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Multi-Disciplinary Engineering Design Conference
Kate Gleason College of Engineering
Rochester Institute of Technology
Rochester, New York 14623
Project Number: 07122
AUTONOMOUS QUADCOPTER
Jason Enslin
Electrical Engineering
Glenn Kitchell
Computer Engineering
Richard Nichols
Electrical Engineering
Courtney Walsh
Mechanical Engineering
ABSTRACT
In this paper, a design for an autonomous quadcopter
is presented. Customer specifications were given for
an autonomous flying machine that requires a load to
be carried. These requirements were translated into
engineering specifications and after several design
concepts were considered, the quadcopter design was
selected. This paper presents the complete design
process for an autonomous quadcopter including lift
calculations using blade element theory as well as a
complex control system invoking both PID and fuzzy
controllers. The electronic systems used for the
quadcopter were tested and experimental results are
presented. The final result of this project is a thorough
design for an autonomous quadcopter including a
detailed test plan for use by subsequent design teams.
INTRODUCTION
This project was conceived by the five authors in the
spring of 2006. As members of the RIT Honors
Program, the authors had the opportunity to apply for
a multidisciplinary design grant. A proposal was
drawn up to design an autonomous flying machine. It
was the authors’ desire to create a project that was an
interesting and challenging engineering design
problem. The honors multidisciplinary grant was
awarded at $1000 and this was the budget for the
project. Dr. Vincent J. Amuso of the electrical
engineering department volunteered to act as the
customer for this project. The autonomous vehicle was
designed to have the ability to carry an antenna for
ground penetrating radar (GPR) research headed by
Dr. Amuso. The rest of this paper is outlined as
follows. The customer needs and resulting engineering
specifications as well as the initial design concepts are
Jeff Welch
Mechanical Engineering
outlined in the Preliminary Design section. A detailed
description of the mechanical, electrical and software
design of the product is presented in the Design
Process section. Testing results and analysis are
provided in the Testing Results section. Conclusions
and future work are given in the Conclusions section.
Finally, detailed equations and large diagrams are
provided in the Appendix.
PRELIMINARY DESIGN
The first step in the preliminary design was to identify
and understand the customer’s needs. GPR research
requires an aircraft that can fly for a specified amount
of time and carry an antenna. The flying vehicle
needed to be reliable and effectively controlled.
Autonomous flight would enhance the “user-friendly”
aspect of the product, but was not deemed essential to
its operation. A summary of the customer needs, with
their level of importance on a scale of one to five is
presented in Table 1.
Customer Need
Reliable flight
Ability to carry a load
Controllability
Two-way communication
Ability to hover
Autonomy
Process data
Importance
5
5
5
4
4
3
2
Table 1. Summary of customer needs, importance
scale: 1 (least important.) to 5 (most important)
These customer needs were then translated into
engineering specifications. It was determined that the
flying machine be able to carry interchangeable loads
to meet possible future customer needs. With this in
Paper Number 07122
Proceedings of the Multi-Disciplinary Engineering Design Conference
mind, the first engineering specification was for a
standard 1 kg pico-satellite load to be incorporated
into the design. This is the load that the METEOR
senior design teams design for and it allows for any
load that is less than 10 cm3 in size and 1 kg in weight
to be carried by the vehicle. The rest of the
engineering specifications followed directly from the
customer needs and are presented in Table 2 along
with their importance to the design on a scale of one to
five.
Engineering Specification
Carry at least 1 kg, 10 cm3 load
Fly & hover 75-125 ft. off ground
Controllable within 0.25 mi. radius (L.O.S.)
Flight time between 10-30 min.
1 channel, bi-directional communication
At least 1 kb/sec data rate
Importance
5
5
5
5
4
3
Table 2.
Summary of engineering specifications,
importance scale: 1 (least important.) to 5 (most
important)
Page 2
been a financial possibility, but the customer needs did
not allow for the selection of this concept. Finally, the
selected concept for this project was attained after
discovering an RC toy called the X-UFO [1]. A fourpropeller flying machine (quadcopter) was the final
considered concept. A sketch of this concept is shown
below in Figure 1. Each propeller is controlled by a
single source, and thus the control of the quadcopter is
easier to understand. Theoretically, if each engine is
spinning each propeller at the same rate, hovering can
be achieved. Lateral movement can be achieved by
spinning one set of engines faster than the others. This
concept had the ability to meet all the customer needs
and was both a challenging and realizable project to
undertake. For these reasons, the quadcopter was
chosen as the product to design.
DESIGN PROCESS
