Department of Aerospace Engineering

International Institute for Aerospace Engineering & Management

Jain Global Campus, Kanakapura Taluk,
Ramanagara District, Karnataka, India -562112

2018-2019

A Dissertation on

VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

Submitted in partial fulfilment for the award of the degree of

MASTER OF TECHNOLOGY

IN

AEROSPACE STRUCTURES ENGINEERING

Submitted by

Ms. ARCHANA M
17MTRAS011

Under the guidance of

Mr. M Jeeva Peter

Department of Aerospace Engineering
International Institute of Aerospace Engineering and Management
JAIN (Deemed-to-be University)
 Department of Aerospace Engineering
International Institute for Aerospace Engineering & Management
Jain Global Campus, Kanakapura Taluk,
Ramanagara District, Karnataka, India -562112

2018-2019

A Dissertation on

VIBRATION ANALYSIS O N MAIN ROTOR BLADE OF

HELICOPTER

SUBMITTED IN PARTIAL FULFILMENT FOR THE AWARD OF THE DEGREE

OF

MASTER OF TECHNOLOGY
In

AEROSPACE STRUCTURES ENGINEERING

Submitted by

Ms. ARCHANA M
17MTRAS011

Under the guidance of

Mr. M Jeeva Peter

Department of Aerospace Engineering
International Institute of Aerospace Engineering and Management
JAIN (Deemed-to-be University)
 Department of Aerospace Engineering

International Institute for Aerospace Engineering & Management

Jain Global Campus, KanakapuraTaluk,
Ramanagara District, Karnataka, India -562112

CERTIFICATE

This is to certify that the dissertation entitled “VIBRATION ANALYSIS ON MAIN

ROTOR BLAE OF HELICOPTER” is carried out by Ms. ARCHANA M (17MTRAS011),
a bonafide student of Master of Technology, at the International Institute for Aerospace
Engineering & Management, JAIN(Deemed-to-be University) in partial fulfilment for the
award of the degree of Master of Technology in Aerospace Structures Engineering,
during the year 2018-2019.

Mr. M Jeeva Peter Dr. Antonio Davis Dr. Manoj Veetil

Department of Aerospace Head of Dept, Director,
Engineering, Department of Aerospace International Institute
International Institute for Engineering, for Aerospace
Aerospace Engineering & International Institute for Engineering &
Management, Aerospace Engineering & Management,
JAIN(Deemed-to- Management, JAIN(Deemed-to-
be University). JAIN(Deemed-to-be University). be University).

Date: Date: Date:

Name of the Examiner Signature of Examiner

1.

2
 DECLARATION

I, Ms. ARCHANA M, a student of fourth semester M. Tech in Aerospace Structures

Engineering, at International Institute for Aerospace Engineering &

Management, JAIN(Deemed-to-be University) hereby declare that the dissertation

entitled “VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER”

has been carried out by me and submitted in partial fulfilment for the award of the degree

of Master of Technology in Aerospace Structures Engineering, during the academic

year 2018-2019. Further, the matter embodied in the dissertation has not been submitted

previously by anybody for the award of any degree or diploma to any University, to the

best of my knowledge and faith.

PLACE: IIAEM, JAIN GLOBAL CAMPUS ARCHANA M

DATE: 17-06-2019 17MTRAS011
 Dedicated
To My Beloved
MOM
 ACKNOWLEDGEMENT

It is a great pleasure for us to acknowledge the assistance and support of a large

number of individuals who have been responsible for the successful completion of this
dissertation work.
First of all, I thank Almighty God for all his blessings and thanks to my parents for
giving encouragement, enthusiasm and invaluable assistance to me.
Second, I take this opportunity to express my sincere gratitude to School of
Engineering & Technology, JAIN (Deemed-to-be University) for providing me with a great
opportunity to pursue my Masters Degree in this institution.
In particular I would like to thank Dr. Hariprasad S. A.,Director,School of
Engineering & Technology, JAIN (Deemed-to-be University) for his constant
encouragement and expert advices.
In particular I would like to thank Dr. Rajasimha A Makaram, Associate Director,
School of Engineering & Technology, JAIN (Deemed-to-be University) for his constant
encouragement and support.
It is a matter of immense pleasure to express my sincere thanks and gratitude to my
guide Prof. M Jeeva Peter, Assistant Professor and In-charge head, Department of Food
Technology, School of Engineering & Technology, JAIN (Deemed-to-be University) for
providing right academic guidance and sparing his valuable time to extend help in every step
of my dissertation work, which paved the way for smooth progress and fruitful culmination of
the project.
I would like to thank all the lecturers and also all the staff members of Food
Technology Department for their support and contributions.
My acknowledgment will never be complete without the special mention of my special
ones, my family and my best friends Prudhviraj, Krishna, Kavya, Varsha, Adithya, Anju,
Anish, Kiran, Srinivas, Mani, Sai and Ramkumar for their constant support, love and
affection.
I would like to thank one and all who directly or indirectly helped me in completing
the Dissertation work successfully.
 CONTENT
CHAPTER DESCRIPTION PAGE NO.
• CONTENT i
• LIST OF FIGURES ii
• ABSTRACT iv
CHAPTER 1: INTRODUCTION 1
CHAPTER 2: LITERATURE SURVEY 7
CHAPTER 3: OBJECTIVE 12
CHAPTER 4: METHODOLOGY 13
CHAPTER 5: DIMENSIONS AND DESIGN OF THE ROTOR
BLADE 14
CHAPTER 6: EQUATIONS OF MOTION AND ITS
CALCULATIONS 16
CHAPTER 7: STATIC STRUCTURAL ANALYSIS 18
CHAPTER 8: MODAL ANALYSIS 22
CHAPTER 9: HARMONIC ANALYSIS 30
CHAPTER 10: FATIGUE CYCLE CALCULATION 34
CHAPTER 11: CONCLUSION 36
REFERENCES 37

