Report; Design of Morphing Wing Micro Air Vehile 'RAA'

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DESIGN OF MORPHING WING MICRO AIR

VEHICLE

A PROJECT REPORT

DEGREE

Submitted by

RAHEEL AHMED, ANAS BIN AQEEL, AHMED ZIA

BACHELORS

IN

MECHANICAL ENGINEERING

Year

2010

PROJECT SUPERVISOR

Dr. MUHAMMAD AFZAL MALIK

COLLEGE OF ELECTRICAL & MECHANICAL ENGINEERING,

PESHAWAR ROAD, RAWALPINDI, PAKISTAN.

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DECLARATION

We hereby declare that no portion of the work referred to in this Project Thesis has been

submitted in support of an application for another degree or qualification of this of any other

university or other institute of learning. If any act of plagiarism found, we are fully responsible

for every disciplinary action taken against us depending upon the seriousness of the proven

offence, even the cancellation of our degree.

COPYRIGHT STATEMENT

Copyright in text of this thesis rests with the student author. Copies (by any process)

either in full, or of extracts, may be made only in accordance with instructions given

by the author and lodged in the Library of NUST College of E&ME. Details may be

obtained by the Librarian. This page must form part of any such copies made. Further

copies (by any process) of copies made in accordance with such instructions may not

be made without the permission (in writing) of the author.

The ownership of any intellectual property rights which may be described in this

thesis is vested in NUST College of E&ME, subject to any prior agreement to the

contrary, and may not be made available for use by third parties without the written

permission of the College of E&ME, which will prescribe the terms and conditions of

any such agreement.

Further information on the conditions under which disclosures and exploitation may

take place is available from the Library of NUST College of E&ME, Rawalpindi.

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ACKNOWLEDGMENTS

Alhamdulillah! We would like to thank Dr. M. Afzaal Malik for his infinite patience, guidance and support

without which this project would never have been completed. Also, our Parents and the rest

of family who have always supported us in everything we have chosen to partake. And, the

bird Eagle which initially inspired us for morphing wing concept.

ABSTRACT

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ABSTRACT

The project objective is to design a micro air vehicle (MAV) having morphing wings, named

as ‘RAA’. Its wings are capable of changing its wing shape during flight to enhance its

performance as compared to conventional fixed-wing aircrafts. This Project is titled as

‘Project RAA’.

A Bio-inspired Micro Air Vehicle of 50cm wingspan with Morphing Wing has been

designed. The morphing wing can change its shape from straight to swept back during

different flight conditions. Sweeping wing equips MAV with improved maneuverability,

agility and endurance. The wings are composed of three feathers which overlap each other

during sweeping. The design is divided into three main modules; Aerodynamic design,

Morphing Mechanism and Control System. Aerodynamic design has been done using

Raymer's approach and CFD analysis. Bio-inspired airfoil shapes have been selected for

‘RAA’. A general computer program of Wing design & Study has been developed using

Engineering Equation Solver (EES) Software. The analyses have shown that drag reduces

and lift-to-drag ratio increases with the sweeping angle during flight. Four mechanisms have

been conceived for the morphing of the wings. After design studies, the best mechanism has

been selected. As wing is composed is composed of three feathers, the efficient morphing

mechanism is able to actuate all feathers only with one feather powered for sweeping.

Working Model has been used for mechanism analysis. Theoretical design study of Control

System has been accomplished to deploy the MAV-RAA for autonomous flight mission.

MEMS sensors have been proposed for this purpose. The MAV has been fabricated with the

available local resources keeping gross weight minimum. Seven test flights have been done

with each one demanding improvements to the MAV.

TABLE OF CONTENTS

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TABLE OF CONTENTS

DECLARATION .......................................................................................................................... 2

ACKNOWLEDGMENTS ........................................................................................................... 3

ABSTRACT................................................................................................................................... 4

TABLE OF CONTENTS ............................................................................................................ 5

LIST OF FIGURES ..................................................................................................................... 9

LIST OF SYMBOLS ................................................................................................................. 12

CHAPTER 1

LITERATURE REVIEW ............................................................................................................ 14

1.1 Aviation History ................................................................................................................. 14

1.2 Current Aircraft Technology ............................................................................................. 17

1.3 Unmanned Air Vehicles, (UAVs) ..................................................................................... 17

1.4 MICRO AIR VEHICLES (MAVs) ................................................................................... 18

1.4.1 Aerodynamics Involved In A MAV: .................................................................... 20

1.4.2 Types of MAV ....................................................................................................... 22

1.5 Morphing Wing; Sweeping in detail ................................................................................. 27

1.5.1 Lentink’s Work: ........................................................................................................... 28

1.5.2 Videler’s Work: ........................................................................................................... 31

1.5.3 Tucker’s Work: ............................................................................................................ 31

1.5.4 Pennycuick’s Work: .................................................................................................... 32

1.5.5 Morphing Wing (by sweeping here) Advantages. ..................................................... 33

1.6 Morphing Standards ...................................................................................................... 33

CHAPTER 2

PROJECT ‘RAA’ ......................................................................................................................... 34

2.1 Project Objective: ............................................................................................................... 34

2.2 Mission ................................................................................................................................ 34

2.3 Design Methodology: ......................................................................................................... 34

2.4 Management: ...................................................................................................................... 35

2.5 Time Management: ............................................................................................................ 35

CHAPTER 3

AERODYNAMIC DESIGN ........................................................................................................ 36

TABLE OF CONTENTS

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3.1 Airfoil Selection ................................................................................................................. 36

3.1.1 Arm-Wing Airfoil: ...................................................................................................... 38

3.1.2 Hand-Wing Airfoil: ..................................................................................................... 45

3.2 Wing Design and Study: .................................................................................................... 46

3.2.1 Aerodynamic Coefficients: ......................................................................................... 46

3.2.2 Lift Calculation: ........................................................................................................... 46

3.2.3 Drag Calculation:......................................................................................................... 49

3.3 Wing Design & Study Program ......................................................................................... 51

3.3.1 Morphing Characteristics ...................................................................................... 55

3.4 Computational Fluid Dynamics (CFD) Analysis: ............................................................ 58

3.4.1 CAD Models: .............................................................................................................. 58

3.4.2 Meshed Geometries: .................................................................................................... 59

3.4.4 Results .......................................................................................................................... 62

3.5 Conclusion: ......................................................................................................................... 62

CHAPTER 4

RAA MORPHING MECHANISM ............................................................................................. 63

4.1 Mechanisms Proposed .................................................................................................. 63

4.1.1 Cam Lobe Mechanism .......................................................................................... 64

4.1.2 Gear Actuated Mechanism .................................................................................... 67

4.1.3 Single Link Actuated Mechanism ........................................................................ 69

4.1.4 Tertiary Pin joint Mechanism ............................................................................... 70

4.2 Results Achieved ........................................................................................................... 80

4.3 Standards Achieved ....................................................................................................... 81

CHAPTER 5

CONTROL SYSTEMS OF RAA................................................................................................ 82

5.1 SENSORS EMPLOYED ON RAA: ................................................................................. 82

5.1.1 GYROSCOPES: .......................................................................................................... 83

5.1.2 GYROSCOPE SUITABLE FOR OUR RAA: .......................................................... 96

5.2 ACCELOROMETERS: ..................................................................................................... 96

5.2.1 What is Acceleration? ................................................................................................. 97

5.2.2 What are some “g” reference points? ......................................................................... 97

TABLE OF CONTENTS

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5.2.3 What are accelerometers useful for? .......................................................................... 97

5.2.4 How do accelerometers work? ................................................................................... 98

5.2.5 What things should One consider when buying an accelerometer? ......................... 98

5.2.6 Why MEMS Based? .................................................................................................. 100

5.3 MAGNETOMETER: ....................................................................................................... 103

5.4 Inertial Navigation System (INS): ................................................................................... 104

CHAPTER 6

THE FABRICATION PHASE .................................................................................................. 106

6.1 The Right Material: .......................................................................................................... 106

6.1.1 Balsa Wood: ......................................................................................................... 106

6.1.2 Jumbo lone Foam: ............................................................................................... 108

6.2.1 The Right Motor: ....................................................................................................... 110

6.2.2 The Right Propeller: .................................................................................................. 111

6.2.3 The Right Battery: ..................................................................................................... 112

6.3 The Ground Control Station: ........................................................................................... 113

6.4 The Actuators: .................................................................................................................. 114

6.4.1Metal Geared Servo Motors:...................................................................................... 115

6.5 Total Weight of our RAA: ............................................................................................... 115

6.6 Flight Tests: ...................................................................................................................... 115

6.7 The Final Fabricated RAA:.............................................................................................. 116

CHAPETR 7

PROJECT LIMITATIONS ........................................................................................................ 119

CHAPTER 8

FUTURE RECOMMENDATIONS OF ‘PROJECT RAA’ .................................................... 120

CONCLUSION ......................................................................................................................... 121

REFERNECES ......................................................................................................................... 122

APPENDEX A .......................................................................................................................... 126

RAA WING DESIGN AND STUDY, EES PROGRAM CODE ........................................... 126

LIST OF TABLES

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LIST OF TABLES

Table 2.1: Management Table ................................................................................................................... 35 Table 3.1: Summary of Airfoils under consideration............................................................................... 44 Table 3.2: RAA Wing parametric Study................................................................................................... 54 Table 3.3: CFD Analysis Results .............................................................................................................. 62 Table 4.1: RAA Gear Parametric Study.................................................................................................... 67 Table 4.2: RAA Feather Parametric Study ............................................................................................... 74 Table 4.3: RAA Feather Parametric Stud ................................................................................................. 80 Table 4.4: RAA Standards Achieved Parametric Study .......................................................................... 81 Table 6.1: Jumbo lone foam Technical Properties .................................................................................109

LIST OF FIGURES

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LIST OF FIGURES

Figure 1.1: Abbas Ibn Firnas flying from the Mosque of Cordoba in 875 AD.[2] ................................. 14 Figure 1.2: Different gliders designed by Cayley. [3] ............................................................................... 15 Figure 1.3: Wright Flyer ............................................................................................................................ 16 Figure 1.4: Boeing 747[4] ........................................................................................................................... 17 Figure 1.5: Air vehicles with their operating Reynolds No. and Mass [5] .............................................. 20 Figure 1.6: (a) ‘Delfly’ Flapping Wing MAV [8], (b) ‘MAVSTAR’ Rotary Wing MAV[9] ................. 23 Figure 1.7: RoboSwift of TU Delft, Netherlands[12] ................................................................................ 26 Figure 1.8: UF MAV[9] .............................................................................................................................. 27 Figure 1.8: Swift sweeping during flight[16] ............................................................................................. 28 Figure 1.9: Swift’ sweeping aerodynamics [17] ......................................................................................... 29 Figure 1.10: Swift’s sweeping wing aerodynamic characteristics [17] .................................................... 30 Figure 1.11: Swift’s Leading Edge Vortex [19] .......................................................................................... 31 Figure 1.12: Mean wing area of the falcon in the wind tunnel at different speeds and glide angles. Wing and tail shape of the falcon at various speeds and glide angles. A, 6-6 m./sec., 6°; B, 8-5 m./sec, 5°; C, 14-3 m./sec., 6°. [20] ........................................................................................................................ 32 Figure 1.13: Bird power consumption curves [21].................................................................................... 33 Figure: 2.1: Mission profile for ‘RAA’ .................................................................................................... 34 Figure 2.2: Gantt Chart of Project RAA ................................................................................................... 35 Figure 3.1: Airfoil shapes of the wing of the Swift bird. [18] ................................................................... 36 Figure 18Figure 3.2: Tucker provided the airfoil shape of Eagle and the data airfoils similar to that of bird. [20] ........................................................................................................................................................ 37 Figure 19Figure 3.3: Eiffel 13 data, at Reynolds Number of 139664, Max Cl = 0.55 Min Cd=.015 and Max L/D = 13 ...................................................................................................................................... 39 Figure 3.4: NPL 4a data at Reynolds Number of 24536, Max Cl = 0.6, Min Cd = .025 and max L/D=13 ........................................................................................................................................................ 39 Figure 3.5: GOE 464 airfoil shape[24] ....................................................................................................... 40 Figure 3.6: GOE 464 Characteristic Curves[24] ........................................................................................ 41 Figure 3.7: Eppler 58 airfoil shape[24] ....................................................................................................... 42 Figure 3.8: Eppler 58 Characteristics Curves[24] ...................................................................................... 43 Figure 3.9: GOE 417a at Re=419000, Max Cl = 1.3, Min Cd=.04, Max L/D=14 ................................ 45 Figure 3.10: Aspect Ratio Effect Upon Lift Curve Slope [25] ................................................................. 46 Figure 3.11: EES program Equation Window ......................................................................................... 51 Figure 3.12: EES Program Formatted Equations Window ..................................................................... 52 Figure 3.13: EES Program Solution Window .......................................................................................... 52 Figure 3.14 : RAA Wing Design and Study EES Program user Interface ............................................. 53 Figure 3.15: Drag Vs Sweeping Angle Plot ............................................................................................. 55 Figure 3.16: Lift Vs Sweeping Angle Plot ............................................................................................... 56 Figure 3.17: Lift-to-Drag ratio Vs Sweeping Angle Plot ........................................................................ 57 Figure 3.18: CFD Analysis Process .......................................................................................................... 58

LIST OF FIGURES

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Figure 3.19: CAD Models of Straight RAA wing and 45o Swept Wing ................................................ 58 Figure 3.20: Meshed Geometries of Straight Wing in Gambit. .............................................................. 59 Figure 3.21: Meshed geometries of 45o Swept wing in Gambit .............................................................. 60 Figure 3.22: Fluent Solution, (a) Velocity magnitude contour plane at Straight Wing. (b) Velocity magnitude contour plane at 45o Swept wing ............................................................................................. 61 Figure 4.1: Pro Engineering Cam Lobe Mechanism ............................................................................... 64 Figure 4.2: Pro Engineering Initial Position Model of Cam Lobe Mechanism ..................................... 65 Figure 4.3: Pro Engineering Final Position Model of Cam Lobe Mechanism....................................... 65 Figure 4.4: Final Fabricated Part of Cam Lobe Mechanism ................................................................... 66 Figure 4.5: Gear Formulas [27] .................................................................................................................. 67 Figure 4.6: Spur Gear and Path of Action Diagram [27] ........................................................................... 68 Figure 4.7: Gear Actuated Train [28] .......................................................................................................... 68 Figure 4.8: Link Actuated Mechanism ..................................................................................................... 70 Figure 4.9: E Quintet Mechanism ............................................................................................................. 71 Figure 4.10: Pro Engineering Wing Design ............................................................................................. 71 Figure 4.11: Pro Engineering Feather Design .......................................................................................... 72 Figure 4.12: Polygon Command in Design bar........................................................................................ 72 Figure 4.13: Properties of Feather 1,2 and 3 ............................................................................................ 73 Figure 4.14: Geometrical Arrangement of Mechanism ........................................................................... 73 Figure 4.15: Motor Placement on Feather 1 ............................................................................................. 74 Figure 4.16: Motor Properties Representation ......................................................................................... 75 Figure 4.17: Rotation, Velocity and Acceleration Graphs of Feather 1 ................................................. 76 Figure 4.18: Rotation, Velocity and Acceleration Graphs of Feather 2 ................................................. 77 Figure 4.19: Rotation, Velocity and Acceleration Graphs of Feather 3 ................................................. 78 Figure 4.20: Rotation Graph of Motor and Time Taken ......................................................................... 79 Figure 4.21: Initial Position of Wing Geometry ...................................................................................... 79 Figure 4.22: Final Position of Wing Geometry........................................................................................ 80 Figure 5.1: Working of Gimbaled Gyroscope .......................................................................................... 85 Figure 5.2: Gimbaled Gyroscope Functioning ......................................................................................... 87 Figure 5.3: Gimbaled Gyroscope, tilting plane......................................................................................... 88 Figure 5.5: Cross Section of Gyroscope [33] .............................................................................................. 91 Figure 5.6: Piezoelectric Plate with orthogonal directions shown .......................................................... 91 Figure 5.7: General Ring Laser Gyro ........................................................................................................ 92 Figure 5.8: Beat Frequency vs. Scatter Position ....................................................................................... 93 Figure 5.9: Rotating Rate found by using Beat Frequency ...................................................................... 94 Figure 5.10 3 axis accelerometer board .................................................................................................... 98 Figure 3.12: (a) Design illustration; (b) physical dimensions and; (c) SEM image of planar piezoresistive accelerometer[38] ................................................................................................................102 Figure 5.12: Piezoelectric Accelerometer is designed for flight vehicles. [40] ......................................103 Figure 5.13: Magnetometer ......................................................................................................................104 Figure 5.14: INS .......................................................................................................................................105 Figure 6.1: (a), (b) Fuselage of Balsa Wood ...........................................................................................107

