design and fabrication of an unmanned aerial vehicle

A Project Report on

“Design and Fabrication of Unmanned Aerial

Vehicle” By

PRASHANTH NATARAJAN PRATEEK JOLLY

USN: 1PI07ME065 USN: 1PI07ME067

VIPUL PAUL

USN: 1PI07ME119

Submitted to

VISVESVARAYA TECHNOLOGICAL UNIVERSITY

BELGAUM-590 014

in partial fulfillment of the requirements for the

award of the degree of

BACHELOR OF ENGINEERING

IN

MECHANICAL ENGINEERING

Project work carried out at

P.E.S. INSTITUTE OF TECHNOLOGY

BANGALORE-560 085

Under the guidance of

Dr.T.S.PRAHLAD

Chair Professor in Fluid Mechanics,

Department of Mechanical Engineering,

P.E.S.Institute of Technology,

Bangalore-560085.

DEPARTMENT OF MECHANICAL ENGINEERING

P.E.S.INSTITUTE OF TECHNOLOGY

BANGALORE-560 085

P.E.S. Institute of Technology

(AUTONOMOUS INSTITUTE UNDER VTU, BELGAUM)

100 ft Ring Road, Banashankari 3rd Stage

,

Bangalore-560085,

Department of Mechanical engineering

CERTIFICATE

Certified that the project work entitled “Design and Fabrication of Unmanned Aerial

Vehicle " carried out by Prashanth Natarajan (1PI07ME065), Prateek Jolly

(1PI07ME067) and Vipul Paul (1PI07ME119) who are bonafide students of P.E.S.Institute

of Technology, in partial fulfilment of Bachelor of Engineering in Mechanical Engineering

of Visvesvaraya Technological University, Belgaum during the year 2011. The project has

been approved as it satisfies the academic requirements in respect of project work prescribed

for the said degree.

Signature of the Guide Signature of the HOD Signature of the Principal

External Viva:

Signature of the Examiner with Date

1.

2.

We, Prashanth Natrajan, Prateek Jolly and Vipul Paul, hereby declare that the Project Work

entitled “Design and Fabrication of Unmanned Aerial Vehicle” has been independently

carried out by us under the guidance of Dr. T. S. Prahlad, Chair Professor in Fluid Mechanics,

Department of Mechanical Engineering, P.E.S.I.T, Bangalore in partial fulfilment of the requirements

of the degree in Bachelor of Engineering in Mechanical Engineering Visvesvaraya Technological

University, Belgaum. We further declare that, we have not submitted this work either in part of full to

any other university for award of any degree.

Place: Bangalore.

Date: 27th May, 2011.

Prashanth Natrajan

USN: 1PI07ME065

B.E. (ME)

P.E.S.I.T

Bangalore

Prateek Jolly

USN: 1PI07ME067

B.E. (ME)

P.E.S.I.T

Bangalore

Vipul Paul

USN: 1PI07ME119

B.E. (ME)

P.E.S.I.T

Bangalore

Contents

Abstract ................................................................................................................................................... 7

Acknowledgements ................................................................................................................................. 9

1) Introduction ................................................................................................................................... 10

Applications ....................................................................................................... 10

Steps in the Design Process ................................................................................. 11

1.2.1) Literature survey: ................................................................................................................ 11

1.2.2) Theoretical Design: ....................................................................................................... 11

1.2.3) Prototype Fabrication and flight tests: .......................................................................... 11

2) Problem Statement ........................................................................................................................ 12

2.1) Design Considerations .................................................................................. 12

2.2) Mission ....................................................................................................... 12

3) Design........................................................................................................................................... 13

3.1) Components ................................................................................................. 14

3.2) Weight of Components ................................................................................. 14

3.3) Wing Loading .............................................................................................. 15

3.4) Wing Geometry............................................................................................ 15

3.5) Lift and Drag ............................................................................................... 15

3.6) Airfoil selection .......................................................................................... 16

3.7) Fuselage drag calculations ............................................................................ 19

3.8) Velocity Correction ..................................................................................... 20

3.9) Mission ....................................................................................................... 23

3.10) Take-Off and Climb ................................................................................... 23

3.11) Energy requirements for mission ................................................................. 24

3.11.1) Take-off ......................................................................................................................... 24

3.11.2) Loiter ............................................................................................................................. 24

3.11.3) Descent Glide ................................................................................................................ 24

3.12) Stability .................................................................................................... 26

