Design and Development of a Micro Air Vehicle (MAV): Test-Bed for Vision-Based Control

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DESIGN AND DEVELOPMENT OF A MICRO AIR VEHICLE (MAV):
TEST-BED FOR VISION-BASED CONTROL

















By

SEWOONG JUNG


A THESIS PRESENTED TO THE GRADUATE SCHOOL
OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT
OF THE REQUIREMENTS FOR THE DEGREE OF
MASTER OF SCIENCE

UNIVERSITY OF FLORIDA


2004

































Copyright 2004

by

Sewoong Jung

































This document is dedicated to my parents.















ACKNOWLEDGMENTS

I would like to thank everyone who participated in this project and supporting

systems. Particular thanks go to Dr. Peter Ifju for providing me with the means and

opportunity to work on this project, and Dr. Richard Lind and Dr. David Jenkins for the

vision and guidance to design and fabricate airplanes. I also wish to thank Kyu-Ho Lee

who not only assisted with numerous test flights, but provided invaluable guidance and

man hours in the manufacture, design and assembly of each UAV test-bed. Mention has

to be given to other graduate and undergraduate students who provided numerous man

hours in manufacturing and assembling each UAV test-bed. Roberto Albertani was

another major contributor to the project, in providing all the wind tunnel test data. Each

student from the Flight Controls Lab, Mujahid Abdulrahim, Jason Jackowski, Kenny

Boothe, and Joeseph Kehoe, assisted tremendously with seeing the project through its

development stages. Lastly, I wish to thank my family who supported me in all my

dreams and endeavors while in the United Sates.
















TABLE OF CONTENTS

page

A C K N O W L E D G M E N T S ................................................................................................. iv

L IST O F T A B L E S .................................................................... .......................... .. vii

LIST OF FIGURES ............. ........................ ................. .... .... ............. viii

A B STR A C T ................................................. ..................................... .. x

CHAPTER

1 IN TR OD U CTION ............................................... .. ......................... ..

1.1 Background of Unmanned Aerial Vehicles..........................................................1
1.2 M icro Air Vehicles ................................... ............. ................. .2
1.3 V ision-B ased C control ............................................ .. .. ........ .......... .......
1.4 Overview of Thesis .................. ............................ .. ..... ................ .5

2 UAV REQUIREMENT AND OVERVIEW OF THE FIRST GENERATION TEST-
BED AIRPLANE FOR VISION-BASED CONTROL..............................................7

2.1 Vision-based control UAV Requirements............................................................7
2.1.1 M mission Profile ............................................... .... ........ .......... .......
2.1.2 U A V R equirem ent .............................................. ............................. 7
2.2 Overview of First Generation Test-bed Airplane ................................................8
2.2.1 Fuselage ................................................................... .......... 9
2 .2 .2 W in g ................................................ .......................... 9
2.2.3 Tail and D rive U nit ............................... .... .... .... ...... ................ .. ..11
2.2.4 Fabrication and A ssem bly ........................................ ....... ............... 12
2.2.5 Observations ................... .... ....................... .......... .. .. ................. 13

3 DEVELOPMENT OF THE SECOND GENERATION TEST-BED AIRPLANE
FOR VISION-BASED CONTROL....................................... ......................... 18

3 .1 A ircra ft D e sig n ............................................................................................... 1 8
3.1.1 Components ................................ ............................ ......18
3.1.2 Take-Off W eight ............................ ................... ... .. ............... 20
3.1.3 A irfoil Selection ................... .... .......... ............. ....... .....2 1
3.1.4 W ing D design .................. .................... ................ .............. 23









3.1.5 Fuselage .................................................................... 24
3.1.6 Control Surface A rea..................................................... ...................25
3.2 M anufacturing......... .................................................................. ..................26
3 .2 .1 R apid P rototy ping ............................................................ .....................2 6
3.2.2 Fabrication and A ssem bly ........................................ ....... ............... 27

4 FLIGHT TESTING AND WIND TUNNEL EXPERIMENT..............................30

4 .1 F lig h t T estin g ................................................. ................ 3 0
4.2 W ind Tunnel Experim ent ............................................................................. 31
4.2.1 W ind Tunnel Setup ..................................................... ....................31
4.2.2 W ind Tunnel R results ................................................ ....... ............... 32
4.2.3 3D Digital Image Correlation Setup........... ..... ..................37
4.2.4 3D Digital Image Correlation Results ................................................39

5 C O N C L U SIO N ......... ......................................................................... ........ .. ..... .. 42

APPENDIX

A TH IN AIRFOIL THEORY ........................................ ..... ................. ...............45

B AIRFOIL DRAWING AND LIFT CURVE MATLAB CODE..............................48

L IST O F R E F E R E N C E S ........................ .. ...................................................................54

B IO G R A PH IC A L SK E TCH ..................................................................... ..................56
















LIST OF TABLES


Table page

2-1 Vision-based control UAV requirements ............ ............................................8

2-2 Description of the first generation test-bed UAV. ..................................................14

2-3 Weight distribution for the first generation vision-based UAV.............................16

3-1 Specification of servos for the control surfaces of the second generation test-bed
M A V ...................................... ....................................................... 1 9

3-2 Component mass for the second generation vision-based UAV............................20
















LIST OF FIGURES


Figure page

1-1 Payload mass of Micro Air Vehicle respect to wingspan compared to other
la rg e r U A V s ..............................................................................................................2

1-2 Vision based control hardware. A) Ground station setup, B) MAV128 with
Furuno G P S. ............................................... ............................. 4

2-1 Vision-based autopilot control UAV mission profile. .............................................7

2-2 Assembly sections: a wing, fuselage, horizontal tail, vertical tail, drive system.......9

2-3 UAV wing tool and flexible wing ......................................... ...............10

2-4 Control surfaces: split elevator and rudder. ............................... ....... ........... 11

2-5 Drive system .............. ...... ........... ....... ............ ...... ......... 12

2-6 The first generation test-bed MAV for vision-based control................................13

3-1 Airfoil shape comparison of Gruven and modified MH30 airfoil. ..........................22

3-2 Lift coefficient comparison between Gruven and modified MH30 airfoil. .............22

3-3 Wing generator output using modified MH30 airfoil. ...........................................24

3-4 CAD drawing showing top and side view of fuselage design ..............................25

3-5 W ing m orphing m echanism ............................................ ............................ 26

3-6 A female wing tool milled by CNC milling machine. ...........................................27

3-7 Morphing mechanism for aileron and a pair of wing supports attached to the
fu se la g e ....................................................................... 2 8

3-8 An assembled second generation test-bed airplane for vision-based control...........29

4-1 W ind tunnel testing setup with articulating balance ............................................. 31

4-2 Comparison of lift curve from preliminary test .....................................................32









4-3 Com prison of lift to drag ratio.................................. ............................................ 34

4-4 Comparisons of pitching and yawing moment coefficient curve.............................35

4-5 Drag and rolling moment coefficient due to % control surface deflection ............36

4-6 Comparison of lift coefficient change respect to elevator deflection....................37

4-8 A pair of 3D Digital Image Correlation cameras on top of wind tunnel..................39

4-9 The second generation right-wing deformation in z direction when aerodynamic
lo ad is ap p lied .................................................. ................ 3 9

4-10 The second generation right-wing deformation when aerodynamic load is
applied with wing maximum right turn morphing. ...............................................40

4-11 The second generation right-wing deformation when aerodynamic load is
applied with maximum left turn morphing. .................................. .................41
















Abstract of Thesis Presented to the Graduate School
of the University of Florida in Partial Fulfillment of the
Requirements for the Degree of Master of Science

DESIGN AND DEVELPOMENT OF A MICRO AIR VEHICLE (MAV):
TEST-BED FOR VISION-BASED CONTROL

By

Sewoong Jung

December 2004

Chair: Peter G. Ifju
Major Department: Mechanical and Aerospace Engineering

This thesis presents the design, fabrication, capabilities, and analysis of a test-bed

Micro Air Vehicle (MAV) used in developing a vision-based flight control algorithm.

Vision-based controlled flight and its further developments are geared towards use on

MAVs of the future. A test-bed airplane was designed from MAV concepts, and was

large enough to accommodate the current commercial components available on the

market.

Two generations of the vision-based MAV were developed and tested. Each

generation had 0.6 m wingspan and used a pusher configuration with a flexible wing

concept. The second generation, however, had an improved fuselage and wing design,

drive system, and a unique morphing mechanism. Both vehicles were tested and

compared during numerous flight and wind tunnel tests. Effects of the morphing

mechanism on the flexible wing were analyzed in the wind tunnel using the digital image

correlation method.














CHAPTER 1
INTRODUCTION

1.1 Background of Unmanned Aerial Vehicles

As time presses on, people are looking more towards the autonomy of aerial

vehicles. Unmanned aerial vehicles (UAV) are considered to be the intelligent robots of

the sky. They can be autonomously guided with on-board navigational systems, or

receive positional instruction from a ground control unit. UAVs provide assistance in two

main arenas, reconnaissance for special operations in military applications, and in

communication relays and earth monitoring for government and commercial users [21].

Small UAVs, operating at lower altitudes and on shorter flight times, have wing

spans ranging from 1-3m in length. Currently used autonomous UAVs are the Pointer,

the Dragon Eye, and the Raven, each developed by Aerovironment [18, 19]. They supply

invaluable video images of extremely toxic, remote, or dangerous regions which would

normally endanger human life. These include forest fires, volcanoes, disaster regions, and

front lines of a war zone. Additionally, small UAVs are also used in applications such as

border patrol, gas and pipeline monitoring, and monitoring and tracking wildlife [17].

An autonomously controlled UAV is designed to navigate along a direct path, from

waypoint A to waypoint B, but it does not account for unexpected objects in the path of

its mission. In such instances, the UAV would fly directly into such an obstacle. These

aircraft, although autonomous, are limited by their inability to maneuver around objects

and return to stable flight. This is particularly important when considering missions in

urban arenas or through a thickly populated forest where size and maneuverability are of









high concern. There is however another breed of aircraft that is fully capable of

maneuvering through these densely populated forests, or through an urban arena. The

Micro Air Vehicles (MAVs) are small, lightweight, maneuverable, but highly sensitive

during flight.

1.2 Micro Air Vehicles

Micro air vehicles (MAVs), as defined by Defense Advanced Research Programs

Agency (DARPA), are miniature aircraft with a maximum wing span of 15 cm (about 6

inches) [10]. This physical size limitation puts the MAVs at least an order of magnitude

smaller than any operational UAVs developed.


106 Predator -

105 Pioneer 'to
Global
104 Hawk

101
103 Hunter
S102 Pointer
1 S1 IMALL
S 10 IIAVs
10


0.1 I
0.1 1 10
Wingspan (m)


Figure 1-1. Payload mass of Micro Air Vehicle respect to wingspan compared to other
larger UAVs.

Figure 1-1 is a plot of vehicle payload verses wingspan. It aids one to better

appreciate the size of the MAV compared to that of currently used UAVs. The Pointer,

developed and operated by Aerovironment, is one of the smaller electric UAVs and is









designed for remote monitoring and surveillance. The Pointer has a 2.7 meter wing span

and approximate payload capacity of 900g [18]. Comparatively MAVs are an order of

magnitude smaller and may display a wide variety of configurations, depending on

specific mission requirements [13]. The MAVs developed by the University of Florida,

in particular, vary in wing span from 4 to 6 inches and can carry payloads of up to 15 g.

Similarly to UAVs, micro air vehicles have been identified as having significant

military potential, primarily for reconnaissance, targeting, surveillance and

communication relays. In civilian applications, it is expected to be useful for

biochemical and hazardous material sensing, sensor implantation, and search and rescue

missions. Commercial applications include traffic monitoring, power line inspections,

real estate aerial photography, and wildlife surveys [7].

MAVs have the ability to navigate through close environments with much more

precision than a UAV with a 1 m wingspan. With their small frames and extremely quiet

operating conditions, MAVs are quickly becoming more suitable for extremely discrete

operations. The long term goal would be to have a flying robot that can navigate through

buildings, around people, and still avoid being detected. Flying in such an arena,

however, would require lightning quick reflexes for control, which often would far

exceed that of human ability. In order to push forward with this technology, some amount

of autonomous sensory control is necessary, not only to avoid objects during flight, but

also to return the aircraft to a stable and level flight configuration.

1.3 Vision-Based Control

MAVs seem to fit the bill for the type of flight operations described above, but

there is currently no sensor or navigational hardware small enough to make this possible.

Steps are however being made to make these lacking pieces available to the autonomy









flight puzzle. An autopilot system is currently being developed by the University of

Florida that utilizes a vision-based horizon tracking system combined with waypoint

navigational software and hardware. The flight control system utilizes vision to stabilize

flight, localize targets, track objects, and to avoid obstacles. GPS supplements the vision

system by providing coordinates for waypoint navigation [7]. This new system avoids

the use of excess payloads, and with added advantages. It is still however not to the point

where it is small enough to be tested on a 15cm wingspan MAV. In order to test the

system, a MAV model with larger wing span, flexible wing, and light weight construction

was designed as the test-bed for the vision-based autopilot system.

There are two main components for the flight control system, the ground station

and the on board computer, as shown in Figures 1-2A and 1-2B. The ground station

receives and transmits visual and control data to and from the airplane. It consists of a

transceiver, a 15-inch laptop, a Sony Video Walkman with receiver, and a USB

converter. The custom on board computer, modified from an Atmel microcontroller,

provides all of the data communication with the ground station. It is only 2 x 1.5 x 0.5

inches in dimension and weights 35 grams. This custom microcontroller is so called

MAV128. There is a 900MHz transceiver interfaced to the MAV128 to provide data link

to the ground station at up to 57.6kbps bidirectional.








AiB
Figure 1-2. Vision based control hardware. A) Ground station setup, B) MAV128 with
Furuno GPS.









The control loop, shown in Figure 1-3, starts with collecting all the sensor and

video data, and they are streamed back to the ground station from on-board computer.

After these data are processed, control commands will be sent back to the airplane by a

custom interface and the trainer function on a Futaba RC controller, which allows

switching between computer and human control instantaneously.


- On-Board Camera
- Transmitter

r\
II I


On-Board Control
Receiver
\
I I I
I I I
I I


Figure 1-3. Control loop diagram

1.4 Overview of Thesis

This thesis documents the design and development of the test-bed for a vision

based controlled airplane. In each chapter, design consideration, fabrication and

assembly, and analysis and future work will be discussed. Chapter 2 reviews the current

test-bed airplane for investigating the vision based autonomous flight. From observation

of the current test-bed, it indicates issues to be improved and to be changed. Chapter 3

discusses new design consideration such as airfoil selection for lift calculation, control

surface, and fuselage in order to satisfy mission requirement. It also describes in more

detail the fabrication and assembly of the test-bed airplane. Chapter 4 discusses flight


On-Board
Computer
/\
I I
I I
I I






6


performance and characteristics determined from test flights and wind tunnel testing. It

also evaluates the validity of the morphing mechanism used on the wing of the second

generation test-bed.














CHAPTER 2
UAV REQUIREMENT AND OVERVIEW OF THE FIRST GENERATION TEST-BED
AIRPLANE FOR VISION-BASED CONTROL

2.1 Vision-Based Control UAV Requirements

2.1.1 Mission Profile

The vision-based controlled UAV will mainly be used for reconnaissance. The

mission profile selected to perform our vision-based control UAV reconnaissance is

essentially a loiter-dominated mission, but with a perhaps equally important cruise

segment to the target. Beginning with the takeoff, the vehicle will then cruise to the

specified low altitude and then loiter at the target. Once the essential data has been

transmitted from the target location, which will include one or several types of

reconnaissance information, the UAV will then cruise back to the launch site. A sketch

of the mission profile is shown in Figure 2-1.




Cruise
Loiter Descent
Climb Cruise


Takeoff

Figure 2-1. Vision-based autopilot control UAV mission profile.

2.1.2 UAV requirement

The University of Florida has designed and produced various test-beds that are

suitable for use with the vision-based autopilot control. The first generation test-bed had a

24 inch wing span, which borrowed its shape from a previous tailless model designed at









the university. To solidify the concept of a vision based control aircraft, several test

flights were completed to get a baseline of the requirements needed to complete the

mission. Table 2-1 below gives the summary of the requirements determined from these

test flights.

Table 2-1. Vision-based control UAV requirements
Requirements Values
Operation Range 500 m 1 km
Mission Altitude 30 m 50m
Maximum Flight Time 15 min
Launch Method Hand Launch
Take-off/Landing Distance 10 m
Camera View Clear front view
Speed 24 Km/h 48 Km/h
Cruise Speed 40 Km/h
Wing Span 0.6 m
Wing Application Morphing Mechanism


2.2 Overview of First Generation Test-bed Airplane

The combination of the design requirements and the flight tests provided enough

information to fabricate a test-bed that allowed for the development of the vision-based

autopilot flight control system. The aircraft was also designed to be light weight, easy to

manufacture, in-expensive, expendable, and durable in crashes. In order to accommodate

the ease of assembly, the sections of the aircraft were reduced to a fuselage, a wing, a

pusher drive unit, and a tail with vertical stab and split elevators. Figure 2-2 shows the

sections of the aircraft before assembly.























