preceding embodiments.

[0044] The first attachment point is offset the blade pitch axis of the first rotor blade and the second attachment point is on the blade pitch axis of the second rotor blade. There is an assymetrical chord offset arrangement of the lead-lag damper. Additionally the first distance from the rotor axis to the first attachment point is bigger than a second distance from the rotor axis to the second attachment point. Thereby there is an assymetrical radial offset arrangement of the lead-lag damper. As a consequence the invention combines an assymetrical radial offset and assymetrical chord offset for all lead lag dampers.

[0045] The interblade lead-lag damper is connecting two adjacent blades with an asymmetrical radial offset. This provides stiffness and damping in the collective drag mode (oscillations in phase) of all blades which are linked in torsion with the entire power chain. This includes the rotor mast, the main gearbox, the tail drive shafts, the tail gearbox and tail rotor and engine. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system.

[0046] The assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades makes it possible to define a sufficient first lever arm and second lever arm (LAI and LA2) to adjust the cyclic lead-lag frequency in order to avoid once projected into the fixed frame of the fuselage from coupling with the natural modes of the fuselage. Therewith fuselage modes on ground for ground resonance and fuselage modes in flight for air resonance can be avoided.

[0047] A further solution of the invention is specified by the features of claim 17. According to the invention a rotor craft comprising an aircraft rotor assembly according to one of the preceding embodiments. There are similar advantages to the preceding solution of independent claim 1 and claim 16.

[0048] In particular the invention combines an assymetrical radial offset and assymetrical chord offset for all lead lag dampers. The interblade lead-lag damper is connecting two adjacent blades with an asymmetrical radial offset. This provides stiffness and damping in the collective drag mode (oscillations in phase) of all blades which are linked in torsion with the entire power chain. This includes the rotor mast, the main gearbox, the tail drive shafts, the tail gearbox and tail rotor and engine. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system. The assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades makes it possible to define a sufficient first lever arm and second lever arm (LA 1 and LA2) to adjust the cyclic lead-lag frequency in order to avoid once projected into the fixed frame of the fuselage from coupling with the natural modes of the fuselage. Therewith fuselage modes on ground for ground resonance and fuselage modes in flight for air resonance can be avoided.

[0049] Other advantageous embodiments and combinations of features come out from the detailed description below and the entirety of the claims.

[0050] Brief description of the drawings

[0051] The drawings show:

[0052] Fig. 1 an aircraft rotor assembly comprising rotor blade mountings connected to the hub by a lead-lag damper known from the prior art;

[0053] Fig. 2 an aircraft rotor assembly comprising rotor blade mountings connected by lead- lag damper known from the prior art;

[0054] Fig. 3 an aircraft rotor assembly according to an embodiment of the invention in a plan view;

[0055] Fig. 4 an aircraft rotor assembly according to a further embodiment of the invention in a perspective view; Fig. 4A an aircraft rotor assembly according to a further embodiment of the invention;

[0056] Fig. 5 a lead lag damper in an aircraft rotor assembly according to an embodiment of the invention;

[0057] Fig. 6A a rotor blade mounting from a blade tip sectional view in different blade pitch values according a further embodiment of the invention;

[0058] Fig. 6B a rotor blade mounting from a blade tip sectional view in different blade pitch values according a further embodiment of the invention; and

[0059] Fig. 6C a rotor blade mounting from a blade tip sectional view having an out-of-plane offset in different blade pitch values according a further embodiment of the invention.

[0060] In the figures, the same components are given the same reference symbols.

[0061] Preferred embodiments

[0062] Figure 1 shows an aircraft rotor assembly comprising rotor blade mountings connected to the hub by a lead-lag damper known from the prior art.

[0063] Figure 2 shows blade holder connected together by lead-lag damper known from the prior art.

[0064] Figure 3 shows an aircraft rotor assembly according to an embodiment of the invention in a plan view.

[0065] The aircraft rotor assembly 100 comprises a rotor head 110 that is configured to rotate with a rotor axis RA. It shows a first rotor blade mounting 120, connecting the rotor head 110 and a first rotor blade 150 (not shown). Additionally a second rotor blade mounting 220 connects the rotor head 110 and a second rotor blade 250 (not shown). Further the aircraft rotor assembly 100 comprises a third rotor blade mounting 320, fourth rotor blade mounting 420 and a fifth rotor blade mounting 520, each connecting the rotor head 110 and a rotor blade.