LIST OF FIGURES

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Figure 6.2: Jumbo lone Foam ..................................................................................................................109 Figure 6.3: RAA wing, Arm-wing of jumbo lone foam.........................................................................109 Figure 6.4: Brushless motor BL1812 ......................................................................................................111 Figure 6.5 The propeller used in RAA ....................................................................................................111 Figure 6.7: LiPo Battery ...........................................................................................................................113 Figure 6.8: WFLY Remote Control used for RAA ................................................................................114 Figure 6.9: Plastic gear mini-Servo motor ..............................................................................................114 Figure 6.10: Metal Gear micro-Servo motors .........................................................................................115 Figure 6.11: (a)Fabricated RAA, top view, (b) Fabricated RAA side view, (c) Fabricated RAA side view, (d) RAA in fields for testing, (e) RAA in fields for testing, one wing swept.............................118

LIST OF SYMBOLS

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LIST OF SYMBOLS

Latin Symbols

CL , Coefficient of Lift of Wing

CD , Coefficient of drag of wing

Cl , Airfoil Coefficient of Lift

Cd , Airfoil Coefficient of Drag

CL,max , Maximum Coefficient of Lift of Wing

Cl,max , Maximum Airfoil Coefficient of Lift

L, Lift

D, Drag

V, Velocity

q, Dynamic Pressure

S, Wing Planform Area

CL,α,rad , Lift curve Slope in per radians for Hand-wing

CL,α , Lift curve slope in per degrees for Hand-wing

CL,arm,alpha , Lift curve slope in per degrees for arm-wing

Cl,arm,α , Lift curve slope in per degrees of airfoil for arm-wing

Aarm , Aspect ratio of arm-wing

earm, oswalt efficiency factor of arm-wing

Sref , Wing reference area

barm , Arm wing span

bhand , Hand wing span

c, wing chord

as , Wing Sweep angle

LIST OF SYMBOLS

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Sexp, Wing exposed area

A, Aspect ratio

M, Mach Number

binitial, total wing span input

Swet, Wetted area of wing

CDo, Parasite drag coefficient of wing

Cf, Skin friction coefficient

FF, Component form factor

Q, Component interference factor

df, fuselage diameter

le, fuselage length

R, Reynolds number

g, Acceleration due to gravity

AOA, Angle of Attack

Greek Symbols

α, Angle of Attack

ρ, Density

λ, wing taper ratio

CH 01. LITERATURE REVIEW

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CHAPTER 1

LITERATURE REVIEW

1.1 Aviation History

Have they not looked at the birds above them, with wings outspread and folded back?

Nothing holds them up but the All-Merciful. He sees all things. (Quran; Surat al-

Mulk:19)[1]

The flight of birds has always challenged human curiosity with the dream of human flight

described already in the Greek myth about the inventor and master craftsman Deadalus, who

built wings for himself and his son Icarus to escape from imprisonment in the Labyrinth of

Knossos on the island of Crete. The leading scholar Abbas Ibn Firnas of the Islamic culture

in Cordoba in Spain studied the mechanics of flight and in 875 AD survived one successful

flight on a pair of wings made of feathers on a wooden frame. Some hundred years later the

great Turkish scholar Al-Djawhari tied two pieces of wood to his arms and climbed the roof

of a tall mosque in Nisabur, Arabia, and announced to a large crowd:

O People! No one has made this discovery before. Now I will fly before your very eyes. The

most important thing on Earth is to fly to the skies. That I will do now. [2]

Figure 1.1: Abbas Ibn Firnas flying from the Mosque of Cordoba in 875 AD.[2]


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Unfortunately, he fell straight to the ground and was killed. It would take 900 years before

the dream of Al-Djawhari became true, after many unsuccessful attempts. One of the more

successful was made by Hezarfen Ahmet Celebi, who in 1638 inspired by work of Leonardo

da Vinci, after nine experimental attempts and careful studies of eagles in flight, took off

from the 183 foot tall Galata Tower near the Bosphorus in Istanbul and successfully landed

on the other side of the Bosphorus. The word Hezarfenmeans expert in 1000 sciences and a

reward of 1000 gold pieces was given to Hezarfen for his achievement.

The understanding of why it is possible to fly has haunted scientists since the birth of

mathematical mechanics in the 17th century. To fly, an upward force on the wing, referred to

as lift L, has to be generated from the flow of air around the wing, while the air resistance to

motion or drag D, is not too big. The mystery is how a sufficiently large ratio L/D can be

created. By elementary Newtonian mechanics, lift must be accompanied by downwash with

the wing redirecting air downwards. The enigma of flight is the mechanism generating

substantial downwash under small drag, which is also the enigma of sailing against the wind

with both sail and keel acting like wings creating lift.

Figure 1.2: Different gliders designed by Cayley. [3]


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The German engineer Otto Lilienthal (1848-1896) expanded Wenham’s work, made careful

studies of the gliding flight of birds recorded in Birdflight as the Basis of Aviation and

designed a series of ever-better hang gliders allowing him to make 2000 successful heavier-

than-air gliding flights starting from a little artificial hill, before in 1896 he broke his neck

falling to the ground after having stalled at 15 meters altitude. Lilienthal rigorously

documented his work, including photographs, and for this reason is one of the best known of

the early pioneers. The first sustained powered heavier-than-air flights were performed by the

two brothers Orwille and Wilbur Wright, who on the windy dunes of Kill Devils Hills at

Kitty Hawk, North Carolina, on December 17 in 1903, managed to get their 400 kg airplane

Flyer off ground into sustained flight using a 12 horse power engine. The modern era of

aviation had started. [0]

Figure 1.3: Wright Flyer [0]


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1.2 Current Aircraft Technology

Modern Aircrafts rely on the design of Fixed Wing aircrafts. A fixed-wing aircraft, typically

called an airplane, aero plane or plane, is an aircraft capable of flight using forward motion

that causes air to pass over its wings to generate lift. Planes include jet engine and propeller

driven vehicles propelled forward by thrust, as well as unpowered aircraft (such as gliders).

Fixed-wing aircraft are distinct from ornithopters in which lift is generated by blades and

rotary-wing aircraft in which wings move relative to the aircraft. Most fixed-wing aircraft are

flown by a pilot on-board the aircraft, but some are designed to be remotely or computer

controlled.

Figure 1.4: Boeing 747[4]

1.3 Unmanned Air Vehicles, (UAVs)

There has been a recent interest by the military to have platforms capable of operating close

to a point of interest without being detected while providing critical surveillance. By

providing information that is not readily available, these platforms could provide a useful


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tool for small unit commanders in potentially life-threatening situations. Highly

maneuverable, slow-flying micro air vehicles could fly under canopies, through alleys, or

indoors to provide such intelligence.

Historically, most airborne missions have been performed by large, expensive, high-

performance piloted aircraft. More recently, Unmanned Aerial Vehicles (UAVs) have taken

over numerous battlefield observation and reconnaissance missions. The use of UAVs has

become very favorable for numerous reasons. First of all, the UAVs are autonomous, or

remotely piloted, thus removing the human from the battlefield and the possibility of human

casualty. Many UAVs are much smaller than conventional piloted military aircraft. Their

reduced size has numerous benefits. Secondly, UAVs are stealthier than their manned

counterparts, using the same technology and radar cross section (RCS) reduction methods

significantly reducing their chance of being detected by the enemy. In addition, the smaller

aircraft also require less logistical support than full size piloted jets. Today’s piloted military

aircraft are equipped with numerous sensor packages, life support systems, and equipment

for pilot interface, which increase weight and cost. Most UAVs can easily be configured for

specific missions and carry only the hardware necessary for completing the mission. UAVs

can be made lighter and cheaper because they do not need to carry life support systems,

ejection seats, video screens, pilot controls, and many other human-interface hardware

components. The performance of conventional tactical aircraft is limited by human

physiology. UAVs can sustain higher g-forces and g-loading without a pilot onboard. By

making the UAV smaller and mission specific, it is a much less expensive liability should a

casualty occur. UAV operators require less training than pilots of conventional aircraft.

Additionally, as the autonomous capabilities of UAVs are increased, the training

requirements of UAV operators will decrease.

1.4 Micro Air Vehicles (MAVs) As the operational demands for UAVs expand, the trend is towards smaller, smarter, and

less-expensive vehicles than those being used today. These new missions demand a new class

of UAVs called Micro Aerial Vehicles (MAVs). The challenge for today’s engineer is to

increase the useful payload and autonomous abilities of these vehicles while reducing cost,

reducing detectability, reducing the amount of training required for the operator, reducing the


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platform size, and reducing ground support requirements. In addition, the MAV must be

robust enough to be brought into and used on the battlefield.

In an attempt to determine how the feasibility and practicality of UAVs is affected by size,

DARPA has funded numerous Micro Air Vehicle (MAV) projects. DARPA’s current

specifications for a MAV include a six-inch package size and a weight of less than four

ounces. Additional specifications would vary depending on the various reconnaissance

missions such a platform would be expected to perform. For example, a hypothetical mission

may require the MAV to fly 1km to a point of interest and loiter within 100ft for 30 minutes

before returning. Depending on the mission, there is an expressed desire for the aircraft 3 to

be stable in winds up to 25mph, perform tight turns around buildings in urban environments,

climb repeatedly to altitudes in the range of 350ft, require minimal pilot training, and have a

low cost, possibly less than $250 to duplicate. For such engineering feats to be accomplished,

advances must be made in all aspects of vehicle design including propulsion, power,

aerodynamics, materials, and electronics.

Low speed MAVs, equipped with cameras, could fly inside buildings or under canopies

giving soldiers a new edge in urban warfare and reconnaissance. Although the initial

missions for MAVs will be for military observation and reconnaissance, the possibilities are

endless. The technology will more than likely expand from the military to other government

organizations. For instance, FBI and Police SWAT teams could use MAVs to evaluate

terrorist or hostage situations. Fire and rescue units could benefit by sending MAVs into

buildings to assess threat or search for survivors before entering themselves. Also, MAVs

could be equipped with sensors to sample the environment for chemical, biological, or

radiation levels. MAV technology could also be used widely for farmers interrogating the

ammonia levels in fields, by the EPA for measuring emissions in industrial smokestacks,

monitoring concentrations of chemical spills, or by the forestry and wildlife services to track

herds of endangered species. MAV sized models are already commercially available for

radio-control hobbyists and as toys.


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1.4.1 Aerodynamics Involved In A MAV: The Reynolds number, an indication of the ratio between inertial and viscous effects, of

MAV flight is very low. Most aircraft fly at Reynolds numbers in the millions, almost 100

million for the Boeing 747.

Figure 1.5: Air vehicles with their operating Reynolds No. and Mass [5]

Small radio controlled aircraft and larger birds of prey fly at Reynolds numbers between

70,000 and 200,000. In this regime, dependant on the particular airfoil, the separation

bubbles may begin to present problems below a certain Reynolds number. MAVs typically

fly at very low Reynolds numbers in the range of 20,000 to 120,000. In addition to their

small size and flight speeds, MAVs are generally required to be as compact as possible. This

design constraint normally leads to very low aspect ratio wings. Most past experimentation

has been done using 2D modeling of infinite span wings. This research is very credible for

large aspect ratio wings; however, the aerodynamics of low aspect ratio wings is very

different from the aerodynamics of high aspect ratio wings. Volumes of data have been

collected for thousands of airfoils for use at high aspect ratio and high Reynolds number

flight. Low aspect ratio wings and airfoil shapes have primarily been ignored, especially at

low Reynolds numbers, until recent years with the increased interest in MAV flight.

Researchers at Notre Dame conducted a program between 1998 and 2000 that consisted of

extensive wind tunnel testing of wings with aspect ratios between 0.5 and 2.0 in a Reynolds

number range from 50,000 to 150,000. In addition, many researchers have found that simple


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modifications, especially aspect ratio corrections, used for larger wings do not have the same

effect when used on smaller wings in low Reynolds number flow, thus making the

aerodynamic design process even more difficult.

Flight under Reynolds number of 50,000 has shown that separation occurs in the laminar

region and transition does not occur in time to reattach the flow. For this reason, researchers

at Notre Dame declared the regime from 50,000 to 70,000 as the most suitable for early

MAV flight and development[6]. Since the time of their research, experimentation and flight

tests have found that boundary layer tripping can be used in this regime to decrease the

critical Reynolds number and maintain attached flow. Stable MAV flight under Reynolds

number of 50,000 has been demonstrated by using flow entrainment in flapping-wing models

and favorable prop-wash effects.

The greatest interest for MAV design is in the range of 20,000 to 70,000. Thousands of small

bird species fit into this region of flight. Thin airfoils are normally selected because the

hysteresis effects caused by transition of airfoils thicker than about 6% can be very

significant. Ideally, an infinitely thin wing should be used. The thickness of the wing drives

the size of the adverse pressure gradient. The thinner the wing used, the less severe the

magnitude of the adverse pressure gradient. In addition, outside disturbances have a much

larger effect on vehicles in low Reynolds number flow.

The first successful design of a micro air vehicle was achieved by AeroVironment. They

designed the Black Widow MAV with funding from DARPA. The Black Widow is a MAV

with a 6 in wingspan, airspeed of about 30 mph and weighs under 100 grams. This vehicle

also has the capability of carrying a video camera which transmits live video to the ground

and has an endurance of 30 minutes. It is also equipped with an autopilot, which is capable of

performing altitude, airspeed, and heading holds as well as a yaw damper. The transmitter

and actuators are some of the smallest and lightest systems available. This design led to

further interest and research in the field of MAVs from several countries and universities.

Although turbulence will make the flight path of a smaller vehicle more unpredictable,

simulations show that outside turbulence helps laminar flow to transition earlier (more

forward) on the wing’s chord.

Some researchers believe MAVs will soon operate at a Reynolds number below 10,000,

where flow is completely laminar. The limited modeling and testing of flow at Reynolds


22

numbers this low has been inconclusive; however, there is much speculation into

phenomenon such as eddy generation and vortex capturing which allow insects to fly at such

low Reynolds numbers.

1.4.2 Types of MAV Broadly, there are two categories that the MAVs can be divided into i.e. Flapping Wing

Micro Air Vehicle and the Morphing Wing Micro Air Vehicle. They are respectively

described as below;

1.4.2.1 Flapping Wing Micro Air Vehicle Let’s start with a statistics. Over one million species of insects fly with flapping wings and

10,000 types of birds and bats flap their wings for locomotion. Indeed, nature has

predominantly selected flapping-wing as its favorite flight-mechanism. However, whether

this choice is one of organic constraints or one of optimal performance is an unsettled matter.

Nevertheless, the fact that birds, bats, insects and many sea-creatures utilize this mechanism

with great success, at least merits a thorough scientific investigation.