The design process for the autonomous quadcopter is
best described by first explaining the mechanical
design and then following up with the electrical
design. The systems have little dependence upon one
another. Where there is an interrelation, it is noted in
the following sections. The following subsections will
describe in detail the steps taken to design the
autonomous quadcopter.
Mechanical Design
The unique nature of a flying machine requires a
frame design that is both strong and lightweight.
However, most materials that exhibit both of these
characteristics are expensive, and also difficult to
process. Therefore, aluminum was selected as the
primary frame material due to its ready availability,
relatively low cost, and reasonable strength to weight
properties. Any components that were not major
structural members were designed to be plastic, which
is generally more expensive but lighter.
Figure 1. Quadcopter initial concept sketch
Several concepts were considered for the autonomous
flying machine. The initial concept that was presented
in the honors multidisciplinary grant proposal was a
helicopter. However, it was determined by the team
and its faculty advisors that this concept would be too
difficult to realize in the twenty week senior design
course. The main issue with the helicopter concept
was that that control of such a vehicle would be too
complicated to be implemented autonomously. A more
simplified concept was then considered – a lighterthan-air vehicle. This concept had the distinct
advantage that the flight aspect would be easily
attainable. The focus of the lighter-than-air vehicle
project would have been the control system and
payload carrier. This concept was explored more
deeply and it was found that a balloon that was
capable of carrying significant loads was well out of
budget for this project. An indoor balloon would have
The original frame design involved mounting all four
engines near the center of the flying machine, and
extending driveshafts out to the four rotors. This
design had the disadvantage of being very heavy
(close to 15 lbs), and the drivetrain design was
complex and difficult to build. The basic quadcopter
design requires two of the four rotors to spin in the
opposite direction as the other two. In the first
generation frame design, this was accomplished
through gearing, adding to the complexity of the
design. The main reason for mounting the engines near
the center was a perceived increase in stability due to
most of the mass of the machine being located near the
center of mass, decreasing the potential moment on the
machine.
In the design review, it was suggested that the increase
in stability offered by this design was outweighed by
Paper Number 07122
Proceedings of the KGCOE Multi-Disciplinary Engineering Design Conference
the design problems presented. A decision was made
to move the engines directly underneath the rotors and
eliminate the drivetrain entirely. Engine testing
verified that the engines were capable of running
reliably in a vertical orientation and in reverse,
eliminating the need for any drivetrain or gear
systems. The frame design was simplified into its
current form, shown below in Figure 2.
Figure 2. 3-D Solidworks model of quadcopter
The frame is symmetric and provides rigidity in the
directions most likely to experience significant load.
The frame is strong vertically, where forces from the
engines and from landing are applied. The X-shaped
design provides this rigidity at a lower weight than a
square shaped design. The X-shaped design lacks
some lateral stiffness compared to a square design, but
the quadcopter would not experience any load in the
directions except in a crash situation, which is difficult
to design for because the magnitude and direction of
the impact loads would be highly unpredictable.
Therefore, the decision was made to design for normal
flight operations, not for catastrophic occurrences such
as crash landings.
The majority of the frame is constructed of aluminum.
The primary frame rails are built from ½” X ½” square
bars of 6061-T3 aircraft quality aluminum. The
engines bolt to the vertical frame rails at the edges of
the structure along with the servos, which provide
throttle control. A custom machined linkage mounts
the rotors to the engine’s crankshaft [2]. Because the
servo and throttle have different angular displacements
across their full range of motion, a four-bar
mechanism was designed to control the throttle
position. The kinematic developments for this analysis
were derived from [3]. In the center of the machine,
three plates provide mounting locations for the
electronic components needed to fly the machine as
well as fuel tanks for the engines. The lower plate
houses the majority of the electronics, including the
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GPU, wireless link, radar altimeter, and PCB boards.
The middle level holds a pair of fuel tanks to supply
the engines. The top level holds the GPS,
accelerometer, gyroscope, and digital compass.
The next task in the mechanical design process was to
determine the size and type of rotors needed as well as
the required horsepower rating of the engines. Before
being able to determine the rotor size needed to lift the