Figure 2-2. Assembly sections: a wing, fuselage, horizontal tail, vertical tail, drive
system.

2.2.1 Fuselage

The test bed for the autopilot systems required that the aircraft have a fuselage that

was large enough to house all the necessary components. A unique shape was

incorporated to reduce profile drag during flight. The vision-based UAV does not have a

landing gear, so some collision is expected on landing. In an effort to reduce damage to

the components and to the aircraft itself, carbon fiber was chosen as the material for

construction.

A hatch was made in the upper section of the fuselage to allow easy access to

components such as the camera, receiver, and speed controller. The GPS and data link

devises used for autonomous control of the aircraft were placed on the hatch to provide

easy access and to avoid damage upon landing.

2.2.2 Wing

The wing design was inspired by previously developed University of Florida

MAVs. They utilized a combination of biological concepts and composite material to

produce a thin undercambered, flexible wing, similar to those of small birds and bats

[14]. Miniature airplanes are extremely sensitive to wind gusts, however, the because of









a flexible wing, UF MAVs are able to maintain stable flight. This is achieved by the

washout effect that takes place as the wing deforms when excessive load is applied

during windy conditions. After experiencing a wind gust, the wing will return to its

original shape and to stable flight [10]. The washout effect also reduces the induced drag

on the wing tips of the wing and creates a higher lift to drag ratio.

Each wing for the first UAV test-bed was shaped from a wing tool designed for a

tailless airplane with a 24 inch wing span, and a 6 inch root chord. Thus, it had a slight

anhedral in the wing, a reflex camber built into the trailing edge for control surface area,

and a reflex cambered airfoil that extends all the way to the tip of the wing. In order to

overcome the reflex camber effects of the wing tool, construction of each UAV test-bed

wing required a forward shift of the construction material on the wing tool. This shift of

construction material increased the leading edge camber and reduced the reflex camber in

the trailing edge. The anhedral effects present in the wing tool were transformed to a

dihedral configuration after the wing was cured. This was accomplished by cutting the

cured wing in half and then re-adjoining them at with a 6 degree dihedral to increase

stability during turns. The wing tool used for construction and the flexible wing used for

the first version test-bed are shown in Figure 2-3A and 2-3B respectively.












Figure 2-3. UAV wing tool and flexible wing. A) UAV wing tool, B) Deformation of
flexible-wing.









2.2.3 Tail and Drive Unit

As shown in Figure 2-4, a combination of split elevator and rudder were added to

provide increased stability and controllability of the flying platform. These were

especially light weight to accommodate the additional weight from the drive unit being in

the rear of the aircraft.















Figure 2-4. Control surfaces: split elevator and rudder.

A key requirement for the vision based test-bed was to have a clear view of the

horizon from the front of the aircraft. As a result, this forced the design of a pusher

airplane. Having a conventional pusher, however, would increase our size envelope and

create issues with propeller clearance during flight. A unique drive system was

developed that avoided these issues. The drive unit was to be geared, and used to turn a

foldable propeller through the tail boom of the aircraft. Besides providing a clear line of

sight of the horizon during flight and avoiding propeller contact with the tail, the pusher

concept has other advantages. It increases lift on the wing by reducing skin friction drag,

and provides channeled airflow over the tail of the aircraft.

The drive unit is comprised of six major components:

* An aluminum propeller hub fitted with a gear









* Ball bearings, placed inside the propeller hub to allow free rotation of the propeller
and hub irrespective of the rest of the system.

* Foldable propeller blades, which reduce any damage to the motor, drive unit or tail
boom because of clearance issues when landing the aircraft.

* A motor mount, which secures the motor in place and reduces vibrations.

* A hollow aluminum shaft which connects the tail boom, the bearings inside the
propeller hub, and the motor mount.

A second gear, affixed to the motor, turns the gear on the propeller hub. The ratio

between the two gears is 1:1.6. This gear reduction increases efficiency and reduces

cruising speed of the test-bed.












Figure 2-5. Drive system A) Foldable propellers with drive system, B) Assembled drive
system with gears.

2.2.4 Fabrication and Assembly

Once design was complete, a foam mold for the fuselage was made. The fuselage

was made from two layers of pre-preg bidirectional carbon fiber. Between the layers of

carbon fiber was one layer of Kevlar. It was placed on the bottom half of the fuselage.

This was done to strengthen the aircraft in sections that receive high impact while

landing. Eventually, it was cooked in the Autoclave for a cycle of four hours at 2600 F.

The end portion of the fuselage was constructed separately. This was called the fuse-tail.

Separation was necessary as this allows ease of removal of the larger portion of the

fuselage from the foam mold once the curing process was complete.









The flexible wing was designed to have unidirectional carbon fiber strips evenly

placed under a high temperature polymer (vacuum bag material), which acts as the wing

skin. This ensures light weight flexibility along with strength. The leading edge was

made to be extremely stiff to maintain the integrity of the airfoil. The horizontal tail of

the aircraft was fabricated in a similar manner as that of the wing. A vertical tail was

constructed from balsa, sandwiched between two layers of carbon fiber.

A jig, as shown in Figure 2-6.A, was constructed to place a wing, tail, a fuselage

and a drive unit for consistency in design and to reduce asymmetry. Once everything was

aligned with the reference lines on the jig, the parts were strategically secured with

CyanoAcrylate adhesive (CA). Figure 2-6.B shows a completely assembled test-bed

airplane for vision-based control.












A B
Figure 2-6. The first generation test-bed MAV for vision-based control. A) An assemble
jig, B) Assembled test-bed MAV.

2.2.5 Observations

The first generation MAV has satisfied most of the requirements. The MAV can be

launched by human hand conveniently. It can carry a payload of 150 grams, has an

operating range of approximately 3 km with a cruising speed of 40km/h, and can be

landed safely within 10 meters of the landing distance. It is powered by a 2500 rpm/V









electric motor with a 6-3" folding propeller. It has an overall length of 0.52 meter and a

0.32 meter fuselage length. Table 2-2 describes the first generation MAV that was

developed for and used for the vision-based autopilot control.

Table 2-2. Description of the first generation test-bed UAV.
Air Vehicle Specifications
Wing Span 0.6 m
Length Overall 0.52 m
Height Overall 0.2 m
Weight 0.53 kg

Air Vehicle Performance
Endurance 10 min
Maximum Flight Time 15 min
Speed 20 km/h 45 km/h
Cruise Speed 40 km/h

Propulsion
Motor Hacker B20 36S
Power Requirement 12V 7Amps
Propeller rpm 2000
Propeller Graupner Folding Prop 6-3"


This aircraft has been flown successfully to test the vision-based autopilot control

algorithm. Through various flight tests, however, there were concerns being raised about

the complexity of the drive unit, the excessive volume in the fuselage, a moisture

sensitive wing skin, the lack of agility, and the extremely high wing loading being

experienced.

Although the drive system was a unique idea, its complexity made it very tedious to

assemble. It also generated a lot of noise, and lacked the required power to perform

maneuvers during flight. The geared system was difficult to assemble, very fragile

during each landing, and highly unreliable after a crash. The propeller was also

hazardous during hand launches because of its position in the tail boom. Because of the









large volume of the fuselage, it made it difficult for one to grab a hold with one hand for

a hand-launch. The initial fuselage design was more than capable of housing all the

components for the on-board computer, but as time went on, the required components

reduced in size. This caused an excess of unused space in the fuselage. Not only is the

fuselage too large, but its design incorporated a long nose which made the aircraft

unstable in yaw flight conditions.

The lack of agility of the aircraft was attributed to two things, lack of power from

the drive system, and the design of the ailevon (combination of elevator and aileron)

deflection surface. These provided the sole source of deflection for the aircraft during

pitching and rolling maneuvers, but they were not large enough to provide enough

authority during a roll. The aircraft often experienced saturated elevator action, and was

unable to effectively pull out from a roll operation, thus causing a lack of agility. Wing

morphing was then employed to increase the aircraft's agility, but was not favorable due

to the design of the morphing mechanism and the wing skin material being used at the

time. The wing skin material of the first generation test-bed was highly sensitive to

moisture. During damp and/or rainy weather conditions, the wing-skin would go from a

taut surface to one with wrinkles and waves. This drastically affected the performance

during flight.

The first generation test-bed had an extremely high wing loading, comparatively,

the wing loading for the Pointer UAV (wingspan 2.7m) is 5.3 kg/mA2, while the wing

loading for the fist generation test-bed (wingspan 0.6 m) was 7.4 kg/mA2. The

excessively high wing loading was credited to the large weight (530g) of the aircraft. In

order to reduce the wing loading, however, it was necessary to either increase the wing









area or reduce the weight of the aircraft. Table 2-3 gives a weight distribution of the first

generation aircraft.

Table 2-3. Weight distribution for the first generation vision-based UAV
Area Components Weight (g)
Airframe Estimation Wing, fuselage, tails, tail boom, etc 140
Motor with speed controller: Hacker b20 36, 50
Propulsion Propeller: Grapuner 6,3-6,3" 9
Drive System: ball bearing, prop and motor hub, etc 26
Receiver: M5 fma 11
R/C components Servo: JR 241 servo x 3 30
Avionics Data link, GPS, MAV128, ANT, Altimeter 105
Avionics
Video camera, transmitter 27
Power Battery: Thunder Power Li-Polymer 132
2100mAh/11.1V
Total 530


It was not only imperative to increase the wing area, but to redesign the airfoil

being used for the wing. The airfoil being used on the first generation test-bed was not

very effective because of the reflex camber and the added dihedral. It was presumed that

having the dihedral closer to the wing tips instead of at the root chord would increase the

effectiveness of the wing.

Reliability of manufacture and assembly was also another big issue with this

aircraft. In order to completely reproduce one airplane required seven to ten days of

labor. The complexity of the assembly made it even more difficult to manufacture and

produce airplanes that were consistent and precise to the one previous to it. The test-bed

was also not very portable. Its 24 inch wing span made it difficult to pack in a compact

fashion as with the other equipment that was being transported to the flight test locations.

Overall, the next generation test-bed airplane for the vision-based autopilot control

system was designed with less volume and drag in the fuselage. The propulsion system

was changed to direct motor power for more efficiency, less noise, and to aid in the






17


reduction of the wing loading. The new wing design had more wing area, no reflex

camber, and its dihedral placed on the wing tips. All of which contributed to increased lift

and more stable flight conditions. With the new wing design, it was now easier to

incorporate a morphing mechanism on the wing for increased agility in flight. The

airplane obviously has to be faster, easier, and cheaper when it is produced and

assembled.














CHAPTER 3
DEVELOPMENT OF THE SECOND GENERATION TEST-BED AIRPLANE FOR
VISION-BASED CONTROL

Along with the new developments being made in the vision based control hardware

and software, there were also developments being made in the test-bed to house this new

equipment. This chapter focuses on the redesign process of the test-bed, which included

aircraft component selection, determination of the take-off weight, airfoil selection for a

new wing design, smaller fuselage and control surfaces, and incorporation of morphing

technology in the wing. Manufacturing of this new test bed was also streamlined to

produce aircraft at a higher rate.

3.1 Aircraft Design

3.1.1 Components

The components required for the development of the vision-based UAV were the

major determining factors in the design of the aircraft. The test-bed was thus fashioned

around them. They are comprised of the following: a propulsion system (motor and

propeller), video camera and transmitter, servos, and a battery.

The propulsion system utilizes a Hacker brushless motor for a direct drive system

rather than a geared one. The motor is designed mainly for R/C models where an ultra

light motor system is required. The Hacker B20-36 model produces a maximum of

80watts, spins at 27,500 RPM, and weights only 40grams. It is coupled with a Graupner

4.7-4.7 propeller to complete the direct drive propulsion system.









The video camera and transmitter are the most significant components utilized for

vision-based control. They must provide vivid view within the mission range, and have

to be as small and light as possible. The smallest available unit on the market is the color

CMOS video camera which supplies a 310 TV line resolution and has a power

requirement of 20mA at 9VDC. It measures 0.84 x 0.84 inches and weighs 15 grams,

which makes it suitable to fit inside of nose of the new fuselage. The transmitter has a

mass of 8 grams and operates on 2.4 GHz at 250 mW. The range of the video signal has

been tested to more than one mile.

There are a total of three servos being used to actuate the control surfaces of the

vision-based test-bed. Two are used to control the tail elevator and rudder, and are housed

inside the fuselage. The remaining servo is housed on the wing of the aircraft, and is

used to control the morphing mechanism. The first generation UAV used JR digital

actuators, which have very high torque but low speed. In order to overcome the lack of

servo speed, the actuator was changed to a regular ball bearing type from the digital

amplifier. GWS NARO servo was chosen to replace the JR digital servos. Table 3-1

describes the speed and torque of the actuators being used to deflect the control surfaces

of the aircraft.

Table 3-1. Specification of servos for the control surfaces of the second generation test-
bed MAV.
ELEVATOR, RUDDER AILERON (WING MORPHING)
Quantity 2 1
Speed 0.09 sec/60deg 0.14 sec/60deg
Torque 19 oz-in (1.40 kg-cm) 31 oz-in (2.20 Kg-cm)
Weight 8.9 g 10g


All electronic components are powered by a Lithium Polymer (LiPo) battery.

Thunder Power 3 cell LiPo batteries offer the highest capacity of all those commercially









available. Even though it weighs 130g, the 2100mAh and 11.1V it provides, gives

enough power to all the electronic components described above.

3.1.2 Take-Off Weight

The calculation of take-off weight can be performed with relatively little use of

empirical data. This is one advantage of designing a MAV or small electrical power

UAV versus a conventional full-scale aircraft design. In a conventional aircraft, there is

fuel consumption during flight, which changes the weight of the aircraft upon landing. In

MAV design, where electric power is being consumed, there will be no change in weight

from take off to landing. Also, all the component weights were known, which eliminated

the need for iterations in finding the take-off weight.

The last item contributing to take-off weight is the airframe of the UAV. The most

effective way to obtain an estimate for structural mass is to approximate it based on the

structures of previous test-bed airplane. Table 3-2 presents a summary of the component

mass for the second generation vision-based control test-bed. The overall take-off mass

is 500 grams.

Table 3-2. Component mass for the second generation vision-based UAV.
Area Components Weight (g)
Airframe Estimation Wing, fuselage, tails, tail boom, etc 130
Motor with speed controller: Hacker b20 36, 50
Propulsion Propeller & Spinner: Grapuner 4,7-4,7" 12
Receiver: M5 fma 11
components Servo: GWS NARO +HP x 1, GWS NARO HP x 2 35
Avionics Data link, GPS, MAV128, ANT, Altimeter 105
Avionics
Video camera, transmitter 25
Power Battery: Thunder Power Li-Polymer 132
2100mAh/11.1V
Total 500









3.1.3 Airfoil Selection

One of the most important choices in UAV or MAV design is the selection of an

airfoil. Since the wing loading for the first generation vision-based UAV was extremely

high at low speed, a more efficient wing design was required. On the MAV scale, the

airfoil section necessary to produce enough lift at all mission segments must have enough

camber such that it produces a high lift curve slope. This latter specification in necessary

to provide an adequate lift coefficient at a reasonable angle of attack since the finite lift

slope for a low aspect ratio wing is much less than for a high aspect ratio wing [13].

The airfoil geometry for the first generation vision-based test-bed was developed

for the tailless airplane, Gruven. Instead of having a tail or stabilizer, tailless airplanes

rely on a reflex camber at trailing edge of the wing to achieve horizontal stability.

However, to achieve greater controllability of the aircraft a tail is required. A tail would

provide more precision and control to the pilot or the control system during banks and

turns, which is highly favorable when trying to acquire video images. Therefore the new

airfoil was designed without the reflex camber to accommodate the stabilizer.

The new airfoil was designed by using a modified version of the upper half of a

MH30 airfoil, provided by Dr. Martin Hepperle. This airfoil is particularly used for

gliders with Reynolds number in the region of 150,000, which is similar to the Reynolds

number of each UAV test-bed. This is similar to the Gruven airfoil, starting at the leading

edge, but the two diverge at the maximum camber. The modifications of the MH30 airfoil

included and increase in camber from 5% to 7 %, and smoothing of the trailing edge to

eliminate any signs of a reflex in the surface. Figure 3-1 shows the comparison between

the modified upper section of the MH30 airfoil and the Gruven airfoil. Maximum

camber position also moved slightly back from 20% to 25%.