3.13) Horizontal Tail Sizing ................................................................................ 27

3.14) Elevator Sizing .......................................................................................... 28

3.15) Vertical Tail Sizing .................................................................................... 28

3.16) Rudder Sizing ............................................................................................ 29

3.17) Aileron Sizing ........................................................................................... 29

3.18) Components Selected ................................................................................. 30

3.18.1) Global Positioning System (GPS) ............................................................. 30

3.18.2) Camera ................................................................................................... 31

3.18.3) Motor ..................................................................................................... 31

3.18.4) Servo Motors .......................................................................................... 31

3.18.5) Propeller ................................................................................................. 32

3.18.6) Receiver ................................................................................................. 32

3.18.7) Electronic Speed Control ......................................................................... 32

4) Fabrication ...................................................................................................... 33

4.1) Fuselage Fabrication .................................................................................... 34

4.1.1) Balsa Fuselage .......................................................................................... 34

4.1.2) Coroplast Fuselage: ................................................................................... 39

4.2) Wing Fabrication ......................................................................................... 40

4.2.1) Material .................................................................................................... 40

4.2.2) Hot Wire Cutter ........................................................................................ 40

4.2.3) Steps in Fabrication of Foam Wing: ........................................................... 43

5) Glider Tests ................................................................................................................................... 50

5.1) Glider test without a spar ............................................................................. 50

5.2) Glider Tests with Balsa Spar ........................................................................ 50

5.3) Glider tests with Carbon Fibre spar ............................................................... 51

6) Maiden Flight ................................................................................................................................ 53

6.1) Setup........................................................................................................... 53

6.2) Summary ..................................................................................................... 53

6.3) Flight Path .................................................................................................. 53

6.4) Objectives Achieved .................................................................................... 53

6.5) Duration of Flight - 21 seconds ................................................ 53

6.6) Comments ................................................................................................... 54

6.7) Damage Reported ......................................................................................... 54

7) Flight Test Number 1: First Flight with Coroplast Fuselage ........................................................ 56

7.1) Setup........................................................................................................... 56

7.2) Summary ..................................................................................................... 56

7.3) Flight Path .................................................................................................. 56


7.5) Duration of Flight - 1 minute ................................................................ 57

7.6) Comments ................................................................................................... 57

7.7) Damage Reported ......................................................................................... 57

8) Flight Test Number 2: Acrobatics and Manoeuvrability Test ...................................................... 59

8.1) Setup........................................................................................................... 59

8.2) Summary ..................................................................................................... 59

8.3) Flight Path .................................................................................................. 59


8.5) Duration of Flight: 1.10 minute .................................................................... 60

8.6) Comments ................................................................................................... 60

8.7) Damage Reported ......................................................................................... 60

9) Flight Test Number 3: Heavy Cross Winds with Aborted Landing .............................................. 61

9.1) Setup........................................................................................................... 61

9.2) Summary ..................................................................................................... 61

9.3) Flight Path .................................................................................................. 61


9.5) Duration of Flight: 2.30 minutes ................................................................... 62

9.6) Comments ................................................................................................... 62

9.7) Damage Reported ......................................................................................... 62

10) Flight Test Number 4: Wing Failure and First Crash ............................................................... 63

10.1) Setup ......................................................................................................... 63

10.2) Summary ................................................................................................... 63

10.3) Flight Path ................................................................................................. 63

10.4) Objectives Achieved .................................................................................. 63

10.5) Duration of Flight: 0.50 minutes ................................................................. 63

10.6) Comments .................................................................................................. 64

10.7) Damage Reported ....................................................................................... 64

10.8) Analysis of the Crash ................................................................................. 66

10.9) Changes in Design ..................................................................................... 66

11) Flight Test Number 5: Gusty Weather Flight ........................................................................... 67

11.1) Setup ......................................................................................................... 67

11.2) Summary ................................................................................................... 67

11.3) Flight Path ................................................................................................. 67


11.5) Duration of Flight: 3 seconds ...................................................................... 67

11.6) Comments .................................................................................................. 68

11.7) Crash Analysis ........................................................................................... 68

11.8) Damage Reported ....................................................................................... 68

11.9) Changes to Design: None ........................................................................... 68

12) Flight Test Number 6: First Long Endurance Flight ................................................................ 69

12.1) Setup ......................................................................................................... 69

12.2) Summary ................................................................................................... 69

12.3) Flight Path ................................................................................................. 69


12.5) Duration of Flight: 6 minutes ...................................................................... 70

