Rotor for an aircraft capable of hovering
阅读说明:本技术 用于能够悬停的飞行器的旋翼 (Rotor for an aircraft capable of hovering ) 是由 阿蒂利奥·科隆博 路易吉·博塔索 保罗·皮萨尼 安德烈亚·法瓦罗托 达里奥·科隆博 费代里 于 2018-06-28 设计创作,主要内容包括:描述了一种用于飞行器(1)的旋翼(3,3’,3”,3”’),其包括:叶毂(5),该叶毂能围绕轴线(A)旋转并且又包括多个桨叶(9);支柱(6),该支柱能连接至飞行器(1)的驱动构件,并连接至叶毂(5)以驱动叶毂(5)围绕轴线(A)旋转;以及阻尼装置(15;20a,20b;21a,21b),用于抑制振动在与轴线(A)正交的平面中传递至支柱(6);阻尼装置(15;20a,20b;21a,21b)包括至少第一质量体(20a,20b)和第二质量体(21a,21b),第一质量体和第二质量体能够分别以第一旋转速度和第二旋转速度((N-1)*Ω;-(N+1)*Ω)围绕轴线(A)偏心地旋转;第一质量体(20a,20b)和第二质量体(21a,21b)操作性地连接至支柱(6),以分别在支柱(6)上产生在轴线(A)的径向方向上具有主要分量的第一阻尼力和第二阻尼力;旋翼(3,3’,3”,3”’)包括传动单元(19,19’,19”,19”’),该传动单元介于支柱(6)与第一质量体和第二质量体(20a,20b;21a,21b)之间,以驱动第一质量体和第二质量体(20a,20b;21a,21b)旋转。(A rotor (3, 3 ', 3 ", 3"') for an aircraft (1) is described, comprising: a hub (5) rotatable about an axis (A) and comprising in turn a plurality of blades (9); a strut (6) connectable to a drive member of the aircraft (1) and to the hub (5) to drive the hub (5) in rotation about the axis (A); and damping means (15; 20a, 20 b; 21a, 21b) for damping the transmission of vibrations to the strut (6) in a plane orthogonal to the axis (A); the damping device (15; 20a, 20 b; 21a, 21b) comprises at least a first mass (20a, 20b) and a second mass (21a, 21b) which are eccentrically rotatable about the axis (A) at a first rotational speed ((N-1) × Ω; - (N +1) × Ω), respectively; the first and second masses (20a, 20b, 21a, 21b) are operatively connected to the strut (6) to generate, respectively, on the strut (6), a first and a second damping force having a main component in a radial direction of the axis (a); the rotor (3, 3 ') comprises a transmission unit (19, 19') interposed between the mast (6) and the first and second masses (20a, 20 b; 21a, 21b) to drive the first and second masses (20a, 20 b; 21a, 21b) in rotation.)
1. A rotor (3, 3 ', 3 "') for an aircraft (1) capable of hovering, comprising:
-a hub (5) rotatable about a first axis (a) and comprising in turn a plurality of blades (9);
-a strut (6) connectable to a drive member of the aircraft (1) and operatively connected to the hub (5) to drive the hub (5) in rotation about the first axis (a);
damping means (15; 20a, 20 b; 21a, 21b) for damping the transmission of vibrations to said strut (6) in a plane orthogonal to said first axis (A);
the damping device (15; 20a, 20 b; 21a, 21b) in turn comprises at least one first mass (20a, 20b) and at least one second mass (21a, 21b) which are eccentrically rotatable about the first axis (A) at a first rotation speed (N-1) omega) and a second rotation speed (- (N +1) omega), respectively, with respect to the strut (6);
the first and second masses (20a, 20b, 21a, 21b) are operatively connected to the strut (6) to generate on the strut (6) a first and second damping force, respectively, having a main component in a radial direction of the first axis (a);
the rotor (3, 3 ', 3 "') further comprises a transmission unit (19, 19 ', 19"') functionally interposed between the mast (6) and the first (20a, 20b) and second (21a, 21b) masses to drive the first (20a, 20b) and second (21a, 21b) masses in rotation in opposite directions to each other;
characterized in that said first mass (20a, 20b) and said second mass (21a, 21b) generate, in use, a resultant force on said strut (6) having a sinusoidal trajectory;
said rotor (3, 3 ', 3 "') further comprising actuating means (80a, 80b, 81a, 81b) selectively operable to vary the amplitude and phase of said resultant force;
the actuation means (80a, 80b, 81a, 81b) comprise:
-a drive member (82);
-a support element (66) operatively connected to one of said first mass (20a, 20b) and said second mass (21a, 21b) and rotatable with respect to said pillar (6); and
-a further transmission unit (65, 67) drivable by said driving member (82) and designed to drive, in use, said support element (66);