0.4


0.3


0.2


0.1


0


-0.1


-0.2


-0.3


22



Airfoil

Modification
Gruven


0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
X axis

Figure 3-1. Airfoil shape comparison of Gruven and modified MH30 airfoil.


Thin airfoil theory lift curve at 13mis








: I 1


E 0.5
-- - - --- -- -- - -- -- -- -- -


^ ------ ------ -- -^------------- ------.--------- ------- I----- --- -


0.5 "/ -- -- ---------- ----- ------ -------. ------.------- ------ ------


-0.5 ---------------------- ------ ............................ .....




Modification1
Gruven
-1.5 -----i i
-10 -8 -6 -4 -2 0 2 4 6 8 10
Angle of Attack, a (deg)

Figure 3-2. Lift coefficient comparison between Gruven and modified MH30 airfoil.









Thin airfoil theory was used to analyze both airfoils, as this was the most

appropriate method available for the analysis. Figure 3-2 shows clearly that the new

airfoil has a much higher lift coefficient than the 2Griven airfoil at the same angle of

attack. The new model also generated 50% more lift which resulted in positive lift at

zero angle of attack. Additional explanation of the thin airfoil theory can be found in

Appendix A and the MATLAB code supporting the graph and calculations is included in

Appendix B.

3.1.4 Wing Design

The University of Florida MAV team developed software allowing a user to design

a surface model of the wing geometry and create it in a CAD program. This tool speeds

up the design iteration process by just entering the following wing design variables.

* Airfoil Shape

* Wing Span

* Root Chord

* Washout Angle

* Sweep Angle

* Dihedral Shape

Figure 3-3 shows the output drawing of the 24 inch wing span from wing generator

software using the modified upper half of the MH30 airfoil. The new wing has a 6.5 inch

root chord, 2 degrees of washout angle, 2 degrees of dihedral angle, and 5 degrees of

sweep angle. As can be seen in the Figure 3-3, the wing was designed as similar as

possible to an elliptical shaped wing. It bears an increasing polyhedral which increases

slightly from the root chord to the wing tips. Also, the wing tips of the newly designed

wing do not posses an airfoil.












Side View




SA.C. = 2.521 in
Top View










Front View

Figure 3-3. Wing generator output using modified MH30 airfoil.

The dihedral is beneficial in self righting the aircraft from a wind gust. This added

stability makes the plane easier to fly because it eliminates the need for constant

adjustment from the pilot to maintain level wings. Unfortunately, the dihedral also

causes the plane to roll away from the direction of the skid during a side wind gust [20].

Incidentally, sweeping a wing back also gives dihedral effect with about five degrees of

sweep being equivalent to one degree of dihedral [3]. Swept wing brings aerodynamic

center away from the leading edge. This allows the center of gravity to be moved back as

well, while still maintaining stability of the airplane.

3.1.5 Fuselage

The new fuselage design reduced the overall volume by making a thinner and

shorter fuselage with a more compact tail section. This new slender design also

minimizes the aerodynamic drag, and provides a better holding surface for hand

launching the aircraft. As can be seen from Figure 3-4, there are fewer curves in the









surface of the fuselage, thus making manufacture of the aircraft more manageable. A

CAD model of the fuselage for top and side view is shown in Figure 3-4.



















Figure 3-4. CAD drawing showing top and side view of fuselage design.

The Hacker B20-36 motor is placed in the top part of trailing fuselage, and bottom

part is designed to connect a tail boom. It does however use the same hatch method as in

the first generation test-bed. Here, all the computer components for the vision-based

control are mounted in the flat rectangular shape of a hatch.

3.1.6 Control Surface Area

There are three different control surfaces on the aircraft: ailerons (wing morphing),

elevator, and rudder. They are designed to change and control the moments about the roll,

pitch, and yaw axes. To maximize the effect of the control surfaces, 90% of the tail span

and about 20 to 50% of the tail chord are designated as the elevator and the rudder [20].

The tail of the second generation test-bed was made 10% smaller in total surface area

than its predecessor. This was to account for the increase in the length of the moment arm

between the center of gravity the control surface area.









Wing morphing has the capability to change the wing in terms of planform, area,

aspect ratio, and camber to optimize flight performance [3]. The wing of the test-bed is

constructed of a flexible membrane made by carbon fiber ribs and a thin high-

temperature polymer film. Hence, it is easier to adopt the morphing technology unto the

wing of the aircraft. The wing morphing is controlled by a very strong servo that is

mounted through the wing of the aircraft. It operates by controlling two push rods that

connect to control horns, which in turn are affixed to the leading edge of the wing. They

operate in a push/pull fashion, thus converting the wing into a massive aileron during

flight. Figure 3-5 shows an unattached wing with a mounted morphing mechanism.

















Figure 3-5. Wing morphing mechanism.

3.2 Manufacturing

3.2.1 Rapid Prototyping

A Lack of sufficient quantity of airplanes was one of the issues of concern during

the development of the vision-based control system. Thus an improved system was

needed to increase and control repeatability, and to reduce fabrication time on airframes.

A rapid prototype wing generating software solved this problem. Once the software

generated a geometrical shape of the desired wing, it was converted to a 3D wire-frame









drawing in a CAD program. This 3D wire-frame drawing file was then translated to a

tool-path for CNC milling machine. Finally, this machine created a female tooling of the

wing. Figure 3-6 shows a female wing tool being milled out of high density tooling foam

by a CNC milling machine.
















Figure 3-6. A female wing tool milled by CNC milling machine.

Designing and building a fuselage mold using the CNC milling machine is quite

conceivable. Unfortunately, to design complex freeform surfaces for a fuselage using a

CAD program is a highly time consuming process. Working with a tool milled out from

the CNC machine, it is also difficult when trying to cover all the edges during the carbon

fiber lay-up process. It is trivial to construct a mold using traditional methods. It is done

by designing the top and side view of a fuselage and sanding down to the desired shape.

3.2.2 Fabrication and Assembly

Composite materials are used for all the airframes due to its durability and light

weight. As described in Chapter 2, the wing is fabricated on a female tool with uni-

directional and bi-directional carbon fiber. A thin high temperature polymer film which

does not react to moisture substitutes as the wing skin. Two layers of bi-directional

carbon fiber are shaped around a male fuselage mold with an extra layer being placed in









the area of the hatch for reinforcement purpose. The horizontal and vertical tails are

constructed on a flat plate from a 2-D drawing placed on top of the plate. They are made

from multiple layers of uni-directional carbon fiber and the high temperature polymer

film. Eventually, these parts are sealed in a vacuum bag and cured at 260 degrees

Fahrenheit for four hours.

Both the horizontal and vertical stabilizers are attached to a 0.35 inch diameter

carbon tube with cyanoacrylate adhesives while the other end of this tube is connected to

the fuselage. Finally, the assembly is completed by attaching the wing to the top of the

fuselage with a pair of support rods as shown in Figure 3-7. Nylon screws affix the wing

with the fuselage. This new wing mounting method increases the ease of assembly and

packability of the second generation vision based test-bed. The entire fabrication and

assembly processes, including motor and servos installation only requires two days.

Figure 3-8 shows a completely assembled airplane ready to fly.


Figure 3-7. Morphing mechanism for aileron and a pair of wing supports attached to the
fuselage.






29

















Figure 3-8. An assembled second generation test-bed airplane for vision-based control.














CHAPTER 4
FLIGHT TESTING AND WIND TUNNEL EXPERIMENT

4.1 Flight Testing

Several flight tests were successfully completed with the first prototype of the

second generation airplane. In the design process, flight testing was essential to the

evaluation of flying characteristics. Throughout the remote control (RC) flight testing,

designers received instant feedback from pilots and observers. This allowed for

modifications to be made to the airplane before further flight tests and wind tunnel tests

were conducted.

The new fuselage shape places the motor further away from the gripping point of

the aircraft during a hand launch. This ensures that the hand is out of the envelope of the

spinning propeller and ensures safety during a hand launch. Each hand launch of the new

aircraft was now done with more ease, due to the slender and more ergonomically shaped

fuselage.

Though the new model uses the same motor, battery, and payload, it has a higher

rate of climb, a higher glide ratio, and improved turning capability. Further, the new

design improved glide performance by increasing lift. However, a better cooling system

was required for the motor. After flight testing, the motor was seen to overheat to the

point of becoming untouchable. A bigger air intake hole with an aluminum motor mount

possessing more area for airflow proved to be the answer to this need.









4.2 Wind Tunnel Experiment

4.2.1 Wind Tunnel Setup

The University of Florida Department of Mechanical and Aerospace Engineering

has an on campus wind tunnel testing facility. The advanced equipment consists of a six-

component, high-sensitivity sting balance that digitally measured lift, drag, and side-force

loads, as well as the three moments about the balance center. The string balance is

connected to an automated PC data acquisition system. The setup of the experiments on

the test-bed UAV is shown in Figure 4-1.


















Figure 4-1. Wind tunnel testing setup with articulating balance

The plane is held by a drive from an arm which is connected to a brushless

servomotor operated by a single axis motion controller. A modified horizontal tail was

fabricated with 5 layers of bidirectional carbon fiber and was used to aid in suspending

the aircraft from the sting balance. The pitch or yaw angle can automatically be set to

any time-variable angle of attack.










4.2.2 Wind Tunnel Results

Wind tunnel data was acquired for both the first and second generation UAV test-

beds. Wind tunnel testing was performed on a full-scale version of each test-bed. This

included the fuselage, the wing, and the vertical and horizontal tail. The first preliminary

wind tunnel testing was performed with a zero deflection angle on all control surfaces at

different angles of attack. As shown in Figure 4-2, the wind tunnel tests confirm that the

second generation wing generates more lift for a given angle of attack. This correlates

well with the results predicted from using 2-D thin airfoil theory as discussed in chapter 3.

Cl vs. AOA
1.4

---------1.2- -- -

I I U- -1- -1 -I 3-









------- -.4 ----- New 24MAV
S- -- -- 24MAV

0.2

-I I--I-I-I-I
I0 -5 I5 to 15 20 25 30
-- -- 0.2.-JL -- -- -- -- -- -- - --- ------

'- - .4 -- - T- - - - - New_24MAV

--D..V1_24MAV

Angle of Attack (deg)
Figure 4-2. Comparison of lift curve from preliminary test

Over a large range of angle of attack, the new version outperforms the previous

version in terms of higher lift. It has a negative zero-lift angle of attack as shown in the

thin airfoil theory, and is overall a more linear function of angle of attack. However, the

new model has a lower stall angle of about 160 compared to the old version at 200. By









placing the servo above the wing of the new aircraft, it is expected to cause the

detachment of air over the wing sooner than expected. It is however inconsequential,

since the neither aircraft will ever have a need to fly at an angle of attack greater than 100

during normal flight conditions.

Given a known take-off weight and estimated airspeed, it is possible to calculate

the lift coefficient required to sustain level flight. That is,

W
CL,
Sreq. pV2S


where Wis the weight of the aircraft, p is the air density at sea-level, Vis the estimated

cruise speed, and S is the wing area [5]. A reasonable estimate of the cruise speed of the

UAV is about 13 m/s. This approximation is based on test flights of first version

airplanes with the same motor. For a weight of 500 grams, at an approximated airspeed

of 13 m/s, and sea-level conditions, the required CL is about 0.6. According to Figure 4-2,

it can be predicted that new MAV needs slightly less angle of attack to accomplish the

required CL to sustain level flight (6 versus 8).

In addition to generating higher lift on the UAV test-bed, reduced drag is another

design objective. From the flight tests, a direct measure of lift to drag ratio (L/D or

Cl/Cd) was presented. The higher L/D ratio suspected during flight testing was

confirmed with wind tunnel data. From the wind tunnel results, the Cl/Cd can be

calculated and compared with the old model as shown in Figure 4-3.

The new MAV design has a higher lift-to-drag ratio through negative and smaller

angle of attack, up to a point of approximately 7. Beyond this point, however, the drag

of the second generation test-bed is slightly higher than the first generation test-bed, and










results in a fall in the lift to drag ratio. This can be attributed to the servo being placed

atop the wing, the pushrods which protrude from the aircraft, and the connecting columns

between the wing and the fuselage.

Lift over Drag Ratio vs. AOA

S N New 24MAV
------i---- 0- -- --- --D-V1 24MAV

S ---








I I. 10 12 14 15



Angle of Attack (deg)
Figure 4-3. Comparison of lift to drag ratio

For longitudinal motion, pitching moment about the y-axis is the most important

consideration for static stability. If an airplane were statically stable, pitching moment

would tend to rotate the airplane back toward its equilibrium point from a gust of wind or

disturbance that causes the angle of attack to increase or decrease. This happens because

the center of gravity of the airplane would be in front of the neutral point, where the lift

acts and thus restores the vehicle to the equilibrium condition. Therefore, the slope of the

pitching moment coefficient with respect to the angle of attack should be negative [4].

Figure 4-4 illustrates that the second generation test-bed airplane exhibits a

negative slope of the curve of pitching moment coefficient (about the center of gravity)

versus angle of attack. The airplane has improved in static stability. The positive

intercept is shown to be about 2.80, at the equilibrium point when the moment coefficient










is zero. The slope of this curve, the pitching-moment-derivative, is approximately -

0.0207 per degree (-1.18 per radian), a value which is typical of transport-type aircraft.

The corresponding static margin shows a relatively high value of 0.24 (24% of mean

aerodynamic chord). Using a mean aerodynamic chord of 4.4 inch, the dimensional static

margin is 1 inch.


Pitching Moment Coefficient Curve

-*-New 24MAV
.------ V1 24MAV

10 14


-- _- _0. ------------5



S- --------- ---- ---- -- ----- ------



-------------0-.25-------------------------------------------
-"' tL- :.. 2








AOA (Deg)
Figure 4-4. Comparisons of pitching and yawing moment coefficient curve

Rolling stability is a critical factor that was compared for both the first and second

generation test-beds. Based on the second graph presented in Figure 4-5, it shows that

the second generation test-bed has a more effective roll coefficient, determined by the

slope of the graph. Having an effective roll coefficient translates to a lower or more

constant drag on the aircraft during turns. The upper graph in Figure 4-5 also points out

the increase in drag experienced by the first generation test-bed, where the drag for the

second generation test-bed remained almost a constant during a roll or turn. Although the











second generation test-bed shows a slightly higher drag with no control, during increased


angles of attack, it does show more agility during turns and rolling maneuvers because of


its low drag. Results from the wind tunnel shown in Figure 4-5 prove that the morphing


mechanism is more beneficial to roll control than the split elevators used on the first


generation test-bed.


0.15



00

0.05









S0.1
o



-0.15



- -0.2


Rolling Moment Coefficient and Drag

--New 24MAV: CRM
-*-New 24MAV: CD
- V1 24MAV: CRM I
--o--V1 24MAV:CD .

----- ----












-- -- ---------- -------
^--^^ 0--------"^----~-------------
T' o
I I
T2 ,
0T2 0T40r6 01

I I I "N'I. .


Split elevator, Morphing % Deflection

Figure 4-5. Drag and rolling moment coefficient due to % control surface deflection.

For control properties, Figure 4-6 shows the lift coefficient as a function of elevator


deflection for both airplanes. Angles of attack of 1 and 10 degree are plotted for first


generation test-bed; angles of attack of 0 and 10 degree are plotted for the second


generation. The angle of attack does not tend to have a strong relationship with the lift


-_ 0)











curve slopes. For the second generation test-bed airplane, the elevator deflection is

effectively symmetrical from -10 to 10 degree and has an average slope of 0.012 per

degree (0.687 per radian). For the first generation airplane, the elevator deflection is

properly working from -5 to 20 degree, and shows an average slope of 0.013 per degree

(0.748 per radian).


Lift Coeffcient vs. Elevator Deflection

1.4
--New 24MAV: AOA=0
-*-New 24MAV: AOA=10
4 .2 - -
V1 24MAV: AOA1=
--0- V1 24MAV: AOA=10
1- -----------------




---- --------
U
-^-t-





------T- ------ -0.2 ------T------- ---- -----


-0 -O0 -0 .0 10 20 30



0.4





4.2.3 3D Digital Image Correlation Setup

Using Digital Image Correlation techniques, the wing deformation was measured in

the wind tunnel. This method of analysis made it possible to visualize the deformations

of the flexible wing during flight conditions due to morphing and aerodynamic loads.

Especially, 3D Digital Image Correlation is a powerful tool used in measuring full-field

3D of an advanced deformation, stress, and strain. Digital Image Correlation is a data

analysis method which uses a proprietary mathematical correlation method to analyze









digital image data taken while samples undergo normal loadings. Consecutive image

captures taken during the testing phase will show a change in surface characteristics as

the specimen is effected by the loadings. It is commonly used for biomedical

applications, aircraft fuselage or wings, rubber tires and crash testing because it is

substantially more robust, easy to use, and has greater dynamic range [21].



