required weight of the quadcopter, calculations for lift
needed to be made. For these calculations, a mixture
of the Momentum Theory and the Blade Element
Theory were used [4]. The Blade Element Theory uses
a more precise method to calculate the lift of the blade
by deriving a more exact equation for the downwash
velocity, which is then used as an integral to determine
the lift along the blade. For simplicity, a rectangular
untwisted blade in hover was assumed. The size of the
blade used in the calculation was eight inches because
it was a manageable size and readily available. This
still resulted in a very complex equation that related
the tip velocity of the blades to the overall lift
(Equation A4). From Equation A4, another equation
was derived to find the power required for the
designated tip velocity denoted as Equation A5. These
equations are too complicated to reproduce in the text,
but they are shown in the appendix. Equations A1
through A3 are provided for reference in Equation A4.
Equations A6 through A8 describe how to calculate
the total power required and the torque. Table 3 was
created correlating the amount of lift, or weight, which
was required to the tip velocity and then correlated
further to the amount of power to generate a certain
amount of lift. Table A1 in the appendix lists the
values of the parameters used for the calculations.
MATLAB® files for these calculations are available at
[5]. With the information in Table 3, it was then
possible to choose an engine that could generate the
correct amount of power and thus the appropriate
amount of lift.
Weight Tip Speed, V t Induced Power
lb
m/s
rpm
W
hp
3.0
152
9344
127
0.170
3.1
154
9499
133
0.178
3.2
157
9650
139
0.187
3.3
159
9800
146
0.196
3.4
161
9947
153
0.205
3.5
164
10093
159
0.214
3.6
166
10236
166
0.223
3.7
168
10377
173
0.233
3.8
180
10516
180
0.242
3.9
173
10654
188
0.252
4.0
175
10790
195
0.261
Profile Power
W
hp
70
0.094
73
0.098
77
0.103
80
0.108
84
0.113
88
0.118
92
0.123
96
0.128
99
0.133
103
0.139
107
0.144
Total Power Torque
W
hp
N-m
196
0.263 0.201
206
0.277 0.207
216
0.290 0.214
226
0.304 0.221
237
0.318 0.227
247
0.332 0.234
258
0.346 0.241
269
0.361 0.247
280
0.375 0.254
291
0.390 0.261
302
0.405 0.268
Table 3. Summary of power and torque calculations
using the Blade Element Theory
With this calculated horsepower, the type of
propulsion device had to be determined. A decision
had to be made between using an electric motor or a
gas engine. Electric motors are easy to start and
control, but do not have a desirable power-to-weight
ratio. In order to attain the necessary horsepower, a
Paper Number 07122
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Proceedings of the Multi-Disciplinary Engineering Design Conference
large battery would be required on board the
quadcopter. Gas engines are more difficult to start and
control their throttle, but it was found that they
provide a lot of power in small physical sizes. A
0.4 HP glow fuel hobby aircraft engine was the choice
to power the quadcopter. Its small physical dimensions
(56.5 x 24.8 x 61 mm) and weight (5 oz) led to the
choice. From observing Table 3, it can be seen that
0.4 HP would allow for one engine to lift 4 lbs.
Therefore, the four engines collectively should be able
to lift 16 lbs. It was decided that this amount of lift
provided sufficient overhead for the weight estimate of
the redesigned frame (10.87 lbs). Electric servos were
used to control the throttle (more on this in the
electrical design section), so the only disadvantage to
using these engines was the tedious starting process.
Electrical Design
A complex electrical system had to be designed in
order to gather and process information in order to
control the quadcopter. A top-level block diagram
containing all of the major components of the
electrical system is shown in Figure 3. All components
shown in Figure 2 except for one of the wireless links
and the PC were designed to be placed on board the
quadcopter.
The accelerometer used for this project was the tripleaxis MMA7260Q accelerometer developed by
Freescale Semiconductor. This single chip allows for
the measurement of acceleration in the x, y, and z
directions. The chip is small and consumed a small
amount of current, making it desirable for this project.
The selected gyroscope has very similar attributes. A
relatively new chip called the IDG-300 Integrated
Dual-Axis Gyroscope by InvenSense was used for this
design. This single chip has the ability to measure
rotation in both the x and y direction. Along with the
accelerometer, this gyroscope produces an analog
output between zero and the supply voltage (3.6 V).
This is where the first electrical design problem
manifested in this project. The GPU can only accept
analog inputs between zero and two volts. Therefore, a
level shifter had to be designed to reduce the
accelerometer and gyroscope outputs to this range. A
simple op-amp circuit was designed to accomplish this
task. This circuit is shown in Figure 4. The resistor
values were determined by setting the ratio of the
feedback resistor to the input resistor equal to the ratio
of the supply voltage to the maximum sensor output
voltage. Such large resistor values were used to reduce
the amount of current drawn and thus the power