i
 LIST OF FIGURES
FIG. NO. DESCRIPTION PAGE NO.
Fig 1 Displaced blades 2
Fig 2 Ground resonance 3
Fig 3 Flowchart of the methodology 13
Fig 4 Design of the main rotor blade with ribs 14
Fig 5 Wireframe view of modeled rotor blade 14
Fig 6 Shape of airfoil at the root 15
Fig 7 Screenshot of the designed rotor blade 15
Fig 8 Meshed rotor blade 18
Fig 9 Direction of lift acting on the main rotor blade 19
Fig 10 Direction of drag acting on the main rotor blade 19
Fig 11 Direction of weight acting on the main rotor blade 19
Fig 12 Direction of centrifugal force acting on the main rotor blade 19
Fig 13 Hinged or pinned supports of the main rotor blade 20
Fig 14 Total deformation obtained by static structural analysis 20

Fig 15 Maximum principal stress obtained by static structural analysis 21

Fig 16 Von Misses stresses obtained by static structural analysis 21
Fig 17 Mode shapes and natural frequencies obtained by modal
analysis of main rotor blade of helicopter 29
Fig.18 Acceleration acting on the main rotor blade 30
Fig 19(a) Frequency response curve with respect to amplitude and
phase angle along the direction of drag 31
Fig.19(b) Frequency response curve along the direction of drag 31
Fig 20(a) Frequency response curve with respect to amplitude and
phase angle along the direction of lift 32
Fig.20(b) Frequency response curve along the direction of lift 32
Fig 21(a) Frequency response curve with respect to amplitude and
ii
 phase angle along the direction of centrifugal force 33
Fig.21(b) Frequency response curve along the direction of
centrifugal force 33
Fig 22 Maximum principal stress obtained by harmonic analysis 33

iii
 ABSTRACT

In this study, the design of the main rotor blade of helicopter is modeled in CREO
software. The study starts with the introduction on the helicopter, vibration caused on the main
rotor blades of helicopter and the various types of vibration analysis. The different types of
vibration analysis and a detailed introduction on the principle of vibration analysis are discussed.
Further, the individual forces acting on the rotor blade of the helicopter model are calculated
using the equations of motion and their directions are defined. Using the values of the forces
obtained on the main rotor blade, a static structural, modal and harmonic analysis is presented
using ANSYS workbench. Then, with the stresses obtained by the analysis on the main rotor
blade, fatigue life cycles are calculated.
Finally, the conclusions and inferences arising in course of the work are presented and
the references used in this work are mentioned.

iv
 Chapter 1

INTRODUCTION
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

`CHAPTER 1
INTRODUCTION
A helicopter is a type of rotorcraft in which lift and thrust are supplied by rotors. This
allows the helicopter to take off and land vertically, to hover, and to fly forward, backward, and
laterally. These attributes allow helicopters to be used in congested or isolated areas where fixed-
wing aircraft and many forms of VTOL (vertical takeoff and landing) aircraft cannot perform.

There are three basic flight conditions for a helicopter:

a) Hover
b) Forward flight
c) Transition from hover to forward flight

a) Hover
Hovering is the most challenging part of flying a helicopter. This is because a helicopter
generates its own gusty air while in a hover, which acts against the fuselage and flight control
surfaces. The end result is constant control inputs and corrections by the pilot to keep the
helicopter where it is required to be. Despite the complexity of the task, the control inputs in a
hover are simple. The cyclic is used to eliminate drift in the horizontal plane, that is to control
forward and back, right and left. The collective is used to maintain altitude. The pedals are used
to control nose direction or heading. It is the interaction of these controls that makes hovering so
difficult, since an adjustment in any one control requires an adjustment of the other two, creating
a cycle of constant correction.

b) Forward flight
In forward flight a helicopter's flight controls behave more like those of a fixed-wing
aircraft. Displacing the cyclic forward will cause the nose to pitch down, with a resultant
increase in airspeed and loss of altitude. Aft cyclic will cause the nose to pitch up, slowing the
helicopter and causing it to climb. Increasing collective (power) while maintaining a constant
airspeed will induce a climb while decreasing collective will cause a descent. Coordinating these
two inputs, down collective plus aft cyclic or up collective plus forward cyclic, will result in
airspeed changes while maintaining a constant altitude. The pedals serve the same function in

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

both a helicopter and a fixed-wing aircraft, to maintain balanced flight. This is done by applying
a pedal input in whichever direction is necessary to center the ball in the turn and bank indicator.

c) Transition from hover to forward flight

As a helicopter moves from hover to forward flight, it enters a state called translational
lift which provides extra lift without increasing power. This state, most typically, occurs when
the airspeed reaches approximately 16–24 knots, and may be necessary for a helicopter to obtain
flight.