Figure 4-7. The second generation wing for Digital Image Correlation test.

As Figure 4-7 shows, half the surface area of the wing from a second generation

test-bed is painted with a random black dotted pattern on a white background. This wing

is then viewed in the wind tunnel by a pair of high resolution, digital CCD camera to

measure the wing's synchronized image of 3D coordinates and 3D deformation. When

aerodynamic load is applied, deformation of the random pattern is recorded by the digital

cameras, and a PC correlates the patterns to the deformation, strain, and stress. Figure 4-

8 shows a pair of CCD cameras on top of the wind tunnel where they are used to measure

the wing deformation and morphing at flight conditions.










n, -


Figure 4-8. A pair of 3D Digital Image Correlation cameras on top of wind tunnel.
4.2.4 3D Digital Image Correlation Results
Using digital image correlation techniques, we were able to prove that the flexible

wing does have the ability to generate washout during aerodynamic loading. As discusses

previously, having washout on a wing reduces the induced drag experienced by the wing,

and in effect raises the lift to drag ratio of the aircraft. The extent of washout is

determined by calculating the negative angle between the root chord and the tip chord.


0.827227


w (mm)


9.92573


Figure 4-9. The second generation right-wing deformation in z direction when
aerodynamic load is applied









Based on the results shown in Figure 4-9, the latter end of the wing experiences a

concave upward deformation, but it is even greater in the trailing edge of the wing tip

area. The maximum deflection of the wing tip area was 10mm in the z direction, which

correlates to a negative 3.4 degree difference between the root and the tip chords of the

wing. Neutral conditions during testing were zero wind speed at 10 degrees angle of

attack, and the deformation testing conditions were 13 m/s wind speed at 10 degrees

angle of attack.

Figures 4-10 and 4-11 show the deformation of the flexible wing during right and

left turns respectively using the morphing mechanism. Each figures gives a more in

depth look at exactly how morphing causes changes in the wing shape during a turning

maneuver. Figure 4-10 shows a simulated maximum right turn in the wind tunnel, using

the morphing mechanism, the tip area of the wing experienced a large negative

deformation from its neutral position. The in-board section of the trailing edge, however,

remained at the neutral position during this exercise.

















-8.41803 w (mm) 5.22582

Figure 4-10. The second generation right-wing deformation when aerodynamic load is
applied with wing maximum right turn morphing.


























-0.388694 w (mm) 10.4858

Figure 4-11. The second generation right-wing deformation when aerodynamic load is
applied with maximum left turn morphing.

Figure 4-11 shows a simulated maximum left turn in the wind tunnel, using the

morphing mechanism, and the deformation experienced by the right wing of the aircraft.

The associated left turn shows more of a uniform distribution in deformation over the

entire wing. As expected, the wing tip area experiences the most deflection in the

positive z-direction of approximately 10.5mm, where the root experiences a slight

deflection in the negative z-direction of only -0.3.














CHAPTER 5
CONCLUSION

The design and manufacturing of the first generation test-bed airplane for vision-

based autopilot control has been done successfully. Vision-based controlled flight, and

its further developments are geared towards use on MAVs of the future. However,

current commercial components are not available for MAV size aircraft, so a bigger test-

bed airplane was designed. The first generation test-bed has a wing span of 24inch with a

pusher configuration by a unique geared drive system. The pusher concept allows clear

view from the front to provide better visual data for the vision-based control. A thin and

flexible wing also added to optimize performance while the plane is flying. Through out

the numerous flight testing and wind tunnel results, issues were addressed for a better

implementation of vision-based control algorithms. The five main issues are following:

* High wing loading.

* Complex drive system generated too much noise.

* There were an excess of unused spaced in the fuselage.

* New wing design with non-reflex cambered airfoil was needed.

* Wing morphing mechanism was required to have a more agile airplane.

The second generation test-bed was similar to its predecessor in wingspan and

overall length, but the process of making each airplane was significantly different. By

using a CNC milling machine, the process of producing a wing was now more repeatable

and with a quicker turn around time for any adjustments or new designs. By reducing the

total airframe and propulsion weight by 30 grams, and increasing wing area by 10%, the









wing loading was decreased by 15%. If the hardware weight for the vision-based system

reduces, the wing loading will be reduced even more.

The propulsion system was changed to a direct drive system from the complex

geared one. Flight testing required a better cooling system for this powerful direct drive

unit. Modification was made by making a pair of larger intake hole and by mounting the

motor in a thin aluminum plate with plenty area to allow airflow. The new fuselage

design reduced the overall volume by making a thinner and shorter fuselage with a more

compact tail section. The new wing design also generates higher lift with new airfoil,

and this was confirmed by the thin airfoil theory, wind tunnel testing, and flight testing.

A wing morphing mechanism was also added into the wing as an aileron control surface.

This new concept was successfully demonstrated through a number of flight tests.

Design of the second generation test-bed also allows for the convenient removal

and replacement of the wing from the aircraft. This comes in handy when the airplane is

to be transported, as the wing can be removed from the fuselage, thus reducing the

amount of space the airplane occupies while not in use. This airplane is a good test-bed

for developing the vision-based control algorithm and can also be used in other

applications. It does however show a slight increase in drag from its previous version.

This is attributed to the servo being mounted on the wing, which affects the stall margins,

and the push rods which protrude from the aircraft. Steps are being taken to reduce the

effect of these drag producing areas of the test-bed by covering them with aerodynamic

caps.

Wing tunnel and flight tests proved to be invaluable tools in evaluating and

comparing both the first and second generation UAV test-beds. Digital Image






44


Correlation was used along with wind tunnel testing to further analyze the flexible wing

concept and the benefits of adding the morphing mechanism to the second generation

test-bed. Results obtained from the Digital Image Correlation confirmed the washout

effect experienced by flexible wings as they deform during the application of

aerodynamic loads. It also showed the wing deformation caused by morphing during

simulated turns and flight conditions in the wind tunnel.














APPENDIX A
THIN AIRFOIL THEORY

For a thin airfoil the distribution of a vortex sheet over the surface of the airfoil

looks almost same as a vortex sheet placed on the camber line when view from a

distance. If the airfoil is thin, the camber line is close to the chord. Thus, we can assume

the vortex sheet appears to fall in the chord line. In other words, the airfoil can be

replaced by a vortex sheet along the chord line from the camber line for an airfoil in a


uniform flow, V.













Assuming small angle, the velocity normal to the camber line is induced by the

vortex sheet,

dz
U(s) U(x)= K (a- ) (1)
dx

Considering the strength of the vortex sheet y varies with the distance along the chord,

the velocity dU at point x induced by the elemental vortex at point x':

Y(x')dx'
dU(x) (2)
2;r(x x')









So, the total induced velocity at the point x is given by:

1 !7(x')dx'
U(x) = 2- xx' (3)
2;rO x-x'

Substituting Equation (3) into (1), we obtain:

1 7(x')dx' dz
= (a ) (4)
2; x x-x' dx

It is convenient to introduce the variable 0

cos = 1- 2x (5)

and to write as a Fourier series:

1+ cosO
y(0) = 2V (A no + A, sin nO) (6)
sin 0 .=,

where Ao,A1,A2,... are constants to be determined in terms of the angle of attack and the

slope of the camber line dz/dx.

Substituting (6) into (4) and carrying out the integration using trigonometric relations:

[cos(n 1)x cos(n + 1)x] = 2 sin x sin nx
cosnO' dO' sin n
----- d6'= n--
Scos 0'- cos sin 0

the equation becomes

co dz
A A cosnO =a-- (7)
n=l dx

Finally, multiply both sides by cos mO and integrate from 0 to r

1 'dz
Ao =a- -dO
(8)
S C dx







47


All aerodynamic quantities can now be calculated from Equation (8) using Equation (6).

For lift and drag coefficient,


1
S= pV.ydx
0
A
C, = 2.r(Ao + )
2

mLE =pV.yxdx
0
4 A2A
Cm= (A +A1 2)
2 2


C --(A-A2)
4 4















APPENDIX B
AIRFOIL DRAWING AND LIFT CURVE MATLAB CODE

%%Airfoil Drawing MATLAB Code

Clear all

Close all

%% First modification airfoil

x=0:0.05:1; %% x position variables

%% coefficients of the airfoil equation

Cl=-0.0118;

C2=1.1415;

C3=-3.3528;

C4=4.117;

C5=-2.7207;

C6=0.812;

C7=0.0004;

%% z position variables

for i=l:length(x)

zl(i)= Cl*x(i)A6 + C2*x(i)A5 + C3*x(i)A4 + C4*x(i)A3 + C5*x(i)A2 + C6*x(i) + C7;

end

zl=double(zl);

%% Gruven airfoil

%% coefficients of the airfoil equation









D1=-1.1588;

D2=4.8332;

D3=-8.2081;

D4=7.3023;

D5=-3.5916;

D6=0.8227;

D7=0.0004;

%% z position variables

for i=l:length(x)

z2(i)= Dl*x(i)A6 + D2*x(i)A5 + D3*x(i)A4 + D4*x(i)A3 + D5*x(i)A2 + D6*x(i) + D7;

end

z2=double(z2);

%% Second modification airfoil

%% coefficients of the airfoil equation

E1=-4.1632;

E2=13.8648;

E3=-17.7259;

E4= 11.6092;

E5=-4.4595;

E6=0.8742;

E7=0.0006;

for i=l:length(x)

z3(i)= El*x(i)A6 + E2*x(i)A5 + E3*x(i)A4 + E4*x(i)A3 + E5*x(i)A2 + E6*x(i) + E7;









end

z3=double(z3);

plot(x,z2,'b',x,z3,'r')

%axis([0 1 -0.1 0.1])

xlabel('\bf{X axis}','fontsize',10)

ylabel('\bf{Y axis }','fontsize',10)

title('\bf{Airfoil }','fontsize',12)

legend('Modification','Gruven')

axis equal



%%Lift Curve MATLAB Code

clear all

close all

%% First modification airfoil

%% coefficients of the airfoil equation

C1=-0.0118;

C2=1.1415;

C3=-3.3528;

C4=4.117;

C5=-2.7207;

C6=0.812;

C7=0.0004;

rho = 1.225; % air density (kg/mA3)









v = 13; % cruise speed (m/s)

pi =3.14;

S = 0.079; % planform area of wing

syms theta;

x=(1-cos(theta))/2;

%%Equation of the Airfoil

z = C1*xA6 + C2*xA5 + C3*xA4 + C4*xA3 + C5*xA2 + C6*x + C7;

%%Derivative of z Respect to x

dzdx = 6*C1*xA5 + 5*C2*xA4 + 4*C3*xA3 + 3*C4*xA2 + 2*C5*x + C6;

cf = vpa(dzdx*(cos(theta)- 1),5);

alpzero= vpa(-1/pi*int(cf,0,pi),5) % zero lift angle

alpha=-10*pi/180:l*pi/180:10*pi/180 % angle of attack

%%Lift Coefficients

for i=l:length(alpha)

Cl(i) = vpa(2*pi*(alpha(i) alpzero),5);

end

Cl=double(C1);

%% Second Modification Airfoil

%% Coefficients of the Airfoil Equation

D1=-1.1588;

D2=4.8332;

D3=-8.2081;

D4=7.3023;









D5=-3.5916;

D6=0.8227;

D7=0.0004;

%% Equation of the Airfoil

z2 = D1*xA6 + D2*xA5 + D3*xA4 + D4*xA3 + D5*xA2 + D6*x + D7;

%% Derivative of z Respect to x

dzdx2 = 6*D1*xA5 + 5*D2*xA4 + 4*D3*xA3 + 3*D4*xA2 + 2*D5*x + D6;

cf2 = vpa(dzdx2*(cos(theta)-1),5);

alpzero2= vpa(-1/pi*int(cf2,0,pi),5); % zero lift angle

%% Lift Coefficient

for i=l:length(alpha)

C12(i) = vpa(2*pi*(alpha(i) alpzero2),5);

end

C12=double(C12);

%% Griven Airfoil

%% Coefficients of the Airfoil Equation

E1=-4.1632;

E2=13.8648;

E3=-17.7259;

E4= 11.6092;

E5=-4.4595;

E6=0.8742;

E7=0.0006;









%% Equation of the Airfoil

z3= E1*x6 + E2*xA5 + E3*xA4 + E4*x3 + E5*x2 + E6*x + E7;

% Derivative of z Respect to x

dzdx3 = 6*E1*xA5 + 5*E2*xA4 + 4*E3*xA3 + 3*E4*xA2 + 2*E5*x + E6;

cf3 = vpa(dzdx3*(cos(theta)-1),5);

alpzero3= vpa(-1/pi*int(cf3,0,pi),5); % zero lift angle

%% Lift Coefficient

for i=l:length(alpha)

C13(i) = vpa(2*pi*(alpha(i) alpzero3),5);

end

C13=double(C13);

%plot(alpha* 180/pi,Cl,'r',alpha* 180/pi,C12,'b',alpha* 180/pi,C13,'g')

plot(alpha* 180/pi,C12,'b',alpha* 180/pi,C13,'g')

xlabel('\bf{Angle of Attack, \alpha (deg)}','fontsize',10)

ylabel('\bf{Lift Coefficeint, Cl}','fontsize',10)

title('\bf{Thin airfoil theory lift curve at 13m/s}','fontsize',12)

legend('Modificationl','Gruven',4)

%legend('Modification 1','MH30','Gruven',4)

grid on

hold
















LIST OF REFERENCES


1. M. Abdulrahim, R. Albertani, P. Bamswell, F. Boria, D. Claxton, J. Clifton, J.
Cocquyt, K. Lee, S Mitryk, and P. Ifju, "Design of the University of Florida
Surveillance and Endurance Micro Air Vehicles," 7th Annual Micro Air Vehicle
Competition Entry, April 2003.

2. R. Albertani, P. Barnswell, F. Boria, D. Claxton, J. Clifton, J. Cocquyt, A. Crespo,
C. Francis, P. Ifju, B. Johnson, S. Jung, K. lee, and M. Morton, "University of
Florida Biologically Inspired Micro Air Vehicel," 8th Annual Micro Air Vehicle
Competition Entry, April 2004.

3. J. D. Anderson, Aircraft Performance and Design, WCB/McGraw-Hill, Boston,
1998.

4. E. Bernard and R. L. Duff, Dynamics of Flight: Stability and Control, John Wiley
& Sons, INC., New York, 1996.

5. M. Burgart, J. Miller, and L. Murphy, "Design of a Micro Air Vehicle for the 2000
MAV Competition," internal progress report, University of Notre Dame, 2000.

6. M. Drela (MIT Aero & Astro), XFOIL v6.91 http://raphael.mit.edu/xfoil/, 20 July
2004.

7. S. Ettinger, M. C. Nechyba, P. G. Ifju and M. Waszak, "Vision-Guided Flight
Stability and Control for Micro Air Vehicles," Proc. IEEE Int. Conf. on Intelligent
Robots and Systems, vol. 3, pp. 2134-40, August 2002.

8. J. M. Grasmeyer and M. T. Keennon, "Development of the Black Widow Micro
Air Vehicle," AIAA APATC, AIAA Paper 2001-0127, January 2001.

9. P. G. Ifju, S. Ettinger, D. A. Jenkins, Y. Lian,W. Shyy and M. R.Waszak,
"Flexible-Wing-Based Micro Air Vehicles," 40th AIAA Aerospace Sciences
Meeting, Reno, NV, AIAA 2002-0705, January 2002.

10. P. G. Ifju, S. Ettinger, D. A. Jenkins and L. Martinez, "Compostie Materials for
Micro Air Vehicles," SAMPE Journal, vol. 37 No. 4, pp. 7-13, July/August 2001.

11. S. M. Kanowitz, "Desgin and Implementation of A GPS-based Navigation System
for Micro Air Vehicles," M.S. Thesis, Electrical and Computer Engineering,
University of Florida, August 2002.









12. A. Lennon, The Basics of R/C Model Aircraft Design, Air Age Inc., Ridgefield,
1999.

13. P. B. Lissaman, "Low-Reynolds-Number Airfoils," Annual Review of Fluid
Mechanics, Vol. 15, pp. 223-239, January 1983.

14. J. M. McMichael and Col. M. S. Francis, "Micro Air Vehicles--Toward a New
Dimension in Flight," http://www.darpa.mil/tto/mav/mav auvsi.html, December
1997.

15. T. J. Muller, "Low Reynolds Number Aerodynamics of Low Aspect Ratio,
Thin/Flat/Cambered-Plate Wings," Journal of Aircraft, vol. 37 No. 5, pp. 825-832,
August 2000.

