Helicopter and method for reducing vibration of a helicopter fuselage

Abstract: A helicopter structure (10; 10 1 ) comprises a fuselage (12; 12 1 ) supporting or supported by a rotating system (11; 11 1 ) parts (12, 17; 12 1 15 1 ) of the structure (10; 10 1 ) being capable of relative motion at an exciting frequency, there being a plurality of actuators (19; 19 1 ) connected between the relatively movable parts (12, 17; 12 1 , 15 1 ) of the structure (10; 10 1 ), means (20, 22; 22 1 ) continuously to oscillate the actuators (19; 19 1 ) at a frequency substantially corresponding to the exciting frequency, and a plurality of sensors (21; 21 1 ) attached to the rotating system (11; 11 1 ) at selected locations on the rotating system (11; 11 1 ), the sensors (21; 21 1 ) being adapted to generate signals representative of dynamic changes at the respective selected locations, processing means (22; 22 1 ) adapted to process the signals from the sensors (21; 21 1 ) and to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators (19; 19 1 ) and for varying the phase and magnitude characteristics of the forces so as to compensate for changes in the dynamic characteristic of the rotating system (11; 11 1 ), whereby the overall level of vibration in the fuselage structure (12; 12 1 ) is reduced.

Description: Description of the Invention The dominant source of vibration in a helicopter in forward flight is that generated by the main sustaining rotor rotating system at the blade passing frequency. Forces and moments are transmitted usually through the transmission via fuselage attachments, to produce vibration in the fuselage. Our previous US patent 4819182 describes a method of reducing vibration of a helicopter fuselage by connecting a plurality of actuators between parts of a helicopter structure capable of relative motion at an exciting frequency, e.g. the blade passing frequency. That arrangement utilises a plurality of accelerometers attached to the helicopter fuselage at a plurality of locations, which generate signals representative of fuselage dynamic accelerations. These signals are processed by on-board processing means which provide output signals for controlling phase and magnitude of the output forces generated by the actuators between the relatively vibrating parts of the structure. However, in an arrangement such as described in US 4819182, the accelerometers and associated cabling are prone to accidental damage. This arrangement provides advantages over other prior proposals for absorbing vibrations, which include mechanical absorbers mounted above a head of the rotor, which are effective because vibrations are "absorbed" at or closely adjacent to the source of those vibrations. However such devices require a parasitic mass of considerable magnitude which provides an unacceptable weight penalty. According to a first aspect of the invention we provide a helicopter comprising a structure including a fuselage supporting or supported by a rotating system, parts of the structure being capable of relative motion at an exciting frequency, there being a plurality of actuators connected between the relatively movable parts of the structure, means continuously to oscillate the actuators at a frequency substantially corresponding to the exciting frequency, and a plurality of sensors attached to the rotating system at selected locations, the sensors being adapted to generate signals representative of dynamic changes at the respective selected locations during rotation, processing means adapted to process the signals from the sensors and to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators and for varying the phase and magnitude characteristics of the applied forces so as to compensate for changes in the dynamic characteristic of the rotating system whereby the overall level of vibration in the fuselage is reduced. Thus the arrangement of the present invention provides many of the advantages of the system described in our prior US patent 4819182, but also, as the sensors are provided on the rotating system, the invention simulates the act of a head mounted mechanical absorber but with a much lower weight penalty. Further, by providing the sensors on the rotating system, rather than on the helicopter fuselage, the sensors are able to sense vibrations occurring in the rotating system i.e. at the source of the excitation. The sensors may be mounted on the rotating system such that the sensors and associated cabling are less prone to accidental damage than would be the case where the sensors are mounted on the fuselage. The invention is particularly applicable where the rotating system is the main sustaining rotor of the helicopter which is the dominant source of vibrations in a helicopter fuselage. One part of the structure which is capable of motion relative to another part of the structure may comprise one of or an assembly of more than one of an engine, a transmission, and a supporting structure of the rotating