But what is so special about flapping-wing MAV (also called ‘entomopter’) compared to its

fixed-wing or rotary-wing counterpart? The answer is superior maneuverability which aids

(a) obstacle avoidance (particularly in urban environment) and (b)effective navigation in

small spaces(e.g. search and rescue operation in debris). The development was accelerated by

advances in several micro technologies like MEMS, CCD cameras and tiny infrared sensors

resulting low detectability, low noise and real-time data acquisition. Bio-fluid-dynamicists

have attempted to explain the underlying physics of flapping wing flight both in the quasi-

steady limit and in the fully unsteady regime. Quasi-steady flow is only applicable to large

birds(like eagles) which soar and glide. When soaring, the wings are fixed and rigid, hence

acting like those of conventional aircrafts. For these fliers, flapping is restricted to limited

operations, such as take-off, landing and stabilization. For other birds and insects unsteady

flow results. Empirical correlation[6] predict the break between quasi-steady and unsteady

flight at approximately 15 cm wingspan. Interestingly 15 cm wingspan is also the arbitrary

design limit set for MAVs[7]. In fact many performance requirements of MAV(like high

maneuverability, very low speed flight capability, high power etc.) are inspired by the flight


23

of small birds and insects. Thus it is imperative that any MAV design must account for

similar sized biological fliers; one where the flow field is unsteady, laminar, incompressible

and occurs at low Reynolds number.(104< Re<105)

In spite of such specialized approach, the complexity of mechanism is so high that it often

extinguishes the boundaries between aerodynamics, structural analysis and control. For

example, bats can control their wing surface by changing the degree of tension in their wing

membrane, thereby effectively changing the wing-camber due to the passive aero elastic

response of the membrane to the aerodynamic loading. Thus wing stiffness distribution plays

an important role for resultant force generation and control.

1.4.2.2 Rotary Wing Micro Air Vehicles: These are the small helicopters with small rotor diameters. These are very useful if able to

successfully hover and fly vertically. For Rotary-Wing MAVs high thrust to weight ratio is

needed. For a meaningful mission completion, high endurance is also needed. For these a

suitable match of battery, electric motor and rotor is a challenging job.

(a)

(b)

Figure 1.6: (a) ‘Delfly’ Flapping Wing MAV [8], (b) ‘MAVSTAR’ Rotary Wing MAV[9]


24

1.4.2.3 Morphing Wing Micro Air Vehicle:

It has been acknowledged that the best flyers are Nature’s flyers; birds and insects. Engineers

can learn these flyers and develop efficient flying machines by copying their techniques. One

of the techniques that bird use for improved flight is to morph their wings during flight. Birds

are forced to alter their wing shapes dramatically in order to accomplish cruise glides, steep

descents, and aggressive maneuvering[10]. “The words “morph” and “morphing” are actually

digressive forms of the word “metamorphosis,” which derives from the Greek “meta” (to

change) and “morfe” (form). This is an apt description of the ability that birds possess to

change the form or geometry of their bodies for increased maneuverability, as well as for

stable flight in a wide variety of ambient conditions”[11]. The very first aircraft developed by

Wright brothers was also inspired by birds as its wing tips could be twisted for lateral control

and stability. After that many aircrafts were developed with some inspiration to their natural

counterparts but they far behind these. Now, in modern era engineers are again inclined

towards nature’s flyers; their impeccable flight. NASA Administrator Dan Goldin stated:

NASA will open the door to a bold and revolutionary era by using technology to mimic

nature. The seemingly efortless flight of birds provides the inspiration for new aircraft

utilizing wings that reconfigure in flight. The vehicle changes—or morphs—from a low-

speed configuration to one more suited for high speed[11].

Types of shape changes include span, chord, camber, area, thickness, aspect ratio and

planform. The morphing can also be applied to a control surface in order to eliminate hinges.

Morphing can be used as a control effector by changing the shape of the aircraft in order to

alter the flight dynamics. The concept of morphing has been looked at by DARPA and

NASA to show the aerodynamic benefits; however, the use of morphing for control design

has not been studied extensively.

Morphing is also used on the F-14 which has a variable sweep on the wing, therefore

changing the shape of the wing during flight. The wings are swept in order to balance the

range and speed by slowing down the increase in drag which develops as velocity increases.

[7]


25

1.4.2.3.1 Different Configurations Of Morphing:

i. Change Of Camber:

NASA has designed a wing which changes the camber of the wing[6]. One of the wings

considered is referred to as a Hyper-Elliptic Cambered Span (HECS) because the curvature

along the span is continuously changing. This provides a larger area which allows for greater

lift. This vehicle uses a hinge-less panel along the trailing edge of the wing as a form of a

control surface for pitch and roll. The simulations demonstrated the aerodynamic benefits but

also show this vehicle has unstable lateral-directional dynamics. The use of smart materials

such as shape memory alloys and piezos have been considered in the design of morphing

wings but there is still a limit in that not enough force can be produced in order to twist large

wings using these mechanisms. However, smart spars have been built which provide

different types of morphing but have not been tested on flight vehicles

ii. Change Of Sweep:

Another mechanism for morphing which has been studied considers changing the sweep of

the wings of a small unmanned air vehicle (UAV) [P. de Marmier, and N.M. Wereley,

“Morphing Wings of a Small Scale UAV Using Inflatable Actuators for Sweep Control,”

AIAA-2003-1802, 2003.]. The morphing on this UAV is done in order to meet changing

mission requirements. Actuation of the morphing is done by using inflatable actuators which

are powered with compressed air. One of the benefits of this project is that the actuation

mechanism used is much lighter than the typical hydraulic systems that are used on full scale

aircraft. The effects of the sweep are then studied considering the change of aspect ratio, lift

and drag. The area of the wing can also be changed by extending the length or trailing edge

as some birds do. The aspect ratio is also affected by the morphing and can be used to

consider lift and drag for aerodynamics. Sweeping wing MAV named, ROBOSWIFT, has

been developed by the TU DELFT University in Netherlands. It is bio-inspired by the bird

Swift. No documented design of RoboSwift is available.


26

Figure 1.7: RoboSwift of TU Delft, Netherlands[12]

iii. Change Of Aspect Ratio:

Similarly, a study has been done considering an inflatable telescopic spar which can be

morphed span wise[13]. This design allows for changes in the aspect ratio while still providing

enough support from the spars for the air loads which are being applied. This is achieved

because the telescopic spar is pressurized and the telescopic skins maintain the geometry of

the airfoil as well as provide effective storing and deployment of the mechanism.

iv. Twist In Wing:

This is currently being used for control on the Active Aero elastic Wing (AAW) as well as

the vehicles in this project. The morphing on the AAW causes the wings to be twisted in

response to the moments induced by the control surfaces. Birds and bats also do this in order

to obtain the

required lift or thrust during flight.

Also for the purpose of considering improved maneuverability and performance, roll

maneuvers have been studied using a flexible wing[14]. Numerical studies were used to

consider the aerodynamic loads on a flexible wing at high speeds. Wing twist is also

considered in order to recover the rolling moment lost but has not been tested in flight. This

is due to the challenges involved in implementing a functioning mechanism for wing twist on

a full scale aircraft.


27

v. Variable Gull-Wing Angling: This mechanism changes between the in-board and out-board wing sections in flight to vary

performance capability. The mechanism uses a linear lead-screw actuator to quasi-statically

vary the gull-wing position[9].

Figure 1.8: UF MAV[9]

vi. Wing Curling: Wing curling is good for high maneuverability. As the actuator adjusts the tension on the

cable, the wing deforms into a twisted form that is appropriate for flight control. Namely, the

resulting shape increases the angle of incidence of the morphed wing and increases the lifting

force produced. When one wing side is morphed, a lift differential is created which causes

the aircraft to incur a roll rate[15].

University of Florida, USA has developed Morphing wing MAVs of flexible wing with the

above stated morphing configurations and some other like Leading-Edge twisting, Multi-

point wing shaping, Wing-Tail folding and forward-backward section sweeping.

1.5 Morphing Wing; Sweeping in detail

As there are several morphing capabilities that birds have but most common is Wing

sweeping configuration. From small birds to large birds, sweeping wing trend have been

seen. Birds sweep their wings during gliding flight to adapt flight conditions. It enables them

to achieve efficient flight. Most of the birds sweep their wings backwards for high speed.

Some birds have multi-joint wings which sweep forward and backward in sections like

falcon and sea-gull. By sweeping, aspect ratio and wing area reduces.


28

Figure 1.8: Swift sweeping during flight[16]

Intrigued scientists have been studying birds specifically the gliding birds for long time.

David Lentink, Videler, Tucker and others have succeeded in documenting aerodynamic data

of the birds concerning wing morphing.

1.5.1 Lentink’s Work:

The Swift bird lives majority of its lifetime in air, eating, forging, mating etc. David Lentink

studied the gliding flight of Swift (Apus apus). The aerodynamic performance of swift

affected by the shape changing wings, morphing wing, was studied.

“Aerodynamic force is proportional to force coefficient x wing area x square of glide speed.

Increasing sweep angle from5deg (fully extended) to50deg…, decreases wing area and shape

(that is, aspect ratio) by roughly one-third.” [17]

The experiments show that “swept wings contribute low drag coefficients at low angles of

attack; extended wings contribute high lift coefficients at high angles of attack”. [17]


29

Figure 1.9: Swift’ sweeping aerodynamics [17]

As shown in the figure above, at high angle of attack extended wings give high lift and drag,

so at low speed extended wings are good as for takeoff and landing. At low angles of attack

the swept wings give low drag coefficient which is good at high speed. The swept wings

provide to manage high speed without additional power consumption. The effects of wing

shape are amplified by wing area, as figure shows, The decrease in wing area with increasing

sweep further enlarges the enveloping polar for a given glide speed, further widening the

performance gap between fixed-shape and morphing wings. [17]

This study by Lentink also shows that Swept wings can bear higher loads during fast turns. at

high speeds, bird can get sharp turn and experiences high wing loads. The swept wings do

not flutter while extended wings cannot bear such loads. The extended wings bent at 15ms-1

and start vibrating violently in this case of swift wing.


30

Figure 1.10: Swift’s sweeping wing aerodynamic characteristics [17]


31

1.5.2 Videler’s Work:

J.J. Videler also studied aerodynamics of the Swift. He provided the wing characteristics,

airfoil shapes and the aerodynamic flow data of the swift in his work; “Leading Edge Vortex

lifts the Swift”. The work shows that by sweeping the wing the lift coefficient decreases but

the lift loss is compensated by the leading edge vortex generated by the sharp leading edge of

the wing. The vortex attached to the leading edge has low pressure which gives an extra lift

which compensated the lift loss but not fully.

Figure 1.11: Swift’s Leading Edge Vortex [18]

1.5.3 Tucker’s Work: The gliding flight of falcon has also been studied by biologists specifically by Tucker and

Parrot. They studied the flight of living falcon in the wind tunnel and were able to acquire the

data of the living flying falcon. The falcon change its wing shape by change sweeping its

wings as depicted by the figures below. The falcon changes its wing shape by sweeping at

different angles for different flight speeds.


32

Figure 1.12: Mean wing area of the falcon in the wind tunnel at different speeds and glide angles. Wing and tail shape of the falcon at various speeds and glide angles. A, 6-6 m./sec.,

6°; B, 8-5 m./sec, 5°; C, 14-3 m./sec., 6°. [20]

“The falcon over its entire speed range flies closer to the speeds where L/D is maximum for

the observed wing span, wetted area and K values. It accomplishes this by adjusting wing

span and wetted area as speed changes. Thus, as air speed increases, the falcon decreases

wing span and wetted area … maximum L/D also increases… The falcon adjusted its wing

span in flight to achieve nearly the maximum possible L/D value over its range of gliding

speeds.” [20]

1.5.4 Pennycuick’s Work: “Family of curves relating speed, tail spread and power required for fight. The result

calculated by the Pennycuick model (Pennycuick, 1989) is recovered at zero tail spread. The

tail can reduce the power required for fight at low speeds by taking some of the load of the

wings. At high speeds the tail should be furled to reduce drag. The bird modeled here is at

here is theoretical swallow. Maximum saving in power due to the tail is just over 25%. The

diagonal line across the family of curves shows the path across the surface that gives the

minimum power at each speed.


33

Family of curves relating speed, power and wing span (where wing span is proportional to

wing area i.e. constant mean cord) for the theoretical swallow. Each curve recovers the

Pennycuick model for a given span. For absolute minimum power the bird should adopt

morphologies on the bold curve.” [21]

Figure 1.13: Bird power consumption curves [21]

1.5.5 Morphing Wing (by sweeping here) Advantages. 1. By sweeping wing at certain angles, speed can be varied without varying power

consumption in contrast to that of fixed non-morphing wing. This will improve the endurance

by adapting the flight conditions.

2. Morphed wings can bear higher loads which are encountered on sharp turns.

3. By asymmetric morphing, sweeping only one wing at a time, sharp turns can be

achieved which conventional airplanes cannot accomplish by conventional control surfaces.

1.6 Morphing Standards Morphing aircraft in accordance to NASA are aircraft utilizing wings which are capable of

changing the shape of planform significantly during aircraft flight. It could be perhaps a

200% change in the aspect ratio, 50% change in wing area and a 20 degree change in wing

sweep of the aircraft wing. By the help of these achievements of the wing shape the design

might be integrated to enhance the operational potential of the aircraft such as reducing the

aircraft‘s required takeoff gross weight and enable an aircraft to fly the required flight which

a fixed wing aircraft cannot follow. [22]

CH 02. PROJECT ‘RAA’

34

•Design Objectives

1• Aerodynamic Design

2•Morphing Mechanism

3 •Control System

•FABRICATION

CHAPTER 2 PROJECT ‘RAA’

2.1 Project Objective: Project objective is to design a Bio-Inspired Morphing Wing Micro Air Vehicle named, RAA. It will be capable of sweeping its wing backward during flight for improved endurance, maneuverability and agility. It will able to adapt flight conditions. Wing will be composed of three feathers which will overlap each other as the wing sweep

2.2 Mission

Figure: 2.1: Mission profile for ‘RAA’

2.3 Design Methodology: The design is broken into three modules; Aerodynamic design, Morphing mechanism, Control System then Fabrication in the last.

CH 02. PROJECT ‘RAA’

35

2.4 Management: The design is broken down into three modules; Aerodynamic design, Morphing mechanism, Control System and Fabrication in the last.

Serial No. Name Assigned Modules 01 Raheel Ahmed Aerodynamic Design, Procurement,

Fabrication 02 Anas Bin Aqeel Morphing Mechanism, Fabrication

03 Ahmed Zia Control System, Fabrication

Table 2.1: Management Table

2.5 Time Management:

Figure 2.2: Gantt Chart of Project RAA

CH 03. AERODYNAMIC DESIGN

36

CHAPTER 3

AERODYNAMIC DESIGN

3.1 Airfoil Selection

The airfoil selection is the very important in the Design of RAA. As the Reynolds Number of

MAVs are low in 50000-150000 range[23] so, Airfoils were selected of low Reynolds

Number. There is a long list of airfoil but some were chosen for further consideration. Bio-

inspiration was kept in mind during airfoil selection, as it is our whole philosophy of the

project. RAA would have two section wing; arm-wing and hand wing. Arm-wing is fixed and

hand wing is comprised of three feathers and these will sweep back during flight. Two

different airfoils were selected for fixed-wing and hand-wing. The fixed wing airfoil is

thicker relative to hand wing. It will also be of higher Coefficient of lift; Cl.

Videler has shown in his paper that of swift has different airfoils for fixed-wing and hand-

wing as shown in the figure.

Figure 3.1: Airfoil shapes of the wing of the Swift bird. [18]


37

Also, Tucker in his paper “Aerodynamics of Falcon and Gliding birds” has shown airfoils

similar to that of birds. Below is the snapshot of Tucker’s work.

Figure 3.2: Tucker provided the airfoil shape of Eagle and the data airfoils similar to that of

bird. [20]

Airfoil number 5 in the above excerpt is that of eagle which is very much similar to swift’s

airfoil. Other birds have similar airfoils too.