16. R.C. Nelson, Flight Stability and Automatic Control, WCB/McGraw-Hill, Boston,
1998.

17. N. Newcome, "News Room" UAV Forum 2001. SRA International, Inc.
http://www.uavforum.com/library/news.htm, 4 October 2003.

18. A. Parsch, "AeroVironment FQM-151 Pointer" Directory of U. S. Military Rockets
and Missiles. 2004. Designation-Systems.Net. http://www. designation-
systems.net/dusrm/m-151.html, 24 March 2004.

19. J. Pike, "Dragon Eye" Intelligence Resources. 2000. GlobalSecurity.org.
http://www.globalsecurity.org/intell/systems/dragon-eye.htm, 21 December 2003.

20. D. P. Raymer, Aircraft Design: A Conceptual Approach, AIAA Education Series
third edition, AIAA, Inc., Reston, 1999.

21. C. Sutton, S. Helm, P. McNeil "Improved 3-D Image Correlation for Surface
Displacement Measurement," Experimental Techniques, vol. 26 No. 5, pp. 39-42,
May/June 2003.

22. S. Waddington, "Commercial and Civil Missions for Public Service Agencies: Are
UAVs a Viable Option?" Unmanned Vehicles Magazine Business Analysis Forecast
Dec. 2002. UAV World. Brass Trading Ltd. http://www.uavworld.com/civil.htm,
29 April 2004.

23. M. R. Waszak, L. N. Jenkins, and P.G. Ifju, "Stability and Control Properties of an
Aeroelastic Fixed Wing Micro Aerial Vehicle," AIAA Atmospheric Flight
Mechanics Conference, Montreal, Canada, AIAA 2001-4005, 6-9, August 2001.















BIOGRAPHICAL SKETCH

Sewoong Jung was born in Seoul, Korea, on Jan 18, 1980. He came to the U.S. to

attend the White Mountain School in Bethlehem, NH, in 1995. After he received his high

school diploma with honor, he attended the Boston University. He received his

bachelor's degree in aerospace engineering in May 2002. He has worked in the Micro

Air Vehicle lab under Dr. Peter Ifju and will receive his Master of Science in aerospace

engineering in August 2004.

PAGE 1

DESIGN AND DEVELOPMENT OF A MICRO AIR VEHICLE (MAV): TEST-BED FOR VISION-BASED CONTROL By SEWOONG JUNG A THESIS PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLOR IDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF SCIENCE UNIVERSITY OF FLORIDA 2004

PAGE 2

Copyright 2004 by Sewoong Jung

PAGE 3

This document is dedicated to my parents.

PAGE 4

ACKNOWLEDGMENTS I would like to thank everyone who participated in this project and supporting systems. Particular thanks go to Dr. Peter Ifju for providing me with the means and opportunity to work on this project, and Dr. Richard Lind and Dr. David Jenkins for the vision and guidance to design and fabricate airplanes. I also wish to thank Kyu-Ho Lee who not only assisted with numerous test flights, but provided invaluable guidance and man hours in the manufacture, design and assembly of each UAV test-bed. Mention has to be given to other graduate and undergraduate students who provided numerous man hours in manufacturing and assembling each UAV test-bed. Roberto Albertani was another major contributor to the project, in providing all the wind tunnel test data. Each student from the Flight Controls Lab, Mujahid Abdulrahim, Jason Jackowski, Kenny Boothe, and Joeseph Kehoe, assisted tremendously with seeing the project through its development stages. Lastly, I wish to thank my family who supported me in all my dreams and endeavors while in the United Sates. iv

PAGE 5

TABLE OF CONTENTS page ACKNOWLEDGMENTS.................................................................................................iv LIST OF TABLES............................................................................................................vii LIST OF FIGURES.........................................................................................................viii ABSTRACT.........................................................................................................................x CHAPTER 1 INTRODUCTION........................................................................................................1 1.1 Background of Unmanned Aerial Vehicles............................................................1 1.2 Micro Air Vehicles.................................................................................................2 1.3 Vision-Based Control.............................................................................................3 1.4 Overview of Thesis.................................................................................................5 2 UAV REQUIREMENT AND OVERVIEW OF THE FIRST GENERATION TEST-BED AIRPLANE FOR VISION-BASED CONTROL................................................7 2.1 Vision-based control UAV Requirements..............................................................7 2.1.1 Mission Profile.............................................................................................7 2.1.2 UAV Requirement........................................................................................7 2.2 Overview of First Generation Test-bed Airplane...................................................8 2.2.1 Fuselage........................................................................................................9 2.2.2 Wing.............................................................................................................9 2.2.3 Tail and Drive Unit.....................................................................................11 2.2.4 Fabrication and Assembly..........................................................................12 2.2.5 Observations...............................................................................................13 3 DEVELOPMENT OF THE SECOND GENERATION TEST-BED AIRPLANE FOR VISION-BASED CONTROL............................................................................18 3.1 Aircraft Design.....................................................................................................18 3.1.1 Components................................................................................................18 3.1.2 Take-Off Weight........................................................................................20 3.1.3 Airfoil Selection.........................................................................................21 3.1.4 Wing Design...............................................................................................23 v

PAGE 6

3.1.5 Fuselage......................................................................................................24 3.1.6 Control Surface Area..................................................................................25 3.2 Manufacturing.......................................................................................................26 3.2.1 Rapid Prototyping.......................................................................................26 3.2.2 Fabrication and Assembly..........................................................................27 4 FLIGHT TESTING AND WIND TUNNEL EXPERIMENT....................................30 4.1 Flight Testing.......................................................................................................30 4.2 Wind Tunnel Experiment....................................................................................31 4.2.1 Wind Tunnel Setup.....................................................................................31 4.2.2 Wind Tunnel Results..................................................................................32 4.2.3 3D Digital Image Correlation Setup...........................................................37 4.2.4 3D Digital Image Correlation Results........................................................39 5 CONCLUSION...........................................................................................................42 APPENDIX A THIN AIRFOIL THEORY.........................................................................................45 B AIRFOIL DRAWING AND LIFT CURVE MATLAB CODE.................................48 LIST OF REFERENCES...................................................................................................54 BIOGRAPHICAL SKETCH.............................................................................................56 vi

PAGE 7

LIST OF TABLES Table page 2-1 Vision-based control UAV requirements...................................................................8 2-2 Description of the first generation test-bed UAV....................................................14 2-3 Weight distribution for the first generation vision-based UAV...............................16 3-1 Specification of servos for the control surfaces of the second generation test-bed MAV.........................................................................................................................19 3-2 Component mass for the second generation vision-based UAV..............................20 vii

PAGE 8

LIST OF FIGURES Figure page 1-1 Payload mass of Micro Air Vehicle respect to wingspan compared to other larger UAVs...............................................................................................................2 1-2 Vision based control hardware. A) Ground station setup, B) MAV128 with Furuno GPS................................................................................................................4 2-1 Vision-based autopilot control UAV mission profile................................................7 2-2 Assembly sections: a wing, fuselage, horizontal tail, vertical tail, drive system.......9 2-3 UAV wing tool and flexible wing............................................................................10 2-4 Control surfaces: split elevator and rudder..............................................................11 2-5 Drive system.............................................................................................................12 2-6 The first generation test-bed MAV for vision-based control...................................13 3-1 Airfoil shape comparison of Gruven and modified MH30 airfoil...........................22 3-2 Lift coefficient comparison between Gruven and modified MH30 airfoil..............22 3-3 Wing generator output using modified MH30 airfoil..............................................24 3-4 CAD drawing showing top and side view of fuselage design..................................25 3-5 Wing morphing mechanism.....................................................................................26 3-6 A female wing tool milled by CNC milling machine..............................................27 3-7 Morphing mechanism for aileron and a pair of wing supports attached to the fuselage.....................................................................................................................28 3-8 An assembled second generation test-bed airplane for vision-based control...........29 4-1 Wind tunnel testing setup with articulating balance................................................31 4-2 Comparison of lift curve from preliminary test.......................................................32 viii

PAGE 9

4-3 Comparison of lift to drag ratio................................................................................34 4-4 Comparisons of pitching and yawing moment coefficient curve.............................35 4-5 Drag and rolling moment coefficient due to % control surface deflection..............36 4-6 Comparison of lift coefficient change respect to elevator deflection.......................37 4-8 A pair of 3D Digital Image Correlation cameras on top of wind tunnel..................39 4-9 The second generation right-wing deformation in z direction when aerodynamic load is applied...........................................................................................................39 4-10 The second generation right-wing deformation when aerodynamic load is applied with wing maximum right turn morphing...................................................40 4-11 The second generation right-wing deformation when aerodynamic load is applied with maximum left turn morphing..............................................................41 ix

PAGE 10

Abstract of Thesis Presented to the Graduate School of the University of Florida in Partial Fulfillment of the Requirements for the Degree of Master of Science DESIGN AND DEVELPOMENT OF A MICRO AIR VEHICLE (MAV): TEST-BED FOR VISION-BASED CONTROL By Sewoong Jung December 2004 Chair: Peter G. Ifju Major Department: Mechanical and Aerospace Engineering This thesis presents the design, fabrication, capabilities, and analysis of a test-bed Micro Air Vehicle (MAV) used in developing a vision-based flight control algorithm. Vision-based controlled flight and its further developments are geared towards use on MAVs of the future. A test-bed airplane was designed from MAV concepts, and was large enough to accommodate the current commercial components available on the market. Two generations of the vision-based MAV were developed and tested. Each generation had 0.6 m wingspan and used a pusher configuration with a flexible wing concept. The second generation, however, had an improved fuselage and wing design, drive system, and a unique morphing mechanism. Both vehicles were tested and compared during numerous flight and wind tunnel tests. Effects of the morphing mechanism on the flexible wing were analyzed in the wind tunnel using the digital image correlation method. x

PAGE 11

CHAPTER 1 INTRODUCTION 1.1 Background of Unmanned Aerial Vehicles As time presses on, people are looking more towards the autonomy of aerial vehicles. Unmanned aerial vehicles (UAV) are considered to be the intelligent robots of the sky. They can be autonomously guided with on-board navigational systems, or receive positional instruction from a ground control unit. UAVs provide assistance in two main arenas, reconnaissance for special operations in military applications, and in communication relays and earth monitoring for government and commercial users [21]. Small UAVs, operating at lower altitudes and on shorter flight times, have wing spans ranging from 1-3m in length. Currently used autonomous UAVs are the Pointer, the Dragon Eye, and the Raven, each developed by Aerovironment [18, 19]. They supply invaluable video images of extremely toxic, remote, or dangerous regions which would normally endanger human life. These include forest fires, volcanoes, disaster regions, and front lines of a war zone. Additionally, small UAVs are also used in applications such as border patrol, gas and pipeline monitoring, and monitoring and tracking wildlife [17]. An autonomously controlled UAV is designed to navigate along a direct path, from waypoint A to waypoint B, but it does not account for unexpected objects in the path of its mission. In such instances, the UAV would fly directly into such an obstacle. These aircraft, although autonomous, are limited by their inability to maneuver around objects and return to stable flight. This is particularly important when considering missions in urban arenas or through a thickly populated forest where size and maneuverability are of 1

PAGE 12

2 high concern. There is however another breed of aircraft that is fully capable of maneuvering through these densely populated forests, or through an urban arena. The Micro Air Vehicles (MAVs) are small, lightweight, maneuverable, but highly sensitive during flight. 1.2 Micro Air Vehicles Micro air vehicles (MAVs), as defined by Defense Advanced Research Programs Agency (DARPA), are miniature aircraft with a maximum wing span of 15 cm (about 6 inches) [10]. This physical size limitation puts the MAVs at least an order of magnitude smaller than any operational UAVs developed. Figure 1-1. Payload mass of Micro Air Vehicle respect to wingspan compared to other larger UAVs. Figure 1-1 is a plot of vehicle payload verses wingspan. It aids one to better appreciate the size of the MAV compared to that of currently used UAVs. The Pointer, developed and operated by Aerovironment, is one of the smaller electric UAVs and is 0.1 1 10 10 6 Predato r 10 5 Pionee r Global Hawk 10 4 10 10 10 Pa y load Mass ( g ) 3 Hunte r Pointer 2 SMALL UAVs MAVs 1 0.1 Wingspan (m)

PAGE 13

3 designed for remote monitoring and surveillance. The Pointer has a 2.7 meter wing span and approximate payload capacity of 900g [18]. Comparatively MAVs are an order of magnitude smaller and may display a wide variety of configurations, depending on specific mission requirements [13]. The MAVs developed by the University of Florida, in particular, vary in wing span from 4 to 6 inches and can carry payloads of up to 15 g. Similarly to UAVs, micro air vehicles have been identified as having significant military potential, primarily for reconnaissance, targeting, surveillance and communication relays. In civilian applications, it is expected to be useful for biochemical and hazardous material sensing, sensor implantation, and search and rescue missions. Commercial applications include traffic monitoring, power line inspections, real estate aerial photography, and wildlife surveys [7]. MAVs have the ability to navigate through close environments with much more precision than a UAV with a 1 m wingspan. With their small frames and extremely quiet operating conditions, MAVs are quickly becoming more suitable for extremely discrete operations. The long term goal would be to have a flying robot that can navigate through buildings, around people, and still avoid being detected. Flying in such an arena, however, would require lightning quick reflexes for control, which often would far exceed that of human ability. In order to push forward with this technology, some amount of autonomous sensory control is necessary, not only to avoid objects during flight, but also to return the aircraft to a stable and level flight configuration. 1.3 Vision-Based Control MAVs seem to fit the bill for the type of flight operations described above, but there is currently no sensor or navigational hardware small enough to make this possible. Steps are however being made to make these lacking pieces available to the autonomy

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4 flight puzzle. An autopilot system is currently being developed by the University of Florida that utilizes a vision-based horizon tracking system combined with waypoint navigational software and hardware. The flight control system utilizes vision to stabilize flight, localize targets, track objects, and to avoid obstacles. GPS supplements the vision system by providing coordinates for waypoint navigation [7]. This new system avoids the use of excess payloads, and with added advantages. It is still however not to the point where it is small enough to be tested on a 15cm wingspan MAV. In order to test the system, a MAV model with larger wing span, flexible wing, and light weight construction was designed as the test-bed for the vision-based autopilot system. There are two main components for the flight control system, the ground station and the on board computer, as shown in Figures 1-2A and 1-2B. The ground station receives and transmits visual and control data to and from the airplane. It consists of a transceiver, a 15-inch laptop, a Sony Video Walkman with receiver, and a USB converter. The custom on board computer, modified from an Atmel microcontroller, provides all of the data communication with the ground station. It is only 2 x 1.5 x 0.5 inches in dimension and weights 35 grams. This custom microcontroller is so called MAV128. There is a 900MHz transceiver interfaced to the MAV128 to provide data link to the ground station at up to 57.6kbps bidirectional. A B Figure 1-2. Vision based control hardware. A) Ground station setup, B) MAV128 with Furuno GPS.

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5 The control loop, shown in Figure 1-3, starts with collecting all the sensor and video data, and they are streamed back to the ground station from on-board computer. After these data are processed, control commands will be sent back to the airplane by a custom interface and the trainer function on a Futaba RC controller, which allows switching between computer and human control instantaneously. Figure 1-3. Control loop diagram On-Board Camera On-Board Control Receiver On-Board Computer Transmitter Futaba Radio Controller Video Receiver Ground Station Transceiver Sony Walkman 15-inch Powerbook 1.4 Overview of Thesis This thesis documents the design and development of the test-bed for a vision based controlled airplane. In each chapter, design consideration, fabrication and assembly, and analysis and future work will be discussed. Chapter 2 reviews the current test-bed airplane for investigating the vision based autonomous flight. From observation of the current test-bed, it indicates issues to be improved and to be changed. Chapter 3 discusses new design consideration such as airfoil selection for lift calculation, control surface, and fuselage in order to satisfy mission requirement. It also describes in more detail the fabrication and assembly of the test-bed airplane. Chapter 4 discusses flight

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6 performance and characteristics determined from test flights and wind tunnel testing. It also evaluates the validity of the morphing mechanism used on the wing of the second generation test-bed.