consumed.
R2
2
3
R2
100k
Figure 3.
system
Top-level block diagram of electrical
The GPU is the most important part of the
quadcopter’s electrical system. It accepts inputs from
the various sensors and produces output by moving the
servos. The GPU also has the ability to send back
position information to the PC over the bi-directional
wireless link. The choice of GPU for this project was
the RCM4000 produced by Rabbit Semiconductor.
The main criterion for the GPU was that it has enough
analog inputs to accommodate the accelerometer and
gyroscope. The RCM4000 provides eight analog
inputs, which is enough for the five analog signals
produced by the accelerometer and gyroscope. The
user friendly aspect of the RCM4000 in addition to its
relatively low cost made the choice a simple decision.
+
B
0
OPA2337
1
6
5
Vout
180k
7
Vin
R1
8
C1
OUT
180k
V+
0
-
C2
R1
V-
4
100k
+3.6 V
0
Figure 4. Level shifter op-amp schematic – shifts
input range (0 to 3.6 V) to output range (0 to 2 V)
A digital compass was incorporated into the design so
that direction can be sensed by the system. The
MicroMag 3 Three-Axis Magnetic Sensor from PNI
was used for this project. Like the other chips, this one
consumes a low amount of power and is small in size.
The digital output was connected directly to a digital
input pin on the GPU. The position sensing module
used for this project was the EM-406 GPS chip
produced by GlobalSat. Like the digital compass, the
output of the GPS chip was connected directly to the
digital input pins of the GPU. The GPS allows for the
quadcopter to measure its position and use that
information for its flight path.
Paper Number 07122
Proceedings of the KGCOE Multi-Disciplinary Engineering Design Conference
There also had to be some circuitry designed for the
four rotation sensors. The rotation sensors are used to
monitor the speed of the engine by “counting” the
number of times that a thin piece of opaque material
attached to the driveshaft passes through the optical
switch. The circuit shown in Figure 4 provides an
output voltage of approximately 3.2 V when the light
source is blocked and approximately 200 mV when
the light source is allowed to pass through. This
difference in voltage is easily read by the digital input
pins on the GPU and the time between rising edges is
measured to determine the speed. The 1:10 ratio for
resistor R1 and R2 in Figure 5 was recommended by
the optical switch’s datasheet. The actual values of
1 kΩ and 10 kΩ were selected as a compromise
between current consumption and switching time.
Large resistance values, while consuming less current,
increase the switching time of the optical sensor. Since
the max RPM of the engine is 17,000 RPM, the
switching time had to be less than 3.5 ms. For the
10 kΩ load resistance, the switching time of the sensor
is 150 μs.
+3.6 V
0
R1
1k
D1
LED
Vout
R2
10k
+3.6 V
0
Figure 5. Optical switch schematic – gives Vout ≈
3.2 V in “off” (blocked) state and Vout ≈ 200 mV in
“on” (transparent) state
The lower right portion of the block diagram in
Figure 2 shows the components that comprise a radar
altimeter. This system was incorporated into the
design of the quadcopter in order to obtain highly
accurate height measurements. The GPS chip can
provide height information, but with a degree of error
of approximately 10 m. This simply would not be
acceptable for autonomous landing. The radar
altimeter can provide height information accurate to a
few inches. This precision makes for much easier
landing. The altimeter is designed to operate at 1 GHz.
This frequency reflects off the ground well, and since
this is the common frequency that cell phones operate
at, parts for this frequency are cheap and readily
available. The radar altimeter operates by sending a
signal out of an antenna and then measuring the time it
takes for the reflected signal from the ground to return.
Page 5
This information is processed by the GPU and a height
value is then calculated.
Data is transmitted between the quadcopter and GPU
through RF wireless transmission. A wireless
transmitter and receiver made by Shenzhen Hac
(HAC-UM12) were purchased for this project. These
links are low power, they operate at 433 MHz, and
they have a maximum range of 1000 ft. The maximum
range required by the customer for this project was
approximately 500 ft, so these links exceed
specification.
The final components shown in Figure 2 are the four
analog servos. These servos are used to control the
throttle on the four engines. A pulse of voltage is
applied by the GPU to the servo and the pulse length
determines how much the servo rotates. The servos
have an experimentally determined range of
approximately 140 degrees.
All of the components described above can be
operated on the 3.6 V battery source except for the
GPS and the four servos. These components require a
voltage in the range of 4.5 to 6 V. A 3.6 V lithium
battery supplying 2.4 Ah was the choice for the main
power source. This battery was chosen because of its
small size (AA size) and cheap price ($3.50 per
battery). It was necessary to obtain a 3.6 V battery
because of a specification for the GPU. Originally, the
design for the main power supply was going to be two
standard AAA batteries, which would supply
