International Journal of Engineering Trends and Technology (IJETT) – Volume 10 Number 5 - Apr 2014
Structural Design and Analysis of Cost Effective Rotorcraft for Recovery
Purposes
Bruce Ralphin Rose J1, Vetrivel S2
1
Assistant Professor, Department of Aeronautical Engineering, Regional Centre of Anna University, Tirunelveli, India, 627007.
2
PG scholar, Department of Aeronautical Engineering, Regional Centre of Anna University, Tirunelveli, India, 627007
Abstract—The structural design and analysis of
rotorcrafts is a complicated process in terms of material
selection and inertial forces acting on it. This article
presents the structural design of reliable highperformance rotorcraft in order to enhance the handling
qualities during recovery purposes. Airframe structure of
the rotorcraft is tailored based on the mission
requirements. A low navigation performance is focused
under the complex circumstances like flood and forest
fire. The cost effective design is proposed with suitable
material selection and design optimization. The operating
altitude of the rotorcraft is limited such that is can be
operated proximate to ground. The proposed model of
recovery vehicle is capable to perform under adverse
conditions with robustness. Selection and design of
structural components for fuselage, rotor blades and
landing gear are carried out in this article. A set of
‘primitive’ design variables form the representative
design space in the present task of minimizing blade
weight under several frequencies. Maintaining the nondimensional frequency and amplitude of the forces
against several lifting conditions are analyzed with FEM
approach. The aerodynamic coefficients throughout the
operational Mach number range are included to assess
the total stresses and strains induced. In the future
aviation industry, these cost effective designs can be
employed in the fire service stations and at critical
locations.
Keywords— Rotorcraft, Finite Element Modeling,
Recovery purposes, Material Optimization.
I. INTRODUCTION
Helicopters are classified as rotorcraft or rotary-wing
aircraft to distinguish it from fixed-wing aircraft. Helicopters
derive its source of required lift force from the rotor blades
rotating around a mast. [1] The word “helicopter” is tailored
from the French hélicoptère, invented by Gustave de Ponton
d’Amécourt in 1861. The primary advantage of the helicopter
is owing to the rotor blades that revolve through the air,
providing lift without forward motion. It eliminates the need
ISSN: 2231-5381
for runways and helping the helicopters able to take off or
land in congested or isolated areas where fixed-wing aircraft
cannot be operated. [1] The lift force created by the rotor
allows the helicopter to hover in a specified location and to
do vertical takeoff and landing, which is not possible with
fixed wing aircraft. The Lift force created by the helicopter's
wings (Main Rotor Blades) is depending on the shape and the
angle of the blades as it moves through the air. As the well
known fact, single main rotor with an anti-torque tail rotor
configuration design is recognized worldwide as the
helicopter. Because of the unique operating characteristics of
the helicopter, it is preferred to conduct intensive tasks that
are not possible with aircraft on the ground. [6] Today,
helicopters are used for transportation, construction,
firefighting, search and rescue, and a variety of other jobs
that require its special capabilities.
The helicopter rotor system is the rotating part of a
helicopter that generates lift. A rotor system may be mounted
horizontally, as main rotors are providing lift vertically. It
can be fixed vertically, such as a tail rotor to provide lift
horizontally as thrust to counteract torque effect. In the case
of tilt rotors, the rotor is mounted on a nacelle that rotates at
the edge of the wing to change the rotor from a horizontal
mounted position, providing lift horizontally as thrust.
Tandem rotor (sometimes referred to as dual rotor)
helicopters have two large horizontal rotor assemblies; a twin
rotor system, instead of one main assembly and a smaller tail
rotor. Single rotor helicopters need a tail rotor to neutralize
the twisting momentum produced by the single large rotor.
[9] This configuration also has the advantage of being able to
hold more weight with shorter blades, since there are two
sets.
A. Tail rotor
The tail rotor is a smaller rotor mounted vertically or with
slight inclination on the tail of a traditional single-rotor
helicopter. The tail rotor either pushes or pulls against the tail
to counter the torque. The tail rotor drive system has a drive
shaft that is powered by the gearbox mounted at the end of
the tail boom. The flexible couplings connected with the
drive shaft allow it to flex with the tail boom. [9] The
gearbox at the end of the tail boom offers an angled drive for
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International Journal of Engineering Trends and Technology (IJETT) – Volume 10 Number 5 - Apr 2014
the tail rotor and includes gearing to adjust the output to the
optimum rotational speed.