1.1 VIBRATIONS IN HELICOPTER

Typically, an unusual amount of vibrations in a helicopter is due to a malfunction in the
aircraft. These malfunctions may include loose hardware, out of track or out-of-balance
conditions or worn bearings. Due to the various moving parts and rotor system stress during
operation, helicopters have a high level of vibrations, which left unchecked will cause machine
failure or other serious damage to the aircraft in a short amount of time.

Different types of vibrations caused in helicopters are:

a) Low to High Frequency
One type of helicopter vibration is a frequency vibration. This type of vibration
may occur as a low, medium or high frequency. A low frequency vibration typically
occurs when the revolution of the rotor is disturbed. A medium frequency vibration is a
common rotor system vibration that occurs due to loose components of the aircraft. A
high frequency vibration typically occurs when the tail rotor gears, tail drive wire and
shaft or the tail rotor engine, fan or shaft assembly vibrates or rotates at an equal or
greater speed than the tail rotor.

Fig 1: Displaced blades

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

b) Ground Resonance
Ground resonance is a type of vibration that is the most destructive and dangerous
of the vibrations and can destroy a helicopter within seconds. Ground resonance never
occurs during flight and only affects grounded helicopters with turning rotors. Grand
resonance is often the result of unbalanced forces in a rotor system that causes an aircraft
to rock on the landing gear when the helicopter is at or near its natural frequency. Other
causes of ground resonance are incorrect tire pressure, defective rotor blade lag
dampeners and incorrect adjustments to landing gear shock struts.

Fig 2: Ground resonance

c) Lateral and vertical

Lateral and vertical vibrations are also a type of vibration that can affect a
helicopter. Lateral vibrations are often the result of worn, loose or cracked parts or a
lateral imbalance such as a span-wise imbalance, a chord-wise imbalance or a
combination of both. Vertical vibrations typically occur when a rotor blade is out of
track.

1.2 EFFECTS OF VIBRATION

To the flight crew and passengers, the vibration of a helicopter can cause physical and
mental effects, ranging from slight worry and annoyance to definite pain and distress. To the
pilot these effects can lead to an impaired ability causing, say, a poorly made landing at the end
of a flight. To the passengers, these effects may cause such a deep distrust of the helicopter that
they will refuse to fly in one again.
Of equal or greater importance are the effects of all vibrations to the helicopter. These
effects include:

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

➢ Accelerated wear
i. In the bearings, control rod ends, cables, pulleys and fairleads and bell-crank
attachments of flight control systems.
ii. In the bearings of all rotating parts
iii. In all instruments
➢ The cracking of fuselage skins, frames and stringers (especially near the tail rotor).
➢ The loosening of rivets and of the attachments for component parts, which in turn leads to
fretting and to corrosion.
➢ Internal damage to electronic equipment.
➢ The reduction in life of the components which is especially danger.

1.3 TYPES OF VIBRATION ANALYSIS

The different types of vibration analysis are:
a) Modal analysis
b) Harmonic analysis
c) Transient analysis
d) Spectrum analysis
e) Random vibration analysis

a) Modal analysis
A modal analysis is a technique used to determine the vibration characteristics of
structures such as:
➢ Natural frequencies: Frequencies at the structure vibrate naturally.
➢ Mode shapes: Shape of the structure at each natural frequency.
➢ Mode participation factors: Amount of mass that participates in a given direction
for each mode.

Benefits of the modal analysis are as follows:

➢ Allows the design to avoid resonant vibrations or to vibrate at a specified
frequency.

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

➢ Gives engineers an idea of how the design will respond to different types of
dynamic loads.
➢ Helps in calculating solution controls (time steps, etc) for dynamic analysis.

b) Harmonic analysis
Harmonic analysis is a technique to determine the steady state response of a
structure to sinusoidal (harmonic) loads of known frequency. Harmonic analysis is done
to make sure that a given design can withstand sinusoidal loads at different frequencies
and to detect resonant response and avoid it if necessary.

c) Transient vibration
Transient structural analysis is needed to evaluate the response of deformable
bodies when inertial effects become significant. Liner or nonlinear static analysis is
performed when the inertia and damping effects are ignored. A harmonic analysis is more
efficient, if the loading is purely sinusoidal and the response is linear. Rigid dynamic
analysis is more cost-effective, if the bodies can be assumed to be rigid and the
kinematics of the system is of interest.

d) Spectrum analysis
A response-spectrum analysis calculates the maximum response of a structure to a
transient loading. It is performed as a fast alternative of approximating a full transient
solution. The maximum response is computed as scale factor times the mode shape.
These maximum responses are then combined to give a total response of the structure.

f) Random vibration analysis

The random process is stationary (does not change with time) and ergodic (one
sample tells us everything about the random process).

1.4 NEED FOR VIBRATION CONTROL

The control of vibration is important for four main reasons:
➢ To improve crew efficiency and hence, safety of operation.

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

➢ To improve the comfort of passengers.

➢ To improve the reliability of avionic and mechanical equipment.
➢ To improve the fatigue lives of airframe structural component.