approximately 3.2 V. However, the input impedance
of the electrical system produced a voltage drop of
about 0.3 V. This drop in voltage led to the GPU being
supplied with about 2.9 V. The GPU enters a reset
mode if its supply voltage drops below 2.93 V.
Therefore, getting a higher voltage supply was a
necessity. For the components that require a higher
voltage, four rechargeable standard lithium AAA
batteries were used. These batteries produce
approximately 5 V together and are rated at 0.9 Ah. It
is assumed that the voltage supplied by the batteries is
constant for the amp-hour ratings that are given. This
assumption generally holds true for lithium batteries.
A power consumption analysis was performed in order
to determine the amount of flight time that could be
achieved with respect to battery life. Table 4 shows
the breakdown of the power consumption of each
component and provides the life of each power supply
while Equations 1 and 2 show the life of each battery.
3.6 V Battery Life 
2.4 Ah
 13.6 h
0.1763 A
(1)
5.0 V Battery Life 
0.9 Ah
 5.3 h
0.170 A
(2)
Paper Number 07122
Proceedings of the Multi-Disciplinary Engineering Design Conference
Item
GPS
Servos (4)
5 V, 0.9 Ah battery
Gyroscope
GPU
Level Shifter (5)
Wireless Link
Accelerometer
Digital Compass
Optical Switch LED (4)
Optical Switch Phototransistor (4)
3.6 V, 2.4 Ah battery
Operating V
5.0
5.0
Current Draw (mA)
70.0
4 x 25.0
Power (mW)
350.0
500.0
Totals:
170.0
850.0
3.6
3.6
3.6
3.6
3.6
3.6
3.6
3.6
9.5
110.0
5 x 0.525
40.0
0.5
0.5
4 x 3.0
4 x 0.3
34.2
396.0
4.5
144.0
1.8
1.8
43.2
4.3
Totals:
176.3
629.8
Table 4. Power consumption analysis summary
All of the previously described electrical components
are used to provide input to the control system. In
order to fly the quadcopter in a stable manner, a very
detailed and complicated control system was needed.
This control system was designed and tested using the
Simulink® software package. The system uses a
combination of PID and fuzzy controllers to match the
altitude, pitch, roll, and heading to the desired values.
The top-level block diagram of the quadcopter control
system is shown in Figure A1 in the appendix. There
are several sub-blocks represented in this block
diagram that could not be presented due to spatial
constraints. All of the control system models are
available at [5].
The control design began by deriving a model for how
the quadcopter moves based on its four servo
positions. Initially in the design, an approximate
model was used. This model is accurate when the
pitch and roll angles are small, but becomes inaccurate
for large angles. It also did not track the current
heading, which was one of the variables to be
controlled. Because of these limitations, a more
accurate flight model was derived. A second
coordinate system was defined to stay fixed with the
vehicle, and the three vectors defining this coordinate
system are tracked. This is done by first finding all of
the rotational velocities and accelerations (shown in
the appendix by Equations A9, A10 and A11) and then
combining them [6]. The coordinate system is then
rotated about the calculated axis of rotation using
matrix rotations [7], resulting in the new coordinate
system. When these operations are combined and
simplified, Equations A12, A13 and A14 are obtained.
The new coordinate system can be used to calculate
the translational acceleration from the rotor forces, as
done in Equation A15. A conversion from engine
speed to lift force (Equation A17) was obtained using
the results of the Blade Element Theory calculations.
An estimate of the engine transfer function was also
used for design and simulation.
The first component of the control system is the
controller to manage the engine speed. A feedback
PID controller was chosen for this application. The
current measured engine speed is subtracted from the
desired engine speed, and this value, its derivative, and
its integral are used to calculate the servo position.
Page 6
Since there are four engine/servo combinations in the
system, this component was duplicated four times.
Blocks were then designed to control altitude, pitch,
roll, and heading. Each of these blocks uses feedback
to determine the error between the desired value and
the actual value of the controlled variable. This error is
fed through a fuzzy controller, and the output is used
to adjust the desired engine speeds. For example, the
controller monitoring altitude will add or subtract
equally from all four desired engine speeds, while the
controller monitoring pitch will subtract from the right
two engines whatever it adds to the left two. When
these four fuzzy controllers are combined with the
four engine controllers, the system is capable of
keeping stable flight. The parameters of the PID and
fuzzy controllers can be optimized to minimize
response time and minimize overshoot.
Software Design
Software had to be developed for the quadcopter on
both the PC end and the GPU end. The PC software
was written in Java while the PC software was
completed in Dynamic C. A brief overview of each
software system is presented in this subsection.
The PC application developed in Java was named
“FlightRabbit.” FlightRabbit is the connecting fiber