II. PROBLEM PREDICTION IN ROTORCRAFT
The issues of structural integrity for rotary-wing
aircraft are completely different from those of fixed-wing
aircraft. In helicopters, the dynamic rotor components are
safe life designs and are replaced at the end of their service
life. Thus, airworthiness concerns of structural integrity for
helicopters are limited but still pose great challenges in
adjusting to changing missions. [4]
Fig 1. Typical helicopter drive train system
B. Landing gear
The function of the landing gear in a helicopter is to
provide a support on which the helicopter can rest under
normal conditions. The main structural requirement for the
landing gear is to absorb the energy caused by descent. [3]
The proposed fixed landing gear design consists of a pair of
skids, wheels, or floats. The design is prepared based on the
skid and float combination to establish the primary role.
C. Damage Tolerance
Damage tolerance is the attribute of a structure that
permits it to retain the required residual strength for a period
of unrepaired usage. It happens after the structure has
sustained levels of fatigue cycles, corrosion, and accidental or
discrete source damage such as [5]
 Unstable propagation of fatigue cracks.
 Unstable propagation of initial or service induced
damage.
 Impact damage from a discrete source.
The focus of damage tolerance is to quantify the level of
damage that a structure can tolerate and repair it before a
catastrophic failure occurs. It assumes that any structure is
essentially imperfect as a result of the Inherent material
structure, material processing, and component design.
D. Safe Life and Fail Safe Structures
Safe-Life of a structure is the number of events such as
flights or flight hours in which the strength decreases below
its ultimate design value. [4] This approach assigns a finite
life to a component; (i.e.,) the baseline strength and its
degradation in operation. Safe Life analysis infers the
difficulties and cost of inspecting complex rotorcraft
structures. Fail-safe is the attribute of a structure that permits
to keep its required residual strength for a period, after the
failure of a Principal Structural Element (PSE). A PSE is an
element of the structure whose integrity is essential for
maintaining the overall structural integrity of rotorcraft. Even
though the fail-safe concept states that residual strength is
essential to achieve redundancy, it does not develop
inspection requirements.
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Fig 2. Causes of structural failures and its Percentage
Structural integrity problems for helicopters are in the
airframe, avionics, and retention hardware for non-airframe
related structures. Thus, the primary challenge for aging
military helicopters is to assure structural integrity with more
capable dynamic components. Damage tolerance design for
helicopters is a “band aid - short-term solution” that helps
only to maintain immediate flight safety and return to safelife operation. [6] The majority of the accidents are caused by
the pilot error followed by an engine failure or power loss.
The remaining possible failures are highlighted in Fig 2.
III. ROTORCRAFT DESIGN PARAMETERS
The structural design of rotorcraft is done by the use of
CATIA V5 design software. The complex geometric sections
are designed with significant accuracy to ensure the physical
results. The rotorcraft structural details are computed from
the design data sheets and existing recent publications related
to the field. The similar geometric and aerodynamic
configurations are compared using the graphs and the region
of maximum points is identified as indicated in Fig 3. It helps
to ensure the possibility of the new design with optimum
parameters. The design parameters are important in the
structural design of rotorcraft for the assigned reference
Mach number. [9] The following graphical representations
are used to compute the optimum parameters values. The
length of the fuselage and height of the main rotor are given
by Fig 3 and Fig 4 respectively. All the 11 design parameters
are computed in this fashion to ensure the feasibility of the
design.
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The Design process is carried out using CATIA V5 part
design module. The optimized design parameters are used
with ½ scaled dimensions in the Finite Element Analysis
(FEA) tool. Virtually every industry now recognizes that a
key strategy for success is to incorporate computer-based
engineering simulation early in the development process. The
Ribs and frames are carefully designed to ensure the
structural integrity of the rotorcraft as illustrated in Fig 5. The
main rotor induced lift force is calculated as the Load on the
model (P) = 8000 N. Therefore, the value of theoretical
normal stress is, σ = 10.864 x106 Pa. The aerodynamic loads
also coupled along with the lifting load and it is being
included in the Numerical simulation process. [11]
Fig 3. Mach number Vs length of
fuselage
Fig 4. Mach no Vs Height to top of main rotor head
TABLE 1
OPTIMUM DESIGN PARAMETERS
Sl.No
Parameters
Values
1.
Cabin max width
1.24 (m)
2.
Main rotor diameter
7 .6 (m)
3.
Cabin length
1.47 (m)
4.
Cabin height
1.31 (m)
5.
Main rotor disc
46
6.