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CHAPTER 2 UAV REQUIREMENT AND OVERVIEW OF THE FIRST GENERATION TEST-BED AIRPLANE FOR VISION-BASED CONTROL 2.1 Vision-Based Control UAV Requirements 2.1.1 Mission Profile The vision-based controlled UAV will mainly be used for reconnaissance. The mission profile selected to perform our vision-based control UAV reconnaissance is essentially a loiter-dominated mission, but with a perhaps equally important cruise segment to the target. Beginning with the takeoff, the vehicle will then cruise to the specified low altitude and then loiter at the target. Once the essential data has been transmitted from the target location, which will include one or several types of reconnaissance information, the UAV will then cruise back to the launch site. A sketch of the mission profile is shown in Figure 2-1. Cruise Descent Loite r Cruise Climb Takeoff Figure 2-1. Vision-based autopilot control UAV mission profile. 2.1.2 UAV requirement The University of Florida has designed and produced various test-beds that are suitable for use with the vision-based autopilot control. The first generation test-bed had a 24 inch wing span, which borrowed its shape from a previous tailless model designed at 7

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8 the university. To solidify the concept of a vision based control aircraft, several test flights were completed to get a baseline of the requirements needed to complete the mission. Table 2-1 below gives the summary of the requirements determined from these test flights. Table 2-1. Vision-based control UAV requirements Requirements Values Operation Range Mission Altitude 500 m 1 km 30 m 50m Maximum Flight Time 15 min Launch Method Hand Launch Take-off/Landing Distance 10 m Camera View Clear front view Speed 24 Km/h 48 Km/h Cruise Speed 40 Km/h Wing Span 0.6 m Wing Application Morphing Mechanism 2.2 Overview of First Generation Test-bed Airplane The combination of the design requiremen ts and the flight tests provided enough information to fabricate a test-bed that allo wed for the development of the vision-based autopilot flight control system. The aircraft was also designed to be light weight, easy to manufacture, in-expensive, expendable, and dur able in crashes. In order to accommodate the ease of assembly, the sections of the ai rcraft were reduced to a fuselage, a wing, a pusher drive unit, and a tail with vertical st ab and split elevators. Figure 2-2 shows the sections of the aircraft before assembly.

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9 Figure 2-2. Assembly sections: a wing, fuselage, horizontal tail, vertical tail, drive system. 2.2.1 Fuselage The test bed for the autopilot systems required that the aircraft have a fuselage that was large enough to house all the necessary components. A unique shape was incorporated to reduce profile drag during flight. The vision-based UAV does not have a landing gear, so some collision is expected on landing. In an effort to reduce damage to the components and to the aircraft itself, carbon fiber was chosen as the material for construction. A hatch was made in the upper section of the fuselage to allow easy access to components such as the camera, receiver, and speed controller. The GPS and data link devises used for autonomous control of the aircraft were placed on the hatch to provide easy access and to avoid damage upon landing. 2.2.2 Wing The wing design was inspired by previously developed University of Florida MAVs. They utilized a combination of biological concepts and composite material to produce a thin undercambered, flexible wing, similar to those of small birds and bats [14]. Miniature airplanes are extremely sensitive to wind gusts, however, the because of

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10 a flexible wing, UF MAVs are able to maintain stable flight. This is achieved by the washout effect that takes place as the wing deforms when excessive load is applied during windy conditions. After experiencing a wind gust, the wing will return to its original shape and to stable flight [10]. The washout effect also reduces the induced drag on the wing tips of the wing and creates a higher lift to drag ratio. Each wing for the first UAV test-bed was shaped from a wing tool designed for a tailless airplane with a 24 inch wing span, and a 6 inch root chord. Thus, it had a slight anhedral in the wing, a reflex camber built into the trailing edge for control surface area, and a reflex cambered airfoil that extends all the way to the tip of the wing. In order to overcome the reflex camber effects of the wing tool, construction of each UAV test-bed wing required a forward shift of the construction material on the wing tool. This shift of construction material increased the leading edge camber and reduced the reflex camber in the trailing edge. The anhedral effects present in the wing tool were transformed to a dihedral configuration after the wing was cured. This was accomplished by cutting the cured wing in half and then re-adjoining them at with a 6 degree dihedral to increase stability during turns. The wing tool used for construction and the flexible wing used for the first version test-bed are shown in Figure 2-3A and 2-3B respectively. A B Figure 2-3. UAV wing tool and flexible wing. A) UAV wing tool, B) Deformation of flexible-wing.

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11 2.2.3 Tail and Drive Unit As shown in Figure 2-4, a combination of split elevator and rudder were added to provide increased stability and controllability of the flying platform. These were especially light weight to accommodate the additional weight from the drive unit being in the rear of the aircraft. Figure 2-4. Control surfaces: split elevator and rudder. A key requirement for the vision based test-bed was to have a clear view of the horizon from the front of the aircraft. As a result, this forced the design of a pusher airplane. Having a conventional pusher, however, would increase our size envelope and create issues with propeller clearance during flight. A unique drive system was developed that avoided these issues. The drive unit was to be geared, and used to turn a foldable propeller through the tail boom of the aircraft. Besides providing a clear line of sight of the horizon during flight and avoiding propeller contact with the tail, the pusher concept has other advantages. It increases lift on the wing by reducing skin friction drag, and provides channeled airflow over the tail of the aircraft. The drive unit is comprised of six major components: An aluminum propeller hub fitted with a gear

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12 Ball bearings, placed inside the propeller hub to allow free rotation of the propeller and hub irrespective of the rest of the system. Foldable propeller blades, which reduce any damage to the motor, drive unit or tail boom because of clearance issues when landing the aircraft. A motor mount, which secures the motor in place and reduces vibrations. A hollow aluminum shaft which connects the tail boom, the bearings inside the propeller hub, and the motor mount. A second gear, affixed to the motor, turns the gear on the propeller hub. The ratio between the two gears is 1:1.6. This gear reduction increases efficiency and reduces cruising speed of the test-bed. A B Figure 2-5. Drive system A) Foldable propellers with drive system, B) Assembled drive system with gears. 2.2.4 Fabrication and Assembly Once design was complete, a foam mold for the fuselage was made. The fuselage was made from two layers of pre-preg bidirectional carbon fiber. Between the layers of carbon fiber was one layer of Kevlar. It was placed on the bottom half of the fuselage. This was done to strengthen the aircraft in sections that receive high impact while landing. Eventually, it was cooked in the Autoclave for a cycle of four hours at 260 F. The end portion of the fuselage was constructed separately. This was called the fuse-tail. Separation was necessary as this allows ease of removal of the larger portion of the fuselage from the foam mold once the curing process was complete.

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13 The flexible wing was designed to have unidirectional carbon fiber strips evenly placed under a high temperature polymer (vacuum bag material), which acts as the wing skin. This ensures light weight flexibility along with strength. The leading edge was made to be extremely stiff to maintain the integrity of the airfoil. The horizontal tail of the aircraft was fabricated in a similar manner as that of the wing. A vertical tail was constructed from balsa, sandwiched between two layers of carbon fiber. A jig, as shown in Figure 2-6.A, was constructed to place a wing, tail, a fuselage and a drive unit for consistency in design and to reduce asymmetry. Once everything was aligned with the reference lines on the jig, the parts were strategically secured with CyanoAcrylate adhesive (CA). Figure 2-6.B shows a completely assembled test-bed airplane for vision-based control. A B Figure 2-6. The first generation test-bed MAV for vision-based control. A) An assemble jig, B) Assembled test-bed MAV. 2.2.5 Observations The first generation MAV has satisfied most of the requirements. The MAV can be launched by human hand conveniently. It can carry a payload of 150 grams, has an operating range of approximately 3 km with a cruising speed of 40km/h, and can be landed safely within 10 meters of the landing distance. It is powered by a 2500 rpm/V

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14 electric motor with a 6-3 folding propeller It has an overall le ngth of 0.52 meter and a 0.32 meter fuselage length. Table 2-2 desc ribes the first generation MAV that was developed for and used for the vi sion-based autopilot control. Table 2-2. Description of the first generation test-bed UAV. Air Vehicle Specifications Wing Span 0.6 m Length Overall 0.52 m Height Overall 0.2 m Weight 0.53 kg Air Vehicle Performance Endurance 10 min Maximum Flight Time 15 min Speed 20 km/h 45 km/h Cruise Speed 40 km/h Propulsion Motor Hacker B20 36S Power Requirement 12V 7Amps Propeller rpm 2000 Propeller Graupner Folding Prop 6-3" This aircraft has been flown successfully to test the vision-based autopilot control algorithm. Through various flight tests, however there were concerns being raised about the complexity of the drive unit, the exce ssive volume in the fuselage, a moisture sensitive wing skin, the lack of agility, and the extrem ely high wing loading being experienced. Although the drive system was a unique idea, its complexity made it very tedious to assemble. It also generated a lot of noise, and lacked the required power to perform maneuvers during flight. The geared system was difficult to assemble, very fragile during each landing, and highl y unreliable after a crash. The propeller was also hazardous during hand launches because of its po sition in the tail boom. Because of the

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15 large volume of the fuselage, it made it difficult for one to grab a hold with one hand for a hand-launch. The initial fuselage design was more than capable of housing all the components for the on-board computer, but as time went on, the required components reduced in size. This caused an excess of unused space in the fuselage. Not only is the fuselage too large, but its design incorporated a long nose which made the aircraft unstable in yaw flight conditions. The lack of agility of the aircraft was attributed to two things, lack of power from the drive system, and the design of the ailevon (combination of elevator and aileron) deflection surface. These provided the sole source of deflection for the aircraft during pitching and rolling maneuvers, but they were not large enough to provide enough authority during a roll. The aircraft often experienced saturated elevator action, and was unable to effectively pull out from a roll operation, thus causing a lack of agility. Wing morphing was then employed to increase the aircrafts agility, but was not favorable due to the design of the morphing mechanism and the wing skin material being used at the time. The wing skin material of the first generation test-bed was highly sensitive to moisture. During damp and/or rainy weather conditions, the wing-skin would go from a taut surface to one with wrinkles and waves. This drastically affected the performance during flight. The first generation test-bed had an extremely high wing loading, comparatively, the wing loading for the Pointer UAV (wingspan 2.7m) is 5.3 kg/m^2, while the wing loading for the fist generation test-bed (wingspan 0.6 m) was 7.4 kg/m^2. The excessively high wing loading was credited to the large weight (530g) of the aircraft. In order to reduce the wing loading, however, it was necessary to either increase the wing

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16 area or reduce the weight of the aircraft. Table 2-3 gives a weight distribution of the first generation aircraft. Table 2-3. Weight distribution for the first generation vision-based UAV Area Components Weight (g) Airframe Estimation Wing, fuselage, tails, tail boom, etc 140 Motor with speed controller: Hacker b20 36, 50 Propeller: Grapuner 6,3-6,3 9 Propulsion Drive System: ball bearing, prop and motor hub, etc 26 Receiver: M5 fma 11 R/C componets Servo: JR 241 servo x 3 30 Data link, GPS, MAV128, ANT, Altimeter 105 Avionics Video camera, transmitter 27 Power Battery: Thunder Power Li-Polymer 2100mAh/11.1V 132 Total 530 It was not only imperative to increase the wing area, but to redesign the airfoil being used for the wing. The airfoil being used on the first generation test-bed was not very effective because of the reflex camber and the added dihedral. It was presumed that having the dihedral closer to the wing tips instead of at the root chord would increase the effectiveness of the wing. Reliability of manufacture and assembly was also another big issue with this aircraft. In order to completely reproduce one airplane required seven to ten days of labor. The complexity of the assembly made it even more difficult to manufacture and produce airplanes that were consistent and precise to the one previous to it. The test-bed was also not very portable. Its 24 inch wing span made it difficult to pack in a compact fashion as with the other equipment that was being transported to the flight test locations. Overall, the next generation test-bed airplane for the vision-based autopilot control system was designed with less volume and drag in the fuselage. The propulsion system was changed to direct motor power for more efficiency, less noise, and to aid in the

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17 reduction of the wing loading. The new wing design had more wing area, no reflex camber, and its dihedral placed on the wing tips. All of which contributed to increased lift and more stable flight conditions. With the new wing design, it was now easier to incorporate a morphing mechanism on the wing for increased agility in flight. The airplane obviously has to be faster, easier, and cheaper when it is produced and assembled.

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CHAPTER 3 DEVELOPMENT OF THE SECOND GENERATION TEST-BED AIRPLANE FOR VISION-BASED CONTROL Along with the new developments being made in the vision based control hardware and software, there were also developments being made in the test-bed to house this new equipment. This chapter focuses on the redesign process of the test-bed, which included aircraft component selection, determination of the take-off weight, airfoil selection for a new wing design, smaller fuselage and control surfaces, and incorporation of morphing technology in the wing. Manufacturing of this new test bed was also streamlined to produce aircraft at a higher rate. 3.1 Aircraft Design 3.1.1 Components The components required for the development of the vision-based UAV were the major determining factors in the design of the aircraft. The test-bed was thus fashioned around them. They are comprised of the following: a propulsion system (motor and propeller), video camera and transmitter, servos, and a battery. The propulsion system utilizes a Hacker brushless motor for a direct drive system rather than a geared one. The motor is designed mainly for R/C models where an ultra light motor system is required. The Hacker B20-36 model produces a maximum of 80watts, spins at 27,500 RPM, and weights only 40grams. It is coupled with a Graupner 4.7-4.7 propeller to complete the direct drive propulsion system. 18

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19 The video camera and transmitter are the most significant components utilized for vision-based control. They must provide vivid view within the mission range, and have to be as small and light as possible. The smallest available unit on the market is the color CMOS video camera which supplies a 310 TV line resolution and has a power requirement of 20mA at 9VDC. It measures 0.84 x 0.84 inches and weighs 15 grams, which makes it suitable to fit inside of nose of the new fuselage. The transmitter has a mass of 8 grams and operates on 2.4 GHz at 250 mW. The range of the video signal has been tested to more than one mile. There are a total of three servos being used to actuate the control surfaces of the vision-based test-bed. Two are used to control the tail elevator and rudder, and are housed inside the fuselage. The remaining servo is housed on the wing of the aircraft, and is used to control the morphing mechanism. The first generation UAV used JR digital actuators, which have very high torque but low speed. In order to overcome the lack of servo speed, the actuator was changed to a regular ball bearing type from the digital amplifier. GWS NARO servo was chosen to replace the JR digital servos. Table 3-1 describes the speed and torque of the actuators being used to deflect the control surfaces of the aircraft. Table 3-1. Specification of servos for the control surfaces of the second generation test-bed MAV. ELEVATOR, RUDDER AILERON (WING MORPHING) Quantity 2 1 Speed 0.09 sec/60deg 0.14 sec/60deg Torque 19 oz-in (1.40 kg-cm) 31 oz-in (2.20 Kg-cm) Weight 8.9 g 10 g All electronic components are powered by a Lithium Polymer (LiPo) battery. Thunder Power 3 cell LiPo batteries offer the highest capacity of all those commercially

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20 available. Even though it weighs 130g, the 2100mAh and 11.1V it provides, gives enough power to all the electronic components described above. 3.1.2 Take-Off Weight The calculation of take-off weight can be performed with relatively little use of empirical data. This is one advantage of designing a MAV or small electrical power UAV versus a conventional full-scale aircraft design. In a conventional aircraft, there is fuel consumption during flight, which changes the weight of the aircraft upon landing. In MAV design, where electric power is being consumed, there will be no change in weight from take off to landing. Also, all the component weights were known, which eliminated the need for iterations in finding the take-off weight. The last item contributing to take-off weight is the airframe of the UAV. The most effective way to obtain an estimate for structural mass is to approximate it based on the structures of previous test-bed airplane. Table 3-2 presents a summary of the component mass for the second generation vision-based control test-bed. The overall take-off mass is 500 grams. Table 3-2. Component mass for the second generation vision-based UAV. Area Components Weight (g) Airframe Estimation Wing, fuselage, tails, tail boom, etc 130 Motor with speed controller: Hacker b20 36, 50 Propulsion Propeller & Spinner: Grapuner 4,7-4,7 12 Receiver: M5 fma 11 R/C componets Servo: GWS NARO +HP x 1, GWS NARO HP x 2 35 Data link, GPS, MAV128, ANT, Altimeter 105 Avionics Video camera, transmitter 25 Power Battery: Thunder Power Li-Polymer 2100mAh/11.1V 132 Total 500

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21 3.1.3 Airfoil Selection One of the most important choices in UAV or MAV design is the selection of an airfoil. Since the wing loading for the first generation vision-based UAV was extremely high at low speed, a more efficient wing design was required. On the MAV scale, the airfoil section necessary to produce enough lift at all mission segments must have enough camber such that it produces a high lift curve slope. This latter specification in necessary to provide an adequate lift coefficient at a reasonable angle of attack since the finite lift slope for a low aspect ratio wing is much less than for a high aspect ratio wing [13]. The airfoil geometry for the first generation vision-based test-bed was developed for the tailless airplane, Gruven. Instead of having a tail or stabilizer, tailless airplanes rely on a reflex camber at trailing edge of the wing to achieve horizontal stability. However, to achieve greater controllability of the aircraft a tail is required. A tail would provide more precision and control to the pilot or the control system during banks and turns, which is highly favorable when trying to acquire video images. Therefore the new airfoil was designed without the reflex camber to accommodate the stabilizer. The new airfoil was designed by using a modified version of the upper half of a MH30 airfoil, provided by Dr. Martin Hepperle. This airfoil is particularly used for gliders with Reynolds number in the region of 150,000, which is similar to the Reynolds number of each UAV test-bed. This is similar to the Gruven airfoil, starting at the leading edge, but the two diverge at the maximum camber. The modifications of the MH30 airfoil included and increase in camber from 5% to 7 %, and smoothing of the trailing edge to eliminate any signs of a reflex in the surface. Figure 3-1 shows the comparison between the modified upper section of the MH30 airfoil and the Gruven airfoil. Maximum camber position also moved slightly back from 20% to 25%.