between the pilot and the quadcopter. It is a userfriendly and highly expandable Java program that
translates the pilot’s input into useful quadcopter
commands. FlightRabbit uses the state pattern to
change the functionality of the program throughout the
different modes in the quadcopter’s operation; Initial,
Engine Start-up, Manual Flight, Autonomous Flight,
Retrieve Flight-Plan, and Abort & Land modes. In
each of these modes, the FlightRabbit program enters
one or more of a series of states with well-defined
state transitions to achieve the simplest and safest
control over the quadcopter. These states each offer a
unique Graphical User Interface that presents the pilot
or user with the detailed information needed. A
screenshot of the Manual Flight GUI is shown in
Figure 6. Beyond simply controlling the quadcopter,
FlightRabbit also maintains flight logs and usage
history so that previous flights can be reviewed and
analyzed. Previous flight plans can be saved, reloaded
and edited across multiple flights.
FlightRabbit utilizes the threading along with the
command and observer patterns to integrate the
layered systems created among the USB Controller,
GUI, and serial communication with the quadcopter.
These patterns eliminate busy waiting and allow each
layer to operate freely and in parallel with the others.
The USB Controller, the GUI threading, and the
wireless serial communication protocol are all custombuilt. They are designed to be reliable but efficient,
Paper Number 07122
Proceedings of the KGCOE Multi-Disciplinary Engineering Design Conference
and are streamlined so that the quadcopter can act on a
minimal but flexible amount of data.
Figure 6. Screenshot of the GUI for Manual Flight
Mode using the FlightRabbit Java software
The GPU program is made up of several threads, each
one performing a specific task. The threads are not
true multi-tasked threads, but use a cooperative multitaking scheme. Cooperative multitasking is better for
this application, since it requires less overhead and
provides more control. There are a total of six threads,
one for each of the following tasks: communicating
with the PC, reading from the GPS, reading from the
compass, reading from the other sensors, writing to the
servos, and determining the desired servo positions
based on the sensor input.
A high-precision timing interrupt was written to allow
for timing more precise than the built-in 1 ms timing.
The interrupt counts every 8.681 μs, which gives
enough precision to control the servos and read from
the sensors. Since the servo control requires a very
precise pulse, the code for writing to the servos was
put directly into the interrupt routine.
TESTING RESULTS
The original goal of this project was to have the
complete quadcopter assembled and thoroughly tested
by the end of Senior Design II. However, unforeseen
delays in the schedule of the project, highlighted by a
complete redesign of the frame following the first
design review forced a revised deliverable list. After
consulting with the customer, it was agreed upon that
the revised deliverable list would consist of a complete
design of the quadcopter, including assembly
instructions and all software to operate the vehicle, as
well as a thorough test plan to verify each step of the
assembly process culminating with a final test flight
plan.
Page 7
There was significant electrical and software testing
that took place during the course of the project. The
first thing that was done was to verify that the level
shifting circuit performed as expected when hooked up
to the accelerometer and gyroscope. Oscilloscope
measurements verified the operation of the level
shifters, and the circuits were assembled on a PCB
board. Four 80 x 54 mm through-hole PCB boards
were designed to be placed on board the quadcopter.
One of these boards consists of the five level shifting
circuits, one contains the circuitry for the four rotation
sensors and the 5 V battery source, one contains the
pin outs for the GPU and the final one is designed to
hold the radar altimeter. The altimeter was designed
completely and the parts that comprise it were
specified but the system was not purchased or built
due to monetary and time constraints. The test plan
document contains information on how to assemble
and test the altimeter.
The GPU software was shown to successfully read the
inputs from the accelerometer and gyroscope. The
FlightRabbit program visually displayed the rotation
sensed by the gyroscope accurately. The program also
output the acceleration in each direction. All of these
readings were made wirelessly. The sensors, the GPU,
and one wireless link were placed in a box and the
readings were received and successfully displayed by
the computer.
The operation of the optical rotation sensors were
verified through breadboard testing. An opaque
material was passed through the sensor and the GPU
was able to read and count the number of passes. The
rotation sensor circuitry was then soldered to one of
the PCB boards described previously.
One of the more difficult tasks in this project was to
test the operation of the servos wirelessly. Software
was written to enable the servo to be moved with a
USB video game controller. Moving the joystick to
different positions resulted in the servo moving to