[0066] Each rotor blade mounting 120,220, 320,... comprises a lead-lag hinge with a lead-lag axis LLA1,LLA2,LLA3,.. that is substantially parallel to the rotor axis RA. A flap hinge with a flap axis FA1,FA2,FA3,... that is substantially perpendicular to and intersects the lead-lag axis LLA 1 ,LLA2,LLA3 , . . . and a blade pitch hinge with a blade pitch axis BPA 1 ,BPA2,BPA3 , . . . that is substantially perpendicular to and intersects both the lead-lag axis LLA and the flap axis FA. Between each adjacent pair of rotor blade mountings 120,220, 320,... damping element is arranged. A first damping element 180 having a first damping axis DAI includes a first end 182 and a second end 184, whereby the first end 182 is attached to the first rotor blade mounting 120 and the second end 184 is attached to the second rotor blade mounting 220. In a plan view the first end 182 of the first damping element 180 is attached to a first attachment point 125 offset the blade pitch axis BPA1 of the first rotor blade 150 and the second end 184 of the first damping element 180 is attached to a second attachment point 245 on the blade pitch axis BPA2 of the second rotor blade 250.

[0067] Since the first attachment point 125 is offset the blade pitch axis BPA 1 of the first rotor blade and the second attachment point 245 is on the blade pitch axis BPA2 of the second rotor blade there is an assymetrical chord offset arrangement of the first damping element 180. A first distance from the rotor axis RA to the first attachment point 125 is bigger than a second distance from the rotor axis RA to the second attachment point 245. Thereby there is an assymetrical radial offset arrangement of the first damping element 180. As a consequence the invention combines an assymetrical radial offset and assymetrical chord offset. The interblade lead-lag damper respectively the first damping element 180 is connecting two adjacent blades - first rotor blade 150 and second rotor blade 250 - with an asymmetrical radial offset. This provides stiffness and damping in the collective drag mode (oscillations in phase) of all blades which are linked in torsion with the entire power chain. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system. The assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades - first rotor blade 150 and second rotor blade 250 - makes it possible to define a sufficient first lever arm and second lever arm (LAI and LA2) to adjust the cyclic lead-lag frequency in order to avoid once projected into the fixed frame of the fuselage from coupling with the natural modes of the fuselage. Therewith fuselage modes on ground for ground resonance and fuselage modes in flight for air resonance can be avoided.

[0068] The aircraft rotor assembly 100 discloses a third rotor blade mounting 320, connecting the rotor head 110 and a third rotor blade (not shown), wherein a second damping element 280 respectively a lead-lag damper has a second damping axis DA2. The second damping element 280 includes a first end 282 and a second end 284, wherein the first end 282 is attached to the third rotor blade mounting 320 and the second end 284 is attached to the first rotor blade mounting 120. In a plan view of the aircraft rotor assembly 100 the first end 282 of the second damping element 280 is attached to a first attachment point 325 that is offset the blade pitch axis BPA3 of the third rotor blade 350. The second end 284 of the second damping element 280 is attached to a second attachment point 145 on the blade pitch axis BPA1 of the first rotor blade 150.

[0069] A first lever LAI between the lead lag axis LLA1 and the first damping axis DAI is bigger than the a second lever LA2 between the lead lag axis LLA1 and the second damping axis DA2. This means that the distance between the lead lag axis LLA1 and the first damping axis DAI is greater than the distance between the lead lag axis LLA1 and the second damping axis DA2. This provides stiffness and damping in the collective drag mode. Therefore, it is no longer required to the engine to introduce damping in the power chain and thus preserve the responsiveness potential of the propulsion system. This effect is additionally increased by the assymetrical chord offset arrangement of the lead-lag damper connection points connecting two adjacent blades. The greater the chord offset at the first attachment point 325, the greater the difference between the first lever LAI and the second lever LA2. The chord offset CO is defined as a distance between the blade pitch axis BPA and the first attachment point. E.g. at the first rotor blade mounting 120 the chord offset CO is defined as the distance between the first blade pitch axis BPA1 and the first attachment point 125. The ratio of the first lever LAI to the chord offset CO is less or equal to 2. In other words, the ratio of the distance between the lead lag axis LLA1 and the first damping axis DAI to the chord offset CO, defined as the distance between the first blade pitch axis BPA1 and the first attachment point 125, is less or equal to 2.