the further transmission unit (65, 67) is configured to only irreversibly transmit motion from an output member (83) of the drive member (82) to the support element (66), and to prevent motion from being transmitted in the opposite direction from the support element (66) to the drive member (82);
said further transmission unit (65, 67) comprising a worm (65) operable by said drive member (82) and helical gear teeth (67) carried by said support element (66) and irreversibly engaged with said worm (65);
-for each of said first (20a, 20b) and second (21a, 21b) masses, said rotor (3, 3 ', 3 "') comprising an associated housing (62, 63);
said associated housing (62, 63) rotating around said first axis (A) integrally with said strut (6) and with said drive member (82) and at least partially housing said further transmission unit (66, 67);
-the associated output member (83) rotates relative to the housing (62, 63) about a second axis (B) transverse to the first axis (a);
the support element (66) is mounted around the first axis (a) in an angularly integral manner with the housing (62, 63).
2. A rotor as claimed in claim 1, characterized in that said transmission unit (19, 19 ', 19 "') comprises a first output member (30, 68; 98c, 98d, 68) connected to said first mass (20a, 20b) and at least a second output member (40, 68; 98a, 98b, 68) connected to said second mass (21a, 21 b);
the transmission unit (19; 19') is configured to drive the first output member (30, 68; 98c, 98d, 68) in a first direction at a third angular speed associated with the first angular speed ((N-1) × Ω), and to drive the second output member (40, 68; 98a, 98b, 68) in a second direction opposite to the first direction at a fourth angular speed associated with the second angular speed ((N +1) × Ω).
3. A rotor as claimed in claim 2, characterized in that said transmission unit (19, 19 ', 19 "') comprises:
-a first level (25) functionally interposed between said pillar (6) and said at least one first mass (20a, 20 b); and
-a second level (26) functionally interposed between said first level (25) and at least one of said second masses (21 a; 21 b).
4. A rotor as claimed in claim 3, characterized in that said first and second levels (25, 26) comprise a first and a second epicyclic gear train, respectively;
the first and second sun gears (29, 45) of the first and second epicyclic gear trains defined by the first and second levels (25, 26) are angularly integral with the first and second mass bodies (20a, 20b, 21a, 21b) respectively about the first axis (A);
said first sun gear (29) being also angularly integral with a second ring gear (41) of said second epicyclic gearing defined by said second level (26);
the first and second epicyclic gear trains defined by the first and second levels (25, 26) further comprise:
-respective first (33, 35) and second (46, 48) planet gears;
-a common planet gear carrier (28) around which the first (33, 35) and second (46, 48) planet gears rotate and which is connected to the fuselage (2) of the aircraft (1);
the first planet gears (33, 35, 37) are in mesh with the first sun gear (29) and the first ring gear (27);
the second planet gears (46, 48) are meshed with the second sun gear (45) and the second ring gear (41).
5. A rotor according to any preceding claim, wherein the first rotation speed is equal to (N-1) × Ω and the second rotation speed is equal to- (N +1) × Ω, where N is the number of blades (9) and Ω is the rotation speed of the mast (6) in a reference frame integral with the fuselage (2);
-said first mass (20a, 20b) is rotatable with respect to said pillar (6) at said first angular velocity ((N-1) × Ω) in the same direction as said pillar (6);
the second mass (21a, 21b) is rotatable with respect to the strut (6) at the second angular velocity ((N +1) × Ω) in a direction opposite to the strut (6).