corresponding positions. This was done for both
testing purposes and as part of a manual flight mode
design. Problems occurred over the wireless
transmission. It was found that often times, bits or
even full bytes were being dropped over the link
resulting in either a slow or inaccurate response by the
servo. To correct this problem, a wireless protocol
with error correction was implemented. This
sufficiently fixed the problem and the servo now
responds reliably to commands from the USB
controller.
The GPS and digital compass both require drivers to
read their digital data. At press time, these drivers are
being completed and tested. It is expected that they
will be included in the final software package.
Paper Number 07122
Proceedings of the Multi-Disciplinary Engineering Design Conference
A test fixture was designed and built in order to test
the functionality of the engine. A picture of this
apparatus is shown below in Figure 7. The motor was
mounted approximately two feet off the ground to
reduce ground effects. A servo was attached to the
fixture and connected to the engine throttle by a wire
linkage. There were two objectives to be accomplished
by running the engine from this test fixture. First, a
measurement of the fuel consumption was to be
determined. Secondly, the test called for this fixture to
be placed on top of a scale. An experimental lift was
to be determined by recording the difference in the
scale when the engine was running. Unfortunately,
there were difficulties getting the engine started. There
was a steel custom machined part that was used to
attach the rotors to the engine. Unbeknownst to the
designers, the part made the engine too top-heavy and
as the starter tried to rotate the engine, the screws from
the engine mount were ripped out of the test fixture.
The piece attaching the rotors was thus redesigned,
and at press time the test fixture is being rebuilt. If this
testing cannot be completed, it will be included into
the comprehensive test plan for a subsequent design
team. Also, to ensure the validity of the design two
important tests were included in the test plan. A finite
element analysis (FEA) to check the performance of
relevant parts of the frame under the anticipated loads
was included. Also, linear elastic fracture mechanics
(LEFM) methods as described in [8] were
incorporated to develop an inspection program to
prevent fatigue failures. These tests were unable to be
completed due to time constraints stemming from the
complete frame redesign.
Page 8
CONCLUSIONS
The quadcopter design currently meets all customer
specifications. The control system demonstrates in
simulation that reliable and controlled flight can be
achieved. Tests have confirmed the functionality and
range of wireless transmission. Theoretical
calculations have determined that the desired amount
of lift can be achieved. Experimental tests must be
completed to verify these calculations. A complete and
thorough test plan has been provided to the customer
and can be used by subsequent design teams to
complete this project. This plan includes assembly
diagrams and instructions as well as a complete bill of
materials including where to purchase the remaining
materials. Reflecting on this project, it was very
aggressive for a five person design team. A larger
mechanical engineering staff would have been
beneficial for the completion of this project. This
project required an extreme amount of engineering to
be done in a short amount of time. With respect to this
aggressive nature of the project, its status at this point
can be deemed a mild success. A solid design
foundation has been provided for a future team.
ACKNOWLEDGMENTS
The authors would like to thank Dr. Vincent J. Amuso
for serving as the customer and advisor for this
project. They would also like to acknowledge Dr.
Jeffrey Kozak, Dr. Frank Sciremammano, and Dr. P.
Venkataraman for their assistance in the design review
process. Finally, the authors would like to thank the
RIT Honors Program for sponsoring this project.
REFERENCES
Figure 7. Test fixture apparatus for engine testing
[1]http://www.gadgetmadness.com/archives/20050927
-review_silverlit_xufo_rc_flying_toy_xufo.php
[2] Shigley, Joseph E, Mischke, Charles R., &
Budynas, Richard G., 2004, Mechanical Engineering
Design, McGraw Hill, New York, NY.
[3] Myszka, David H., 2005, Machines &
Mechanisms, Prentice Hall, Columbus, OH, Chap. 4.
[4] Keys C.N. & Stepniewski, W.Z., 1984, Rotary
Wing Aerodynamics, Dover Publications, New York,
NY.
[5] https://edge.rit.edu/content/P07122/public/Home
[6] Hibbeler, R.C., 2004, Engineering Mechanics
Dynamics, Prentice Hall, Upper Saddle River, NJ.
[7] Tekalp Murat A., 1995, Digital Video Processing,
Prentice Hall, Upper Saddle River, NJ, Chap. 2.
[8] Grandt, Alten F. Jr., 2004, Fundamentals of
Structural Integrity, John Wiley & Sons, Hoboken, NJ.
Paper Number 07122
Page 9
Proceedings of the KGCOE Multi-Disciplinary Engineering Design Conference
APPENDIX
Nomenclature for Blade Element theory equations:
Ar – Area
A,B – Constants
Pind – Induced power in axial translation
Ppr – Profile power in hovering
Ptot – Total power
Q – Torque
R – Radius of blades
S – Blade area
Tk – Rotor thrust in hover
Vc – Climb velocity
Vt – Tip speed