[0070] Figure 4 shows an aircraft rotor assembly according to a further embodiment of the invention in a perspective view. The aircraft rotor assembly 100 comprises the rotor head 110 that is configured to rotate with the rotor axis RA. It shows the first rotor blade mounting 120, connecting the rotor head 110 and the first rotor blade 150. Additionally the second rotor blade mounting 220 connects the rotor head 110 and the second rotor blade 250. Again the aircraft rotor assembly 100 comprises the third rotor blade mounting 320, the fourth rotor blade mounting 420 and the fifth rotor blade mounting 520, each connecting the rotor head 110 and a rotor blade.

[0071] Each rotor blade mounting 120,220, 320,... comprises the lead-lag hinge with the lead-lag axis LLA1,LLA2,LLA3,.. that is substantially parallel to the rotor axis RA. The flap hinge with the flap axis FA1,FA2,FA3,... that is substantially perpendicular to and intersects the lead-lag axis LLA 1 ,LLA2,LLA3 , . . . and the blade pitch hinge with the blade pitch axis BPA 1 ,BPA2,BPA3 , . . . that is substantially perpendicular to and intersects both the lead-lag axis LLA and the flap axis FA. Between each adjacent pair of rotor blade mountings 120,220, 320, ... a damping element is arranged. The first damping element 180 is having the first damping axis DAI including the first end 182 and the second end 184, whereby the first end 182 is attached to the first rotor blade mounting 120 and the second end 184 is attached to the second rotor blade mounting (220). A description of the other identical features to the preceding embodiment is not repeated here.

[0072] The first attachment point 125 of the first damping element 180 is fitted at the first rotor blade mounting 120 with high radial offset and a chord wise offset to the blade pitch axis BPA1 to define a lever arm of the first damping element 180 adequate to get enough stability margin for ground and air resonance. The second end 184 of the first damping element 180 is attached to a second attachment point 245 on the blade pitch axis BPA2 of the second rotor blade 250.

[0073] Additionally to the preceding embodiment the first attachment point 125 of the first damping element 180 has an out-of-plane offset that is adjusted to minimize the Lead-lag Damper displacement of the first damping element 180.

[0074] In this context, the plane is defined as the rotor plane in which the rotor blades move geometrically around the axis of rotation RA. Out-of-plane therefore means either above or below the rotor plane. The second attachment point 245 of the first damping element 180 is directly on the blade pitch axis BPA2 of the second rotor blade 250, sufficiently close to the flap and lead-lag hinge. As a consequence a high damping to the torsional dynamics of the engine and drive train system of the helicopter and a low coupling with flap of the blade can be provided.

[0075] In other words the first end 182 of the first damping element 180 is attached to a first attachment axis 126 offset the blade pitch axis BPA1 of the first rotor blade 150 and the second end 184 of the first damping element 180 is attached to a second attachment axis 246 on the blade pitch axis BPA2 of the second rotor blade 250. The first rotor blade mounting 120, comprises an upper blade holder 121 and a lower blade holder 122 connected by the first attachment axis 126 and supporting the first rotor blade 150 between the upper blade holder 121 and the lower blade holder 122. Between the upper blade holder 121 and the lower blade holder 122 there is a first bolt 135 that is in line with the first attachment axis 126. A first distance Bl between the first end 182 of the first damping element 180 and the upper blade holder 121 is different from a second distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. So this embodiment includes the three attachment point parameters the radial offset, the chord offset and the out-of-plane offset. In this embodiment the rotor head 110 comprises the first rotor blade mounting 120, connecting the rotor head 110 and a first rotor blade 150 (not shown). Additionally the second rotor blade mounting 220 connects the rotor head 110 and a second rotor blade 250 (not shown), the third rotor blade mounting 320 connects the rotor head 110 and a third rotor blade 350 (not shown), the fourth rotor blade mounting 420 connects the rotor head 110 and a fourth rotor blade 450 (not shown), and the fifth rotor blade mounting 520 connects the rotor head 110 and a fifths rotor blade 550 (not shown). The connection between one of the rotor blade mountings 120, 220, 320, 420, 520 and a respective rotor blade 150, 250, 350, 450, 550 (not shown) is supported via a first blade attachment bolt 123 and a second blade attachment bolt 124. The first blade attachment bolt 123 and the second blade attachment bolt 124 are arranged at a distance from each other in a blade pitch axis direction.

