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and the ability to cater for variations in rotor speed can be controlled by the
amount of damping built into the DAVI units.
(iv) The Bell ‘Nodamatic’ gearbox mounting system3
This method (Fig. 8.12) interposes a beam mounting arrangement between the
gearbox and the fuselage and is configured such that the fuselage is effectively
suspended from the node points of the beam system when it is vibrating in
response to the rotor hub forcing loads and moments. In fact, the system is
mathematically equivalent to the DAVI principle, with the beam system providing
the stiffness and inertia properties of the DAVI units.
8.5 Dynamic response of the fuselage
Due to the fundamental design requirements of the helicopter primarily in terms of
the requirements of crew and equipment, and aerodynamic drag, the basic shape of
the fuselage will be determined by considerations other than its vibrational
characteristics. In the early stages of design, the fuselage can be represented by an
assemblage of beams of defined bending and torsional stiffness properties (Fig. 8.13)
together with the appropriate mass distribution.
Calculations can then be performed which will:
(i) indicate the rough proximity of any major beam bending mode frequency to
the bΩ rotor forcing frequency; and
(ii) indicate the sensitivity of the forced response to changes in the stiffness of
structural components which may be amenable to significant alteration, e.g.
DETAIL FROM A
TYPICAL DESIGN
Elastomeric Sandwich
(Rubber/Steel Sandwich)
Gearbox Mounting
Strut
Vibratory
Loads
Beam Nodes
Location on Helicopter
Airfame
Spherical Bearing
Elastomeric
Bearing
302 Bramwell’s Helicopter Dynamics
Excitation
87
13 12 11
6
5 4 9
19 20
14
16
17
1 2
Pilot’s
position
P18 3
10
15
Fig. 8.13 Fuselage beam (stick) model
Fig. 8.14 Army Lynx with struts on tailcone
main rotor head stiffness relative to the fuselage, and tail boom stiffness.
Thus, if ground vibration or flight tests indicate evidence of a major forced
response problem, it may be possible to significantly reduce the response by
making an acceptable change to the structural stiffness in a particular area
(Figs 8.14 and 8.15).
Rotor induced vibration 303
Fig. 8.15 Lynx cockpit 4R vibration as a function of strut stiffness
Fig. 8.16 Structural manipulation
An extension of this approach referred to as ‘structural manipulation’ has been developed
whereby an order of ranking of areas of structural sensitivity is determined (Fig.
8.16). The method4 then optimises the values of stiffness changes required in a
chosen number of structural elements to minimise a defined ‘index of performance’
which could be in the form of the vibrational response in a designated area of the
fuselage due to specified forcing components applied at a defined location.
AREA OF EACH STIFFENING STRUT (IN2)
0.1 0.2 0.3 0.4 0.5 0.6
0.01
0.0075
0.005
0.0025
0
VIBRATION
LEVEL (IN)
BASIS:
The locus of a single
point response to
a single point
excitation when a
single structural
parameter is
varied is a circle.
This property
forms the basis of
optimisation
routine.
Locus of Response
Vector as K Varies
Sensitivity Using
‘Circle’ Criteria
Response
Lynx Stick Model
K
Excitation
Imag.
K
Real
Fuselage Location
Gearbox Area
Engines
Sides,
Roof
Floor
Tail Boom
304 Bramwell’s Helicopter Dynamics
When the structural design of the fuselage is at a more advanced stage, a finite
element method of dynamic analysis such as NASTRAN5 will be used to predict
more accurately the dynamic response characteristics of the airframe. Figure 8.17
indicates a typical finite element model.
Due to difficulties in assessing the stiffness properties of structural joints, the
results of such calculations will be used to predict trends in the dynamic behaviour
of the fuselage as a result of possible changes in stiffness and mass distribution.
Following manufacture and vibration test of a prototype airframe, the finite element
model would be empirically adjusted to match the test results. The use of the adjusted
model will increase confidence in subsequent analytical predictions, from which
beneficial design changes may result.
8.6 Vibration absorbers
Having made all the beneficial stiffness and mass distribution changes which are
　
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