a – Section lift-curve slope
b – Number of blades
bs – Blade span
c – Chord
cd – Drag coefficient
re – Non-dimensional effective radial distance
rpm – Revolutions per minute
Ω – Rotor rotational speed
θo – Blade section pitch angle
ρ – Air density
σ – Rotor Solidity Ratio
Vt
R
(A1)
 60s
rpm  Vt 
 1 min
 1rev 


 2R m 
(A2)



2
2
1 3 4 A 2 A  3Bre A  Bre
2 2

Th  4R Vt A re  Bre 

3
15B 2

2
Vt 
Pind
2
Th



4 A 2 A 2  3Bre A 2  Bre
1
3
2  2 2
4R  A re  Bre 

3
15 B 2





 3Br  2 A 2 A 2  Br 3 A 2
2
2
2 5 A  Bre
e
e
 4R Vt  

2
5
35
B

2
3
  7A 
7
2
B





3

3
A 2  Bre
(A3)
(A4)




  2A r
5
3
e
2
2

3
 ABre 


(A5)
1
2
Ppr  R 2 Vt c d
8
(A6)
Ptot  Pind  Ppr
(A7)
Q
Ptot

(A8)
Radius of Rotors (r)
0.155
m
Climb Velocity (Vc)
3
Density of Air (ρ)
Blade Section Pitch Angle (θ0)
Non-dimensional effective radial distance (re)
Wing Span (bs)
10
0.9355
0.13
kg/m
degrees
m
Wing Area (Ar)
Approximate Section Lift-Curve Slope (a)
0.0026
6.2832
m2
-
1
.
2
9
0.5
m/s
Chord (c)
0.02
Rotor Solidity Ratio (σ)
Number of Blades (b)
2
A Constant 0.03226
m
B Constant 0.01126
-
Table A1. List of parameters used for Blade Element Theory calculations
Paper Number 07122
0
.
0
8
2
1
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Proceedings of the Multi-Disciplinary Engineering Design Conference
Page 10
Nomenclature for control system design equations:
Fk – Force provided by kth rotor
I – Mass moment of inertia of quadcopter about x or y axis
Iblades – Mass moment of inertia of rotors about z axis
Iframe – Mass moment of inertia of quadcopter about z axis
a – Translational acceleration from rotor forces
d – Half distance between two rotors
m – Mass of quadcopter
x,y,z – Axes of the second coordinate system
αk – Angular acceleration of kth rotor
ωk – Angular velocity of kth rotor
d
F1  F2  F3  F4 
I
d
 y  F1  F2  F3  F4 
I
I Blades
1  2  3  4 
z 
I Frame
x 
(A9)
(A10)
(A11)
dx
  y x Z   z xY  z x X   x x Z  x xY   y x X
dt
dy
  y y Z   z yY  z y X   x y Z  x yY   y y X
dt
dz
  y z Z   z zY  z z X   x z Z  x zY   y z X
dt
z
a  F1  F2  F3  F4 
m
N
CB  1.5e5
(rad / s) 2
Fk  C Bk
2
(A12)
(A13)
(A14)
(A15)
(A16)
(A17)
Figure A1. Simulink® block diagram of control system
Paper Number 07122