Helicopter and method for reducing vibration of a helicopter fuselage

A helicopter structure comprises 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 on the rotating system, the sensors being adapted to generate signals representative of dynamic changes at the respective selected locations, 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, whereby the overall level of vibration in the fuselage structure is reduced.

BACKGROUND 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 U.S. Pat. No. 4,819,182 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 U.S. Pat. No. 4,819,182, 
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. 
SUMMARY OF THE INVENTION 
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 U.S. Pat. No. 4,819,182, 
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.

DESCRIPTION OF THE PREFERRED EMBODIMENTS 
Referring to FIGS. 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 21 Hz and for a five bladed helicopter, 17 Hz. 
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 21 Hz, preferably 
the electro-hydraulic actuators 19 are continuously oscillated at a 
frequency of about 21 Hz. 
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. 
FIG. 3 shows an alternate embodiment with similar parts of the helicopter 
of FIG. 1 indicated by the same reference numerals, but with a prime sign 
added. The FIG. 3 embodiment does not have a raft like raft 17 in FIG. 1. 
Rather, a helicopter structure 10.sup.1 comprises a fuselage 12.sup.1 
carrying a gearbox 15.sup.1 which transmits drive to a rotating system 
comprising a main sustaining rotor 11.sup.1, to drive the rotor 11.sup.1 
about an axis A. The gearbox 15.sup.1 is attached to the fuselage 12.sup.1 
by means of a plurality of resilient strut assemblies 18.sup.1, each of 
which includes an actuator 19.sup.1. The actuators 19.sup.1 are, like 
actuators 19 in FIG. 1, of electro-hydraulic construction receiving 
operating signals 26.sup.1 from a processing means 22.sup.1 located in the 
fuselage 12.sup.1. The processing means 22.sup.1 receives input signals 
from sensors 21.sup.1 again comprise accelerometers mounted on the rotor 
blades of the rotor 11.sup.1, or otherwise on the rotor 11.sup.1, to vary 
the phase and magnitude of exciting forces, provided by the actuators 
19.sup.1. Thus in this arrangement, the parts of the helicopter structure 
which are capable of relative motion comprise the gearbox 15.sup.1 and the 
fuselage 12.sup.1, and the resilient strut assemblies 18.sup.1 are 
arranged at about 45.degree. to the vertical, with the helicopter in level 
flight. 
It will be appreciated that an outer embodiment described, the invention 
may be utilised to reduce vibration levels in the fuselage 12 or 12.sup.1, 
by connecting actuators such as 19, 19.sup.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.sup.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 (FIG. 1) D.sup.1 
(FIG. 3) of the rotor 11, 11.sup.1, to measure rotor drive shaft D, 
D.sup.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.sup.1 to generate signals representative of 
fuselage dynamic accelerations, such signals being fed to the computer 22, 
22.sup.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.sup.1 between the relating 
vibrating parts of the structure. 
Thus the actuators 19, 19.sup.1 output forces cater not only for changing 
dynamic characteristics of the rotor 11, 11.sup.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.