[0076] Figure 4A shows a lead lag damper of an aircraft rotor assembly according to a further embodiment of the invention. This embodiment comprises almost all the features of the preceding embodiment of Figure 4, why it is not described again. The difference between this embodiment and the embodiment shown in figure 4 is in the arrangement of the first blade attachment bolt 123 and the second blade attachment bolt 124. According this embodiment the first blade attachment bolt 123 and the second blade attachment bolt 124 are arranged parallel at a distance from each other in a blade pitch axis direction. In other words, a connecting axis between the first blade attachment bolt 123 and the second blade attachment bolt 124 is arranged orthogonally to the blade pitch axis direction. The parallel arrangement of the first blade attachment bolt 123 and the second blade attachment bolt 124 is advantageous compared to the arrangement in a blade pitch axis direction in figure 4, because the rotor blade 150, 250, 350, 450, 550 can be controlled more precise due to a more direct torque transmission from the rotor blade mounting to the rotor blade. This embodiment provides a blade attachment with increased stiffness in torsion and lead lag, as well as a better distribution of centrifugal forces within the blade.

[0077] Figure 5 shows a lead lag damper of an aircraft rotor assembly according to an embodiment of the invention. The first damping element 180 is having the first damping axis DAI including the first end 182 and the second end 184, whereby the first end 182 is attached to the first rotor blade mounting 120 and the second end 184 is attached to the second rotor blade mounting 220. The first end 182 of the first damping element 180 is attached to the first attachment point 125 offset the blade pitch axis BPA 1 of the first rotor blade 150 and the second end 184 of the first damping element 180 is attached to the second attachment point 245 on the blade pitch axis BPA2 of the second rotor blade 250.

[0078] In the perspective view it becomes clear that the first attachment point 125 can be described as a point in the plan view. In the perspective view the first attachment point 125 is lying on a first attachment axis 126. Similar to this the second attachment point 245 is lying on a second attachment axis 246. So the first end 182 of the first damping element 180 is attached to the first attachment axis 126 offset the blade pitch axis BPA1 of the first rotor blade 150 and the second end 184 of the first damping element 180 is attached to the second attachment axis 246 on the blade pitch axis BPA2 of the second rotor blade 250. The first rotor blade mounting 120, comprises the upper blade holder 121 and the lower blade holder 122 connected by the first attachment axis 126 and supporting the first rotor blade 150 between the upper blade holder 121 and the lower blade holder 122. Between the upper blade holder 121 and the lower blade holder 122 there is the first bolt 135 that is in line with the first attachment axis 126. The first distance Bl between the first end 182 of the first damping element 180 and the upper blade holder 121 is different from the second distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. The first attachment point 125 lying on the first attachment axis 126 comprises a spherical function bearing 130 for mounting the first end 182 of the first damping element 180 and the second attachment point 245 lying on the second attachment axis 246 comprises a spherical function bearing 230 for mounting the second end 184 of the first damping element 180.

[0079] Figure 6A shows a rotor blade mounting 120, 220, 320,... from a blade tip sectional view in different blade pitch values according a further embodiment of the invention. The rotor blade mounting 120, 220, 320,... comprises the upper blade holder 121 and the lower blade holder 122. Between the upper blade holder 121 and the lower blade holder 122 there is the first bolt 135 comprising an upper sleeve 136 defining the first distance Bl between the first end 182 of the first damping element 180 and the upper blade holder 121. Additionally the first bolt 135 comprising an lower sleeve 137 defining the second distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. The first distance B 1 between the first end 182 of the first damping element 180 and the upper blade holder 121 is identical to the distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. The first attachment point 125 lying on the first attachment axis 126 comprises a spherical function bearing 130 for mounting the first end 182 of the first damping element 180. According to the first illustration in figure 6A the pitch angle of the rotor blade is deflected upwards so that the first attachment point 125 is located above the rotor plane RP.

[0080] According to the second illustration in figure 6A the pitch angle of the rotor blade is not deflected so that the first attachment point 125 is located directly on the rotor plane RP.

[0081] According to the third illustration in figure 6A the pitch angle of the rotor blade is deflected downwards so that the first attachment point 125 is located below the rotor plane RP.