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 Chapter 2

LITERATURE SURVEY
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 2
LITERATURE SURVEY

Niranjan Roy, Ranjan Ganguli(1) presented the structural damage in materials evolves
over time due to growth of fatigue cracks in homogenous materials and a complicated process of
matrix cracking, delamination, fiber breakage and fiber matrix debonding in composite materials.
In this study, a finite element model of the helicopter rotor blade is used to analyze the effect of
damage growth on the modal frequencies in a qualitative manner. Phenomenological models of
material degradation for homogenous and composite materials are used. Results show that
damage can be detected by monitoring changes in lower as well as higher mode flap (out of-
plane bending), lag (in-plane bending) and torsion rotating frequencies, especially for composite
materials where the onset of the last stage of damage of fiber breakage is most critical.
A.R.SBramwell, G.Done(2) derived thoroughly the various equations governing the
motion of the rotor blades and the various stresses acting on different parts of the blade as well as
on the individual components of the helicopter.
A. Paternoster, R. Loendersloot, A. de Boer and R. Akkerman (3) presented that many
active concepts are being studied but they all face a large number of challenges to be
successfully integrated within a helicopter blade. The rotation speed generates critical loads on
the blade and any system within it. With the helicopter blade being the component providing
both lift and control in a helicopter, any mechanism influencing its behaviour is required to be
durable, reliable and safe. Actuation of the active system is the most critical aspect of a smart
adaptive blade. Piezoelectric actuators have the potential to provide compelling actuation for
these systems. They are actively tested for many of these concepts. Their toughness, size and
reliability make them especial candidates for delivering the required mechanical power.
Robert G. Loewy(4) presented that the designing helicopters for low vibration levels is so
difficult and complex that determined efforts should be made at the earliest possible stages in an
aircraft's development to ensure that all the basic factors are in the proper range. This involves
rotor blade, drive system and fuselage natural frequencies; providing adequate fuselage/rotor
aerodynamic clearance. As regards ensuring proper placement of fuselage modes and
frequencies, finite element analyses and correlative ground shake tests are mandatory. The same
may be said for blade designs where branched structures are encountered. Rotor blade natural

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

frequencies should be calculated with all the known couplings, including aeroelastic effects,
represented. If a linear theory with small perturbations is assumed, they should always be taken
about large, mean deflected positions. Nonlinearities as, for example, post buckling behavior in
fuselage structures and geometric nonlinearities in rotor blade dynamics should be incorporated
on the basis of the kind of 'ordering analyses' which are beginning to show which effects are
commensurate with a particular level of accuracy.
Castillo-Rivera, S., Tomas-Rodriguez, M., Marichal-Plasencia(5) presented the
different modelling aspects of helicopter aerodynamics. The helicopter model is on Sikorsky
configuration, main rotor in perpendicular combination with a tail rotor. The rotors are
articulated and their blades are rigid. The main rotor implementation takes into account flap, lag
and feather degrees of freedom for each of the equi-spaced blades as well as their dynamic
couplings. The model was built by using VehicleSim, software specialized in modeling
mechanical systems composed by rigid bodies. Appearing vibrations due to the rotating
behaviour of the rotors are studied in here. This work presents an aerodynamic model that allows
to simulate hover flight. The aerodynamic model has been built up using blade element theory.
The aerodynamic load creates vibrations on the helicopter and these are analyzed on the fuselage
by using short time Fourier transform processing to study the vibrations spectrum.
James Sathya Kumar(6) presented that the helicopters are versatile for a variety of roles
& applications in both military and civil applications and their utility has grown many folds, over
the years. The use of helicopters for relatively new applications such as air ambulance, Heli
tourism, fire fighting, monitoring of traffic or forests and fires or condition monitoring of remote
installations like transmission lines, oil pipe lines and so on has further given impetus to
advancements in helicopter design and production. The conventional roles, as well as the new
roles also demand reduced vibrations to achieve greater human comfort, better reliability of
structures and systems and efficiency. Helicopters are inherently prone to higher vibration levels
than the conventional Aircraft. Therefore it essential that the inherent helicopter vibrations are
not only addressed in the conceptual & design stages but also in service by way of either
modifications or continuous monitoring & rectification.
Phuriwat Anusonti-Inthra, Farhan Gandhi(7) examined the influence of cyclic
variations in flap-, lag-, and torsion-stiffness of the blade root region (at harmonics of the
rotational speed), for reduction of vibratory hub loads of a helicopter in forward flight. The

