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Engineering Notes
Development of a Micro Twin-Rotor
Cyclocopter Capable of
Autonomous Hover
Moble Benedict,∗Elena Shrestha,†Vikram Hrishikeshavan,∗
and Inderjit Chopra‡
University of Maryland, College Park, Maryland 20742
DOI: 10.2514/1.C032218
I. Introduction
GROWING interest in highly portable versatile flying platforms
and recent advancements in microelectronics have led to the
development of a scaled-down class of unmanned aerial vehicles
known as micro air vehicles (MAVs) [1,2]. The potential applications
of MAVs could range from reconnaissance, terrain mapping, and
search and rescue in both military and civilian settings. For these
types of missions, hover/low-speed flight capability, high endurance,
maneuverability, and the ability to tolerate environmental distur-
bances such as wind gusts are critical requirements for MAVs.
Because MAVs operate in a unique aerodynamic regime (low
Reynolds numbers) with a different set of mission requirements and
challenges as compared to a full-scale aircraft, it is important to
explore novel out-of-the-box vehicle concepts that might have the
potential for superior performance at these scales. This note describes
the vehicle design and control system development of one such MAV
concept: the cyclocopter (shown in Fig. 1). The cyclocopter uses
cycloidal rotors (cyclorotors), a horizontal axis propulsion concept
that has many advantages such as higher aerodynamic efficiency [3],
maneuverability, and high-speed forward flight capability [4,5]
when compared to a conventional helicopter rotor. A cyclorotor is
essentially a rotating-wing system where the span of the blades runs
parallel to the axis of its rotation. The pitch angle of each blade
is varied cyclically by mechanical means such that the blade
experiences positive geometric angles of attack at both the top and
bottom halves of the azimuth cycle (Fig. 2). Varying the amplitude
and phase of the cyclic blade pitch is used to change the magnitude
and direction of the net thrust vector produced by the cyclorotor.
Although many breakthroughs in cyclorotor research have
occurred in recent years, attempts to develop a cycloidal rotor-based
aircraft date back to the early 20th century [6,7]. Numerous full-scale
manned aircrafts were designed and built, but none of the attempts
was successful in achieving flight. However, recently, the feasibility
of this concept was demonstrated at the University of Maryland by
developing two cyclocopter configurations (a hybrid twin-rotor
cyclocopter [8] and quad-rotor cyclocopter [9]) capable of free hover.
The only other cyclocopter capable of controlled free flight was
developed at the Seoul National University [10]. The main focus of
the present work is to develop an improved version of the cyclocopter
built in [8] to demonstrate efficient hover capability and to develop
and implement a closed-loop control strategy that can be used to
autonomously stabilize and control the vehicle in hover without a
human pilot.
II. Twin-Rotor Cyclocopter Vehicle Design
As shown in Fig. 1, a twin-rotor cyclocopter (twin-cyclocopter)
weighing 210 g was designed and built. The vehicle has a lateral
dimension (rotor tip-to-tip) of 35 cm (14 in.), longitudinal dimension
of 30 cm (12 in.) and a height of 18 cm (7 in.). Most of the design
specifications are provided in Table 1. In the present vehicle con-
figuration, both the cyclorotors spin in the same direction, producing
a large nose-up reaction moment that is counterbalanced by the thrust
produced by the horizontal tail rotor.
A. Cyclorotor Design
Systematic experimental and computational studies were
performed in the past to optimize the performance of MAV-scale
cyclorotors [3,7,11–14]. Several blade kinematics and rotor
geometric parameters (blade-pitching amplitude, location of pitch
axis, chord/radius ratio, blade airfoil, planform, etc.) were varied to
improve overall rotor performance in hover. Using the understanding
obtained from these studies, the present cyclorotors are designed
for maximum thrust-to-power ratio (power loading). The design
specifications of the cyclorotor are given in Table 1. While optimizing
the rotor parameters for maximum aerodynamic performance,
emphasis was also placed on the blade and rotor structural design to
reduce the overall rotor weight. Details on the design of the cyclorotor
structure is provided in [8].