6. A rotor as claimed in any one of the foregoing claims, characterized in that said actuating means (80a, 80b, 81a, 81b) are selectively operable to vary a first angle between said first masses (20a, 20b) and/or a second angle between said second masses (21a, 21b) with respect to said first axis (a).
7. A rotor as claimed in any one of the foregoing claims, characterized in that said first mass (20a, 20b) and said second mass (21a, 21b) are movable with respect to said mast (6) in a radial direction of said first axis (a) and are in contact with respective guides (23') rotating integrally with said mast (6) about said first axis (a).
8. A rotor as claimed in any one of the foregoing claims, characterized in that it comprises an air flow conveyor (10) connected to said hub (5) and designed to direct the air flow generated, in use, by the rotation of said blades (9) according to a predetermined path;
the first mass body (20a, 20b) and the second mass body (21a, 21b) are housed inside the air flow conveyor (10);
said rotor (3, 3') being characterized in that said first and second epicyclic gearing defining said first and second levels (25, 26; 90a, 90b, 90c, 90d) are housed inside said mast (6); alternatively, the entire transmission unit (19 "') is housed inside the air flow conveyor (10).
9. An aircraft (1) capable of hovering, characterized in that it comprises a rotor (3, 3 ', 3 "') according to any one of the preceding claims.
10. A rotor (3 ") for an aircraft (1) capable of hovering, comprising:
-a hub (5) rotatable about a first axis (a) and comprising in turn a plurality of blades (9);
-a strut (6) connectable to a drive member of the aircraft (1) and operatively connected to the hub (5) to drive the hub (5) in rotation about the first axis (a);
damping means (15; 20a, 20 b; 21a, 21b) for damping the transmission of vibrations to said strut (6) in a plane orthogonal to said first axis (A);
the damping device (15; 20a, 20 b; 21a, 21b) in turn comprises at least one first mass (20a, 20b) and at least one second mass (21a, 21b) which are eccentrically rotatable about the first axis (A) at a first rotation speed (N-1) omega) and a second rotation speed (- (N +1) omega), respectively, with respect to the strut (6);
-said first and second masses (20a, 20b, 21a, 21b) are operatively connected to said strut (6) to generate on said strut (6) a first and a second damping force, respectively, having a main component in a radial direction of said first axis (a);
the rotor (3 ") further comprises a transmission unit (19") functionally interposed between the mast (6) and the first (20a, 20b) and second (21a, 21b) masses, to drive the first (20a, 20b) and second (21a, 21b) masses in rotation in opposite directions to each other;
the transmission unit (19') comprises:
-a housing (120) fixed with respect to the fuselage (2);
-a first level (90a, 90b) functionally interposed between said pillar (6) and said second mass (21a, 21 b); and
-a second level (90c, 90d) functionally interposed between said pillar (6) and said first mass (20a, 20 b);
the first level (90a, 90b) and the second level (91a, 91b) comprise a first and a second epicyclic gear train, respectively;
-said first and second epicyclic gearing having a third and a fourth sun gear (97a, 97b, 97c, 97d), respectively, said third and fourth sun gears (97a, 97b, 97c, 97d) being operatively connected to the respective second and first mass bodies (21a, 21b, 20a, 20b), respectively, and defining a first and a second motion output, respectively;
the first epicyclic gearing further comprises:
-a third ring gear (96 a; 96b) defining a first motion input and rotating integrally with said strut (6);
-a plurality of third planet gears (99a, 99b) meshing with said third ring gear (96a, 96b) and said third sun gear (97a, 97 b); and
-a third planetary gear carrier (103a, 103b), the third planetary gears (99a, 99b) being rotatably mounted with respect to the third planetary gear carrier;
the second epicyclic train further comprises:
-a fourth ring gear (96c, 96d),
-a plurality of fourth planet gears (99c, 99d) meshing with said fourth ring gear (96 c; 96d) and said fourth sun gear (97c, 97 d); and