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

results indicate that considerable reduction in hub vibrations is possible using small-to-moderate
amplitude cyclic variations in stiffness (no greater than 15% of the baseline stiffness value).
Torsion stiffness variations produced moderate reductions in vertical hub force, lag stiffness
variations produced substantial reductions in all hub forces and the hub yaw moment, and flap
stiffness variations produced very significant reductions in all hub forces and the hub roll and
pitch moments. The amplitude of the cyclic stiffness variations required generally increase with
increasing forward speed, for comparable reductions in vibration. At any given forward speed, if
the amplitudes of cyclic stiffness variation are too large, the hub vibrations can actually increase.
The stiffness variations that reduce the vibratory hub loads could produce increases in certain
vibratory blade root load harmonics. Vibration reductions are achieved due to a decrease in the
inertial contribution to the hub loads, or a change in relative phase of various contributions.
M. Giglio, A. Manes , F. Vigano(8) presented a complete experimental and numerical
analysis necessary to certify the fail safe and eventually the damage tolerant behavior of the
component. In this paper the fail safe behavior of a composite titanium-graphite Rotor Hub is
analyzed. The component, with artificial technological defects, was tested with complex
contingent fatigue load in order to cause the failure of the titanium section. The failure started in
proximity of a high stressed area and it propagated quickly in the whole titanium section but
without involve the surrounding filament winding graphite, that is a fail safe device. The
behavior of the hub during the whole test is simulated, with good accuracy, by means of a
complete FE model that reproduces also the 3D propagation of the crack in the titanium section.
Prof. Rafiq A. Kanai, Dr. S. P. Chavan(9) presented the analysis on the benefits of
applying multiple alternatives for mock-up and compression of helicopter harmonic motion.
Multiple mock-up approaches, along with a weighted-average mechanism, are approximated so
that quicksand affiliated with only applying alone better alternate for the rotor blade vibration-
reduction dilemma are bypassed. A harmonic motion external function matching to a flight
circumstance in which blade-vortex collaboration drives amplitudinous levels of harmonic
motion is accounted. The design variables consist of cross-sectional areas of the architectural
component of the blade and non-architectural masses. The optimized considerations are matched
with a baseline consideration looking like a UF-60 reference blade. The aftereffects demonstrate
that at relatively insufficient additional cost matched with optimizing a single alternate, multiple
alternates can be applied to connect diverse reduced-vibration concepts that would be excused if

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

only a single mock-up method was exercised, and the most precise alternate may not control to
the superior design.
Henning Mainz, Berend G. van der Wall, Philippe Leconte(10) explained that ABC is
the acronym for ‘’Active Blade Concept’’ and represents a 38% Mach scaled model rotor of the
Advanced Technology Rotor (ATR) of Eurocopter Germany (ECD, [4]). In contrast to the ATR
the model rotor is fully articulated. Specifically, it is equipped with a flap at the trailing edge of
each blade, which is driven by a piezoelectric actuator. The ABC project is a cooperation
between the French ONERA and the German DLR within the research concept ‘’The Active
Rotor’’. This rotor will be used for investigations of the effect of different flap positions on
noise, vibrations and performance. ONERA was responsible for the structural design of the blade
and the manufacturing of a prototype blade. This paper deals with the particularities of the
mould, the build up of the blades including the mechanism to drive the flaps, the test of the flap
units in laboratory and in the S3MA wind tunnel, the manufacturing of the prototype blade and
its testing and the series blades manufacturing and laboratory test of these.
C. C. Crawford, Jr. R. L. Carlson p. R. Bates(11) presented that although safe-life
design concepts .have served the rotorcraft industry and its users well, since the invention of the
helicopter, the need for damage tolerant design is ever increasing. This paper discusses the issues
that have impaired such concepts to date with emphasis on improved crack growth data base for
small cracks, understanding crack growth near threshold together with retardation effects, and
characterization of composites under delamination. Of equal importance is the development of
simple but accurate inspection techniques for small cracks at field maintenance level. An
initiative comparable to those for ballistic tolerance of the 70's to establish and implement simple
damage tolerant criteria warrants consideration.
Mikael Amuraa, Lorenzo Aiellob, Mario Colavitac(12) explained the crash of a military
helicopter. The aircraft was flying at cruise speed in clear sky at 1000 feet over a flat area. The
crew suddenly lost the control of the helicopter that crashed in the immediately causing fatal
injuries to the whole crew. Four main rotor blades were found close to the impact point, while
the fifth blade was found about 900 m before the wreck. Therefore efforts were directed to the
failure of this blade that had apparently separated in the air. The rotor blade comprised a long
hollow 6061-T6 aluminum alloy extrusion and 25 thin metallic pockets that provided the trailing
edge airfoil shape. Visual examination of the fracture surface of the aluminum extrusion

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

indicated fatigue crack growth followed by ductile overload separation. Examination by optical
and electronic microscopy of the fatigue fracture revealed an abnormal incision that appeared to
be the fracture origin site. The incision was about 2,3 cm long and 190 μm deep. Fatigue failure
growth time was determined using fracture mechanics. Electronic microscopy equipped with X-
EDS analyzer revealed the presence of iron in the incision. This evidence allowed to as certain
that the incision at the crack origin resulted from the use of an inappropriate tool used to remove
pockets during maintenance activities. In addition to IBIS and in order to improve the safety of
flights, NDTs were developed and then established every 200 flight hours.

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Chapter 3

OBJECTIVE
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 3
OBJECTIVE
Objective is to create a main rotor blade model of helicopter and to study the vibrations
caused in the main rotor blades of helicopter and also to do analysis on vibrations with the
application of various steady and unsteady loads on main rotor blades. Also, to calculate the
fatigue life cycles of the designed main rotor blade of helicopter and to study the effects and
proper ways of minimizing the vibration in the helicopter main rotor blade.

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 Chapter 4

METHODOLOGY
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 4
METHODOLOGY
CAED model of main rotor
blade is created using CREO
software

The created model is

extracted into ANSYS
workbench

In ANSYS workbench, the

model is assigned with initial
conditions for static analysis

The model is then meshed

and simulated for modal and
harmonic analysis

The modal analysis results

are obtained in terms of
natural frequencies and total
deformations

The harmonic results are

obtained in terms of
frequency response curves

Fatigue cycles are calculated

based on the maximum and
minimum stress obtained
from the analysis

Conclusion
Fig 3: Flowchart of the methodology.