B. Blade Structural Design
One of the biggest disadvantages of a cyclocopter is that rotor
weight forms a significant fraction of the empty weight of the vehicle.
The rotor weight is directly related to the blade weight because it
governs the centrifugal force, which is the predominant structural
load on a cyclorotor. Designing lightweight blades for the cyclorotor
is challenging because the centrifugal force acts in the transverse
direction, producing large blade deformations and even structural
failure of the blades. Previous studies have shown that large bending
and torsional deformations degrade the thrust producing capability
and efficiency of the cyclorotor [7,12,14]. Therefore, the emphasis of
the present work was to design and fabricate extremely lightweight
blades with large stiffness-to-weight ratio.
The blades used on the twin-cyclocopter were fabricated mostly
out of foam with carbon-fiber reinforcement inside and a single-layer
0∕90 deg carbon composite prepreg skin wrapped around the foam
core at the blade tips, as shown in Figs. 1 and 3. Foam helped maintain
the required airfoil shape for the blades, whereas most of the bending
and torsion stiffness was provided by the carbon-fiber structure
embedded inside the foam. Each of the finished foam blades was
around 1.5 g. Therefore, with the foam blades, there is almost 50%
reduction in blade weight compared to the carbon-fiber blade design
(1.5 g versus 3 g) used on the previous generation twin-cyclocopter
[8], which is a huge advantage because each gram of blade produces
almost 200 g of centrifugal force at the operating rotational speed.
C. Blade-Pitching Mechanism
One of the key requirements for the success of a cyclocopter is a
simplified lightweight blade-pitching mechanism. Modeled after a
Received 28 November 2012; revision received 12 July 2013; accepted for
publication 15 August 2013; published online 11 March 2014. Copyright ©
2013 by the American Institute of Aeronautics and Astronautics, Inc. All
rights reserved. Copies of this paper may be made for personal or internal use,
on condition that the copier pay the $10.00 per-copy fee to the Copyright
Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include
the code 1542-3868/14 and $10.00 in correspondence with the CCC.
*Assistant Research Scientist, Department of Aerospace Engineering.
Member AIAA.
†Graduate Research Assistant, Department of Aerospace Engineering.
Student Member AIAA.
‡Alfred Gessow Professor and Director, Alfred Gessow Rotorcraft Center.
Fellow AIAA.
672
JOURNAL OF AIRCRAFT
Vol. 51, No. 2, March–April 2014
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four-bar linkage system, the present pitching mechanism enables
passive blade pitching as the blades move about the circular
trajectory. The schematic of the mechanism is depicted in Fig. 4,
where the four bars of the linkage system are labeled L1,L2,L3, and
L4.L1, also referred to as rotor radius, is the distance between the
blade-pitching axis and the horizontal axis of rotation. The pitch links
(of length L3) are connected to the end of the offset link on one end,
and the other end is connected to point B, which is at a distance L4
behind the pitching axis. The connections at both ends of the pitch
link are through pin joints to allow the rotational degree of freedom.
With this arrangement, as the rotor rotates, the blade automatically
pitches cyclically, where the pitching amplitude depends on the offset
length L2when the other linkage lengths remains fixed. The rotation
of the offset link changes the phasing of the cyclic pitching and
thereby changes the direction of the thrust vector. More details on the
implementation of this pitch mechanism on the actual vehicle are
provided in [8].
Once the rotors were built, systematic tests were conducted for a
range of rotational speeds where the thrust and electrical power were
measured. Variation of thrust and power with rotational speed for
each of the cyclorotors used on the twin-cyclocopter are shown in
Figs. 5a and 5b, respectively. At the operating rotational speed of
1800 rpm, each rotor produced around 80 g of thrust providing
enough thrust for the twin-cyclocopter to hover along with the
tail rotor.