Length of fuselage
6.50 (m)
7.
Empty Weight
230 (kg)
8.
Height to top of main rotor
head
2.40 (m)
9.
Skid track
1.80 (m)
10.
Max. Take-off weight
480 (kg)
11.
Length of rotor turning
8.60 (m)
Fig 5. Rotorcraft Model Diagram (All Dimensions are in mm)
IV. RESULTS AND DISCUSSION
The lifting load acts on the top portion of the cabin
structure across each node. The upper portion of cabin
structure consists of four nodes and each node is located at
the equal distances. The nodes are having the same geometric
dimensions, and the lower surface of the structure is
constrained, (i.e., Uz =0). [12] So the total deformation of the
structure is quantified by varying the material properties.
(m2)
Fig 6. Total deformation contour for Al-Alloy 1100 material
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The minimum deformation will be at the bottom shell in
the cabin structure and the maximum deformation will be at
the tail boom structure. For 2000 N load acting on each node,
the maximum deformation will be in 113  (micro strain)
of the tail boom structure. The minimum deformation is one
third of the magnitude of the tail deformation occurs at the
cabin ring frame. Figure (7) presents the cabin structure with
the nodes are designed in circular shape, because this shape
has more strength to weight ratio. The nodes are located at
various places and will have equal distance across the cabin
structure. The material property of the structure is changed to
Al-Alloy 2024 material. The same load is applied on each
node, and the minimum deformation is observed in the cabin
frame portion. [14] The maximum deformation is obtained at
the end of the tail boom structure with 108  (micro
strain).
For Al-Alloy 7075 material, the maximum deformation is
attained in the similar frames and its magnitude is 111  .
Therefore, the influence of similar alloy material properties
on the total deformation magnitude of a rotorcraft is almost
very small. The combined lift and horizontal aerodynamic
force configuration also revealed the similar properties.
Hence, the material based optimization for the rotorcraft
design allows choosing light weight materials for its cabin
design.
Fig 8. Total deformation contour for Al-Alloy 7075 material
The total deformation Vs axial distance is plotted between
for the average structural deformations for the three materials
being analyzed. The different matrial properties are used in
the structural analysis, the total deformation vary from
material to material. The graphical representations are used to
quantify the maximum and minimum deformations exist in
the structure.
The maximum deformation occurs in the middle and end
of the tail boom structure, and the material selection should
address this issue. However, for a failsafe design the loads
induced by the tail rotor and the directional moments also
must be included. The geometric configuration (truss) should
carry the variety of bending and torsional loads induced by
abrupt operations.
Fig 9. Total average deformation (  ) Vs axial distance (m) for
three materials
The loads acting on each node on the upper surface of the
cabin structure induce the maximum stress distributed among
the structural frames. The minimum stress is acting on the
bottom surface of the cabin structure, and also at the tail
boom structure. [15] The Maximum tensile stress is
computed as 14 MPa and the compressive stress is 12 MPa.
For the load 2000 N acting on each node of the upper surface,
the maximum stress computed occurs on the total diameter of
the cabin structure.
Fig 7. Total deformation contour for Al-Alloy 2024 material
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International Journal of Engineering Trends and Technology (IJETT) – Volume 10 Number 5 - Apr 2014
addressed with extra care. As a final point, the experimental
prototype should be prepared to verify the influence of
selected design parameters on the performance of rotorcraft.
ACKNOWLEDGEMENTS
Anna university support for the work of the authors is
greatly acknowledged. It has provided extensive computing
resources and materials for the completion of this project
work successfully.
REFERENCES
Fig 10. Total average stress contour for Al-Alloy 7075 material
Fig 11. Total average stress (MPa) Vs axial distance (m) for three
materials
The maximum and minimum stress levels are in one third
of the ultimate stress of the material used as displayed in Fig
(11). The factor of safety is ensured through this stress
evaluation process even at severe hovering missions.
V.
CONCLUSIONS
The selected design parameters are satisfied the loading
condition on the structural design of the model. The stresses
and deformations computed for three different materials are
revealed the suitability of the design for intense recovery
purposes. The experimental and theoretical values of the
structural design of the model are consistent and reliable. The
many loading conditions are applied on the structure to
evaluate the efficiency of the airframe. Even at the maximum
loads applied on the structure, the structure didn’t experience
any permanent deformations. Hence, the proposed airframe
model is fully capable to handle variety of load conditions.
Also, focusing on low navigation performance of rotorcraft
under complex circumstances, the tail moments must be
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