THE CH-47D “CHINOOK” HELICOPTER SHAKE TEST
Tom Rosenberger
Gary Foss
Boeing Defense 8 Space Group
Box 16859, MS P32-15
Philadelphia PA
Boeing Defense & Space Group
Box 3999, MS 86-12
Seattle WA 98124
combination of structural modifications, and both passive
and active vibration absorbers will be used to improve the
CH-47D “Chinook”.
A series of vibration tests of the Boeing Ct%47D tandem
rotor helicopter was performed to vertfy analytical models
and assist in the design and evaluation of efforts to
These efforts included
reduce in-flight vibration.
structural modifications, and the addition of active and
passive vibration absorbers. The test setup and conduct
are described, as wsll as some of the results.
&&,&er Vibration Basics
Helicopter vibration is distinguished by the forced
structural response to periodic inputs. Broadband inputs
are negligible. The periodic inputs are integer multiples
of the rotor speed, typically around 3-4 HZ, times the
number of blades on the rotor. For the CH-47D these
primary input frequencies are three times per rotor
revolution, 3P, six times per rotor revolution, 6P, nine
times per rotor revolution, 9P, and twelve times per rotor
revolution, 12P. The responses of interest are those of
the blades, fuselage bending, engine and power train,
cargo floor, and fuel bladder. Of particular interest are
modes which cause high vibration levels in the cockpit.
Introduction
One of the greatest challenges facing the designers of
rotary wing aircraft is vibration. These aircraft have huge
flexible rotors with tip speeds approaching Mach 1. The
blades are subject to dramatically different aerodynamic
conditions depending on whether they advance or recede
into the wind. To account for this, blade pitch is cycled
for each revolution, further complicating the aerodynamic
Fuselage structures are
forces and responses.
lightweight and flexible, subjected to hub forces produced
by vibrating blades, and exposed to a huge downwash of
air from each passing blade. The combination of these
features almost guarantees vibration problems.
The blade modes are important because they react down
through the rotor head into the airframe. They are fairly
well understood and predictable. Rigid blade flapping
occurs just over once per revolution. Rigid lead lag
happens at less than once per revolution. The first
elastic flapwise bending mode is usually just under 3P.
The first elastic chordwise bending mode is typically
between 2P and 6P, and the first elastic torsion mode is
usually between 3P and t3P. [l]
Flapwise bending
frequencies are sensitive to the centrifugal stiffening
effect. This means a problem with a mode near 3P
doesn’t diminish much with a change in rotor speed,
because the mode will follow right along. In contrast,
chordwise bending and torsion are relatively insensitive
to rotor speed.
Vibration has a detrimental effect on the crew, cargo, and
the aircrafl structure and systems. Sustained vibration
causes crew fatigue, limiting available flight time.
Airframe vibration leads to fatigue cracking of the
structure and component failure which results in high
maintenance costs and aircraft unavailability. The payoff
for helicopter vibration reduction is substantial.
Many helicopters in present service were designed in the
1950’s and 60’s when analytical tools and test equipment
consisted of slide rules and Brush recorders. The Boeing
CH-47D “Chinook” was originally designed in the late
fifties and has undergone a series of modifications and
upgrades, but is structurally very similar to the original
design. The Boeing Company recognizes that successful
future marketing of this venerable helicopter will depend
on reducing fleet operation costs. The most direct route
to this goal is to minimize vibration, Modern
analytical tools such as NASTRAN, and large test
systems are being brought to bear on the effort, A
The helicopter designer’s goal is to make the regularly
spaced peaks of the input force spectrum coincide with
the valleys of the structural frequency response functions,
rather like lining up two picket fences so the pickets don’t
match up. In practice, perfect spacing is never achieved,
but a particular design may be optimized over a range of
rotor RPM and airspeed. For existing designs such as
the Chinook, radical structural changes are not practical,
so we try to add stiffening at critical locations, coupling
this effort with active and passive damping techniques.
497
Test Goals
helicopter (empty of cargo and fuel) was hung from a
spring bank at each end. Suspension frequencies were
verified to be less than two Hertz. A pair of 1200 pound
electrodynamic
shakers were suspended over each rotor
plate and used one or two at a time to apply veriical,
lateral, and longitudinal forces and moments. Fig. 1
shows the suspended helicopter with the shakers
hanging overhead.