III. Control System Development
Figure 6a shows the pitch, roll, and yaw axes definition for the
twin-cyclocopter. With the present controls strategy, pitch, roll, and
yaw moments are completely decoupled other than through
gyroscopic effects. The control strategies for pitch, roll, and yaw are
shown in Figs. 6b–6d, respectively. The red dotted arrows show the
thrust vectors for trimmed flight and the green solid arrows show the
new thrust vectors for generating a control moment. The tail rotor is
used to control the pitch by varying its rotational speed. For instance,
a positive pitching moment can be obtained by decreasing the tail
rotor rotational speed, and vice versa for negative pitch (Fig. 6b). Roll
is directly controlled by differential rotational speed variation of the
Fig. 1 Twin-cyclocopter.
Fig. 2 Blade-pitching kinematics.
Table 1 Design specifications for the twin-cyclocopter
Design features Value
Cyclorotor diameter 5 in.
Blade span 4 in.
Blade chord 1.3 in.
Blade airfoil section NACA 0015
Blade pitch amplitude 45 deg
Blade pitch axis 45% from LE
Cyclorotor motor ELE AD-100 1850 KV Outrunner (13 g)
Tailrotor motor 18-11 2000 KV Outrunner (10 g)
Tailrotor propeller EP5030 (5-in.-diam)
Cyclorotor gear ratio 6∶1
Thrust vectoring servo Blue Bird BMS 306 (7 g)
Processor-sensor board GINA 2.2 (2 g)
Battery 25 C three-cell 350 mA ·h
Fig. 3 Blade structure.
Fig. 4 Schematic of the blade-pitching mechanism.
J. AIRCRAFT, VOL. 51, NO. 2: ENGINEERING NOTES 673
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cyclorotors. Positive roll is executed when the rotational speed ofthe
left cyclorotor is greater than the right (Fig. 6c). Finally, yaw is
controlled by differentially rotating the two thrust vectors of the
cyclorotors (Fig. 6d). A positive yawing moment is produced by
tilting the thrust vector of rotor 1 forward and rotor 2 backward.
A. Avionics and Telemetry
A feedback control system is required to provide sufficient attitude
damping and stiffness to achieve stable hover. This was implemented
using a telemetry setup. A 2.4 GHz Atmel AVR transceiver was
attached to a base station. This was used to wirelessly update (IEEE
802.15.4 protocol) the feedback gains, trim inputs, and attitude
reference commands to the vehicle in flight. Due to the lack of
damping, an aggressive high-bandwidth attitude feedback control is
required. This was made possible by incorporating a lightweight (2 g)
processor-sensor board (GINA2.2 developed by U.C.Berkeley [15])
on the vehicle. The principal components of this board are a TI
MSP430 microprocessor for onboard computation tasks, ITG3200
tri-axial gyros, KXSD9 tri-axial accelerometer, and an Atmel radio
and antenna for wireless communication tasks. The wireless
communication has a latency less than 20–30 ms. The time-critical
inner-loop feedback occurs at an update rate of 3 ms. The user
communicates with the vehicle using a LabVIEW interface.
B. Inner-Loop Feedback Control System
The gyros measure the pitch q, roll p, and yaw rattitude rates,
while the accelerometers record the tilt of the gravity vector. The
vehicle attitude can be extracted by integrating the gyro measure-
ments with time. However, it is known that this leads to drift in
attitude measurements [16]. Accelerometers, on the other hand, offer
stable bias but are sensitive to vibrations and, in general, offer poor
high-frequency information [17]. Therefore, a complementary filter
600 1000 1400
0
20
40
60
80
Rotational speed (rpm)
Thrust (grams)
Cyclorotor 1
Cyclorotor 2
1800
Operating rpm
a) Resultant thrust
600 1000 1400
0
5
10
15
20
25
Rotational speed (rpm)
Electrical power (W)
Cyclorotor 1
Cyclorotor 2
1800
Operating rpm
b) Input electrical power
Fig. 5 Variation of thrust and input electrical power with rotational speed for the two cyclorotors used on the twin-cyclocopter.
Fig. 6 Control strategy for the twin-cyclocopter.