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 Chapter 5

DIMENSIONS AND DESIGN OF THE

ROTOR BLADE
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 5
DIMENSIONS AND DESIGN OF THE ROTOR BLADE
The dimensions of the main rotor blade are as follows:
➢ Length : 10m
➢ Root chord : 298.28mm
➢ Thickness of the airfoil at the root : 44.742mm
➢ Tip chord : 238.4mm
➢ Thickness of the airfoil at the tip : 35.76mm
➢ Surface area of blade : 1.4 sq.mts
➢ Main root radius : 132.5mm
Thus the main rotor blade of helicopter is modeled according to the above dimensions using
CREO software and these models are as shown in Fig4 and Fig5.

Fig 4: Design of the main rotor blade with ribs.

Fig 5: Wireframe view of the modeled rotor blade.

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

Fig 6: Shape of airfoil at the root

Fig 7: Screenshot of the designed rotor blade.

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 Chapter 6

EQUATIONS OF MOTION AND ITS

CALCULATIONS
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 6
EQUATIONS OF MOTION AND ITS CALCULATIONS
i. Area of the rotor disc
𝑆 = 𝜋𝑅 2
Where, R is the radius of the rotor disc
𝑆 = 𝜋 × 102
𝑺 = 𝟑𝟏𝟒𝒎𝟐
ii. Lift on the rotor blade
1
𝐿 = 𝜌𝑣 2 𝑆𝐶𝐿
2
Where, L is the lift on the rotor blade
𝜌 is the density
𝑣 is the velocity in m/s
𝑆 is the area of the rotor blade
𝐶𝐿 is the co-efficient of lift i.e, 0.8
143
𝑣= × 2 × 𝜋 × 10
60
𝒗 = 𝟏𝟒𝟗. 𝟕𝟒𝟗𝒎/𝒔
Therefore,
1
𝐿= × 1.29 × 149.7492 × 314 × 0.8
2
𝑳 = 𝟑𝟔𝟑. 𝟑 × 𝟏𝟎𝟒 𝑵
iii. Drag on the rotor blade
1
𝐷 = 𝜌𝑣 2 𝑆𝐶𝐷
2
Where, D is the drag on the rotor blade
𝜌 is the density
𝑣 is the velocity in m/s
𝑆 is the area of the rotor blade
𝐶𝐷 is the co-efficient of drag i.e, 0.052

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

1
𝐷= × 1.29 × 149.7492 × 314 × 0.052
2
𝑫 = 𝟐𝟑𝟗. 𝟏𝟔𝟕 × 𝟏𝟎𝟑 𝑵
iv. Centrifugal force acting on the rotor blade
𝐶𝐹 = 𝑀𝑏 𝑟 2 𝑅
Where, 𝑀𝑏 is the mass of the rotor blade
𝑟 is the revolution of rotor blade per second
R is the radius of the rotor disc
143 2
𝐶𝐹 = 12193 × ( ) × 10
60
𝑪𝑭 = 𝟔𝟗𝟐. 𝟓𝟗𝟔𝟑 × 𝟏𝟎𝟑 𝑵
v. Lift to Drag ratio of the rotor blade
𝐿 363.3 × 104
=
𝐷 239.167 × 103

𝑳
= 𝟏𝟓. 𝟏𝟗
𝑫

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 Chapter 7

STATIC STRUCTURAL ANALYSIS

 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 7
STATIC STRUCTURAL ANALYSIS
Steps involved in static structural analysis are:
i. Pre processing (setting up model):
Engineering Data module is used to define the material properties. Geometry
module opens the Design Modeler application, which can be used to import CAD models
from Creo. Model, Setup, Solution, and Results modules opens the Mechanical
application, which can be used to set up and solve the simulation (includes meshing, load
and boundary condition applications, solving, and results).
ii. Boundary conditions:
The geometric model is given with the properties of Aluminium2014-T6 alloy.
Mass and volume of the geometric model when given with the properties of this alloy is
12193Kg and 4.5666m3.
The material properties of this alloy are:
➢ Density: 2.80 g/cm3
➢ Young's modulus: 73 GPa
➢ Ultimate tensile strength: 190 to 480 MPa
➢ Thermal Conductivity: 130 to 190 W/m-K.
➢ Thermal Expansion: 23 μm/m-K.
iii. Meshing:
The extracted geometric model is meshed with tetrahedron elements. The meshed
element has 17279 nodes and 63318 elements. The meshed model is shown in Fig 8.

Fig 8: Meshed rotor blade.

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iv. Static structural analysis:

The meshed model is given with the boundary conditions such as lift, drag,
weight, centrifugal force and fixed or hinged supports. The application of these boundary
condition are shown in Fig 9, Fig 10, Fig 11, Fig 12 and Fig 13 respectively.

Fig 9: Direction of lift acting on the main rotor blade.

Fig 10: Direction of drag acting on the main rotor blade.

Fig 11: Direction of weight acting on the main rotor blade.

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Fig 12: Direction of centrifugal force acting on the main rotor blade.