[0082] Figure 6B shows a rotor blade mounting 120, 220, 320,... from a blade tip sectional view in different blade pitch values according a further embodiment of the invention. Again the rotor blade mounting 120, 220, 320, . . . comprises the upper blade holder 121 and the lower blade holder 122. In contrast to the above embodiment, the upper blade holder 121 and the lower blade holder 122 have a different cross-sectional geometry, as a result of which the length of the upper sleeve 136 and the lower sleeve 137 can be formed differently. Overall, the upper sleeve 136 and the lower sleeve 137 are shorter than in the above embodiment, as a result of which the deflection of the first attachment point 125 has less of an effect at varying pitch angles relative to rotor plane RP. Between the upper blade holder 121 and the lower blade holder 122 there is the first bolt 135 comprising the upper sleeve 136 defining the first distance Bl between the first end 182 of the first damping element 180 and the upper blade holder 121. Additionally the first bolt 135 comprising the lower sleeve 137 defining the second distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. The first distance B 1 between the first end 182 of the first damping element 180 and the upper blade holder 121 is identical to the distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. The first attachment point 125 lying on the first attachment axis 126 comprises a spherical function bearing 130 for mounting the first end 182 of the first damping element 180.

[0083] According to the first illustration in figure 6B the pitch angle of the rotor blade is deflected upwards so that the first attachment point 125 is located above the rotor plane RP.

[0084] According to the second illustration in figure 6B the pitch angle of the rotor blade is not deflected so that the first attachment point 125 is located directly on the rotor plane RP.

[0085] According to the third illustration in figure 6B the pitch angle of the rotor blade is deflected downwards so that the first attachment point 125 is located below the rotor plane RP. Figure 6C shows a rotor blade mounting 120, 220, 320,... from a blade tip sectional view in different blade pitch values according a further embodiment of the invention. The rotor blade mounting 120, 220, 320,... comprises the upper blade holder 121 and the lower blade holder 122. Between the upper blade holder 121 and the lower blade holder 122 there is the first bolt 135 comprising an upper sleeve 136 defining the first distance Bl between the first end 182 of the first damping element 180 and the upper blade holder 121. Additionally the first bolt 135 comprising an lower sleeve 137 defining the second distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122.

[0086] The first distance Bl between the first end 182 of the first damping element 180 and the upper blade holder 121 is smaller compared to the distance B2 between the first end 182 of the first damping element 180 and the lower blade holder 122. The first attachment point 125 lying on the first attachment axis 126 comprises a spherical function bearing 130 for mounting the first end 182 of the first damping element 180.

[0087] According to the first illustration in figure 6C the pitch angle of the rotor blade is deflected upwards so that the first attachment point 125 is located above the rotor plane RP.

[0088] According to the second illustration in figure 6C the pitch angle of the rotor blade is not deflected but the first attachment point 125 is still located above the rotor plane RP due to the difference of the distance Bl compared to the distance B2. So this embodiment shows an example to realize the out-of-plane offset.

[0089] According to the third illustration in figure 6C the pitch angle of the rotor blade is deflected downwards so that the first attachment point 125 is located below the rotor plane RP.

[0090] List of reference symbols

Claims

Claims1. Aircraft rotor assembly ( 100) comprising: a rotor head (110) configured to rotate with a rotor axis (RA); a first rotor blade mounting (120), connecting the rotor head (110) and a first rotor blade (150); a second rotor blade mounting (220), connecting the rotor head (110) and a second rotor blade (250); each rotor blade mounting (120,220) comprises a lead-lag hinge with a lead-lag axis (LLA1,LLA2,LLA3,..) that is substantially parallel to the rotor axis (RA), a flap hinge with a flap axis (FA1,FA2,FA3,...) that is substantially perpendicular to and intersects the lead-lag axis (LLA1,LLA2,LLA3), and a blade pitch hinge with a blade pitch axis (BPA1,BPA2,BPA3,...) that is substantially perpendicular to and intersects both the lead-lag axis (LLA) and the flap axis (FA); a first damping element (180) having a first damping axis (DAI) including a first end (182) and a second end ( 184), the first end ( 182) being attached to the first rotor blade mounting (120) the second end (184) being attached to the second rotor blade mounting (220); wherein in a plan view the first end (182) of the first damping element (180) is attached to a first attachment point (125) offset the blade pitch axis (BPA1) of the first rotor blade (150) and the second end (184) of the first damping element (180) is attached to a second attachment point (245) on the blade pitch axis (BPA2) of the second rotor blade (250); wherein a first distance from the rotor axis (RA) to the first attachment point (125) is bigger than a second distance from the rotor axis (RA) to the second attachment point (245).