674 J. AIRCRAFT, VOL. 51, NO. 2: ENGINEERING NOTES
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was incorporated to extract the pitch and roll Euler angles using a
high-pass filter for the gyros (4 Hz cutoff) and a low-pass filter for
accelerometers (6 Hz cutoff). The rotor vibrations were filtered out
because it was sufficiently higher than the body dynamics.
The onboard inner-loop feedback was implemented using a
proportional–derivative (PD) controller as shown in Fig. 7. The
feedback states were the pitch and roll Euler angles (θ,ϕ) and the
attitude rates (p,q, and r). An outer-loop feedback capability was
provided for translational positioning by a human pilot or a position
tracking system such as VICON. The final control inputs to the
vehicle actuators are the individual rotational speeds for the two
cyclorotors and tail rotor and the two servo inputs as shown in Fig. 7.
IV. Flight Testing in Hover
Prior to free flight testing, it was necessary to investigate the
closed-loop attitude stability of the vehicle on a constrained setup.
This was achieved by mounting the vehicle on a spherical gimbal,
which restricted the vehicle in translation but allowed free rotation in
pitch, roll, and yaw. The proportional and derivative gains were tuned
using the Ziegler–Nichols approach. The gains that offered
acceptable stiffness and damping to reject external disturbance with
minimal oscillations were chosen. Once repeatability in vehicle
stability was established with a given set of trim and gain values, free
flight tests were conducted. It must be noted that achieving stable
attitude in the gimbal setup was an important necessary condition to
ensure stable free flight. It enabled quick troubleshooting with
minimal damage to the vehicle. However, the trim values would
change because the position of the center of lift (of the entire vehicle
based on the relative contribution from each rotors) is not known
exactly a priori, and therefore they would have to be determined in
free flight.
The flight tests were conducted by providing a pure throttle
command and ensuring that the vehicle comes out of ground effect
sufficiently quick. The flight performance was determined by
observing if the vehicle assumed a stable hover attitudewith minimal
drift. Now, a pure throttle command simultaneously increases the
rotational speed of all the rotors such that all of the moments are
cancelled and the center of lift is at the center of gravity of the vehicle.
However, it is possible that the rate of increase of tail rotor rotational
speed with throttle input does not match that of the cyclorotors. This
results in a travel of the center of lift, which results in the vehicle
quickly going out of trim in pitch mode, leading to undesirable
crashes. Therefore, an appropriate rotational speed ratio must be
determined beforehand on the gimbal stand. Roll was trimmed by
differentially adjusting the rotational speed of the cyclorotors. Pitch
trim was done by varying the rotational speed of the tail rotor. These
trim values were determined through multiple systematic free-flight
tests and was considered to be a key step. Once trim was achieved and
the proportional and derrivative gains of the inner-loop control
system were correctly tuned, the vehicle successfully executed stable
hover flight without a human pilot (shown in Fig. 8).
V. Conclusions
This note discusses one of the pioneering studies that successfully
culminated in the development of a flying vehicle using the
cyclorotor concept. The present study showed that the key elements
to the success of a cyclocopter are 1) developing an extremely
lightweight cyclorotor, which is only possible through innovative
lightweight and high strength-to-weight ratio blade and rotor
structural design along with a simplified passive blade-pitching
mechanism, and 2) choosing the right vehicle configuration and
attitude control strategy. Another notable achievement is the fact that
the present vehicle could be controlled completely without a human
pilot using aggressive inner-loop feedback stabilization. Attitude
disturbances were satisfactorily damped out without exhibiting
unstable flight modes.
Acknowledgment
This research was supported by the U.S. Army’s Micro
Autonomous Systems and Technology (MAST), Collaborative
Technology Alliance (CTA), Center for Microsystem Mechanics
with Brett Piekarski Army Research Lab (ARL), and Chris
Kroninger Army Research Lab Vehicle Technology Directorate
(ARL-VTD), as Technical Monitors. The authors would also like to
thank Shane Boyer for his valuble insights during flight testing.
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