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22 Figure 3-1. Airfoil shape comparison of Gruven and modified MH30 airfoil. Figure 3-2. Lift coefficient comparison between Gruven and modified MH30 airfoil.

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23 Thin airfoil theory was used to analyze both airfoils, as this was the most appropriate method available for the analysis. Figure 3-2 shows clearly that the new airfoil has a much higher lift coefficient than the 2Griven airfoil at the same angle of attack. The new model also generated 50% more lift which resulted in positive lift at zero angle of attack. Additional explanation of the thin airfoil theory can be found in Appendix A and the MATLAB code supporting the graph and calculations is included in Appendix B. 3.1.4 Wing Design The University of Florida MAV team developed software allowing a user to design a surface model of the wing geometry and create it in a CAD program. This tool speeds up the design iteration process by just entering the following wing design variables. Airfoil Shape Wing Span Root Chord Washout Angle Sweep Angle Dihedral Shape Figure 3-3 shows the output drawing of the 24 inch wing span from wing generator software using the modified upper half of the MH30 airfoil. The new wing has a 6.5 inch root chord, 2 degrees of washout angle, 2 degrees of dihedral angle, and 5 degrees of sweep angle. As can be seen in the Figure 3-3, the wing was designed as similar as possible to an elliptical shaped wing. It bears an increasing polyhedral which increases slightly from the root chord to the wing tips. Also, the wing tips of the newly designed wing do not posses an airfoil.

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24 Figure 3-3. Wing generator output using modified MH30 airfoil. The dihedral is beneficial in self righting the aircraft from a wind gust. This added stability makes the plane easier to fly because it eliminates the need for constant adjustment from the pilot to maintain level wings. Unfortunately, the dihedral also causes the plane to roll away from the direction of the skid during a side wind gust [20]. Incidentally, sweeping a wing back also gives dihedral effect with about five degrees of sweep being equivalent to one degree of dihedral [3]. Swept wing brings aerodynamic center away from the leading edge. This allows the center of gravity to be moved back as well, while still maintaining stability of the airplane. 3.1.5 Fuselage The new fuselage design reduced the overall volume by making a thinner and shorter fuselage with a more compact tail section. This new slender design also minimizes the aerodynamic drag, and provides a better holding surface for hand launching the aircraft. As can be seen from Figure 3-4, there are fewer curves in the

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25 surface of the fuselage, thus making manufacture of the aircraft more manageable. A CAD model of the fuselage for top and side view is shown in Figure 3-4. Figure 3-4. CAD drawing showing top and side view of fuselage design. The Hacker B20-36 motor is placed in the top part of trailing fuselage, and bottom part is designed to connect a tail boom. It does however use the same hatch method as in the first generation test-bed. Here, all the computer components for the vision-based control are mounted in the flat rectangular shape of a hatch. 3.1.6 Control Surface Area There are three different control surfaces on the aircraft: ailerons (wing morphing), elevator, and rudder. They are designed to change and control the moments about the roll, pitch, and yaw axes. To maximize the effect of the control surfaces, 90% of the tail span and about 20 to 50% of the tail chord are designated as the elevator and the rudder [20]. The tail of the second generation test-bed was made 10% smaller in total surface area than its predecessor. This was to account for the increase in the length of the moment arm between the center of gravity the control surface area.

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26 Wing morphing has the capability to change the wing in terms of planform, area, aspect ratio, and camber to optimize flight performance [3]. The wing of the test-bed is constructed of a flexible membrane made by carbon fiber ribs and a thin high-temperature polymer film. Hence, it is easier to adopt the morphing technology unto the wing of the aircraft. The wing morphing is controlled by a very strong servo that is mounted through the wing of the aircraft. It operates by controlling two push rods that connect to control horns, which in turn are affixed to the leading edge of the wing. They operate in a push/pull fashion, thus converting the wing into a massive aileron during flight. Figure 3-5 shows an unattached wing with a mounted morphing mechanism. Figure 3-5. Wing morphing mechanism. 3.2 Manufacturing 3.2.1 Rapid Prototyping A Lack of sufficient quantity of airplanes was one of the issues of concern during the development of the vision-based control system. Thus an improved system was needed to increase and control repeatability, and to reduce fabrication time on airframes. A rapid prototype wing generating software solved this problem. Once the software generated a geometrical shape of the desired wing, it was converted to a 3D wire-frame

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27 drawing in a CAD program. This 3D wire-frame drawing file was then translated to a tool-path for CNC milling machine. Finally, this machine created a female tooling of the wing. Figure 3-6 shows a female wing tool being milled out of high density tooling foam by a CNC milling machine. Figure 3-6. A female wing tool milled by CNC milling machine. Designing and building a fuselage mold using the CNC milling machine is quite conceivable. Unfortunately, to design complex freeform surfaces for a fuselage using a CAD program is a highly time consuming process. Working with a tool milled out from the CNC machine, it is also difficult when trying to cover all the edges during the carbon fiber lay-up process. It is trivial to construct a mold using traditional methods. It is done by designing the top and side view of a fuselage and sanding down to the desired shape. 3.2.2 Fabrication and Assembly Composite materials are used for all the airframes due to its durability and light weight. As described in Chapter 2, the wing is fabricated on a female tool with uni-directional and bi-directional carbon fiber. A thin high temperature polymer film which does not react to moisture substitutes as the wing skin. Two layers of bi-directional carbon fiber are shaped around a male fuselage mold with an extra layer being placed in

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28 the area of the hatch for reinforcement purpose. The horizontal and vertical tails are constructed on a flat plate from a 2-D drawing placed on top of the plate. They are made from multiple layers of uni-directional carbon fiber and the high temperature polymer film. Eventually, these parts are sealed in a vacuum bag and cured at 260 degrees Fahrenheit for four hours. Both the horizontal and vertical stabilizers are attached to a 0.35 inch diameter carbon tube with cyanoacrylate adhesives while the other end of this tube is connected to the fuselage. Finally, the assembly is completed by attaching the wing to the top of the fuselage with a pair of support rods as shown in Figure 3-7. Nylon screws affix the wing with the fuselage. This new wing mounting method increases the ease of assembly and packability of the second generation vision based test-bed. The entire fabrication and assembly processes, including motor and servos installation only requires two days. Figure 3-8 shows a completely assembled airplane ready to fly. Figure 3-7. Morphing mechanism for aileron and a pair of wing supports attached to the fuselage.

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29 Figure 3-8. An assembled second generation test-bed airplane for vision-based control.

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CHAPTER 4 FLIGHT TESTING AND WIND TUNNEL EXPERIMENT 4.1 Flight Testing Several flight tests were successfully completed with the first prototype of the second generation airplane. In the design process, flight testing was essential to the evaluation of flying characteristics. Throughout the remote control (RC) flight testing, designers received instant feedback from pilots and observers. This allowed for modifications to be made to the airplane before further flight tests and wind tunnel tests were conducted. The new fuselage shape places the motor further away from the gripping point of the aircraft during a hand launch. This ensures that the hand is out of the envelope of the spinning propeller and ensures safety during a hand launch. Each hand launch of the new aircraft was now done with more ease, due to the slender and more ergonomically shaped fuselage. Though the new model uses the same motor, battery, and payload, it has a higher rate of climb, a higher glide ratio, and improved turning capability. Further, the new design improved glide performance by increasing lift. However, a better cooling system was required for the motor. After flight testing, the motor was seen to overheat to the point of becoming untouchable. A bigger air intake hole with an aluminum motor mount possessing more area for airflow proved to be the answer to this need. 30

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31 4.2 Wind Tunnel Experiment 4.2.1 Wind Tunnel Setup The University of Florida Department of Mechanical and Aerospace Engineering has an on campus wind tunnel testing facility. The advanced equipment consists of a six-component, high-sensitivity sting balance that digitally measured lift, drag, and side-force loads, as well as the three moments about the balance center. The string balance is connected to an automated PC data acquisition system. The setup of the experiments on the test-bed UAV is shown in Figure 4-1. Figure 4-1. Wind tunnel testing setup with articulating balance The plane is held by a drive from an arm which is connected to a brushless servomotor operated by a single axis motion controller. A modified horizontal tail was fabricated with 5 layers of bidirectional carbon fiber and was used to aid in suspending the aircraft from the sting balance. The pitch or yaw angle can automatically be set to any time-variable angle of attack.

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32 4.2.2 Wind Tunnel Results Wind tunnel data was acquired for both the first and second generation UAV test-beds. Wind tunnel testing was performed on a full-scale version of each test-bed. This included the fuselage, the wing, and the vertical and horizontal tail. The first preliminary wind tunnel testing was performed with a zero deflection angle on all control surfaces at different angles of attack. As shown in Figure 4-2, the wind tunnel tests confirm that the second generation wing generates more lift for a given angle of attack. This correlates well with the results predicted from using 2-D thin airfoil theory as discussed in chapter 3. Cl vs. AOA-0.6-0.4-0.200.20.40.60.811.21.4-10-5051015202530Angle of Attack (deg)Lift Coefficient 35 New_24MAV V1_24MAV Figure 4-2. Comparison of lift curve from preliminary test Over a large range of angle of attack, the new version outperforms the previous version in terms of higher lift. It has a negative zero-lift angle of attack as shown in the thin airfoil theory, and is overall a more linear function of angle of attack. However, the new model has a lower stall angle of about 16 compared to the old version at 20. By

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33 placing the servo above the wing of the new aircraft, it is expected to cause the detachment of air over the wing sooner than expected. It is however inconsequential, since the neither aircraft will ever have a need to fly at an angle of attack greater than 10 during normal flight conditions. Given a known take-off weight and estimated airspeed, it is possible to calculate the lift coefficient required to sustain level flight. That is, SVWCL2req. 21 where W is the weight of the aircraft, is the air density at sea-level, V is the estimated cruise speed, and S is the wing area [5]. A reasonable estimate of the cruise speed of the UAV is about 13 m/s. This approximation is based on test flights of first version airplanes with the same motor. For a weight of 500 grams, at an approximated airspeed of 13 m/s, and sea-level conditions, the required C L is about 0.6. According to Figure 4-2, it can be predicted that new MAV needs slightly less angle of attack to accomplish the required C L to sustain level flight (6 versus 8). In addition to generating higher lift on the UAV test-bed, reduced drag is another design objective. From the flight tests, a direct measure of lift to drag ratio (L/D or Cl/Cd) was presented. The higher L/D ratio suspected during flight testing was confirmed with wind tunnel data. From the wind tunnel results, the Cl/Cd can be calculated and compared with the old model as shown in Figure 4-3. The new MAV design has a higher lift-to-drag ratio through negative and smaller angle of attack, up to a point of approximately 7. Beyond this point, however, the drag of the second generation test-bed is slightly higher than the first generation test-bed, and

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34 results in a fall in the lift to drag ratio. This can be attributed to the servo being placed atop the wing, the pushrods which protrude from the aircraft, and the connecting columns between the wing and the fuselage. Lift over Drag Ratio vs. AOA-4-2024681012-4-20246810121416Angle of Attack (deg)Cl/Cd ratio New_24MAV V1_24MAV Figure 4-3. Comparison of lift to drag ratio For longitudinal motion, pitching moment about the y-axis is the most important consideration for static stability. If an airplane were statically stable, pitching moment would tend to rotate the airplane back toward its equilibrium point from a gust of wind or disturbance that causes the angle of attack to increase or decrease. This happens because the center of gravity of the airplane would be in front of the neutral point, where the lift acts and thus restores the vehicle to the equilibrium condition. Therefore, the slope of the pitching moment coefficient with respect to the angle of attack should be negative [4]. Figure 4-4 illustrates that the second generation test-bed airplane exhibits a negative slope of the curve of pitching moment coefficient (about the center of gravity) versus angle of attack. The airplane has improved in static stability. The positive intercept is shown to be about 2.8, at the equilibrium point when the moment coefficient

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35 is zero. The slope of this curve, the pitching-moment-derivative, is approximately -0.0207 per degree (-1.18 per radian), a value which is typical of transport-type aircraft. The corresponding static margin shows a relatively high value of 0.24 (24% of mean aerodynamic chord). Using a mean aerodynamic chord of 4.4 inch, the dimensional static margin is 1 inch. Pitching Moment Coefficient Curve-0.25-0.2-0.15-0.1-0.0500.050.1-6-4-202468101214AOA (Deg)Moment Coefficient New_24MAV V1_24MAV Figure 4-4. Comparisons of pitching and yawing moment coefficient curve Rolling stability is a critical factor that was compared for both the first and second generation test-beds. Based on the second graph presented in Figure 4-5, it shows that the second generation test-bed has a more effective roll coefficient, determined by the slope of the graph. Having an effective roll coefficient translates to a lower or more constant drag on the aircraft during turns. The upper graph in Figure 4-5 also points out the increase in drag experienced by the first generation test-bed, where the drag for the second generation test-bed remained almost a constant during a roll or turn. Although the

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36 second generation test-bed shows a slightly higher drag with no control, during increased angles of attack, it does show more agility during turns and rolling maneuvers because of its low drag. Results from the wind tunnel shown in Figure 4-5 prove that the morphing mechanism is more beneficial to roll control than the split elevators used on the first generation test-bed. Rolling Moment Coefficient and Drag-0.25-0.2-0.15-0.1-0.0500.050.10.150.200.20.40.60.811.2Split elevator, Morphing % DeflectionRolling Moment Coefficient Drag Coeffcient New_24MAV: CRM New_24MAV: CD V1_24MAV: CRM V1_24MAV: CD Figure 4-5. Drag and rolling moment coefficient due to % control surface deflection. For control properties, Figure 4-6 shows the lift coefficient as a function of elevator deflection for both airplanes. Angles of attack of 1 and 10 degree are plotted for first generation test-bed; angles of attack of 0 and 10 degree are plotted for the second generation. The angle of attack does not tend to have a strong relationship with the lift

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37 curve slopes. For the second generation test-bed airplane, the elevator deflection is effectively symmetrical from -10 to 10 degree and has an average slope of 0.012 per degree (0.687 per radian). For the first generation airplane, the elevator deflection is properly working from -5 to 20 degree, and shows an average slope of 0.013 per degree (0.748 per radian). Lift Coeffcient vs. Elevator Deflection-0.4-0.200.20.40.60.811.21.4-30-20-10010203040Elevator Deflection (deg)Lift Coefficient, Cl New_24MAV: AOA=0 New_24MAV: AOA=10 V1_24MAV: AOA=1 V1_24MAV: AOA=10 Figure 4-6. Comparison of lift coefficient change respect to elevator deflection 4.2.3 3D Digital Image Correlation Setup Using Digital Image Correlation techniques, the wing deformation was measured in the wind tunnel. This method of analysis made it possible to visualize the deformations of the flexible wing during flight conditions due to morphing and aerodynamic loads. Especially, 3D Digital Image Correlation is a powerful tool used in measuring full-field 3D of an advanced deformation, stress, and strain. Digital Image Correlation is a data analysis method which uses a proprietary mathematical correlation method to analyze

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38 digital image data taken while samples undergo normal loadings. Consecutive image captures taken during the testing phase will show a change in surface characteristics as the specimen is effected by the loadings. It is commonly used for biomedical applications, aircraft fuselage or wings, rubber tires and crash testing because it is substantially more robust, easy to use, and has greater dynamic range [21]. Figure 4-7. The second generation wing for Digital Image Correlation test. As Figure 4-7 shows, half the surface area of the wing from a second generation test-bed is painted with a random black dotted pattern on a white background. This wing is then viewed in the wind tunnel by a pair of high resolution, digital CCD camera to measure the wings synchronized image of 3D coordinates and 3D deformation. When aerodynamic load is applied, deformation of the random pattern is recorded by the digital cameras, and a PC correlates the patterns to the deformation, strain, and stress. Figure 4-8 shows a pair of CCD cameras on top of the wind tunnel where they are used to measure the wing deformation and morphing at flight conditions.