Fig 13: Hinged or pinned supports of the main rotor blade.

v. Analysis:
The model is solved using the solved and the respective deformations and stress
variations on the main rotor blade are obtained as shown in Fig 14, Fig 15 and Fig 16.

Fig 14: Total deformation obtained by static structural analysis.

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Fig 15: Maximum Principal Stress obtained by static structural analysis.

Fig 16: Equivalent (Von Misses) Stress obtained by static structural analysis.

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 Chapter 8

MODAL ANALYSIS
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 8
MODAL ANALYSIS
Steps involved in modal analysis are:
i. Modal analysis is carried out in order to obtain the mode shapes and natural frequency of
the model i.e., main rotor blade of helicopter.
ii. There will no pre-stresses acting on the model as this analysis is done to find the natural
frequency of the component.
iii. Analysis setting is set to the maximum modes up to the natural frequency of 500Hz.
iv. The supports in the main rotor blade are hinged and solved using solver.
v. The solution of the component is obtained and is shown in the table (Fig 17) below.

MODE MODE SHAPES NATURAL

NO. FREQUENCY
1 6.6425Hz

2 17.036Hz

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3 27.112Hz

4 45.574Hz

5 58.474Hz

6 70.705Hz

7 89.055Hz

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

8 90.354Hz

9 104Hz

10 116.38Hz

11 134.65Hz

12 150.08Hz

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

13 170.06Hz

14 226.37Hz

15 233.44Hz

16 237.66Hz

17 277.25Hz

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

18 303.29Hz

19 318.58Hz

20 367.13Hz

21 368.38Hz

22 370.87Hz

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

23 384.78Hz

24 395.6Hz

25 406.38Hz

26 417.55Hz

27 433.98Hz

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

28 442.78Hz

29 446.42Hz

30 448.52Hz

31 450.97Hz

32 456.95Hz

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

33 468.67Hz

34 473.74Hz

35 483.67Hz

36 487.74Hz

37 492.81Hz

Fig 17: Mode shapes and natural frequencies obtained by modal analysis of main rotor blade of helicopter.

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 Chapter 9

HARMONIC ANALYSIS
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 9
HARMONIC ANALYSIS
Steps involved in harmonic analysis:
a) Analysis settings are set for the frequency ranging from 1Hz to 500Hz with the damping
ratio of 2e-002.
b) The model is given with an acceleration of 9.801m/s2 and the component is set free to
displace along the rotation of the rotor blade.

Fig 18: Acceleration acting on the main rotor blade.

c) Then it solved using the solver for the harmonic analysis and the obtained results along x,
y and z i.e. drag, lift and centrifugal force directions respectively for the frequency
ranging from 1Hz to 500Hz is shown in Fig17, Fig 18 and Fig 19 respectively.

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Fig 19(a): Frequency response curve with respect to amplitude and phase angle along direction of drag.

19(b): Normalized frequency response curve along the direction of drag.

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Fig 20(a): Frequency response curve with respect to amplitude and phase angle along the direction of lift.

20(b): Frequency response curve along the direction of lift.

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Fig 21(a): Frequency response curve with respect to amplitude and phase angle along the direction of
centrifugal force.

21(b): Frequency response curve along the direction of centrifugal force.

d) The maximum principal stress obtained by the harmonic analysis is 6.7156e7Pa and this
is shown in the Fig 22 below.

Fig 22: Maximum principal stress obtained by harmonic analysis.

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 Chapter 10

FATIGUE CYCLE CALCULATION

 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 10
FATIGUE CYCLE CALCULATION
a) Factor of safety:
𝑌𝑖𝑒𝑙𝑑 𝑠𝑡𝑟𝑒𝑠𝑠
𝐹𝑂𝑆 =
𝐸𝑞𝑢𝑖𝑣𝑎𝑙𝑒𝑛𝑡 𝑠𝑡𝑟𝑒𝑠𝑠

Yield strength of the selected material i.e, Aluminium2014-T6 is 414MPa

Also from static structural analysis, the equivalent stress obtained is 384.33MPa
414
FOS =
384.33
FOS = 1.077
b) Using Goodman’s Diagram,
𝜎𝑎𝑙𝑡𝑒𝑟𝑛𝑎𝑡𝑖𝑣𝑒
𝜎𝑓′ = 𝜎𝑚𝑒𝑎𝑛
1−
𝜎𝑢𝑙𝑡𝑖𝑚𝑎𝑡𝑒

From harmonic analysis we have got,

Maximum Principal Stress = 90MPa
Ultimate stress = 496.422MPa
Therefore, 𝜎𝑚𝑎𝑥 = 90MPa and 𝜎𝑚𝑖𝑛 = 0MPa
➢ Mean stress is given by,
𝜎𝑚𝑎𝑥 + 𝜎𝑚𝑖𝑛
𝜎𝑚𝑒𝑎𝑛 =
2
90+0
=
2
𝝈𝒎𝒆𝒂𝒏 = 𝟒𝟓𝑴𝑷𝒂
➢ Alternative stress is given by,
𝜎𝑚𝑎𝑥 − 𝜎𝑚𝑖𝑛
𝜎𝑎𝑙𝑡𝑒𝑟𝑛𝑎𝑡𝑖𝑣𝑒 =
2
90−0
=
2
𝝈𝒂𝒍𝒕𝒆𝒓𝒏𝒂𝒕𝒊𝒗𝒆 = 𝟒𝟓𝑴𝑷𝒂
Therefore,