-a fourth planetary gear carrier (103c) defining a second motion input and rotating integrally with said struts (6), and said fourth planetary gears (99c, 99d) being mounted in a rotatable manner with respect to said fourth planetary gear carrier;
the third planetary gear carriers (103a, 103b) and the fourth ring gear (96 c; 96d) are in sliding contact with the housing (120);
characterized in that said first mass (20a, 20b) and said second mass (21a, 21b) generate, in use, a resultant force on said strut (6) having a sinusoidal trajectory;
said rotor (3 ") further comprising actuating means (80a, 80b, 81a, 81b) selectively operable to vary the amplitude and phase of said resultant force;
the actuation means (80a, 80b, 81a, 81b) comprise:
-a drive member (82 ");
-a support element (66 ") operatively connected to one of said first mass (20a, 20b) and said second mass (21a, 21b) and rotatable with respect to said pillar (6); and
-a further transmission unit (65, 67) drivable by said driving member (82 ") and designed to drive, in use, said support element (66");
the further transmission unit (65, 67) is configured to only irreversibly transmit motion from an output member (83) of the drive member (82 ") to the support element (66") and to prevent motion from being transmitted in the opposite direction from the support element (66 ") to the drive member (82");
said further transmission unit (65, 67) comprising a worm (65) operable by said drive member (82) and helical gear teeth (67) carried by said support element (66) and irreversibly engaged with said worm (65);
the rotor (3 ") further comprises:
-a first said drive member (82 ") designed to cause the rotation of said second mass (21a, 21b) about said first axis (a) according to a predetermined angle and a first said support element (66"), said first said support element (66 ") being operatively connected to the first said drive member (82") through a first said further transmission unit (66 ", 67); the first support element (66 ") is carried by the third planetary gear carrier (103a, 103 b); operation of the first drive member (82 ") causes relative movement of the third planetary gear carrier (103a, 103b) relative to the housing (120) and about the first axis (a); and/or
-a second said drive member (82 ") designed to cause the first mass (20a, 20b) to rotate about the first axis (a) according to a predetermined angle, and a second said support element (66") operatively connected to the second said drive member (82 ") through a second said further transmission unit (66", 67); said second support element (66 ") being carried by said fourth ring gear (96c, 96 d); operation of the second drive member (82') in use causes relative movement of the fourth ring gear (96c, 96d) relative to the housing (120) and about the first axis (A).
11. A rotor as claimed in claim 10, characterized in that said transmission unit (19 ") comprises a first output member (30, 68; 98c, 98d, 68) connected to said first mass (20a, 20b) and at least one second output member (40, 68; 98a, 98b, 68) connected to said second mass (21a, 21 b);
the transmission unit (19') is configured to drive the first output member (30, 68; 98c, 98d, 68) in a first direction at a third angular speed associated with the first angular speed ((N-1) × Ω), and to drive the second output member (40, 68; 98a, 98b, 68) in a second direction opposite to the first direction at a fourth angular speed associated with the second angular speed (- (N +1) × Ω).
12. A rotor as claimed in claim 10 or 11, characterized in that said first rotation speed is equal to (N-1) × Ω and said second rotation speed is equal to- (N +1) × Ω, where N is the number of blades (9) and Ω is the rotation speed of said mast (6) in a reference frame integral with said fuselage (2);
-said first mass (20a, 20b) is rotatable with respect to said pillar (6) at said first angular velocity ((N-1) × Ω) in the same direction as said pillar (6);
the second mass (21a, 21b) is rotatable with respect to the strut (6) at the second angular velocity ((N +1) × Ω) in a direction opposite to the strut (6).
13. A rotor as claimed in any one of claims 10 to 12, characterized in that said actuating means (80a ", 80 b", 81a ", 81 b") are selectively operable to vary a first angle between said first masses (20a, 20b) and/or a second angle between said second masses (21a, 21b) with respect to said first axis (a).