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39 Figure 4-8. A pair of 3D Digital Image Correlation cameras on top of wind tunnel. 4.2.4 3D Digital Image Correlation Results Using digital image correlation techniques, we were able to prove that the flexible wing does have the ability to generate washout during aerodynamic loading. As discusses previously, having washout on a wing reduces the induced drag experienced by the wing, and in effect raises the lift to drag ratio of the aircraft. The extent of washout is determined by calculating the negative angle between the root chord and the tip chord. 0.827227 w ( mm ) 9.92573 Figure 4-9. The second generation right-wing deformation in z direction when aerodynamic load is applied

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40 Based on the results shown in Figure 4-9, the latter end of the wing experiences a concave upward deformation, but it is even greater in the trailing edge of the wing tip area. The maximum deflection of the wing tip area was 10mm in the z direction, which correlates to a negative 3.4 degree difference between the root and the tip chords of the wing. Neutral conditions during testing were zero wind speed at 10 degrees angle of attack, and the deformation testing conditions were 13 m/s wind speed at 10 degrees angle of attack. Figures 4-10 and 4-11 show the deformation of the flexible wing during right and left turns respectively using the morphing mechanism. Each figures gives a more in depth look at exactly how morphing causes changes in the wing shape during a turning maneuver. Figure 4-10 shows a simulated maximum right turn in the wind tunnel, using the morphing mechanism, the tip area of the wing experienced a large negative deformation from its neutral position. The in-board section of the trailing edge, however, remained at the neutral position during this exercise. -8.41803 w ( mm ) 5.22582 Figure 4-10. The second generation right-wing deformation when aerodynamic load is applied with wing maximum right turn morphing.

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41 -0.388694 w ( mm ) 10.4858 Figure 4-11. The second generation right-wing deformation when aerodynamic load is applied with maximum left turn morphing. Figure 4-11 shows a simulated maximum left turn in the wind tunnel, using the morphing mechanism, and the deformation experienced by the right wing of the aircraft. The associated left turn shows more of a uniform distribution in deformation over the entire wing. As expected, the wing tip area experiences the most deflection in the positive z-direction of approximately 10.5mm, where the root experiences a slight deflection in the negative z-direction of only -0.3.

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CHAPTER 5 CONCLUSION The design and manufacturing of the first generation test-bed airplane for vision-based autopilot control has been done successfully. Vision-based controlled flight, and its further developments are geared towards use on MAVs of the future. However, current commercial components are not available for MAV size aircrafts, so a bigger test-bed airplane was designed. The first generation test-bed has a wing span of 24inch with a pusher configuration by a unique geared drive system. The pusher concept allows clear view from the front to provide better visual data for the vision-based control. A thin and flexible wing also added to optimize performance while the plane is flying. Through out the numerous flight testing and wind tunnel results, issues were addressed for a better implementation of vision-based control algorithms. The five main issues are following: High wing loading. Complex drive system generated too much noise. There were an excess of unused spaced in the fuselage. New wing design with non-reflex cambered airfoil was needed. Wing morphing mechanism was required to have a more agile airplane. The second generation test-bed was similar to its predecessor in wingspan and overall length, but the process of making each airplane was significantly different. By using a CNC milling machine, the process of producing a wing was now more repeatable and with a quicker turn around time for any adjustments or new designs. By reducing the total airframe and propulsion weight by 30 grams, and increasing wing area by 10%, the 42

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43 wing loading was decreased by 15%. If the hardware weight for the vision-based system reduces, the wing loading will be reduced even more. The propulsion system was changed to a direct drive system from the complex geared one. Flight testing required a better cooling system for this powerful direct drive unit. Modification was made by making a pair of larger intake hole and by mounting the motor in a thin aluminum plate with plenty area to allow airflow. The new fuselage design reduced the overall volume by making a thinner and shorter fuselage with a more compact tail section. The new wing design also generates higher lift with new airfoil, and this was confirmed by the thin airfoil theory, wind tunnel testing, and flight testing. A wing morphing mechanism was also added into the wing as an aileron control surface. This new concept was successfully demonstrated through a number of flight tests. Design of the second generation test-bed also allows for the convenient removal and replacement of the wing from the aircraft. This comes in handy when the airplane is to be transported, as the wing can be removed from the fuselage, thus reducing the amount of space the airplane occupies while not in use. This airplane is a good test-bed for developing the vision-based control algorithm and can also be used in other applications. It does however show a slight increase in drag from its previous version. This is attributed to the servo being mounted on the wing, which affects the stall margins, and the push rods which protrude from the aircraft. Steps are being taken to reduce the effect of these drag producing areas of the test-bed by covering them with aerodynamic caps. Wing tunnel and flight tests proved to be invaluable tools in evaluating and comparing both the first and second generation UAV test-beds. Digital Image

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44 Correlation was used along with wind tunnel testing to further analyze the flexible wing concept and the benefits of adding the morphing mechanism to the second generation test-bed. Results obtained from the Digital Image Correlation confirmed the washout effect experienced by flexible wings as they deform during the application of aerodynamic loads. It also showed the wing deformation caused by morphing during simulated turns and flight conditions in the wind tunnel.

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APPENDIX A THIN AIRFOIL THEORY For a thin airfoil the distribution of a vortex sheet over the surface of the airfoil looks almost same as a vortex sheet placed on the camber line when view from a distance. If the airfoil is thin, the camber line is close to the chord. Thus, we can assume the vortex sheet appears to fall in the chord line. In other words, the airfoil can be replaced by a vortex sheet along the chord line from the camber line for an airfoil in a uniform flow, V h Assuming small angle, the velocity normal to the camber line is induced by the vortex sheet, ()()()dzUsUxVdx (1) Considering the strength of the vortex sheet varies with the distance along the chord, the velocity at point x induced by the elemental vortex at point x: dU (')'()2(') x dxdUx x x (2) 45

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46 So, the total induced velocity at the point x is given by: 101(')()2' x dxUx x x (3) Substituting Equation (3) into (1), we obtain: 101(')'()2' x dxdzV x xd x (4) It is convenient to introduce the variable cos12 x (5) and to write as a Fourier series: 011cos()2(sin)sinnnVAAn (6) where A 0 ,A 1 ,A 2 are constants to be determined in terms of the angle of attack and the slope of the camber line dz/dx. Substituting (6) into (4) and carrying out the integration using trigonometric relations: 0[cos(1)cos(1)]2sinsincos'sin'cos'cossinnxnxxnnnd x the equation becomes 01cosnndzAAndx (7) Finally, multiply both sides by cos m and integrate from 0 to 0012cosondzAddxdzAndx d (8)

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47 All aerodynamic quantities can now be calculated from Equation (8) using Equation (6). For lift and drag coefficient, 101010201122()2()(2244lLElmlVdxACAmVxdxCACAAA )A (9)

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APPENDIX B AIRFOIL DRAWING AND LIFT CURVE MATLAB CODE %%Airfoil Drawing MATLAB Code Clear all Close all %% First modification airfoil x=0:0.05:1; %% x position variables %% coefficients of the airfoil equation C1=-0.0118; C2=1.1415; C3=-3.3528; C4=4.117; C5=-2.7207; C6=0.812; C7=0.0004; %% z position variables for i=1:length(x) z1(i)= C1*x(i)^6 + C2*x(i)^5 + C3*x(i)^4 + C4*x(i)^3 + C5*x(i)^2 + C6*x(i) + C7; end z1=double(z1); %% Gruven airfoil %% coefficients of the airfoil equation 48

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49 D1=-1.1588; D2=4.8332; D3=-8.2081; D4=7.3023; D5=-3.5916; D6=0.8227; D7=0.0004; %% z position variables for i=1:length(x) z2(i)= D1*x(i)^6 + D2*x(i)^5 + D3*x(i)^4 + D4*x(i)^3 + D5*x(i)^2 + D6*x(i) + D7; end z2=double(z2); %% Second modification airfoil %% coefficients of the airfoil equation E1=-4.1632; E2=13.8648; E3=-17.7259; E4=11.6092; E5=-4.4595; E6=0.8742; E7=0.0006; for i=1:length(x) z3(i)= E1*x(i)^6 + E2*x(i)^5 + E3*x(i)^4 + E4*x(i)^3 + E5*x(i)^2 + E6*x(i) + E7;

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50 end z3=double(z3); plot(x,z2,'b',x,z3,'r') %axis([0 1 -0.1 0.1]) xlabel('\bf{X axis}','fontsize',10) ylabel('\bf{Y axis}','fontsize',10) title('\bf{Airfoil}','fontsize',12) legend('Modification','Gruven') axis equal %%Lift Curve MATLAB Code clear all close all %% First modification airfoil %% coefficients of the airfoil equation C1=-0.0118; C2=1.1415; C3=-3.3528; C4=4.117; C5=-2.7207; C6=0.812; C7=0.0004; rho = 1.225; % air density (kg/m^3)

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51 v = 13; % cruise speed (m/s) pi = 3.14; S = 0.079; % planform area of wing syms theta; x=(1-cos(theta))/2; %%Equation of the Airfoil z = C1*x^6 + C2*x^5 + C3*x^4 + C4*x^3 + C5*x^2 + C6*x + C7; %%Derivative of z Respect to x dzdx = 6*C1*x^5 + 5*C2*x^4 + 4*C3*x^3 + 3*C4*x^2 + 2*C5*x + C6; cf = vpa(dzdx*(cos(theta)-1),5); alpzero= vpa(-1/pi*int(cf,0,pi),5) % zero lift angle alpha=-10*pi/180:1*pi/180:10*pi/180 % angle of attack %%Lift Coefficients for i=1:length(alpha) Cl(i) = vpa(2*pi*(alpha(i) alpzero),5); end Cl=double(Cl); %% Second Modification Airfoil %% Coefficients of the Airfoil Equation D1=-1.1588; D2=4.8332; D3=-8.2081; D4=7.3023;

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52 D5=-3.5916; D6=0.8227; D7=0.0004; %% Equation of the Airfoil z2 = D1*x^6 + D2*x^5 + D3*x^4 + D4*x^3 + D5*x^2 + D6*x + D7; %% Derivative of z Respect to x dzdx2 = 6*D1*x^5 + 5*D2*x^4 + 4*D3*x^3 + 3*D4*x^2 + 2*D5*x + D6; cf2 = vpa(dzdx2*(cos(theta)-1),5); alpzero2= vpa(-1/pi*int(cf2,0,pi),5); % zero lift angle %% Lift Coefficient for i=1:length(alpha) Cl2(i) = vpa(2*pi*(alpha(i) alpzero2),5); end Cl2=double(Cl2); %% Griven Airfoil %% Coefficients of the Airfoil Equation E1=-4.1632; E2=13.8648; E3=-17.7259; E4=11.6092; E5=-4.4595; E6=0.8742; E7=0.0006;

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53 %% Equation of the Airfoil z3 = E1*x^6 + E2*x^5 + E3*x^4 + E4*x^3 + E5*x^2 + E6*x + E7; % Derivative of z Respect to x dzdx3 = 6*E1*x^5 + 5*E2*x^4 + 4*E3*x^3 + 3*E4*x^2 + 2*E5*x + E6; cf3 = vpa(dzdx3*(cos(theta)-1),5); alpzero3= vpa(-1/pi*int(cf3,0,pi),5); % zero lift angle %% Lift Coefficient for i=1:length(alpha) Cl3(i) = vpa(2*pi*(alpha(i) alpzero3),5); end Cl3=double(Cl3); %plot(alpha*180/pi,Cl,'r',alpha*180/pi,Cl2,'b',alpha*180/pi,Cl3,'g') plot(alpha*180/pi,Cl2,'b',alpha*180/pi,Cl3,'g') xlabel('\bf{Angle of Attack, \alpha (deg)}','fontsize',10) ylabel('\bf{Lift Coefficeint, Cl}','fontsize',10) title('\bf{Thin airfoil theory lift curve at 13m/s}','fontsize',12) legend('Modification1','Gruven',4) %legend('Modification 1','MH30','Gruven',4) grid on hold

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LIST OF REFERENCES 1. M. Abdulrahim, R. Albertani, P. Barnswell, F. Boria, D. Claxton, J. Clifton, J. Cocquyt, K. Lee, S Mitryk, and P. Ifju, Design of the University of Florida Surveillance and Endurance Micro Air Vehicles, 7th Annual Micro Air Vehicle Competition Entry, April 2003. 2. R. Albertani, P. Barnswell, F. Boria, D. Claxton, J. Clifton, J. Cocquyt, A. Crespo, C. Francis, P. Ifju, B. Johnson, S. Jung, K. lee, and M. Morton, University of Florida Biologically Inspired Micro Air Vehicel, 8th Annual Micro Air Vehicle Competition Entry, April 2004. 3. J. D. Anderson, Aircraft Performance and Design, WCB/McGraw-Hill, Boston, 1998. 4. E. Bernard and R. L. Duff, Dynamics of Flight: Stability and Control, John Wiley & Sons, INC., New York, 1996. 5. M. Burgart, J. Miller, and L. Murphy, Design of a Micro Air Vehicle for the 2000 MAV Competition, internal progress report, University of Notre Dame, 2000. 6. M. Drela (MIT Aero & Astro), XFOIL v6.91 http://raphael.mit.edu/xfoil/ 20 July 2004. 7. S. Ettinger, M. C. Nechyba, P. G. Ifju and M. Waszak, Vision-Guided Flight Stability and Control for Micro Air Vehicles, Proc. IEEE Int. Conf. on Intelligent Robots and Systems, vol. 3, pp. 2134-40, August 2002. 8. J. M. Grasmeyer and M. T. Keennon, Development of the Black Widow Micro Air Vehicle, AIAA APATC, AIAA Paper 2001-0127, January 2001. 9. P. G. Ifju, S. Ettinger, D. A. Jenkins, Y. Lian,W. Shyy and M. R.Waszak, Flexible-Wing-Based Micro Air Vehicles, 40th AIAA Aerospace Sciences Meeting, Reno, NV, AIAA 2002-0705, January 2002. 10. P. G. Ifju, S. Ettinger, D. A. Jenkins and L. Martinez, Compostie Materials for Micro Air Vehicles, SAMPE Journal, vol. 37 No. 4, pp. 7-13, July/August 2001. 11. S. M. Kanowitz, Desgin and Implementation of A GPS-based Navigation System for Micro Air Vehicles, M.S. Thesis, Electrical and Computer Engineering, University of Florida, August 2002. 54

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55 12. A. Lennon, The Basics of R/C Model Aircraft Design, Air Age Inc., Ridgefield, 1999. 13. P. B. Lissaman, Low-Reynolds-Number Airfoils, Annual Review of Fluid Mechanics, Vol. 15, pp. 223-239, January 1983. 14. J. M. McMichael and Col. M. S. Francis, Micro Air Vehicles--Toward a New Dimension in Flight, http://www.darpa.mil/tto/mav/mav auvsi.html, December 1997. 15. T. J. Muller, Low Reynolds Number Aerodynamics of Low Aspect Ratio, Thin/Flat/Cambered-Plate Wings, Journal of Aircraft, vol. 37 No. 5, pp. 825-832, August 2000. 16. R.C. Nelson, Flight Stability and Automatic Control, WCB/McGraw-Hill, Boston, 1998. 17. N. Newcome, News Room UAV Forum 2001. SRA International, Inc. http://www.uavforum.com/library/news.htm, 4 October 2003. 18. A. Parsch, AeroVironment FQM-151 Pointer Directory of U.S. Military Rockets and Missiles. 2004. Designation-Systems.Net. http://www.designation-systems.net/dusrm/m-151.html 24 March 2004. 19. J. Pike, Dragon Eye Intelligence Resources. 2000. GlobalSecurity.org. http://www.globalsecurity.org/intell/systems/dragon-eye.htm 21 December 2003. 20. D. P. Raymer, Aircraft Design: A Conceptual Approach, AIAA Education Series third edition, AIAA, Inc., Reston, 1999. 21. C. Sutton, S. Helm, P. McNeil Improved 3-D Image Correlation for Surface Displacement Measurement, Experimental Techniques, vol. 26 No. 5, pp. 39-42, May/June 2003. 22. S. Waddington, Commercial and Civil M issions for Public Service Agencies: Are UAVs a Viable Option? Unmanned Vehicles Magazine Business Analysis Forecast Dec. 2002. UAV World. Brass Trading Ltd. http://www.uavworld.com/civil.htm 29 April 2004. 23. M. R. Waszak, L. N. Jenkins, and P.G. Ifju, Stability and Control Properties of an Aeroelastic Fixed Wing Micro Aerial Vehicle, AIAA Atmospheric Flight Mechanics Conference, Montreal, Canada, AIAA 2001-4005, 6-9, August 2001.

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BIOGRAPHICAL SKETCH Sewoong Jung was born in Seoul, Korea, on Jan 18, 1980. He came to the U.S. to attend the White Mountain School in Bethlehem, NH, in 1995. After he received his high school diploma with honor, he attended the Boston University. He received his bachelors degree in aerospace engineering in May 2002. He has worked in the Micro Air Vehicle lab under Dr. Peter Ifju and will receive his Master of Science in aerospace engineering in August 2004. 56