38

3.1.1 Arm-Wing Airfoil: Low Reynolds number Airfoils selected for consideration are as follows:

3.1.1.1 Eiffel 13: From NACA TR-93; Aerodynamic Characteristics of Airfoils, Report No. 93:


39

Figure 3.3: Eiffel 13 data, at Reynolds Number of 139664, Max Cl = 0.55 Min Cd=.015 and Max L/D = 13

3.1.1.2 N.P.L 4a: From NACA TR-93; Aerodynamic Characteristics of Airfoils, Report No. 93:

Figure 3.4: NPL 4a data at Reynolds Number of 24536, Max Cl = 0.6, Min Cd = .025 and

max L/D=13


40

3.1.1.3 Gottingen 464[24]

Figure 3.5: GOE 464 airfoil shape[24]

Thickness: 7.7%

Camber: 9.9%

Trailing edge angle: 6.6o

Lower flatness: 20.0%

Leading edge radius: 3.6%

Max CL: 2.096

Max CL angle: 15.0

Max L/D: 284.023

Max L/D angle: 0.0

Max L/D CL: 0.923

Stall angle: 0.0

Zero-lift angle: -8.5


41

Figure 3.6: GOE 464 Characteristic Curves[24]


42

3.1.1.4 EPPLER 58[24]

Figure 3.7: Eppler 58 airfoil shape[24]

Thickness: 5.6%

Camber: 6.5%

Trailing edge angle: 6.0o

Lower flatness: 68.9%

Leading edge radius: 2.5%

Max CL: 1.81

Max CL angle: 7.0

Max L/D: 68.374

Max L/D angle: -0.5

Max L/D CL: 1.004

Stall angle: -0.5

Zero-lift angle: -9.5


43

Figure 3.8: Eppler 58 Characteristics Curves[24]


44

Airfoil Cl max Cd min L/Dmax

Eiffel 13 0.55 .015 13

N.P.L 4a .6 .025 13

Gottingen 464 1.6 .05 12

Eppler 58 1.6 .025 60

Table 3.1: Summary of Airfoils under consideration

NACA 4704 and NACA 4402 were also brought to consideration. Because of non

availability of their data, these were analyzed with a software application “XFOIL”. It gives

that

NACA 4704 has max Cl =1.028 and Cd=.05

NACA 4402 has max Cl = 1.08 and Min Cd = .019

Gottingen 464 is very much close in shape to that of birds, especially swift and it also has

higher Cl max and L/D max as our main concern is to get high lift with less dimension.

Eppler 58 has max Cl at AOA below 10 o which can result early stall. So, Gottingen 464 is

selected for fixed wing.


45

3.1.2 Hand-Wing Airfoil:

For Hand wing, Gottingen 417a is selected because it is the only available airfoil which is

similar in shape to that of Swift’s hand wing. Its characteristics found from NACA report No.

TR. 286 is shown below,

Figure 3.9: GOE 417a at Re=419000, Max Cl = 1.3, Min Cd=.04, Max L/D=14


46

3.2 Wing Design and Study:

The requirement of the current project is to design MAV which will be of 50cm wingspan, as

it is first time at this level. This work would be improved in the future and wingspan would

be smaller. Airfoils have been selected, now these airfoils would be used to design the wing

with 50cm wingspan and appropriate chord.

3.2.1 Aerodynamic Coefficients: Lift ad drag forces are usually treated as nondimensional coefficients. The wing reference

area Sref or simply S is the full trapezoidal area extending to the aircraft centerline. The

dynamic pressure of the free stream air is called “q” as defined

q= 1/2ῥV2 (1)

L = qSCL (2)

D= qSCD (3)

By definition, the lift force is perpendicular to the flight direction while the drag is parallel to

the flight direction. Remember that the 2-D airfoil characteristics are denoted by lowercase

subscripts (i.e., Cl) whereas the 3-D wing characteristics are denoted by uppercase subscripts

(i.e., CL and CD).

3.2.2 Lift Calculation: The graph of CL vs Angle of attack ‘α’ is a straight line up to some maximum value of CL.

Figure 3.10: Aspect Ratio Effect Upon Lift Curve Slope [25]


47

The lift curve slope CL α reduces as aspect ratio reduces. This is due to the downwash angle

produced, which reduces the effective angle of attack of the wing. Relation of CL α with the

sweeping angle as is, [25]

where,

Reference Wing Area:

(6)

Exposed Wing planform Area:

(7)

bhand and barm is the wingspan of arm-wing and hand –wing respectively.

Total wing span will be; b = bhand + barm (8)

(9)

where binitial is the wingspan to be kept initially at zero swept angle.

Aspect ratio:

(10)

Maximum CL of the hand wing is related to the sweeping angle as

(13)

C L,,ad = 2 ·

0.32 + 0.16 · A

cos ( a s ) · ( 1 – ( M · cos ( as ) ) 2 )

0.5

Lift Curve s lope in degrees:

C L, = · CL,,ad

180

S ref = b arm · c + bhand · c + c ·

2

Sexp = S ref – barm · c

bhand = ( b initial – barm ) · cos ( a s )

A = b 2

S ref

C L,max = 0.9 · Cl max · cos ( a s )

AOA at max CL

max = CL,max

C L, + o

C L = C L, · ( – o )

(4)

(5)

(11)

(12)


48

Similarly, For Arm-Wing section, from Reference 25

where Lift-curve-slope CL arm α of the arm-wing and Clarm α is the lift curve slope of the airfoil

curve, earm is the Oswald efficiency factor which is normally 0.95 and df is the fuselage

diameter.

ΔCLmax is zero for M≤0 and no data available for ΔαCLmax at M≤0.

Now the Lift will be

(22)

where, q is the dynamic pressure.

CL,arm,alpha = Clarm,

1 + 57.3 · Clarm,

· earm · Aarm

Aarm = b arm

c

e arm = 0.95

CL,arm,max = Clarm · 0.9

S ref ,arm = barm · c

Sexp,arm = S ref ,arm – d f · c

arm,max = CL,arm,max

CL,arm,alpha + arm,o

CL,arm = CL, · ( – arm,o )

L = q · S ref · CL + q · b arm · c · CL,arm

(14)

(15)

(16)

(17)

(18)

(19)

(20)

(21)


49

3.2.3 Drag Calculation:

There are two types of drag, Parasite drag and the other is Induced Drag. Total drag would be

the sum of these drag forces.

3.2.3.1 Parasite Drag: For parasite drag calculation, Component Build up Method is used as described in [26].

퐶 = ∗ ∗ ∗ + 퐶 + 퐶 & (23)

where CDo is the parasite drag coefficient, Cf , FF, Q and Swet is the Skin friction drag,

Component form factor, Component interference factor and wetted area respectively of each

component that is wing and the fuselage. Q can be of value 1.0 in this case. CD misc and CD L&P

are special features drag and protuberances contributions. These two factors have been

neglected in this case. Drag because of wing and fuselage have considered in this case.

For Wing variables have been subscripted by 1 and that of fuselage by 2 as follows,

Swet, w is the wetted area of the wing.

Swet, f is the wetted area, df is diameter and le is length of fuselage.

Component form factor of wing is given by, [26]

(28)

Where, tc and xc are ‘maximum thickness to chord ratio of airfoil’ and ‘chord wise location

of airfoil maximum thickness point’ respectively.

tc = .015 and xc = .4 for our airfoil.

Mach number, M= V/333

Component form factor of fuselage is given by, [26]

C D,o,1 = C f ,1 · FF1 · 1 · Swet,w

S ref

Swet,w = 2 · ( S ref – d f · c )

CD,o,2 = C f ,2 · FF2 · 1 · Swet,f

S ref

Swet,f = · d f · le

FF1 = 1 + 0.6 · tcxc

+ 100 · tc 4 · 1.34 · M 0.18 · cos 0.28 ( a s )

(24)

(25)

(26)

(27)


50

(29)

where

, (30)

Skin friction depends on the Reynolds number. In this case laminar flow has been

considered. For wing,

Similarly for fuselage,

Total Parasite drag coefficient,

(35)

3.2.3.2 Induced Drag: Induced drag is due to lift and is proportional to square of lift coefficient with a

proportionality factor called the “drag-due-to-lift factor” or “K”. [26]

(36)

There are two methods to estimate K. One is the Oswald span efficiency method and other is

Leading edge Suction method. Former is the classical method and latter is based upon the

concept of leading edge suction and provides a better estimate of K, one that includes the

effects of the change in viscous separation as lift coefficient is changed. Ref 7

Here, Leading edge Suction Method has been used by which K is inverse of Lift curve slope

(in radians) which is calculated in the lift calculations,

FF2 = 1 + 60

fs3

+ fs

400

fs = led f

C f ,1 = 1.328

Rw0.5

Rw = 1.2 · V · c

C f ,2 = 1.328

R f0.5

R f = 1.2 · V · le

CD,o = CD,o,1 + CD,o,2

C D,i = K · CL2

(31) (32)

(33) (34)


51

퐾 =

(37)

Now, Total drag coefficient

(38)

and Drag Force

(39)

3.3 Wing Design & Study Program

Using Engineering Equation Solver (EES), a general computer program has been developed

named, "RAA Wing Design and Study”. All the above equations have been coded in EES

and a window has been designed for user interface. Following are the some of the snapshots

of the EES program windows,

Figure 3.11: EES program Equation Window

CD = CD,o + CD,i

D = q · Sref · CD


52

Figure 3.12: EES Program Formatted Equations Window

Figure 3.13: EES Program Solution Window


53

Figure 3.14 : RAA Wing Design and Study EES Program user Interface

As shown in the above “Diagram Window”, User enters the values of the “Input Parameters”

then by clicking “Calculate” button the “Output Parameters” are calculated and shown in the

window. The parameters are shown in the window as defined the previously discussed

equations. For parametric study the variable “as” or “V” is set to output then by clicking

“Update Tables” all the parametric tables are updated. “Show Plots” button shows Plots of

different parameters.

By using “RAA Wing Design & Study EES Program” following design parametric results

have been achieved for un-swept condition to select the right wing characteristics. Total

wingspan is set to 50cm. Arm wing airfoil is GOE 464 and that of Hand wing is GOE 417a.

All the required data is entered in the input parameters and get the output parameters. The

AOA ‘α’ is set to 5deg.


54

Arm Wing

Span,

‘barm’ (m)

Hand Wing

Span,

‘bhand’ (m)

Chord, ‘c’

(m)

Speed, V

(m/s)

Lift ‘L’

(N)

Drag ‘D’

(N)

Reynolds

Number

‘Re’

.1 .4 .1 10 2.244 0.2493 70225

.1 .4 .15 10 4.503 0.4771 105337

.15 .35 .1 10 2.68 0.2622 70225

.15 .35 .15 10 5.364 0.5014 105337

.1 .4 .1 15 5.052 .5524 105337

.1 .4 .15 15 10.14 1.064 158006

.15 .35 .1 15 6.033 0.5813 105337

.15 .35 .15 15 12.08 1.119 158006

.2 .3 .1 10 3.132 0.2754 70225

Table 3.2: RAA Wing parametric Study

Increase in chord length and speed, a considerable increase in Lift and Drag has been seen.

Increase in Arm-wing length also contribute to the Lift and Drag increase. As initial

estimation of the weight of RAA is 200g, so we select the wing with 15cm Arm-Wing and

10cm chord, which is capable of lifting 268g at 10m/s speed and 5deg AOA. The Reynolds

Number is 70225 for this wing.


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3.3.1 Morphing Characteristics “RAA Wing Design & Study EES Program” provided the plots to study relation of

aerodynamic forces by sweeping the wing during flight for the selected wing.

Figure 3.15: Drag Vs Sweeping Angle Plot

The Plot shows that Drag reduces from .268N to 0.155N by wing sweeping from zero to 60

degrees; a 42% reduction. At 45deg, Drag is 0.2N which is 25% reduction. This proves that

by sweeping, drag reduces and speed can be increased without additional power

consumption.


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Figure 3.16: Lift Vs Sweeping Angle Plot

Lift also reduces by sweeping the wing but it is compensated by the leading edge vortex

generated by sharp leading edges.


57

Figure 3.17: Lift-to-Drag ratio Vs Sweeping Angle Plot

The Lift to Drag ratio increases with the increase in sweeping angle, that we desired.


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3.4 Computational Fluid Dynamics (CFD) Analysis: The CFD analysis of Straight wing and 45o Swept Wing has been done. The CAD models of

the wings have been designed using Pro-Engineer Software. The CAD models have been

imported to Gambit for Pre-processing and then Fluent has been used for solving the meshed

file.

Figure 3.18: CFD Analysis Process

3.4.1 CAD Models:

Figure 3.19: CAD Models of Straight RAA wing and 45o Swept Wing


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3.4.2 Meshed Geometries:

3.4.2.1 Straight Wing: Tet/Hybrid mesh has been used. The Size function is; Size limit: 30 Growth rate 1.2

generating from the meshed wing faces.

Total Number of Nodes: 581415

Total Elements: 3164647

Figure 3.20: Meshed Geometries of Straight Wing in Gambit.


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3.4.2.1 45o Swept Wing: Similarly, Tet/Hybrid mesh has been used with Size function: growth rate=1.3, max size=25

generating from meshed wing faces.

Total Nodes: 505217

Total Elements: 2678768

Figure 3.21: Meshed geometries of 45o Swept wing in Gambit


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3.4.3 Fluent Solution:

(a)

(b)

Figure 3.22: Fluent Solution, (a) Velocity magnitude contour plane at Straight Wing. (b) Velocity magnitude contour plane at 45o Swept wing


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3.4.4 Results

At 0o of AOA, V=10m/s, using Laminar Model

After 1000 iterations

Wing Lift (N) Drag (N)

Straight Wing 1.5071 0.1186

45o Swept Wing 1.1114 0.0994

Table 3.3: CFD Analysis Results

3.5 Conclusion:

A bio-inspired wing has been design with two sections of different airfoils, like birds have.

The theoretical study using aerodynamic relations with the help of EES program has

validated the drag reduction by increase in sweeping angle. The CFD analysis has also

validated the reduction of drag by sweeping wing. With the help of morphing wing, which

change shape by sweeping in this case, drag reduces to adapt different speed conditions

without additional power consumption. It will save power and increase the endurance of the

air vehicle. Moreover, by sweeping sharp turns can be accomplished which is the demand of

the micro air vehicles.

CH 04. RAA MORPHING MECHANISM

63

CHAPTER 4

RAA MORPHING MECHANISM Before attempting to build various components of the wings and its design, the kinematics of

different parts are being investigated to test the theories of discrete approximation and how

we could reduce system order.

In accordance to our project the best mechanism of all would be the one that could provide us

with a maximum degree of sweep during our flight and could control all components of wing

with two differently actuated motors to control the wings individually.

A number of mechanisms have been proposed to achieve the combination of sweeping with

actuators, so that a maximum degree rotation could be achieved. Our Wing area is 156cm2

and wingspan is 50cm.

4.1 Mechanisms Proposed The number of mechanisms was proposed for achieving the NASA Standards that were

angular degree, plan form area and aspect ratio for this E Quintet mechanism.

Following are the mechanisms that were proposed for our MAV named RAA:

1) Cam Lobe Mechanism

2) Gear Actuated Mechanism

3) Single Link Actuated Mechanism

4) Tertiary Pin joint Mechanism

From all of these mechanisms, we come across different results and difficulties that occurred

during testing of mechanisms such as:

Mechanism restriction

Weight concern

Return back mechanism

Aerodynamic concern

Wing feathers geometry


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4.1.1 Cam Lobe Mechanism

The first mechanism approached to sweep the wings is a simple cam lobe mechanism to

achieve the maximum amount of degree. The wing of the system was designed in pro

engineering software and the feathers were tried to sweep and restrict with the help of cam

lobe mechanism so that it not only sweeps forward but could also return to its initial position

of the wings from where it started sweeping. As we were trying to achieve the return motion

of the wing with the same motor.

Following are some of the diagrams with explanation about the working of this mechanism

Figure 4.1: Pro Engineering Cam Lobe Mechanism

In these diagrams the cam lob mechanism is shown such as a grove and circular block is

made along individual feather of the wing and for restriction the grove is blocked from both

sides of the feather, so that its movement restricts and utilization of single servo motor could

be achieved.

In this mechanism a grove and circle of very small thickness is made as to achieve

aerodynamic stability.