A finite element (FE) model of the Chinook was originally
built and correlated for NASA in the early 1960’s. This
early FE model and shake test correlation showed the
existence of two fuselage modes just above the primary
3P forcing frequency of the rotor. This model was
recently used to determine where the fuselage stiffening
would most efficiently increase the frequency of these
two modes and therefore reduce the fuselage response
at the 3P forcing frequency. Stiffening was analytically
applied, and a loads analysis indicated a vibration
reduction of up to 50% was possible. The purpose of the
shake test in the fall of 1995 was to verify the baseline
condition and further correlate the NASTRAN model for a
few higher order modes not clearly identified in the
previous shake test ten years artier.
Since helicopter dynamics is driven by strictly periodic
forces of fairly high magnitude, sine excitation was used
for most of the testing. Step sine software was available
but was judged too time consuming for the sweep width
required (3.5-50 HZ) and the high channel count.
Continuous sweeps from a sweep generator were
processed into frequency response functions using flattop
windowed, high overlap block averaging. This gave good
data quality with a turnaround time of about 45 minutes
per complete data set, including mode animation. A
typical fuselage bending mode at 7.7 HZ is shown in
Figure 2. Figure 3 shows a typical frequency response
function from a response point near the forward
transmission.
After the first test, structural stiffening was applied, and a
second shake test was performed to check the analysis,
develop a new set of response models and predict
whether the expected in-flight vibration reduction would
be achieved.
An important goal of these tests was to develop response
models based on directly measured frequency response
functions [Z]. Given the flight hardware to test in the lab,
we needed to accurately estimate the flight responses at
the measured degrees of freedom for various fuel and
cargo weights and vibration absorber configurations. It
was decided that a directly measured response model of
a complex airframe would be more accurate than the
manipulation of a modal model approximation, particularly
at frequencies well separated from the poles.
All data was gathered with single reference excitation.
The analysis required an accurate response model for
each excitation degree of freedom, contributing to a total
predicted response from multiple simultaneous inputs.
This contrasts with the usual test approach of measuring
multiple inputs to calculate single reference frequency
response functions.
In addition to the sine sweeps for the second test,
customized software was developed in LMS user
programming language to generate and measure
complex excitations and responses. The frequencies of
most interest to the Chinook analysts are one, three, six,
nine. and twelve times the rotor speed of 3.75 HZ. The
significance of these frequencies is clear from the fact
that the Chinook has a three bladed rotor system. The
software allows a user to generate a waveform by dialing
in selected amounts of each harmonic at selected phase
relationships. The resulting function is downloaded to a
recycling digital to analog converter (DAC), and the u.ser
may dial in a fundamental output frequency. On
command, the software captures a block of data for the
force channel and thirty one response channels. A listing
of the actual applied force at each frequency, and a bar
graph of FRF amplitude and phase plotted against
channel number, one graph per harmonic, is displayed.
Another section of the software is designed to suddenly
double the amplitude of the complex excitation and
capture time histories of each response channel. This
was to simulate a rapid change in airspeed for the study
of the transient behavior of the active vibration
suppresser system. The control panel for this software is
shown in Fig. 4.
Recognizing that the required response model need only
be known for certain discrete harmonically related
frequencies, some software was custom developed for
the second shake test. This software allowed a shaker at
the rotor hub to be driven with a complex signal
representing flight-like forces out to the 12th harmonic.
The inputs and responses were efficiently processed
using discrete Fourier transforms at only the harmonics of
interest. This allowed for rapid observation of nonlinearity with amplitude. response sensitivity to small
changes in rotor RPM, and effectiveness and transient
behavior of damping devices.
Test Conduct
Both tests were pedormed using LMS software on an HP
715 workstation with a 64 channel HP 3565 front end.
Two hundred and forty accelerometers and 16 strain
gages were mounted to the aircraft and measured 64 at a
time.