2. Aircraft rotor assembly (100) according claim 1, wherein a third rotor blade mounting (320), connecting the rotor head (110) and a third rotor blade (350), wherein a second dampingelement (280) having a second damping axis (DA2) including a first end (282) and a second end (284), the first end (282) being attached to the third rotor blade mounting (320) the second end (284) being attached to the first rotor blade mounting (120).

3. Aircraft rotor assembly (100) according claim 2, wherein in a plan view the first end (282) of the second damping element (280) is attached to a first attachment point (325) offset the blade pitch axis (BPA3) of the third rotor blade (350) and the second end (284) of the second damping element (280) is attached to a second attachment point (145) on the blade pitch axis (BPA1) of the first rotor blade (150).

4. Aircraft rotor assembly (100) according to one of the preceding claims, wherein a first lever (LAI) between the lead lag axis (LLA1) and the first damping axis (DAI) is bigger than the a second lever (LA2) between the lead lag axis (LLA1) and the second damping axis (DA2).

5. Aircraft rotor assembly (100) according to claim 4, wherein a ratio between LAI to LA2 is bigger than 1.5.

6. Aircraft rotor assembly (100) according to one of the preceding claims, wherein a distance between the lead lag axis (LLA1) and the second attachment point (145) is smaller than a distance between the lead lag axis (LLA1) and an intersection point between the first damping axis (DAI) and the blade pitch axis (BPA1).

7. Aircraft rotor assembly (100) according to one of the preceding claims, wherein a rotational movement of the first damping element (180) is blocked.

8. Aircraft rotor assembly (100) according to one of the preceding claims, wherein the first attachment point (125) comprises a spherical function bearing (130) for mounting the first end (182) of the first damping element (180) and the second attachment point (245) comprises a spherical function bearing (230) for mounting the second end (184) of the first damping element (180).

9. Aircraft rotor assembly (100) according to one of the preceding claims, wherein the ratio of a first distance (oc2), from the first attachment point (325) of the third rotor blade mounting (320) to the flap axis (FA1) of the first rotor blade mounting (120), to a second distance ( ?2)- from the intersection point between the blade pitch axis (BPA1) of the first rotor blademounting (120) and the second damping axis (DA2)to the flap axis (FA1) of the first rotor blade mounting (120), is between 0 and 0,3.

10. Aircraft rotor assembly (100) according to one of the preceding claims, wherein the ratio of a first distance (o ), from the second attachment point (245) of the second rotor blade mounting (220) to the flap axis (FA1) of the first rotor blade mounting (120), to a second distance (.ftp. from the flap axis (FA1) of the first rotor blade mounting (120) to the intersection point between the blade pitch axis (BPA1) of the first rotor blade mounting (120) and the first damping axis (DAI), is between 0 and 0,3.

11. Aircraft rotor assembly (100) according to one of the preceding claims, wherein each damping element (180, 280,...) has a nominal stiffness between 0,5 kN / mm and 10 kN / mm, in particular each damping element (180, 280,...) has a nominal stiffness between 1 kN / mm and 4 kN / mm.

12. Aircraft rotor assembly (100) according to one of the preceding claims, wherein each damping element (180, 280,...) is fully elastomeric.

13. Aircraft rotor assembly (100) according to one of the preceding claims, wherein the lead-lag hinge with a lead-lag axis (LLA1), the flap hinge with a flap axis (FA1) and the blade pitch hinge with a blade pitch axis (BPA1) are arranged by means of a single laminated spherical abutment.

14. Aircraft rotor assembly (100) according to one of the preceding claims, wherein each rotor blade ( 150, 250, . . . ) is attached to each corresponding rotor blade mounting (120,220, . . . ) by at least two bolts.

15. Aircraft rotor assembly (100) according to one of the preceding claims, wherein the aircraft rotor assembly (100) comprises at least three rotor blade mountings (120,220,320,...).

16. Aircraft (500) comprising an aircraft rotor assembly (100) according to one of the preceding claims.

17. Rotor craft (700) comprising an aircraft rotor assembly (100) according to one of the preceding claims 1 to 15.

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