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Initial Position

Figure 4.2: Pro Engineering Initial Position Model of Cam Lobe Mechanism

Final Position

Figure 4.3: Pro Engineering Final Position Model of Cam Lobe Mechanism


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Fabricated Part

Figure 4.4: Final Fabricated Part of Cam Lobe Mechanism

4.1.1.1 Advantages The benefits for this mechanism are given as under:

It was light in weight due to less weight returning mechanism.

It provides good aerodynamic properties as of attachment between the feathers.

It provides one of the best returning mechanisms

4.1.1.2 Disadvantages Following are the disadvantages occurred in cam lob mechanism:

It provides a restriction to sweep angle of 18 degree which do not fulfill the NASA

standards of minimum 20 degree sweep.

It has a restriction of chord of every individual feather which was the main problem

in the mechanism during sweeping.

It provides a very complex geometry to calculate the exact pattern of the groves made

inside the feathers.


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4.1.2 Gear Actuated Mechanism In this mechanism gears were used to move the feathers in both directions, so that it provides

the returning mechanism as well.

For this purpose we used five different gears to move and to control the motion of the

feathers with the help of spars. For this we used three forced gears and two idle gears to

increase the speed of the second and third feather, so that the planform area of the wings

reduces to the NASA standards.

Figure 4.5: Gear Formulas [27]

For this purpose we used different gears having different sizes as mentioned below

No of Gear Gear 1 Gear 2 Gear 3 Gear 4 Gear 5 Diameter

Pitch 1.5

inch 1

Inch 1.25 Inch

1 inch

1 inch

Material Plastic Plastic Plastic Plastic Plastic

Table 4.1: RAA Gear Parametric Study

All of the gears used were spur gears and adjusted in the manner from big gear on first

feather and smallest gear on the third feather.

The diagram shows the spur gear and the path of action between two gears


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Figure 4.6: Spur Gear and Path of Action Diagram [27]

Out of these five gears, gear 2 and gear 4 were used as the idle gears and were used to control

the speed of feathers with the help of spars, so that the feathers could overlap each other and

reduce the planform area and change the aspect ratio.

Figure 4.7: Gear Actuated Train [28]

The good thing about this mechanism was that it was providing us a 25 degree angle sweep

and could be increased a lot by changing the gear ratios.


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4.1.2.1 Advantages The benefits for this mechanism are given as under:

It provides good aerodynamic properties as of attachment between the feathers.

It provides good returning mechanisms

It could provide a lot of sweep angle compared to any other mechanism by changing

the gear ratios.

4.1.2.2 Disadvantages Following are the disadvantages occurred in gear actuated mechanism:

It provides a very complex geometry.

It was heavy in weight due to gear weight.

It needs gear to be manufactured of steel as iron, so that it withstand the air resistance

of about 10 m/s

Spars were used which were adding an additional weight

4.1.3 Single Link Actuated Mechanism

We tried to accomplish this mechanism but the greatest flaw in this mechanism was a single

link movement was used by placing the pin joints of other three feather links with it at a

calculated distances, so that the movement of link one could adjust the motion of other three

links.

Following diagram shows the movement and mechanism used


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Figure 4.8: Link Actuated Mechanism

The diagram shows the moment of the link 1 that is upward and link 2,3 and 4 were attached

as a pin joint. When link 1 moves upward link 4 displace more angular velocity compared to

the link 3 and link 2. The distance between link 2 and link 3 is 2.5 cm whereas between link

3 and link 4 is 1.25cm.

So the total sweep angle get from this is 22 degree.

4.1.4 Tertiary Pin joint Mechanism

To solve the problem of mechanism our study went to E quintet planner mechanism as

shown under


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Figure 4.9: E Quintet Mechanism

First of all the wings were designed in the Pro Engineering software and then divided into

elliptical feathers as shown in the under diagram

Figure 4.10: Pro Engineering Wing Design


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Figure 4.11: Pro Engineering Feather Design

To do the calculation software is used known as Working Model. From this software we have

calculated the position of the linkages at where we get the maximum sweep angle and

maximum reduction in planform area.

Firstly we imported the feather shapes from the Pro Engineering software with the help of the

polygon command as shown under

Figure 4.12: Polygon Command in Design bar

The following are the first, second and third feathers imported from the Pro E which can be

changed with every single point shown under and could be changed from command

Window --- Geometry


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Figure 4.13: Properties of Feather 1,2 and 3

Figure 4.14: Geometrical Arrangement of Mechanism

This Diagram shows the feathers along with tertiary link in red color attached to the feathers

at points with the help of command point element and pin joint at a distance of 2.5 cm from

each other.


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Part

Feather 1

Feather 2

Feather 3

Tertiary Link

Weight

4 Grams

3 Grams

1 Gram

7 Gram

Span

12 cm

10 cm

9 cm

6cm

Material

Wood

Wood

Wood

Steel

Table 4.2: RAA Feather Parametric Study

After doing the calculations of individual feathers, we move to its simulation, for that purpose we have placed the feathers at a distance of 2.5cm from each other with the help of

point element and pin joint command.

After placing the feathers into the right position along with tertiary pin joint, a motor is

placed at the first feather as shown:

Figure 4.15: Motor Placement on Feather 1


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From Window command bar properties of motor is selected by scrolling down and a velocity

of 10 degree/sec is set for the analysis.

Figure 4.16: Motor Properties Representation

From Menu toolbar, we select different output properties to look at the performance and the

maximum sweep angle for every individual wing and for the whole mechanism.

Following graphs are showing the rotation, velocity and acceleration of individual feathers

along with its numerical values.


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Figure 4.17: Rotation, Velocity and Acceleration Graphs of Feather 1


77

These diagrams show the graphs for second feather



78

These diagrams show the graphs for third feather



79

Following are the graphs and timing for the mechanism along with motor rotation graph

Figure 4.20: Rotation Graph of Motor and Time Taken

Following are the diagrams showing the initial position and final position of the mechanism

where at initial position the angle is at 3.343 degree and at the final position where the

mechanism restricts is 37.343.

Figure 4.21: Initial Position of Wing Geometry


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Figure 4.22: Final Position of Wing Geometry

4.2 Results Achieved

Part

Feather 1

Feather 2

Feather 3

Tertiary Link

Initial Position

(Degree)

363

360.2

1083.4

707

Final Position

(Degree)

397

400.9

1130.4

701.8

Angle Swept

(Degree)

34

40.7

47

5.2

Velocity

(Degree/sec)

10

12.7

15.6

3.8

Acceleration (Degree/sec2)

3.3 e-12

0.435

1.68

2.497

Table 4.3: RAA Feather Parametric Stud


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4.3 Standards Achieved

Thus we have achieved the following NASA standards

Parameters

NASA Standards

Initial

Final

Achieved

Angle Swept

Minimum 20 Degree

3

Degree

37

Degree

34

Degree

Change in Area

Minimum of 50%

76 cm2

156 cm2

54 % area reduced

Table 4.4: RAA Standards Achieved Parametric Study

CH 05. CONTROL SYSTEM OF RAA

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CHAPTER 5

CONTROL SYSTEMS OF RAA

The control systems employed in our RAA constitute a theoretical approach to developing

the control of the MAV. In the ongoing report regarding the control systems of our RAA, we

would be presenting a solid approach as in how and what feasible control systems could have

been placed on the MAV to be considered as a “Fully Autonomous” Micro Air Vehicle. The

reason we cannot practically develop the Autonomous Micro Air Vehicle are the very high

costs involved in purchasing the necessary MEMS related equipment and the great amount of

complexity involved in electronics related task that was not possible due to limited amount of

time.

However, we will be going through the necessary equipments and their functions in making

our RAA completely autonomous so as to be a useful information for future undergraduate

students willing to undertake this project and work on it solely to make it semi-autonomous

or completely autonomous Micro Air vehicle.

As a start, we will be looking at what sensors are basically employed on a Micro Air Vehicle

which help in determining the basic transfer functions of a MAV like RAA. Following that,

we will be presenting the basic transfer functions of various control functions of RAA like

turning, sweeping, pitch and throttle so that various stability characteristics of the vehicle can

be thoroughly studied and subsequently help designing a better model for the vehicle. After

studying the basic transfer functions involved, we will be entering towards a phase for

making our RAA a semi-autonomous or an autonomous vehicle that could navigate on its

own given trajectory.

5.1 SENSORS EMPLOYED ON RAA: For formulating the basic transfer functions for our MAV, we would need to install certain

sensors on it that would directly respond to the various movements of the vehicle. They

would determine the way a vehicle behaves for example when making a turn when the

controller from the ground moves the input stick of the remote control to the right, or how the

vehicle behaves when the controller from the ground moves the input stick backwards on the


83

remote control. These sensors sense the vehicle’s movement and relay the transmitted

information to the Ground Control Station that reads the data sensed and hence tells the

controllers about the various movements and responses of the aircraft. The various sensors

involved include:

1. Accelerometers;

2. Gyroscopes;

3. Magnetometer;

4. Barometer.

The various kinds of each of the above sensors available in the market today are briefly

described below and then the most feasible of all of them used for the mounting on our RAA

is proposed.

5.1.1 GYROSCOPES: First of all, we will start off with what actually gyroscopes do in terms of functions related to

aerodynamics, and then move on to their various types and our preferable selection.

A gyroscope is a device for measuring or maintaining orientation, based on the principles of

conservation of angular momentum.[1] A mechanical gyroscope is essentially a spinning

wheel or disk whose axle is free to take any orientation. This orientation changes much less

in response to a given external torque than it would without the large angular momentum

associated with the gyroscope's high rate of spin. Since external torque is minimized by

mounting the device in gimbals, its orientation remains nearly fixed, regardless of any

motion of the platform on which it is mounted. Solid state devices also exist, such as the ring

laser gyroscope.

Applications of gyroscopes include navigation (INS) when magnetic compasses do not work

(as in the Hubble telescope) or are not precise enough (as in ICBMs) or for the stabilization

of flying vehicles like radio-controlled helicopters or UAVs. [29]

In order to discuss MEMS gyroscopes we must first understand gyroscopes in general and

what role they play in science. Technically, a gyroscope is any device that can measure

angular velocity. As early as the 1700.s, spinning devices were being used for sea navigation

in foggy conditions. The more traditional spinning gyroscope was invented in the early


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1800.s, and the French scientist Jean Bernard Leon Foucault joined the term gyroscope in

1852. In the late 1800.s and early 1900.s gyroscopes were patented for use on ships. Around

1916, the gyroscope found use in aircraft where it is still commonly used today. Throughout

the 20th century improvements were made on the spinning gyroscope. In the 1960.s, optical

gyroscopes using lasers were first

introduced and soon found commercial success in aeronautics and military applications. [30]

This is the prime purpose of us using the gyroscope as the means of measuring pitch, roll and

yaw attitude angles in our Micro Air Vehicle.

5.1.1.1 TYPES OF GYROSCOPES IN MARKET TODAY: 1. Traditional Gimbaled Types Gyroscope:

Gimbaled systems have a platform in the device that is mounted in gimbals. This device has

2 or more mechanical gyroscopes (not likely there are more than 3) that keep this platform

level. It does not need accurate gyroscope orientation sensing, they only need mechanical

gyroscopes to keep a platform level -- a much less demanding task for the gyroscopes.

However, they do have their disadvantages. One is reliability -- the spinning gyros and

gimbals all move, so you have wear and tear and they can fail or lose their accuracy.

However, a more prominant issue is with "gimbal lock." This is when two of those three

gimbals align. Since they both function about one axis, and the other only does one axis too,

you only have two axes -- any rotation about the last axis cannot occur, so the platform is

swung and misaligned with rotation about that axis. There are two main solutions: navigate

around it or a fourth gimbal. On Apollo 11 there were only three gimbals, so they planned

their maneuvers around gimbal lock. Their computers told them where not to go to keep two

gimbals from aligning. The second, but more complex, solution is a fourth gimbal. This

gimbal is motorized to keep it always oriented away from the other gimbals, so you keep

three independent axes at all times. However, this is mechanically more complex. The fourth-

gimbal system is used more in more recent gimballed inertial navigation systems. [31]


85

2. Strapdown Gyroscopes:

Strapdown systems have all their sensors mounted on a platform that changes orientation like

the plane. Instead of mechanical gyros to hold it level, it has three more accurate gyros that

sense the orientation of the system. Additionally, it has the same three acceleration sensors.

Whereas the gimballed system just senses the orientation of the platform to get the aircraft's

attitude, the strapdown systems have three gyroscopes that sense the rate of roll, pitch, and

yaw. It integrates them to get the orientation, then calculates the acceleration in each of the

same axes as the gimballed system.

Due to the sensing of the rate of rotation, rather than just holding a platform level, very

accurate and sensitive gyroscopes are needed. There are several sensitive gyroscope types

available, but the most commonly used on is the ring laser gyro. This basically is a circular

path (actually, usually triangular) that laser light travels around. This light goes both ways.

When it is rotated, one direction appears to go faster than the other -- and when they get

around and meet, they interfere. This creates a pattern that is picked up by sensors. These

gyroscopes are so accurate the alignment for strapdown systems consists of finding true north

by sensing the earth's rotation.

Strapdown systems have fewer moving parts, so they are more reliable and simpler than

other systems. However, they do need more accurate gyroscopes and better computers, so

they are a more recent development.

5.1.1.2 Working of Gyroscopes: 1. How a Gimbaled Gyroscope Works:

Figure 5.1: Working of Gimbaled Gyroscope[31]


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Instead of a complete rim, four point masses, A, B, C, D, represent the areas of the rim that

are most important in visualizing how a gyro works. The bottom axis is held stationary but

can pivot in all directions. When a tilting force is applied to the top axis, point A is sent in an

upward direction and C goes in a downward direction. FIG 1. Since this gyro is rotating in a

clockwise direction, point A will be where point B was when the gyro has rotated 90 degrees.

The same goes for point C and D. Point A is still traveling in the upward direction when it is

at the 90 degrees position in FIG 2, and point C will be traveling in the downward direction.

The combined motion of A and C cause the axis to rotate in the "precession plane" to the

right FIG 2. This is called precession. A gyro's axis will move at a right angle to a rotating

motion. In this case to the right. If the gyro were rotating counterclockwise, the axis would

move in the precession plane to the left. If in the clockwise example the tilting force was a

pull instead of a push, the precession would be to the left. When the gyro has rotated another

90 degrees FIG 3, point C is where point A was when the tilting force was first applied. The

downward motion of point C is now countered by the tilting force and the axis does not

rotate in the "tilting force" plane. The more the tilting force pushes the axis, the more the rim

on the other side pushes the axis back when the rim revolves around 180 degrees. Actually,

the axis will rotate in the tilting force plane in this example. The axis will rotate because

some of the energy in the upward and downward motion of A and C is used up in causing the

axis to rotate in the precession plane. Then when points A and C finally make it around to the

opposite sides, the tilting force (being constant) is more than the upward and downward

counter acting forces. The property of precession of a gyroscope is used to keep monorail

trains straight up and down as it turns corners. A hydraulic cylinder pushes or pulls, as

needed, on one axis of a heavy gyro. Sometimes precession is unwanted so two counter

rotating gyros on the same axis are used. Also a gimbal can be used.


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Figure 5.2: Gimbaled Gyroscope Functioning[31]

The property of precession represents a natural movement for rotating bodies, where the

rotating body doesn’t have a confined axis in any plane. A more interesting example of

gyroscopic effect is when the axis is confined in one plane by a gimbal. Gyroscopes, when

gimbaled, only resist a tilting change in their axis. The axis does move a certain amount with

a given force. Here I attempt to show how much the axis will rotate around a gimbaled axis.