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER
𝜎𝑎𝑙𝑡𝑒𝑟𝑛𝑎𝑡𝑖𝑣𝑒
𝜎𝑓′ = 𝜎
1 − 𝑚𝑒𝑎𝑛
𝜎𝑢𝑙𝑡𝑖𝑚𝑎𝑡𝑒
45
=
45
1 − 496.422

𝝈′𝒇 = 𝟒𝟗. 𝟒𝟖𝟓𝑴𝑷𝒂

c) Using Modified Goodman Diagram,
Number of fatigue cycles is given by,
𝜎𝑓′ = 𝜎𝑢𝑙𝑡𝑖𝑚𝑎𝑡𝑒 (2𝑁𝑓 )𝑏
The value of ‘b’ for Aluminium2014-T6 is (-0.106)
49.485 = 496.422(2𝑁𝑓 )−0.106
𝑵𝒇 = 𝟏. 𝟑𝟗𝟗𝟑𝟓 × 𝟏𝟎𝟗 𝒄𝒚𝒄𝒍𝒆𝒔

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Chapter 11

CONCLUSION
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

CHAPTER 11
CONCLUSION
In this work, an attempt is made to design and analyze the vibration in the main rotor
blade of helicopter. The dynamic forces are evaluated and analysis is carried out in ANSYS
workbench. The following results are observed:
a) For the given rotor specification, the lift generated is found to be 3.633e6N. This lift
force is found to be greater than the weight of the helicopter, which is the necessary
condition for take-off. This result is confirmed for the take off of the helicopter.
b) The ANSYS static structural analysis yields the principal stresses which ranges from
3.8443e8Pa maximum to the 1.6542e5Pa minimum and maximum displacement of
0.33944m.
c) The ANSYS modal analysis yields the mode shapes of the natural frequency ranging
from 0Hz to 500Hz.
d) The ANSYS harmonic analysis yields the principal stresses on the main rotor blade
ranging from maximum of 9.9348e7Pa to the minimum of 49002Pa.
e) The fatigue life cycles are calculated according to the obtained analysis results and are
obtained as 1.39935e9cycles and the material is said to fail after these many number of
stress cycles.

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REFERENCES
 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

REFERENCE
1. Niranjan Roy, Ranjan Ganguli, “Helicopter rotor blade frequency evolution with damage
growthand signal processing”, Journal of the Department of Aerospace Engineering,
Indian Institute of Science, Bangalore 560012, India.
2. A.R.SBramwell, G.Done, D.Balmford, (2001) “Bramwell’s Helicopter Dynamics”,
Second Edition, Butterworth-Heinemann: Oxford, ISBN: 978-0-7506-5075-5.
3. A. Paternoster, R. Loendersloot, A. de Boer and R. Akkerman, “Smart Actuation for
Helicopter Rotorblades”,
4. Loewy, G , 9 , “Helicopter Vibrations: A Technological Perspective,” J. of the American
Helicopter Society, 29(4)
5. Castillo-Rivera, S., Tomas-Rodriguez, M., Marichal-Plasencia, G., N. (2014) “Helicopter
Main Rotor Vibration Analysis with Varying Rotating Speed”. XXXV Jornadas de
Automatica. pp. 34-41. ISBN-13: 978-84-697-5089-6.
6. Hooper W E,( 1984) “The Vibratory Air-loading of Helicopter Rotors Vertica”, Vol. 29,
No. 4, pp 4-30.
7. Phuriwat Anusonti-Inthra, Farhan Gandhi, (2000) “Helicopter Vibration Reduction
through Cyclic Variations in Rotor Blade Root Stiffness”, Rotorcraft Center of
Excellence, Department of Aerospace Engineering, The Pennsylvania State University,
233 Hammond Building, University Park, PA 16802
8. M. Giglio, A. Manes , F. Vigano, “Experimental and Numerical Investigation on Fatigue
Failure of Composite Helicopter Main Rotor Hub”, Politecnico di Milano, Dipartimento
di Meccanica, Italy Corresponding author.
9. Prof. Rafiq A. Kanai, Dr. S. P. Chavan, (2012) “Helicopter Rotor Vibration Reliability
Control Analysis”, International journal of scientific & technology research volume 1,
issue 10.
10. Henning Mainz, Berend G. van der Wall, Philippe Leconte, (2005) “ABC rotor blades:
design, manufacturing and testing”, J. of 31th European rotorcraft forum florence, Italy.
11. C. C. Crawford, Jr. R. L. Carlson p. R. Bates, (1991) “Damage tolerance analysis for
rotorcraft, what the issues are”, seventeenth European rotorcraft forum, Georgia tech
research institute Atlanta, Georgia, united states, ERF 91-32.

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 VIBRATION ANALYSIS ON MAIN ROTOR BLADE OF HELICOPTER

12. Mikael Amuraa, Lorenzo Aiellob, Mario Colavitac, (2014) “Failure of a helicopter main
rotor blade”, 20th European Conference on Fracture (ECF20), Procedia Materials Science
3, pp 726-731.

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