14. A rotor as claimed in any one of claims 10 to 13, characterized in that it comprises an air flow conveyor (10) connected to said hub (5) and designed to direct the air flow generated in use by the rotation of said blades (9) according to a predetermined path;
the first mass body (20a, 20b) and the second mass body (21a, 21b) are housed inside the air flow conveyor (10);
the rotor (3') being characterized in that the first and second epicyclic gearing defining the first and second levels (25, 26; 90a, 90b, 90c, 90d) are housed within the mast (6).
15. Aircraft (1) capable of hovering, characterized in that it comprises a rotor (3 ") according to any one of claims 10 to 14.
Technical Field
The present invention relates to a rotor for an aircraft capable of hovering, in particular for a helicopter.
Background
Helicopters are known that basically comprise a fuselage, a main rotor located on top of the fuselage and rotatable about its own axis, and a tail rotor located at the end of the fuselage.
In more detail, the rotor in turn substantially comprises a hub rotatable about said axis and equipped with a plurality of blades radially fixed to and projecting from said hub, and a mast connectable to the drive member and operatively connected to the hub to drive the rotation thereof.
In use, operation of the rotor causes the generation of high and low frequency vibrations. More specifically, low frequency vibrations are generated by impacts that separate from the center of the blade and hub. This separation occurs in the center of the hub and affects all the vertical and horizontal aerodynamic surfaces of the tail and tail rotors.
In use, rotation of the blades at high angular velocities causes the generation of further high frequency vibrations which are transmitted to the struts and hence to the fuselage, thereby reducing the comfort of occupants within the fuselage.
More specifically, a vibration load acts on the hub and the strut in the axial direction and in a direction orthogonal to the rotation axis of the strut.
It is known in the industry that in a rotating reference frame (thus integral with the rotor, hub and mast), the vibratory loads acting in the rotor plane have pulses equal to (N +1) × Ω, (N-1) × Ω and multiples thereof, where Ω is the rotational speed of the mast and N represents the number of blades of the rotor. It is also known that the vibration loads acting in the rotor plane shift in frequency when transmitted from the rotating system to the fixed system of the fuselage, and have pulses on the fixed system equal to N x Ω and their associated multiples. In other words, the hub and the struts transfer the oscillating aerodynamic load pulses acting in the plane of the blade to the above-mentioned pulses.
From the foregoing, it is evident in the industry that the need to limit the transmission of vibrations from the mast to the fuselage is felt in the case where the above-mentioned pulses are equal to the product of the rotation speed of the mast and the number of blades of the rotor.
Passive and active damping devices are known for this purpose.
Passive damping devices basically comprise a mass resiliently suspended from a strut or hub. The vibrations of these suspended masses cause the vibrations on the struts and the hub to be at least partially dissipated.
While they are easy to construct and install and do not require an energy source external to the rotor, passive damping devices have the greatest limitation in the performance they can provide.
Active damping devices are fundamentally actuators that exert a sinusoidal damping force on the hub or strut that counteracts the force generated by the vibrations.
Examples of such active damping devices are shown in the applicant's patent application EP- cA-2857313.
This patent application shows the use of a pair of actuators operatively connected to the strut and controlled so as to generate on the strut a respective damping force having a component in a plane orthogonal to the axis of rotation of the strut.
In more detail, the actuators are shaped like rings, one on top of the other, driven in rotation by the strut and mounted rotatably about an axis of rotation with respect to the strut.
The actuators rotate counter-rotationally about the rotation axis of the mast and generate a corresponding force on the mast equal to the product of the rotation speed of the mast and the number of blades of the rotor to counteract the above-mentioned angular frequency.
Further examples of active damping devices are described in patent application US-A-2016/0325828 and patent US-B-8,435200.
The active damping device has the advantage of being able to change its damping characteristics according to the development of the vibration conditions of the hub and the strut.
However, active damping devices are inherently more complex than passive damping devices in terms of weight, power and volume. Moreover, they need to be powered with considerable energy.
The industry recognizes the need to dampen the flexural vibrations of the strut and hub, so as to maintain the effectiveness of the active damping devices described above, and at the same time limit the need for dedicated drive components as much as possible, in order to reduce weight, volume and overall cost.