system. Conveniently the one part of the structure is attached to the another part of the structure by plurality of resilient attachment means, such as elastomeric units. The plurality of resilient attachment means and actuators are preferably equal in number, and if desired, each actuator may be an integrated unit with a resilient attachment. Whereas any suitable kind of actuator which is able to apply force between the relatively movable parts of the structure may be utilised, preferably each actuator is an electro-hydraulic actuator. In one arrangement, where the rotating system comprises a rotor head to which are affixed a plurality of rotor blades, at least one of the sensors may be mounted on the rotor head of the rotating system, although could alternatively be mounted on a rotor driveshaft, which commonly is generally vertical, or on a rotor blade which is generally horizontal. Although different kinds of sensors may be utilised, such as strain gauges and displacement transducers, preferably the sensors comprise accelerometers which are compact, reliable and easily installed. According to a second aspect of the invention we provide a method of reducing vibration in a helicopter structure which includes a fuselage which is supporting or supported by a rotating system, parts of the structure being capable of relative motion at an exciting frequency, the method comprising the steps of connecting a plurality of actuators between the relatively movable parts of the structure, oscillating the actuators at a frequency substantially corresponding to the exciting frequency, generating signals representative of dynamic changes at a plurality of locations on the rotating system during rotation and feeding the signals to processing means which are adapted to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators and for varying the phase and magnitude characteristics of the forces so as to compensate for changes in the dynamic characteristic of the rotating system, whereby the overall level of vibration in the fuselage is reduced. According to a third aspect of the invention we provide an apparatus for reducing vibration of a structure of a helicopter which comprises a fuselage supported by a rotating system, and in which parts of the structure are capable of relative motion at an exciting frequency, the apparatus comprising a plurality of actuators adapted to be connected between the relatively movable parts of the structure, means in use continuously to oscillate the actuators at a frequency substantially corresponding to the exciting frequency, and a plurality of sensors adapted to be attached to the rotating system at selected locations, the sensors in use, being adapted to generate signals representative of dynamic changes at the respective selected locations during rotation, the apparatus further comprising processing means adapted to process the signals from the sensors and to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators and for varying the phase and magnitude characteristics of the forces so as to compensate for changes in the dynamic characteristic of the rotating system. The invention will now be described with reference to the accompanying drawings in which:- Figure 1 is an illustrative fragmentary side elevation of a helicopter fitted with a vibration reducing apparatus according to the invention; Figure 2 is a simplified block diagram showing operation of the apparatus of Figure 1; and Figure 3 is a fragmentary illustrative view similar to Figure 1 but of an alternate embodiment. Referring to figures 1 and 2 of the drawings, a helicopter structure generally indicated at 10 includes a rotating system comprising a main sustaining rotor 11 supported by a supporting structure comprising a raft 17, and a helicopter fuselage 12, having rearwardly extending tail part 13 carrying an anti torque rotor 14. The main sustaining rotor 11 is driven through a transmission (gear box) 15 by one or more engines 16, the transmission 15 and engine(s) 16 being mounted on the raft 17. The raft 17 is in this example, generally rectangular and is attached to the fuselage 12 by, in this example, four resilient attachments 18, each of which comprises an elastomeric unit 18 through which operational forces are transmitted from the raft 17 to the fuselage 12, located adjacent a corner of the rectangle. Thus the fuselage 12 and raft 17 comprise parts of the helicopter structure which are capable of relative motion at a frequency corresponding substantially with a vibration exciting frequency. An electro-hydraulic actuator 19 is connected between the fuselage 12 and raft 17, adjacent each of the four elastomeric attachment units 18, and each actuator 19 is adapted during operation, to apply a force to the fuselage 12 that is reacted by an equal and opposite force on the raft 17. In a preferred arrangement, the actuators 19 are each an integrated unit with a respective elastomeric attachment unit 18, such that the