That is to say, how fast it rotates in the direction of a tilting force. In figure 4, the precession

plane in the gimbaled example functions differently than in the above example of figures 1-3,

and it has been renamed as "stop the tilting force plane". The point masses at the rim are the

only mass of the gyro system that is considered. The mass and gyroscope effect of the axis is

ignored. At first consider only ½ of the rim, the left half. The point masses inside the "stop

the tilting force plane" share half their mass on either side of the plane, and add their

combined, 1/4kg, mass to point mass A of 1/2kg. So then the total mass on the left side is ½

the total mass of all 4 point masses, or 1kg. The tilting force will change the position of point

mass B and D very little and change the position of point mass A the most. So we must use

the average distance from the axis of all the mass on the left-hand side.


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Figure 5.3: Gimbaled Gyroscope, tilting plane[31]

The mass on the left side is 1kg. The average distance the mass is from the "stop the tilting

force" plane is 1/2 meter. Figure 5 shows a profile of the average mass in the tilting plane and

the average distance from the axis that the mass is situated. We are concerned at how far the

mass at the average distance will rotate within the tilting plane when a given force is applied

to the axis in the direction indicated. Point mass A is rotating at 5 revolutions per second.

This means that it is exposed to the tilting force for only .1 seconds. The tilting force of 1

Newton, if applied for .1 second, will cause the mass at the average distance to move .005

meter in an arc, in the tilting force plane. Since the length of the axis is twice as long as the

average distance of the rim’s mass, the axis will move .01 meter in an arc. At the end of .1

second the point mass will be in the "stop the tilting force plane" and all the energy

transferred to point mass A is lost in the physical restraint of the gimbal bearings. The same

thing happens when point mass A is on the right side of figure 4. Only now, the tilting force

will move point mass A down, and the axis will advance another .01meter. .01 meter every .1

second is not the whole story because the mass on the right side of the gyro hasn’t been

considered. The right side has the same mass as the left and has the same effect on the axis as

the left side does. So the axis will advance half as much, half of .01 meter, or .005meters.

Both halves of the rim mass will pass through the stop the tilting force plane 10 times in one

second. Each time a half of the rim passes though the "stop the tilting force plane", it losses

all its momentum that was added by the tilting force. The mass has to undergo acceleration

again so we continually calculate the effect that 1 Newton has for .1 second on the rim mass

at the average distance, 10 times a second. So then; at the point that the 1 Newton force is

applied, the axis will move 5cm per second along an arc. The gyro will rotate at .48 RPM

within the tilting force plane. [32]


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5.1.1.2 MEMS Gyroscopes:

These are the types of gyroscopes that are currently in use around the world and that are what

we propose to be used for making the RAA vehicle an autonomous Micro Air Vehicle for the

replacement of the conventional Gimbaled Gyroscopes.

In the last ten to fifteen years, MEMS gyroscopes have been introduced and advancements

have been made to create mass-produced successful products with several advantages over

traditional macro-scale devices.

There are basically four types of MEMS Gyroscopes in use today, namely;

1. Tuning Fork Gyroscope;

2. Vibrating Ring Gyroscope;

3. Macro Laser Ring Gyroscope;

4. Piezoelectric Plate Gyroscope.

1. Draper Tuning fork Gyroscope:

One of the most widely used micro-machined gyroscopes is the tuning fork design from the

Charles Stark Draper Lab (Fig 1). The design consists of two tines connected to a junction

bar which resonate at certain amplitude. When the tines rotate, Coriolis force causes a force

perpendicular to the tines of the fork. The force is then detected as

bending of the tuning fork or a torsional force (Fig 2). These forces are proportional to the

applied angular rate, from which the displacements can be


90

Figure 5.4: Tuning Fork Physics [30]

Since the development of their first tuning fork gyroscope in 1993, the Draper Laboratory

has made significant improvements to the device. Their first gyroscope was developed for

the automobile industry. The gyroscope had command of 1 degree/hr drift, and possessed

4000 deg/hr resolution. [31] These devices eventually functioned as the yaw rate sensor for

skid control in anti-lock braking applications. Tests run on these sensors involve the

examining the change in bias and error of such over a number of variables. Proper data could

be retrieved in 0.8 s and sent to the necessary actuator to cause proper breaking in due time.

These systems need to operate in a range of temperatures, specifically from -40 to 80 degrees

Celsius. Over this range, both the bias error and the scale factor error are both quite stable.

The bias error is approximately 2200 deg/h. Scale factor error was approximately 0.08%.

2. Piezoelectric Plate Gyroscope:

While vibrating ring gyroscopes and tuning fork gyroscopes were the first successful MEMS

gyroscopes and are still the most widely produced, other successful MEMS gyroscopes have

since been created. One of these gyroscopes is the Piezoelectric Plate Gyroscope which uses

a PZT plate as its base. This method, which in the past has been used to try to build macro-

scale gyroscopes, is actually ideal for micro devices. At micro levels, an entire plate can be

made of piezoelectric material. It has


91

advantages over the common vibrating gyroscopes in that it requires a much smaller drive

voltage to create readable outputs.

The piezoelectric plate gyroscope is very simple in its design. In fact, it is much simpler than

the ring or fork gyroscopes. There is a piezoelectric plate, which has a length and width much

larger than its depth. The plate has electrical leads connected to all 6 sides and sits on top of a

thin membrane of a cavity in a silicon wafer. The cavity allows more freedom for the PZT to

vibrate and deform. The leads provide the driving voltage and measure the output.

Figure 5.5: Cross Section of Gyroscope [33]

Like other MEMS gyroscope the piezoelectric plate gyroscope works on the principle of a

vibrating body. In this case, the vibrating body is a piezoelectric sheet. The sheet does not

vibrate like a plate or fork. Instead the thickness vibrates which oscillates with time. This

requires an AC driving voltage applied vertically across the plate, which uses the electro-

mechanical properties of the PZT to create the vibration. Any piezoelectric material can be

used, but PZT has high piezoelectric constants, and can be added at a precise thickness.

Figure 5.6: Piezoelectric Plate with orthogonal directions shown[34]

The piezoelectric plate gyroscope is a feasible alternative to traditional MEMS gyroscopes.

One of its advantages is a lower required drive voltage. However, the sensitivity is only

about 38 microvolts, whereas the sensitivity of a ring gyroscope is around 200 microvolts. [34]

Also, when there is no rotation, traditional gyroscopes come


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much closer to the ideal zero volts output than the piezoelectric plate gyroscope, which still

outputs up to 100 millivolts. A major advantage and the one that could prove most practical

is the versatility of the piezoelectric plate gyroscope. It can measure rotation in two

directions. In addition, if the driving voltage direction is switched, the same device can

measure rotation in the third direction, although with much less sensitivity. Since this device

is easily incorporated into other IC chips, it could be controlled to do more things than a ring

or tuning fork gyroscope, which require three gyroscopes to measure three rotation

directions.

3. Laser Ring Gyroscopes:

In order to discuss the difficulties in creating a laser ring gyroscope in the micrometer

scale, the theory behind the macroscopic gyroscope must be derived. Below is a picture of a

simple laser gyroscope that happens to be in the shape of a triangle rather than a ring. A laser

source outputs two beams traveling in an opposite direction around the ring until they reach

the detector. The detector counts the beat frequency of the combined light wave. This beat

frequency is directly proportional to the angle of rotation of the gyroscope.

Figure 5.7: General Ring Laser Gyro[35]

There are two main sources of error for laser ring gyroscopes. They are varying offset bias

and a dead band at very small rotation rates. The offset bias is due to different indices of

refraction for the beam pairs. This is caused by small differences in the degrees of saturation

in the original beams. [35] Diddams, Atherton and Diels experimented with a light scatterer


93

placed in the laser pulses at different distances from the detector while keeping the gyroscope

at rest. This scatterer represented a reflection of light particles back with the beam traveling

the other direction. They found that when the scatterer was more than 500 microns away

from the detector, the beat frequency was constant and stable. The width of the dead band

also showed good consistency through many tests. When the scatterer was within 100

microns, the beat frequency became non sinusoidal and therefore very hard to measure.

When the scatterer was placed within 10-30 microns of the detector, the beat frequency was

erratic and noncontinuous.[35]

Figure 5.8: Beat Frequency vs. Scatter Position[35]

The dead band region is another limiting factor for this type of sensor. When you are at very

small turning rates, the frequencies of the two light waves are very close to each other. When

these frequencies are within a critical value, it creates a phenomenon

where the frequencies converge toward each other until they are the same. This gives you a

false reading of a zero turning speed when you are actually moving at a small angular

velocity.


94

Figure 5.9: Rotating Rate found by using Beat Frequency[35]

Micro-Laser Gyro

The miniaturization of a macrolaser ring gyroscope poses many problems. The largest

problem is with scaling the device down to the micrometer level. The following equation is

the equation relating the beat frequency to the angular velocity of the gyroscope. [35]

[30]

4. Vibrating Ring Gyroscope: An extremely simple gyroscope for use in a strapdown system having a vibrating ring as the

gyroscopic element. The ring is supported and vibrated by an electrostatic field, established

by a voltage supplied to case-fixed electrodes. An additional pair of fixed electrodes senses

the position of the nodes of vibration of the ring and an eddy current drive forces rotation of

the ring to move the nodes back to a reference position. Optical sensing means provide the

pick-off, sensing the direction and amount of rotation required to move the ring back to the

nodal reference position.

This invention relates to gyroscopes and more particularly to gyroscopes where a vibrating

ring is the inertial reference element. In more gyroscopes a gyro wheel is rotated at a high

rate of speed to provide an angular momentum vector along the axis of the wheel. The wheel

will maintain a constant direction in space, and thus can be used as an angular spatial

reference, provided that the wheel can be supported and driven by means that have

substantially zero friction and error torques. In practice this has proved to be very difficult to


95

do the necessary accuracy, leading to expensive instruments with limited life and reliability.

Furthermore, many of the techniques in current use, such as enclosing the wheel in a

container which is then floated in a fluid of the same average density, limit the maximum

angle between the wheel and its supports to some small value. This is impractical for a

strapdown system in which unlimited angular inputs are required.

These problems may be avoided by using a vibrating ring in place of the rotating wheel. It

can be mathematically proven that if a ring is set into vibration in the out-of-round mode and

then rotated through some angle, the vibration pattern as indicated by the nodes will rotate

exactly four-tenths of this angle. This effect may be used to measure the rotation of the

vehicle by the following method: the ring is supported by means to be described below so

that it is free to rotate about its axis of symmetry. It is set into oscillation and the position of

the nodes is determined by a sensor attached to the vehicle. If the vehicle now rotates, the

sensor will move away from the nodes. The ring is now forced to rotate so as to carry the

nodes back to the sensor. The rotation of the ring necessary to accomplish this is a direct

measure of the vehicle rotation and may be determined by sensing the angle between the ring

and the vehicle. There are a number of problems associated with the support, rotation, and

readout means which must be overcome to provide a practical embodiment.

One of these problems is to provide means for supporting the ring in a manner which will

prevent unacceptable drift. Previously disclosed support means include thin-walled

cylindrical or hemispherical shells upon which the ring is supported. A second means

includes a plurality of support rods extending from a stationary base and connected to the

ring by thin, necked-down portions which constitute a spring. A third means, utilized in the

preferred embodiment of the present invention, involves the use of an electrostatic or electro-

magnetic field to support the ring.

Another necessary feature of such constructions is the means for maintaining the ring in

sustained vibration. Previously suggested means include a radial magnetic or electrostatic

field exerting a force on the ring as a function of its displacement from the nominal circular

condition. While this has the advantage of great symmetry, the actual power deliverable to

the ring will be extremely small, requiring the use of very high magnetic or electrostatic

fields. Another means, preferred for use in the present invention, employs an electrostatic

force supplied from electrodes arranged in two or more pairs of diametrically opposed


96

segments. The local velocity of the ring under each segment is measured and a voltage

applied to the opposite electrode segment to produce a force in phase with the velocity. All

the energy required may be provided without unduly high fields since the effect does not

depend on field non-linearities.

The third problem associated with the use of vibrating rings as gyroscopic elements is the

pick-off or sensing means. A previously suggested mode of operation for performing the

sensing involves fixing the ring to the vehicle and measuring the angle between the

instantaneous nodal position and the vehicle axis by means of a set of variable capacitance or

variable inductance pick-offs. However, to be of use in a navigation system, the angular

accuracy must be of the order of ten seconds of arc, and this accuracy must be maintained

over the full 360° of rotation. This implies a pick-off accuracy and resolution of the order of

50 parts per million. Providing a pick-off means with this degree of accuracy approaches or

exceeds the present state of the art. Thus, the use of pick-offs with this mode of operation has

severely limited the prospects of constructing a practical vibrating ring gyroscope. [36]

5.1.2 Gyroscope Suitable For Our RAA: Although previously companies used macroscale gyroscopes, specifically, the ring laser

gyroscope. This gyroscope was implemented in the HG1700 system which could function

properly and beyond the original requirements for which it was planned. However, this

device was too costly, too large, and too high performance for the emerging smaller, gun-

launched systems. By acquiring Draper.s tuning fork gyroscope, a new inertial measurement

unit could be developed that was cheaper and smaller. These systems functioned similarly to

their predecessors, with the exception that they could be made smaller at an economical cost. [30]

Thus Tuning Fork Gyroscope is the most suitable MEMS GYROSCOPE recommended for

implementation in our RAA.

5.2 ACCELOROMETERS: An accelerometer is an electromechanical device that will measure acceleration forces. These

forces may be static, like the constant force of gravity pulling at your feet, or they could be

dynamic - caused by moving or vibrating the accelerometer.

Before going onto accelerometers, let us discuss the basics of acceleration very briefly;


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5.2.1 What is Acceleration?

Definition: the time rate of change of velocity–A.K.A.: the time rate of change of the time

rate of change of distance What are the units?–Acceleration is measured in (ft/s)/s or

(m/s)/s•What is a “g”?–A “g” is a unit of acceleration equal to Earth’s gravity at sea

level•32.2 ft/s2 or 9.81 m/s2

5.2.2 What are some “g” reference points?

Description“g” level [33]

Earth’s gravity1g

Passenger car in corner2g

Bumps in road2g

Indy car driver in corner3g

Bobsled rider in corner5g

Human unconsciousness7g

Space shuttle10g

5.2.3 What are accelerometers useful for?

By measuring the amount of static acceleration due to gravity, we can find out the angle the

device is tilted at with respect to the earth and also sensing the amount of dynamic

acceleration, you can analyze the way the device is moving.

Measuring tilt and acceleration doesn't seem all that exciting. However, engineers have come

up with many ways to make really useful products using them. In the computing world, IBM

and Apple have recently started using accelerometers in their laptops to protect hard drives

from damage. If you accidentally drop the laptop, the accelerometer detects the sudden

freefall, and switches the hard drive off so the heads don't crash on the platters. In a similar

fashion, high g accelerometers are the industry standard way of detecting car crashes and

deploying airbags at just the right time.


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Figure 5.10 3 axis accelerometer board[37]

5.2.4 How do accelerometers work? There are many different ways to make an accelerometer! Some accelerometers use the

piezoelectric effect - they contain microscopic crystal structures that get stressed by

accelerative forces, which cause a voltage to be generated. Another way to do it is by sensing

changes in capacitance. If you have two microstructures next to each other, they have a

certain capacitance between them. If an accelerative force moves one of the structures, then

the capacitance will change. Add some circuitry to convert from capacitance to voltage, and

you will get an accelerometer. There are even more methods, including use of the

piezoresistive effect, hot air bubbles, and light.

5.2.5 What things should One consider when buying an accelerometer? Analog vs digital - First and foremost, you must choose between an accelerometer with

analog outputs or digital outputs. This will be determined by the hardware that you are

interfacing the accelerometer with. Analog style accelerometers output a continuous voltage

that is proportional to acceleration. E.g. 2.5V for 0g, 2.6V for 0.5g, 2.7V for 1g. Digital

accelerometers usually use pulse width modulation (PWM) for their output. This means there

will be a square wave of a certain frequency, and the amount of time the voltage is high will

be proportional to the amount of acceleration.

If you are using a BASIC Stamp, or any other microcontroller with purely digital inputs, you

will most likely need to go for a digital output accelerometer. The disadvantage here is that it

requires you to use the timing resources of the microcontroller to measure the duty cycle, as


99

well as performing a computationally intensive division operation.