There is also a recognized need in the industry to make the values of the damping forces exerted on the strut as accurate and repeatable as possible in order to make the damping action as accurate and repeatable as possible.
Additionally, it is recognized that there is a need to be able to adjust the phase and amplitude of the damping forces applied to the strut based on different vibration conditions of the fuselage.
Finally, it is recognized in the industry that in the event of failure of one drive member, it is desirable to suppress the occurrence of undesirable vibrational loads.
GB-a-1120193 discloses a rotor for an aircraft according to the preamble of claim 1.
US-A-2014/360830 discloses A rotor for an aircraft, the rotor comprising A hub, A mast, A damping device having at least one first mass and at least one second mass, and A transmission unit driving the first and second masses to rotate in opposite directions.
The above object is achieved by the present invention as far as it relates to a rotor for an aircraft capable of hovering as defined in claim 1.
The invention also relates to a rotor for an aircraft capable of hovering according to
Disclosure of Invention
It is an object of the present invention to provide a rotor for an aircraft capable of hovering that can satisfy at least one of the above specified needs in a simple and inexpensive manner.
The above object is achieved by the present invention as far as it relates to a rotor for an aircraft capable of hovering as defined in claim 1.
Drawings
For a better understanding of the invention, four preferred embodiments are described hereinafter, by way of non-limiting example only, with reference to the accompanying drawings, in which:
figure 1 is a side view of a helicopter comprising a rotor according to a first embodiment of the present invention;
figure 2 shows in section a damping device incorporated in the rotor of figure 1, in which the various parts are not fully shown for the sake of clarity;
figure 3 shows a first component of the rotor in figure 2 on an enlarged scale;
figure 4 shows, on a further enlarged scale, a second component of the damping device in figure 3;
figure 5 shows, on a highly enlarged scale and in cross section, some details of the damping device in figure 2;
figure 6 is a top view of the first part of figure 3;
figure 7 is a top view of the second part of figure 4;
figure 8 is an operational schematic of the first component of figure 5;
figure 9 is a section along the line IX-IX of figure 8;
figure 10 is a section along the line X-X in figure 8;
figure 11 shows a rotor according to a further embodiment of the invention in cross section, wherein the various parts are not fully shown for the sake of clarity;
FIG. 12 is a sectional view taken along line XII-XII of FIG. 11;
figure 13 shows a rotor according to a third embodiment of the invention in cross-section, with some parts removed for clarity; and
figure 14 shows a rotor according to a fourth embodiment of the invention in cross section, with some parts removed for clarity.
Detailed Description
With reference to fig. 1, numeral 1 indicates an aircraft capable of hovering, in particular a helicopter, essentially comprising a
In more detail, the
With reference to fig. 2,
More specifically, the
More specifically, the
In particular, the
The
In more detail, the
The
The
The
More specifically, when travelling from the
The
The
In more detail, the
The
Thus, the
The first damping force and the second damping force correspond to the
These first and second damping forces have sinusoidal trajectories, the amplitude depending on the
The first and second damping forces generate on the
Preferably, the
In this specification, the term angular frequency refers to the frequency multiplied by 2 π.
The
Without being explicitly indicated, in the following of the description, it is understood that the angular velocity (N-1) × Ω has the same meaning as the angular velocity Ω of the
In this way, the first and second damping forces each have an angular frequency N Ω in a frame of reference integral with the
The first damping forces are equal in magnitude to each other and the second damping forces are equal in magnitude to each other.
The
In the illustrated case, the
In a further embodiment, not shown, the
Advantageously, the
In this way, the
Preferably, the
a
a
In particular,
The
As can be seen in fig. 2, in the illustrated case the
The
In more detail, the
a
a
a
The
a plurality (three in the case shown) of planet gears 33 rotatable about their own axes C parallel to axis a and having respective
a plurality (three in the case shown) of planet gears 35 rotatable about their own axes D parallel to axis A, C and having respective
The
Due to the presence of the planet gears 33, 35, the
In other words,
the
the
The planet-
The
a
a
The
a plurality (three in the case shown) of planet gears 46 rotatable about their own axes E parallel to axis a and having respective
a plurality (three in the case shown) of planet gears 48, angularly integral with the planet gears 46 and having respective
In particular, axis E is radially interposed between respective axes C and D.