number of actuators 19 is equal to the number of resilient attachment units 18. A supply of pressurised hydraulic fluid 20 is connected to each of the actuators 19. A plurality of sensors 21, are attached at various locations to the rotor 11. In this example, four sensors 21 are provided, one on each of the four rotor blades of the rotor 11, although only two can be seen in the drawings. The sensors 21 comprise in this arrangement, accelerometers which are adapted to generate signals representative of dynamic changes at the respective selected locations as the rotor 11 rotates. In another arrangement in which the rotating system 11 has other than four rotor blades, one sensor may be provided for each rotor blade. The accelerometers 21 are operatively connected to a processing means comprising a computer 22 which is carried in the fuselage 12, the computer 22 being adapted to provide output signals for controlling the phase and magnitude of the applied forces generated by the actuators 19 and for varying the phase and magnitude characteristics of the forces. In operation, vibratory forces are produced by the main sustaining rotor 11 due, for example, to asymmetric air flow in forward flight. Such vibratory forces arising as the rotor 11 rotates are, in the absence of any compensating systems, transmitted from the raft 17 to the fuselage 12. These forces and related moments are predominantly at the blade passing frequency and harmonics thereof, the blade passing frequency being a product of the rotational speed of the rotor 11 and the number of rotor blades. Typically, a blade passing frequency for a four-bladed helicopter would be in the order of 21Hz and for a five bladed helicopter, 17Hz. The actuators 19 are oscillated at a frequency controlled by the computer 22, so as to apply a set of forces between the fuselage 12 and raft 17 to counteract the vibrations arising in the raft 17 due to the rotation of the rotor 11. Thus where the blade passing frequency is 21Hz, preferably the electro-hydraulic actuators 19 are continuously oscillated at a frequency of about 21Hz. In addition, dynamic changes occurring in the rotor 11 at the locations of the accelerometers 21 cause adjustment of the phase and magnitude of the output signals from the computer 22. The computer 22 is adapted to analyse vibration signals received from the accelerometers 21 for example, by a digital signal processor 23 based on a discrete Fourier transform theory for rotor blade passing frequency and phase information. This measured frequency domain vibration data is then fed to a parameter estimator 24 which utilises the information to construct a linear transfer relationship for the vibration responses and the output forces from the actuators 19. The calculation of the transfer relationship estimates are based on discrete Kalman filter theory. The resultant estimates are fed to an optimal controller 25 which calculates the optimal control forces which minimise a quadratic performance function that comprises the weighted sum of the squares of the measured vibrations and the actuator output forces and produces appropriate output signals 26 to operate the actuators 19. This control procedure continues on a cyclic basis wherein during one cycle a constant set of oscillatory forces are applied to the fuselage 12 as the calculations are conducted. The use of a parameter estimator 24 in the control loop of the computing apparatus 22 ensures that the phase and magnitude response characteristics of the set of output signals 26 to the actuators 19, and therefore the actuator output forces, are continuously varied to cater for changing dynamic characteristics in the rotor 11 or fuselage 12. Thus in the event of a change in the linear transfer relationship, the parameter estimator 24 detects an error between the predicted vibration level and the measured vibration level. The estimates for the relationship are adjusted accordingly and substituted into the optimal controller 25 which then calculates a new set of actuator signals 26. Thus the method and apparatus of the invention imposes changes in the phase and magnitude of the exciting frequency oscillations of the electro-hydraulic actuators 19 simultaneously to cancel or at least substantially reduce the level of vibration transmitted from the raft 17 to the fuselage 12 thereby providing a significant improvement in the overall vibration level of the helicopter fuselage 12. Of course, some means such as a slip ring system, would be required to transfer the signals from the accelerometers 21 which are on the rotor 11 to the computer 22 in the fuselage 12. By virtue of the accelerometers 21 being on the rotor 11, the accelerometers 21 sense vibration at the source of excitation and so simulate an arrangement in which a parasitic mass is provided on a rotor head mounted mechanical vibration absorber, although of course the accelerometers 21 impose a much lower weight penalty. The invention is capable of