曝光臺 注意防騙 網曝天貓店富美金盛家居專營店坑蒙拐騙欺詐消費者

gearbox and fuselage interface is very effective, and Fig. 8.28 shows a method of
gearbox attachment which combines the mounting struts with the actuators.
The basic equation relating the various parameters is identical in form to that for
HHC,
i.e. Y = TX + B
where Y is the measured fuselage vibration
X is the vector of actuator forces
Airframe vibration
Fig. 8.27 Concept of ACSR
314 Bramwell’s Helicopter Dynamics
Fig. 8.29 ACSR frequency domain control strategy
T is the transfer matrix relating the actuator forces to the fuselage vibration
B is the background, uncontrolled, vibration
In general, with N control forces, the response at N locations in the fuselage can
be reduced to zero, provided that the T matrix is non-singular. However, it is
considered preferable to attempt to reduce the vibration at a larger number of
locations to acceptably low levels rather than to attempt to achieve zero vibration
at a few positions.
Two categories of control algorithm are applicable to the implementation of
ACSR. These may be classified as either frequency or time domain in nature.
Emphasis has been placed on the frequency domain approach in the early applications
of this technique.
The general arrangement for the frequency domain control strategy is shown in
Fig. 8.29. This indicates that the primary functions of the controller are signal processing,
parameter estimation and control.
As a very broad statement, it appears that active control techniques can produce
vibration levels which are in the region of about one-half of the levels achieved by
passive methods, and this has indicated that the long-term goal of the ‘jet smooth
ride’ helicopter may at last be a possibility.
Figure 8.30 shows a comparison of the vibration levels of the Westland W30
helicopter without a vibration reduction system, and when fitted with a Flexispring
rotor head absorber, and an ACSR system.
8.8 Vibration at frequencies other than bΩ
The minimisation of forcing frequency components from the rotor system which are
Helicopter flight conditions Main
rotor
Rotor head forces
Fuselage Airframe vibration
dynamics
Controlling forces
Actuators Accelerometers
Measured
vibration
Actuator
demand
Actuator
loop
closure
Frequency/
time
Optimal
controller
Dynamics
estimator
Signal
processor
time/
frequency
Feedback
force
Adaptive controller
Rotor induced vibration 315
normally self-cancelling within the rotor is very much related to the ability to
manufacture identical blades and lag dampers. The major residual is usually at 1Ω
frequency, and this is minimised by blade balancing and tracking procedures on a
whirl tower and on the helicopter during ground running, and if necessary also in
hovering and high-speed forward flight.
Reasonably identical lag plane damper performance characteristics are required so
that undesirable vibration under conditions of 1Ω blade flapping (which results in
large lag plane oscillations) can be avoided.
A quite distinct type of vibration problem which is of aerodynamic origin is due
to the effects of vortices shed from the region near the main rotor head which can,
under certain flight conditions, strike the tail rotor and the fixed vertical and horizontal
tail surfaces. This phenomenon has been observed mainly as a response at the fuselage
fundamental lateral bending frequency and is particularly dependent on the sideslip
angle in flight. This problem may be intensified by the addition of excrescences such
as a radome to the upper surface of the fuselage aft of the rotor.
This problem is often referred to as the ‘lateral shakes’ or the ‘shuffle’. Solutions
to this problem have been found by the addition of suitable fairings to the main rotor
head and to the airframe aft of the head. Figures 8.31 and 8.32 indicate the origin of
Head absorber
Baseline
ACSR
Forward speed (knots)
0.5
0.4
0.3
0.2
0.1
0
Average cabin/cockpit vibration (g)
Fig. 8.30 Comparison of Westland W30 vibration levels
40 60 80 100
Turbulent wake from pylon/rotor head
area affects fin/tail rotor
Fig. 8.31 Origin of the ‘shuffle’ problem
316 Bramwell’s Helicopter Dynamics
Flow curvature caused by beanie
Fig. 8.32 Effect of ‘beanie’ on the turbulent wake
Fig. 8.33 Spectra of vibration amplitudes
0.5
0.4
0.3
0.2
0.1
0
mm
1 2 3 4 5 6
Lateral
　
中國航空網 www.k6050.com
航空翻譯 www.aviation.cn
本文鏈接地址：Bramwell’s Helicopter Dynamics(155)