Further, the diameter of the
The
In other words, the
the
the
the
The
a
a connection group 61a, 61b for connecting the
With particular reference to fig. 6, the
Since these
The
a
a
In more detail, the
a
a
a
The
The
The
In particular, the rotation of the
From the above, it follows that rotation of the
It is important to note that the condition of irreversible movement existing between the
Advantageously,
In more detail, the
The
With particular reference to fig. 3-7, each
an
a
The
The
As a result, the
Conversely, the switching on of the
Rotation of the
In other words, each
The
a first slip ring (not shown) for electrically connecting the power supply carried on the
a second slip ring (also not shown) for electrically connecting the power source with the
Finally, the
a plurality of sensors 85 (schematically shown in fig. 2) able to generate a plurality of signals associated with the acceleration state of the
a control unit 86 (also schematically shown in fig. 2) configured to generate control signals for the
In use, the
More specifically, with reference to a frame of reference integrated with
The rotation of hub 5 and blades 9 generates vibrations that will be transmitted to
For the fixed system of the fuselage, these vibrations mainly have an angular frequency equal to N x Ω, where N is the number of blades 9 and Ω is the angular velocity of rotation of the
In order to reduce these vibrations, the
Due to the centrifugation, the
More specifically, the first and second damping forces are sinusoidal and have angular frequencies equal to (N-1) × Ω and (N +1) × Ω, respectively, in a reference frame integral with the
These first and second damping forces counteract the load due to these vibrations in a plane orthogonal to the axis a.
Furthermore, these first and second damping forces have an angular frequency equal to N Ω in a reference frame integrated with the
The first and second damping forces generate respective first and second resultant forces on the
These first and second resultant forces have a mass
It is important to note that the
And
Hereinafter, the
The
The
In particular, the
A
In more detail, the shaft 30(40) drives in rotation, integrally with itself, about the axis a, the assembly formed by the
In the
In this case, the
As a result, in the above-described situation, the
Rotation of the
In the following, the function of
In case it is detected that the phase and amplitude of the first and second resultant forces need to be changed, the
One or
Rotation of the worm screw or screws 65 causes the disc or
With reference to fig. 11 and 12, reference numeral 3' denotes a rotor according to a second embodiment of the present invention.
Rotor 3' is similar to
In particular, the rotor 3' differs from the
Furthermore, the rotor 3' differs from the
More specifically, the rollers 22 ' extend about respective axes parallel to the axis a and have respective axial ends 24 ' opposite each other, sliding radially in corresponding radial slots 18 ' defined by the free ends of the
The function of the rotor 3' is completely similar to that of the
In particular, the
Since the roller 22 'can slide in a radial direction of the axis a with respect to the associated
With reference to fig. 13,
In more detail, each
a
sun gears 97a, 97b connected to
a plurality (three in the case shown) of
planet-
Each
-
a
a plurality (three in the case shown) of
a
The planet gears 99c, 99d are rotatable about their own axis I relative to the
Each
In the case shown, the
The
In the case shown, the
The
The
Furthermore, the
Axis H is radially interposed between axis a and axis I.
The
The
In the case shown, the
Each
an
a
The function of the
In particular, the
In more detail, the
With
Under these conditions, the
If the
The rotation of the planet-
With reference to fig. 14,
The
In particular,
The function of the
The advantages that can be achieved by examining the characteristics of the
In particular, the
In this way, the respective
Since the
The coupling between each
Thus, when the
Conversely, when the
The
The
The
The
Finally, it is also clear that modifications and variants can be made to the
In particular, the
Further, instead of the
Finally, instead of the
Finally, the rotor according to the invention may be the tail rotor 4 of the helicopter 1 instead of the
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