A custom built CMOS multiplexer was used to
switch data sets into the front end. The helicopter blades
were removed and a steel plate of similar mass was
mounted to each rotor hub. The entire 20,000 lb
498
Each of the two tests required three to four weeks. The
test matrix included different configurations of fuel. cargo
weight, and rotor ineltia (by removing weights from the
rotor plate). Gross vehicle weight was tested at a
minimum 22,500 lb., and a maximum 54,000 Ibs. Each
test generated about 12,000 frequency response
functions. The data system was networked with the
analyst’s workstation so data could be delivered within
minutes after the completion of each run. Data was
written to universal files, which were imported into SDRC
IDEAS for analyst review and use.
shake test. Unfortunately, these trials proved that the
analytic model was correct, and that even with a
significant shift of the near-6P modes, the response of
the accelerometers and strain gages was a mixed result.
Although the reduction of the near-6P modal responses
was unsuccessful, the test showed that there were only
two primary force inputs that contributed to the high
response of the accelerations and stresses. Because of
this it was decided that hub mounted pendulum
absorbers were a better solution at 6P. reducing the
input forces rather than juggling the responses.
Test Results:
Pendulum absorbers, or “pendabs”, are absorbers
mounted in the rotating system of the rotor. They are
passive devices designed to use a combination of
pendulum mass and centrifugal stiffness to produce
forces opposite to those being generated by the blades.
They are tuned to the frequency of interest by adjusting
the length of the pendulum. much in the same way the
pendulum of a clock is tuned.
The first or baseline shake test provided a great deal of
important data for both FE model correlation and
There was good
verification of stiffening design.
correlation of the neardP mode shapes between test and
model.
However, the frequency response function
amplitudes were not as well correlated due to the
insufficient amount of damping in the model, a factor that
is always difficult to predict in aillrame structures. The
strain gages were very useful in determining which forces
produced the highest levels of stress, allowing a further
check of the model. This portion of the test gave us the
confidence to proceed with the fuselage stiffening to
increase the frequency of the near-3P modes. T h e
baseline shake test also helped us to correlate and better
define the near-6P modes which, until this test, were not
It was hoped that an improved
clearly understood.
understanding of the near-6P modes would suggest a
solution in the form of structural stiffening, just as it had
for the near-3P modes.
Although the stiffening just completed will reduce the
vibration in the cockpit, even the reduced levels will still
be undesirable for long duration flights.
There are
currently three passive vibration absorbers under the
floor of the cockpit (locked out for most of the FRF
testing). Due to their passive nature they only respond to
the base motion which can provide a certain level of
vibration reduction. However, Boeing is developing an
active vibration suppresser,
AVS: a closed loop system
using accelerometers and reciprocating weights which
could theoretically drive the vibration level to zero. This
system was also tried during the second shake test, to
prove the concept. Although zero vibration was not
attained, the levels were reduced by more than 50% over
that which could be achieved with passive absorbers.
The second shake test showed that the fuselage
stiffening used to raise the neardP mode frequencies
performed as expected in both shifting of the natural
frequencies, and more impoltantly. reducing the response
of the fuselage at the 3P forcing frequency. Another goal
of the second test was to demonstrate a method of
reducing the 6P response of the fuselage. The analytical
methods to reduce 6P response, which were explored
during the installation of the 3P fuselage stiffening.
showed a successful shift in the natural frequencies of
the near-6P modes. Unfortunately, the modal density
and mole complicated behavior at 6P resulted in some
response increases as well as decreases, for little net
gain.
To double check this analysis, some mass and
stiffness perturbations were tried during the second
The aircraft will be flight tested late in 1996 and we will
learn the effectiveness of our modeling, stiffening, testing,
and absorber development.
We will learn what
percentage reduction in vibration and strain can
eventually be achieved on this old venerable workhorse.
References:
[l] Prouty. R.W., Helicopter Vibration, Sound and
Vibration, Nov. 1968, pg 34
[Z] Ewins, D.J., Modal Testing: Theory and Practice..
Research
Study
Press,
1966,
207.
PS
499
Figure 1. Photo showing helicopter suspended from spring banks and location of shakers
Figure 2. Typical fuselage bending mode at 7.7 HZ
500
Figure 3. Typical sine sweep frequency response fun&m lrom a point near foreward transmission.
Figure 4. Control panel for software to simulale flight-like forces.
501