If you are using a PIC/AVR/OOPIC/Javelin with analog inputs, or a completely analog based

circuit, analog is almost always the best way to go. Depending on the compiler, measuring

analog acceleration can be as simple as acceleration=read_adc(); and can be done in a few

microseconds.

Number of axis - For most projects, two is enough. However, if you want to attempt 3d

positioning, you will need a 3 axis accelerometer, or two 2 axis ones mounted at right angles.

Maximum swing - If you only care about measuring tilt using earth's gravity, a ±1.5g

accelerometer will be more than enough. If you are going to use the accelerometer to

measure the motion of a car, plane or robot, ±2g should give you enough headroom to work

with. For a project that experiences very sudden starts or stops, you will need one that can

handle ±5g or more.

Sensitivity - Generally speaking, the more sensitivity the better. This means that for a given

change in acceleration, there will be a larger change in signal. Since larger signal changes are

easier to measure, you will get more accurate readings.

Bandwidth - This means the amount of times per second you can take a reliable acceleration

reading. For slow moving tilt sensing applications, a bandwidth of 50Hz will probably

suffice. If you intend to do vibration measurement, or control a fast moving machine, you

will want a bandwidth of several hundred Hz. [38]

Accelerometer Sensor Terminology:

Sensitivity:A measure of how much the output of a sensor changes as the input acceleration

changes. Measured in Volts/g

•Vcc:The voltage supplied to the input of the sensor

–5.000 ±0.005V for CAS device


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•%Vcc:Readings are often represented as a % of the supply voltage. This allows for

correction due to supply voltage variances between readings. [33]

There are different types of accelerometers available that will measure a variety of things

useful for many electronic projects. These include: Acceleration; Tilt; Rotation; Gravity;

Vibration & Incline. Each of these measurements are the result of different characteristics of

the device. The make-up and components differ wildly between accelerometers.

1. Traditional Accelerometers;

2. MEMS BASED accelerometers:

1. Piezoresitive;

2. Piezoelectric;

3. Capacitive.

4. Hall Effect;

5. Magnetoresistive;

6. Heat Transfer.

5.2.6 Why MEMS Based? The field of MEMS accelerometers can be segregated into two dominant microsystem

architectures: capacitive and piezoresistive. Although both types of accelerometers employ

internal proof masses that are excited by acceleration, the architectural differences are in the

transduction mechanism used to correlate the movement of the internal proof mass to

acceleration. Capacitive accelerometers employ a differential capacitor whose balance is

disrupted by the movement of the proof mass. Piezoresistive accelerometers generally rely on

strain induced within a flexural element that attaches the proof mass to the sensor housing for

identification of the mass movement. Capacitive-based MEMS accelerometers, such as the

ADXL iMEMS series ~Analog Devices, Norwood, Mass.! have enjoyed more commercial

success than piezoresistive designs. This is a direct result of piezoresistive accelerometers

having not been capable of keeping pace with the reduced fabrication costs associated with


101

capacitive architectures ~Maluf 2000!. Problems associated with temperature coefficients

and drift properties of piezoresistive materials have necessitated careful packaging and

compensation circuitry that have added to piezoresistive accelerometer costs.

The high costs associated with commercial monitoring systems can be eradicated through the

adoption of new and revolutionary technologies from related engineering fields. For

example,

the adoption of MEMS sensors could reduce overall system costs.

Various MEMS-based accelerometers are commercially available that can serve as accurate

substitutes for traditional force-balance accelerometers at substantially reduced costs.

Accelerometers fabricated by a MEMS process are mechanically similar to traditional

accelerometers but only fabricated on a micrometer scale. An additional advantage of MEMS

sensors is their ability to monolithically fabricate signal conditioning circuitry on the same

die, resulting in improved sensor performance and reduced sensor costs ~Judy 2001!. [39]

1. Piezoresitive Acceleormetrs:

The piezoresistive effect in semiconductors is manipulated for sensor devices, also known as

strain gauge accelerometers, used for measurement of mechanical stress. They utilize

different semiconductor materials to convert mechanical strain to a DC output; examples

include: germanium, polycrystalline silicon, amorphous silicon, and single crystal silicon.

The gauges are mounted to a spring or between the seismic mass and the containing

framework, applied stress generated by accelerations correlates with the DC Voltage output;

allowing for useful measurement. Interestingly these accelerometers also give information on

Steady-State accelerations. [34]

In a similar version, a spring element is made of metal or ceramic material, and

semiconductor strips are bonded on the deforming surfaces of the element, similar to the foil

or wire gages. The deflection of the spring element under the acceleration causes deformation

of the gages, producing an electrical output.


102

Figure 5.11: Piezoresistive accelerometers with diffused (a) and cemented (b) semiconductor

strain gages, a = acceleration, 1 = seismic mass, 2 = semiconductor beam, 3 = ceramic or

metal beam, 4, 5, 6, and 7 = diffused semiconductors, 8, 9, 10, and 11 = cemented

semiconductor strips, 12 = case. [40]

The piezoresistive accelerometer possesses superior performance characteristics including

low noise densities when measuring local structural responses characterized by high-

frequency content. Advances made in MEMS fabrication processes require a revisit to the

piezoresistive accelerometer. With new processes available and old processes improved, a

low-cost, high performance piezoresistive accelerometer is now possible. In particular, deep

reactive ion etching ~DRIE! fabrication techniques could be used in the fabrication of

MEMS accelerometers. The low-noise property of piezoresistive accelerometers at high

frequencies,

compared to those of capacitive accelerometers, is additional motivation for a renewed

interest in the piezoresistive.

Figure 3.12: (a) Design illustration; (b) physical dimensions and; (c) SEM image of planar

piezoresistive accelerometer[39]


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2. Piezoelectric Accelerometer:

The piezoelectric effect is a quality of some materials where they produce electricity when

force is applied by the contained seismic-mass. [34]

Figure 5.12: Piezoelectric Accelerometer is designed for flight vehicles. [41]

5.3 MAGNETOMETER: A magnetometer is a scientific instrument used to measure the strength and/or direction of

the magnetic field in the vicinity of the instrument.

Magnetometers can be divided into two basic types:

Scalar magnetometers measure the total strength of the magnetic field to which they

are subjected, and

Vector magnetometers have the capability to measure the component of the

magnetic field in a particular direction, relative to the spatial orientation of the device.

The use of three orthogonal vector magnetometers allows the magnetic field strength,

inclination and declination to be uniquely defined. Examples of vector magnetometers are

fluxgates, superconducting quantum interference devices (SQUIDs), and the atomic SERF

magnetometer. A magnetograph is a special magnetometer that continuously records data.


104

Figure 5.13: Magnetometer[42]

5.4 Inertial Navigation System (INS): An inertial navigation system (INS) is a navigation aid that uses a computer, motion

sensors (accelerometers) and rotation sensors (gyroscopes) to continuously calculate via dead

reckoning the position, orientation, and velocity (direction and speed of movement) of a

moving object without the need for external references. It is used on vehicles such as ships,

aircraft, submarines, guided missiles, and spacecraft. Other terms used to refer to inertial

navigation systems or closely related devices include inertial guidance system, inertial

reference platform, inertial instrument, and many other variations. An inertial navigation

system includes at least a computer and a platform or module containing accelerometers,

gyroscopes, or other motion-sensing devices. The INS is initially provided with its position

and velocity from another source (a human operator, a GPS satellite receiver, etc.), and

thereafter computes its own updated position and velocity by integrating information

received from the motion sensors. The advantage of an INS is that it requires no external


105

references in order to determine its position, orientation, or velocity once it has been

initialized.

INSs have angular and linear accelerometers (for changes in position); some include a

gyroscopic element (for maintaining an absolute angular reference). Angular accelerometers

measure how the vehicle is rotating in space. Generally, there's at least one sensor for each of

the three axes: pitch (nose up and down), yaw (nose left and right) and roll (clockwise or

counter-clockwise from the cockpit).

Linear accelerometers measure non-gravitational accelerations[31] of the vehicle. Since it can

move in three axes (up & down, left & right, forward & back), there is a linear accelerometer

for each axis. A computer continually calculates the vehicle's current position. First, for each

of the six degrees of freedom (x,y,z and θx, θy and θz), it integrates over time the sensed

amount of acceleration, together with an estimate of gravity, to calculate the current velocity.

Then it integrates the velocity to figure the current position.

Inertial guidance is difficult without computers. The desire to use inertial guidance in the

Minuteman missile and Project Apollo drove early attempts to miniaturize computers.

Inertial guidance systems are now usually combined with satellite navigation systems

through a digital filtering system. The inertial system provides short term data, while the

satellite system corrects accumulated errors of the inertial system.

An inertial guidance system that will operate near the surface of the earth must incorporate

Schuler tuning so that its platform will continue pointing towards the center of the earth as a

vehicle moves from place to place.

Figure 5.14: INS

CH 06. THE FABRICATION PHASE

106

CHAPTER 6

THE FABRICATION PHASE After the complete design process, we thoroughly searched the market for possible parts and

equipments that met our requirement and suited to our goals in creating a Sweeping Wing

RAA. After going through the available products in the market, we weighed each of the

separate servo motors and motors needed for or RAA. After having an approximate weight of

each of the separate components, we got an approximate idea of the strength of the material

needed for the fuselage to hold these materials inside while keeping the weight to the

minimum. After choosing the perfect material for our RAA, we chose the right motor,

propeller and battery to provide the required amount of thrust for our MAV. Following this,

we chose the correct ground control station that could remotely control our RAA from a

distance. Finally we selected the right actuators for our vehicle.

6.1 The Right Material: For selecting the right material, we needed a material that had the highest strength to weight

ratio suitable for our aircraft. We required minimum weight so as to acquire ample lift and

attain maximum stability.

There were actually two options available for us to build our MAV from- the conventional

Balsa Wood and the Jumbo lone Foam.

6.1.1 Balsa Wood: Balsa lumber is very soft and light with a coarse, open grain. The density of dry balsa wood

ranges from 40–340 kg/m³ (2.5-21 lb/ft³), with a typical density of about 160 kg/m³ (10

lb/ft³).The light weight of the wood derives from the fact that the tree has large cells that

contain water. As it is low-density but high in strength, balsa is a very popular material to use

when making light, stiff structures in model bridge tests and for the construction of model

aircraft as well as full-sized light wooden aero planes, most famously the World War II de

Havilland Mosquito. [43]

Balsa Wood has been used for the main fuselage body and plus as a strengthening material.


107

(a)

(b)

Figure 6.1: (a), (b) Fuselage of Balsa Wood


108

6.1.2 Jumbo lone Foam: After an exhaustive survey of the market, we came across a very new material that was most

suitable for our small sized MAV. It is the Jumbo lone Foam. It is easily available in

Pakistan, with an added advantage that it is cheaper than Balsa Wood. Moreover it had lesser

density than the Balsa Wood.

TECHNICAL PROPERTIES [44]

ITEM UNIT RESULT TESTING METHOD

Color __ Pink __

Cell structure __ Closed very

fine

__

Density (Check

Density Table)

kg/m3 32 ~ 40 BS4370: Method 2

Thermal

conductivity

BTU in/ft2.hr.°F

W/mk

0.19

0.026

ASTM C-518

KS M 3808-’05

Compressive

strength

kN/m2

(KPA)

300 ~ 400 BS4370: Method 3

KS M 3808-'05

Bending Strength N/cm2 52 KS M 3808-’05

Coefficient of

linear thermal

expansion

mm/mk 0.07 BS4370: Method 13

Temperature limits °C -50+75 __

Fire classification __ A BS3837: Part 1 1986

Size: Length

Width

mm

mm

1250 - 2500

600 - 900

__

Thickness mm 20 ~ 75 __

Edge profile __ Butt Ship lap

Tongue &

groove

__

Surface __ Planed

Shaved

__


109

Grooved

Table 6.1: Jumbo lone foam Technical Properties

We used Jumbo lone foam to manufacture the Feathers of our MAV and the Fixed Part of the

Wing. They were strengthened on top by a thin layer of Vinier sheet, providing them with the

required strength along with needed lightness in weight.

Figure 6.2: Jumbo lone Foam[45]

Figure 6.3: RAA wing, Arm-wing of jumbo lone foam


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6.2 The Propulsion System:

To choose the correct propulsion system, after several design iterations through CFD

analysis, we found the required thrust for our RAA to be approximately 0.1 Newtons or more

conventionally said as 100 grams. For attaining the required thrust we needed the correct

motor to power the propeller, the correct propeller size and the correct battery to power the

propeller. Thus three factors influenced the required thrust – Right Motor, Right Propeller

and the Right Battery.

6.2.1 The Right Motor:

For the right motor we chose the Brushless DC Motor.

Why did we choose a Brushless DC motor instead of brushed one?

The problems of the Brushed motor are:

The brushes eventually wear out.

Because the brushes are making/breaking connections, you get sparking and electrical

noise.

The brushes limit the maximum speed of the motor.

Having the electromagnet in the center of the motor makes it harder to cool.

The use of brushes puts a limit on how many poles the armature can have.

The advantages of Brushless DC motor are:

Because a computer controls the motor instead of mechanical brushes, it's more

precise. The computer can also factor the speed of the motor into the equation. This makes

brushless motors more efficient.

There is no sparking and much less electrical noise.

There are no brushes to wear out.

With the electromagnets on the stator, they are very easy to cool.

You can have a lot of electromagnets on the stator for more precise control.

The only disadvantage of a brushless motor is its higher initial cost, but you can often

recover that cost through the greater efficiency over the life of the motor. [46]

For our MAV we chose initially a 1000 RPM/Volt Brushless DC motor was selected .


111

After test flights, Motor was changed to 2000 RPM/Volt to provide sufficient RPM and thus

THRUST!

“Yinyan Model Tech. BL 1812” BLDC selected that was easily available Locally!

Figure 6.4: Brushless motor BL1812[47]

6.2.2 The Right Propeller:

For finding the right size of propeller that could provide us with the required thrust, we did

initial test with Propeller size of 6 inches .A prop of 6 pitch selected to provide a balance

between required Speed and Torque to run the Propeller at certain RPM.

However, after performing several test flights the propeller size was increased to 7 inches

with 6 pitch.

Figure 6.5 The propeller used in RAA


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6.2.3 The Right Battery:

For the Battery we chose a Lithium Polymer Battery that provided us with the right amount

of power needed to propel the aircraft forwards.

Lithium-ion polymer batteries, polymer lithium ion, or more commonly lithium polymer

batteries (abbreviated Li-poly, Li-Pol, LiPo, LIP, PLI or LiP) are rechargeable batteries

(secondary cell batteries). Normally batteries are composed of several identical secondary

cells in parallel addition to increase the discharge current capability. [48]

Lithium batteries are the preferred power sources for most electric modelers today. They

offer high discharge rates and a high energy storage/weight ratio. However, using them

properly and charging them correctly is no trivial task. There are many things to consider

before using lithium cells for e-flight. But none is more important than safety.

The advantages and limitations of using LiPo battery over other conventional batteries was:

Advantages

Very low profile — batteries that resemble the profile of a credit card are feasible.

Flexible form factor — manufacturers are not bound by standard cell formats. With high

volume, any reasonable size can be produced economically.

Light weight – gelled rather than liquid electrolytes enable simplified packaging, in some

cases eliminating the metal shell.

Improved safety — more resistant to overcharge; less chance for electrolyte leakage.

Limitations

Lower energy density and decreased cycle count compared to Li-ion — potential for

improvements exist. Expensive to manufacture — once mass-produced, the Li-ion polymer

has the potential for lower cost. Reduced control circuit offsets higher manufacturing

costs.[49]

Initially we selected a 2 cell 7.4 V/RPM LiPo Battery. However, after thrust considerations

from test flights, Battery size was increased to 3 Cell 11.1 V/ RPM LiPo Battery. This

provided us with a total RPM of about 22000 RPM which was more than sufficient for our

RAA aircraft MAV.