reducing vibrations transmitted from the raft 17 to the fuselage 12 caused by other mechanisms of the rotating system e.g. the higher harmonics of blade passing frequency. Various modifications may be made without departing from the scope of the invention. For example, the number of actuators 19 provided between the raft 17 and fuselage 12 can be varied. In the present example, the actuators 19 are oriented generally vertically (with the helicopter flying in a level plane) although may be otherwise oriented. The number of accelerometers 21 provided may be altered as is necessary but preferably the number of accelerometers 21 will not be less than the number of actuators 19. One or more or all of the accelerometers 21 may be mounted otherwise on the rotor 11 than described e.g. on a rotor head H of the rotor 11, or on a ( usually generally vertically oriented) drive shaft D. Figure 3 shows an alternate embodiment with similar parts of the helicopter of figure 1 indicated by the same reference numerals, but with a prime sign added. The figure 3 embodiment does not have a raft like raft 17 in figure 1. Rather, a helicopter structure 10 1 comprises a fuselage 12 1 carrying a gearbox 15 1 which transmits drive to a rotating system comprising a main sustaining rotor 11 1 , to drive the rotor 11 1 about an axis A. The gearbox 15 1 is attached to the fuselage 12 1 by means of a plurality of resilient strut assemblies 18 1 , each of which includes an actuator 19 1 . The actuators 19 1 are, like actuators 19 in figure 1, of electro-hydraulic construction receiving operating signals 26 1 from a processing means 22 1 located in the fuselage 12 1 . The processing means 22 1 receives input signals from sensors 21 1 again comprise accelerometers mounted on the rotor blades of the rotor 11 1 , or otherwise on the rotor 11 1 , to vary the phase and magnitude of exciting forces, provided by the actuators 19 1 . Thus in this arrangement, the parts of the helicopter structure which are capable of relative motion comprise the gearbox 15 1 and the fuselage 12 1 , and the resilient strut assemblies 18 1 are arranged at about 45° to the vertical, with the helicopter in level flight. It will be appreciated that in another embodiment, the invention may be utilised to reduce vibration levels in the fuselage 12 or 12 1 , by connecting actuators such as 19, 19 1 between any parts of the helicopter structure which are capable of relative motion at an exciting frequency. In each case, the sensors preferably comprise accelerometers 21, 21 1 , but one or more of all of the sensors could alternatively comprise for example, a strain gauge mounted on a rotor drive shaft D (figure 1) D 1 (figure 3) of the rotor 11, 11 1 , to measure rotor drive shaft D, D 1 , bending for example, or displacement transducers. Accelerometers are preferred as they are compact, reliable and easily installed, whereas strain gauges may be too vulnerable to external and fatigue damage to enable them to form a viable sensor, whilst displacement transducers could be subject to wear and reliability problems in use. If desired, one or more additional sensors may be provided on the helicopter fuselage 12, 12 1 to generate signals representative of fuselage dynamic accelerations, such signals being fed to the computer 22, 22 1 where they are used in the calculations which result in the output signals for controlling the phase and magnitude of the output forces generated by the actuators 19, 19 1 between the relating vibrating parts of the structure. Thus the actuators 19, 19 1 output forces cater not only for changing dynamic characteristics of the rotor 11, 11 1 but of the fuselage too. If desired, the invention may alternatively or additionally be applied to a rotating system comprising the anti-torque rotor 14 of the helicopter 10, in which case sensing would be attached to that rotating system 14.

Claims: A helicopter comprising a structure (10; 10 1 ) including a fuselage (12; 12 1 ) supporting or supported by a rotating system (11; 11 1 ), parts (12, 17; 12 1 15 1 ) of the structure (10; 10 1 ) being capable of relative motion at an exciting frequency, there being a plurality of actuators (19; 19 1 ) connected between the relatively movable parts (12, 17; 12 1 , 15 1 ) of the structure (10; 10 1 ), means (20, 22; 22 1 ) continuously to oscillate the actuators (19; 19 1 ) at a frequency substantially corresponding to the exciting frequency, and a plurality of sensors (21; 21 1 ), processing means (22; 22 1 ) adapted to process the signals from the sensors (21; 21 1 ) and to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators (19; 19 1 ) characterised in that the plurality of sensors (21,21') are attached to the rotating system (11; 11 1 ) at selected locations and are adapted to generate signals representative of dynamic changes at the respective selected locations during rotation, and in that the processing means (22; 22 1 ) provides output signals to the actuators (19; 19 1 ) which vary the phase and magnitude characteristics of the applied forces of the actuators (19; 19 1 ) so as to compensate for changes in the dynamic characteristic of the rotating system (11; 11 1 ) whereby the overall level of vibrations of the fuselage (12; 12 1 ) is reduced. A helicopter according to claim 1 characterised in that the rotating system (11; 11 1 ) is the main sustaining rotor (11; 11 1 ). A helicopter according to claim 1 or claim 2 characterised in that one part of the structure (17; 15 1 ) which is capable of motion relative to another part (12; 12 1 ) of the structure (10; 10 1 ) comprises one of or an assembly of more than one of an engine (16), a transmission (15; 15 1 ), and a supporting structure (17) of the rotating system (11; 11 1 ). A helicopter according to claim 3 characterised in that the one part (17; 15 1 ) of the structure (10; 10 1 ) is attached to the another part (12; 12 1 ) of the structure (10; 10 1 ) by a plurality of resilient attachment means (18; 18 1 ), the plurality of resilient attachment means (18; 18 1 ) and the actuators (19; 19 1 ) being equal in number and/or each actuator (19; 19 1 ) is an integrated unit with a resilient attachment (18; 18 1 ). A helicopter according to any one of the preceding claims characterised in that each of the actuators (19; 19 1 ) is an electro-hydraulic actuator. A helicopter according to any one of the preceding claims characterised in that the rotating system (11; 11 1 ) comprises a rotor head (11) to which are affixed a plurality of rotor blades, and at least one of the sensors (21; 21 1 ) is mounted on the rotor head (H) for example on a rotor drive shaft (D) of the rotating system (11; 11 1 ). A helicopter according to any one of the preceding claims characterised in that the rotating system (11; 11 1 ) comprises a rotor head (H) to which are affixed a plurality of rotor blades, and at least one of the sensors (21; 21 1 ) is mounted on a rotor blade, preferably there being a sensor (21; 21 1 ) provided for each rotor blade of the rotating system (11; 11 1 ). A helicopter according to any one of the preceding claims characterised in that in addition to the sensors (21; 21 1 ) of the rotating system (11; 11 1 ), one or more sensors are provided on a fuselage of the helicopter to generate signals representative of fuselage dynamic accelerations. A method of reducing vibration in a helicopter structure (10; 10 1 ) which includes a fuselage (12; 12 1 ) which is supporting or supported by a rotating system (11; 11 1 ), parts (12, 17; 12 1 , 15 1 ) of the structure (10; 10 1 ) being capable of relative motion at an exciting frequency, the method comprising the steps of connecting a plurality of actuators (19; 19 1 ) between the relatively movable parts (12, 17; 12 1 , 15 1 ) of the structure (10; 10 1 ), oscillating the actuators (19; 19 1 ) at a frequency substantially corresponding to the exciting frequency, generating signals representative of dynamic changes at a plurality of locations on the rotating system (11; 11 1 ) during rotation and feeding the signals to processing means (22; 22 1 ) which are adapted to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators (19; 19 1 ) and for varying the phase and magnitude characteristics of the forces so as to compensate for changes in the dynamic characteristic of the rotating system (11; 11 1 ), whereby the overall level of vibration in the fuselage (12; 12 1 ) is reduced. An apparatus for reducing vibration of a structure (10; 10 1 ) of a helicopter which comprises a fuselage (12; 12 1 ) supporting or supported by a rotating system (11; 11 1 ), and in which parts (12, 17; 12 1 , 11 1 ) of the structure (10; 10 1 ) are capable of relative motion at an exciting frequency, the apparatus comprising a plurality of actuators (19; 19 1 ) adapted to be connected between the relatively movable parts (12, 17; 12 1 , 11 1 ) of the structure (10; 10 1 ), means in use, continuously to oscillate the actuators at a frequency substantially corresponding to the exciting frequency, and a plurality of sensors (21; 21 1 ) adapted to be attached to the rotating system (11; 11 1 ) at selected locations on the rotating system, the sensors (21; 21 1 ) in use, being adapted to generate signals representative of dynamic changes at the respective selected locations during rotations, the apparatus further comprising processing means (22; 22 1 ) adapted to process the signals from the sensors (21; 21 1 ) and to provide output signals for controlling the phase and magnitude of applied forces generated by the actuators (19; 19 1 ) and for varying the phase and magnitude characteristics of the forces so as to compensate for changes in the dynamic characteristic of the rotating system (11; 11 1 ).

Inventors: VINCENT ALAN HENRY, VINCENT, ALAN HENRY
Applicants: GKN WESTLAND HELICOPTERS LTD, GKN WESTLAND HELICOPTERS LIMITED
Application year: 1996
Main classification: B64C27
IPC codes: F16F15, B64C27, F16F7