113

Figure 6.7: LiPo Battery[50]

6.3 The Ground Control Station:

We used a Six Channel remote Control for our Ground control station with each of the

channels independently operating:

Channel 1= Left Feather

Channel 2= Right Feather

Channel 3= Propeller

Channel 4= Elevator

Channel 5= Rudder

We used a 72 MHz 6 Channel Remote Control by ShenZen Wfly Tech FT06-A Transmitter


114

Figure 6.8: WFLY Remote Control used for RAA[51]

6.4 The Actuators: The Actuators used were the Plastic geared Servo motors and the Metal Geared Servo

Motors.

Plastic Geared Motors:

Two Mini Servo Plastic Geared by “ART-TECH” operate the Elevator and the Rudder with 1

kg/cm of Torque

Figure 6.9: Plastic gear mini-Servo motor[52]


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6.4.1Metal Geared Servo Motors:

Two Micro Servos are Metal Geared by “TOWER PRO Electron Co.” to operate the Two

Hand Wings independently with sufficient Torque of 2 kg/cm

Figure 6.10: Metal Gear micro-Servo motors[53]

6.5 Total Weight of our RAA:

The Total weight of our fabricated RAA came out to be approximately 250 grams that is very

light and very stable design.

6.6 Flight Tests: We performed around 7 test flights and after each test flight applied an improvement to our

model until it flew successfully! The final fabricated piece is shown in the subsequent pages.


116

6.7 The Final Fabricated RAA:

(a)

(b)


117

(c)

(d)


118

(e)

Figure 6.11: (a)Fabricated RAA, top view, (b) Fabricated RAA side view, (c) Fabricated RAA side view, (d) RAA in fields for testing, (e) RAA in fields for testing, one wing swept

CH 07. PROJECT LIMITATIONS

119

CHAPETR 7

PROJECT LIMITATIONS

Throughout the Project, because of the novelty of the MAV concept and its design in

Pakistan, we faced some limitations that either forced us to choose another less superior

alternative or hindered the progression of the project from completion in the required form.

The limitations that we faced were as follows:

1. The computational hardware limitations faced during meshing and subsequent analysis in

ANSYS. Since our aircraft analysis was complex, so the design could be improved by using

greater number of iterations than we performed.

2. Fabrication of MAV was relatively a newer concept in Pakistan, so there was a lack of

skilled personnel required to fabricate the complex and intricate geometries involved in the

RAA’s design.

3. The construction material and the electronics involved demanded a lot of money. Since we

were not funded, we could not afford buying carbon fiber and expensive equipments

involved and thus this served as a major limitation.

CH 08. FUTURE RECOMMENDATIONS OF ‘PROJECT RAA’

120

CHAPTER 8

FUTURE RECOMMENDATIONS OF ‘PROJECT RAA’

Since this MAV concept of a Morphing Wing MAV was a relatively new one, so we kicked

off with a successful start and this is ought to continue in future degrees, considering the

wide applications It can serve to Pakistan in the current scenario. Thus, we will lay down

some recommendations for those aspiring to continue this project. Those are:

1. First of all, if proper funding is available, the material for the construction of the fuselage and

the wings can be made from carbon fiber that can be imported from abroad. This would

considerably lighten the aircraft, RAA, and hence make it more stable to fly. Also, RAA

could be designed of smaller wingspans. A successful RAA team would be able to participate

in International Micro Air Vehicle Competition.

2. Again, in case of proper funding, MEMS sensors like accelerometers, gyroscopes and INS

can be installed on the fabricated MAV and the vehicles behavior during various turns and

attitudes noted down. This would help in formulating the transfer functions specific to RAA

and help better understand the dynamics involved.

3. After the aforementioned task is done, the RAA can be made an autonomous one by coding

the trajectory of the vehicle through different terrains and environments and would thus

render the vehicle completely autonomous.

4. If RAA is manufactured by carbon fiber then it will be capable to lift more payloads. Also, a

proper mechanism could be installed for ground locomotion which would make it Micro Air

and Land Vehicle.

5. Project RAA could be extended to MAVs with other morphing configurations. Projects of

min-UAVs and long distance UAVs with morphing wing should also be started. For UAVs,

Shape Memory Alloys can be used for morphing wing actuations.

REFERENCES

121

CONCLUSION

A Biologically-inspired Morphing Wing Micro Air Vehicle named ‘RAA’ has been designed

which is capable of sweeping its three-feathers-wings backward during flight. RAA of 50cm

wingspan has been fabricated with the locally available economical materials and

components controlled by Radio Control.

With no previous similar documented design available, RAA design is self-accomplished

work after studying biologists’ studies of gliding birds & aircraft literature and brain

storming. Design was divided into three modules. RAA wing is bio-inspired two section

wing; arm-wing and hand-wing. In aerodynamic design, airfoils, similar in shape to that of

birds, have been selected which are giving high lift-drag ratio. Through wing design & study

program developed using EES, wing configuration capable of lifting 250g has been selected.

Also, the program shows that drag reduces with increase in sweeping angle which will

increase endurance of flight. The CFD analysis of straight wing and 45o swept wing also

validates the drag reduction. For sweeping three-feathers-wing, four morphing mechanisms

have been conceived for RAA. After comparisons and detail study, the E-quinted mechanism

has been finalized. Detail analysis of proposed mechanism using Working Model software

shows the conformity with NASA morphing standards. A theoretical study of control

systems has also been accomplished for Autonomous flight. MEMS components have been

proposed. Fabrication of RAA has been done using local materials like Balsa wood, jumbo

lone foam. It is controlled by Radio Control of 72MHz frequency. The test flights show room

for many improvements in MAV.

Morphing wing concept has been introduced through this project. Now, ‘Project RAA’ would

be continued by the students, improving it to achieve goal of mission capable MAV.

REFERENCES

122

REFERNECES

0. John Hoffman, Johan Jansson and Claes Johnson (2008), The Secret of Flight Applied Mathematics Body and Soul Vol 6, pp. 3-9

1. Harun Yahya (2004), Design in Nature, Second Edition, Ta-Ha Publishers Ltd,

London, United Kingdom, 2004.

2. http://www.aerialcontrivanceworkshop.com/images/5-abbas-ibn-firnas-2.jpg 13-Jun-

2010

3. http://sln.fi.edu/wright/again/Images/85-18307.gif 26-Apr-2010

4. http://airliners.net 26-Apr-2010

5. http://dcb.larc.nasa.gov/DCBStaff/dlr/BioFlight_Oct02.pdf 06-Jun-2010

6. J.B. Davidson,P. Chwalowski and B.S. Lazos, “Flight Dynamic Simulation

Assessment of a Morph able Hyper-Elliptic Cambered Span Winged Configuration,”

AIAA-2003-5301, 2002.

7. W. Shyy, M. Berg and D. Ljungqvist, “Flapping and Flexible Wings For Biological

And Micro Air Vehicles,” Progress in Aerospace Sciences, Vol. 35, No. 5,1999, pp.

455-506.

8. http://www.delfly.nl 26-Apr-2010

9. Mark Whitty, Lin Chi Mak, Derek Taprell, Anselm Ma, Steven Lee, Cho Ying Ng,

Nicholas Robinson, Harry Xiao. “ Development of the MAVSTAR Micro Aerial

Vehicle and Base Station for IMAV09” School of Mechanical and Manufacturing

Engineering The University of New South Wales, Australia, 2009, pp.6-9.

10. Mujahid Abdulrahim “Dynamic Characteristics of Morphing Micro Air Vehicles”

University of Florida, 2004, pp.10-25

11. Jerry Fausz (2010), Morphing Flight: Beyond Irreducible Complexity, Volum 30,

No.3, January 2010, pp.1-2.

12. http://www.roboswift.nl/ 02-Jun-2010

REFERENCES

123

13. J. Blondeau,J. Richeson and D.J. Pines, “Design, Development and Testing Of A

Morphing Aspect Ratio Wing Using An Inflatable Telescopic Spar, ” AIAA-2003-

1718, 2003

14. N.S. Khot, F.E. Eastep, and R.M. Kolonay, “Method for Enhancement of the

Rolling Maneuver of a Flexible Wing,” Journal of Aircraft, Vol. 34, No. 5, 1997, pp.

673-678

15. Rick Lind, Mujahid Abdulrahim, Kenneth Boothe and Peter Ifju, Department of

Mechanical and Aerospace Engineering University of Florida, “Morphing for Flight

Control of Micro Air Vehicles” pp. 2-3.

16. ]http://3.bp.blogspot.com/_RDHIfo3Wq9Q/S3BPfYnXVTI/AAAAAAAAAiU/eKek

l-9EWaQ/s1600-h/Fastest+Bird+in+the+World+Spine+tailed+swift+171+KMPH-

788015.jpg 29-May-2010

17. D. Lentink, U. K. Muller, E. J. Stamhuis, R. de Kat, W. van Gestel, L. L. M.

Veldhuis, P. Henningsson, A. Hedenstrom, J. J. Videler & J. L. van Leeuwen, “

How swifts control their glide performance with morphing wings” Vol 446, 26 April

2007, pp. 1082-1083.

18. J. J. Videler, et al. “Leading-Edge Vortex Lifts Swifts” Vol 306, 2004, pp. 1960-

1962

19. J. J. Videler, et al. “Leading-Edge Vortex Lifts Swifts” Vol 306, 2004, pp. 1960-

1962

20. Vance A. Tucker and G. Christian Parrott. “Aerodynamics of gliding flight in a

Falcon and other birds” Department of Zoology, Duke University, Durham, North

Carolina, Vol 52,1969 pp. 345-350.

21. Adrian L.R. Thomas “The Flight of Birds that have Wings and a Tail: Variable

Geometry Expands the Envelope of Flight Performance” University of Lund\

Department of Animal Ecology, 1996, pp. 237-245.

22. http://www.webster.com 10-May-2010

23. Thomas J. Muller (2004), Fixed and Flapping Wing Aerodynamics for Micro Air

Vehicle Applications, American Institute of Aeronautics and Astronautics, Virginia,

2004 24. Airfoil database: http://worldofkrauss.com/ 30-March-2010

REFERENCES

124

25. http://www.pilotfriend.com/training/flight_training/aero/lift 26-Apr-2010

26. Daniel P. Raymer. “Aircraft Design : A Conceptual Approach” First Edition,

American Institute of Aeronautics and Astronautics Publishing Company,

Washington, 2002, pp. 257-259

27. http://en.wikipedia.org/wiki/Gear 20-May-2010

28. http://www.coeshow.com/shop/index.php?act=viewProd&productId=6 26-May-2010

29. B. Sanders, F.E. Eastep and E. Forster, ”Aerodynamic and Aeroelastic

Characteristics of Wings with Conformal Control Surfaces for Morphing Aircraft”,

Journal of Aircraft, Vol 40, No. 1, 2003, pp. 94-99

30. Aaronburg, Azeemmeruani, Bobsandheinrich and Michael wickmann “MEMS

Gyroscopes And Their Applications, A study of the Advancements in the Form,

Function and Use of MEMS Gyroscopes” pp. 1-20.

31. http://wiki.flightgear.org/index.php/Inertial_Navigation_System 17-Jan-2010

32. www.gyros.freeserve.co.uk 18-Jan-2010

33. Texas Instruments-Accelerometers, Accelerometers and How they work.ppt 03-Feb-

2010

34. AccSol- Accelerometer Solution -http://www.accsol.info/accelerometers.html 24-

Feb-2010

35. http://en.wikipedia.org/wiki/Gyroscope 10-Jan-2010

36. http://www.freepatentsonline.com/3924475.pdf 25-Jan-2010

37. Everobotics.org 30-Jan-2010

38. www.dimensionengineering.com/accelerometers.htm 13-Feb-2010

39. Jerome P. Lynch,Aaron Partridge, Kincho H. Law, Thomas W. Kenny, Anne S.

Kiremidji Ed Carryer “Design of Piezoresitive MEMS based Accelerometer for

Integration with Wireless Sensing Unit for Structural Monitoring” pp. 1-20.

40. National Instruments -http://zone.ni.com/devzone/cda/ph/p/id/183 03-Mar-2010

41. http://news.thomasnet.com/fullstory/Piezoelectric-Accelerometer-is-designed-for-

flight-vehicles-832842 26-Mar-2010

42. http://en.wikipedia.org/wiki/Magnetometer 27-May-2010

43. en.wikipedia.org/wiki/Ochroma pyramidale 20-Apr-2010

44. http://www.jumbolon.com/board_technical_properties.htm 26-Apr-2010

REFERENCES

125

45. Jumbolon.ja.com 26-Apr-2010

46. http://electronics.howstuffworks.com/brushless-motor.htm 27-Apr-2010

47. Ec21.com 28-Apr-2010

48. http://en.wikipedia.org/wiki/Lithium-ion_polymer_battery 28-Apr-2010

49. http://e-articles.info/e/a/title/Advantages-and-Limitations-of-the-Lithium-Polymer-

Battery 29-Apr-2010

50. Diytrade.com 30-Apr-2010

51. Ariesrc.gr 1-May-2010

52. Valuehobby.com 15-May-2010

53. http://www.raidentech.com/2xtopromgmeg.html 25-Jun-2010

APPENDEX A

126

APPENDEX A

RAA WING DESIGN AND STUDY, EES PROGRAM CODE

"RAA WING DESIGN and STUDY" "Hand Wing:" C_L_alpha_rad= 2*Pi/((.32+(.16*A/Cos(a_s)))*(1-(M*Cos(a_s))^2)^.5) "Lift Curve slope in degrees:" C_L_alpha = Pi*C_L_alpha_rad/180 "Reference Area:" S_ref = b_arm*c + b_hand*(c+c*lamda)*0.5 "Exposed Area:" S_exp = S_ref - b_arm*c "Total Wing Span:" b = b_hand + b_arm d_f = .05 "Aspect Ratio" A = b^2/S_ref b_hand = (b_initial-b_arm)*cos(a_s) "Max Coefficient of lift:" C_L_max = .9*Cl_max*cos(a_s) "AOA at max C_L" alpha_max = C_L_max/C_L_alpha + alpha_o C_L = C_L_alpha*(alpha - alpha_o) q=0.5*1.2*V^2 "Arm wing" C_L_arm_alpha = Cl_arm_alpha/((1+57.3*Cl_arm_alpha)/(Pi*e_arm*A_arm)) A_arm=b_arm/c e_arm=0.95 C_L_arm_max = Cl_arm*0.9 S_ref_arm = b_arm*c S_exp_arm = S_ref_arm - d_f*c alpha_arm_max = C_L_arm_max/C_L_arm_alpha + alpha_arm_o C_L_arm = C_L_alpha*(alpha - alpha_arm_o) "Total Lift" L= q*S_ref*C_L + q*S_ref_arm*C_L_arm "For Drag Calculation:" "Parasite drag by Component Buildup Method" C_D_o_1 = C_f_1*FF_1*1*S_wet_w/S_ref S_wet_w = 2*(S_ref-d_f*c) "Component Form Factor of wing:" FF_1 = (1+ .6*tc/xc +100*tc^4)*(1.34*M^.18*(cos(a_s)^.28)) M = V/333 tc = .015 xc = .4 C_f_1 = 1.328/R_w^.5 R_w = 1.2*V*c/mu "Component Form Factor of fuselage:"

APPENDEX A

127

C_D_o_2 = C_f_2*FF_2*1*S_wet_f/S_ref S_wet_f=Pi*d_f*le FF_2=1+60/f_s^3+f_s/400 f_s=le/d_f C_f_2 = 1.328/R_f̂ .5 mu = 1.78e-5 R_f = 1.2*V*le/mu le = .3 C_D_o = C_D_o_1 + C_D_o_2 "Induced Drag by Leading Edge Suction Method:" C_D_i = K*C_L^2 K=1/C_L_alpha_rad "Total Coefficient of Drag" C_D = C_D_o + C_D_i "Drag:" D = q*S_ref*C_D LD_ratio = L/D P=V*D Re = 1.25*V*c/mu

Report; Design of Morphing Wing Micro Air Vehile 'RAA'

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