12.6) Comments .................................................................................................. 70

12.7) Damage Reported ....................................................................................... 70

13) Flight Test Number 7: Flight Test for Manoeuvrability ........................................................... 71

13.1) Setup ......................................................................................................... 71

13.2) Summary ................................................................................................... 71

13.3) Flight Path ................................................................................................. 71



13.6) Comments .................................................................................................. 72

13.7) Damage Reported ....................................................................................... 72

14) Flight Test Number 8: First Flight in PESIT with On-board Camera ....................................... 73

14.1) Setup ......................................................................................................... 73

14.2) Summary ................................................................................................... 73

14.3) Flight Path ................................................................................................. 73



14.6) Comments .................................................................................................. 74

14.7) Damage Reported ....................................................................................... 74

15) Flight test number 9: 20 Minute Flight ..................................................................................... 79

15.1) Setup ......................................................................................................... 79

15.2) Summary ................................................................................................... 79

15.3) Flight Path ................................................................................................. 79


15.5) Duration of Flight: 20 minutes and 36 seconds ............................................. 79

15.6) Comments .................................................................................................. 83

15.7) Damage Reported: None ............................................................................. 83

16) Comparison of Google Earth Snapshots with Snapshots Taken From the P❺ ........................ 84

17) What Makes the P❺ Different? .............................................................................................. 87

17.1) Cost Split Up: ............................................................................................ 87

18) Future Work .............................................................................................................................. 90

18.1) Autopilot ................................................................................................... 90

18.2) Live telemetry and Video feed .................................................................... 90

18.3) Extension of Flight time ............................................................................. 91

18.4) Portability ................................................................................................. 91

19) Conclusion ................................................................................................................................ 92

Appendix A: Software Used................................................................................................................... 93

Appendix B: Original Design .................................................................................................................. 95

Appendix C: References: ..................................................................................................................... 125

Table of Figures Figure 1: Airfoil Sections ....................................................................................................................... 17

Figure 2: Combined Drag Polar ............................................................................................................. 17

Figure 3: Combined Cl-Alpha Graph ..................................................................................................... 17

Figure 4: PT - 40 Section ........................................................................................................................ 18

Figure 5: Cl - Alpha and Drag Polar ....................................................................................................... 18

Figure 6: Graph to Obtain Optimum Cruise Velocity ............................................................................ 21

Figure 7: Mission Profile ....................................................................................................................... 23

Figure 8: Graph to Obtain Optimum Rate of Climb .............................................................................. 24

Figure 9: Variation of Static Margin with Wing Leading Edge Position ................................................ 27

Figure 10: Elevator Sizing ...................................................................................................................... 28

Figure 11: Tail Volume Ratios ............................................................................................................... 29

Figure 12: The Fuselage Sections Being Cut Out from the Balsa Piece ................................................. 35

Figure 13:The Plywood Pieces Being Positioned Along With the Balsa Side Panels in the Vice and

Being Glued Using Fevicol ..................................................................................................................... 35

Figure 14: The curve of the aft section of the fuselage being done ..................................................... 36

Figure 15: The fuselage during fabrication placed in the vice (Top View) ............................................ 36

Figure 16: Leading edge of the empennage being sanded into shape ................................................. 37

Figure 17: Fuselage after completion ................................................................................................... 37

Figure 18: Fuselage with flap open ....................................................................................................... 38

Figure 19: Fuselage with flap open ....................................................................................................... 38

Figure 20: Hot Wire Cutter .................................................................................................................... 42

Figure 21: The plywood template being placed on the foam block ..................................................... 45

Figure 22: The foam cutter being run on the template. Cutting the leading edge of the wing ........... 45

Figure 23: The foam cutter being run on the template. Cutting the trailing edge ............................... 46

Figure 24: Breaking away the excess foam around the cut wing ......................................................... 46

Figure 25: the wing being removed from the block ............................................................................. 47

Figure 26: The cut wing removed and placed on a table ...................................................................... 47

Figure 27: Applying the Monokote layer using the hot iron ................................................................. 48

Figure 28: A number of attempts to get the correct linkage between the Aileron Servo and the

Ailerons ................................................................................................................................................. 48

Figure 29: Placing the Plastic protective sheath around the wing to hot glue into place .................... 49

Figure 30: Completed Wing (Only MonoKote application left) ............................................................ 49

Figure 31: Broken wing without spar .................................................................................................... 50

Figure 32: Glider with full scale wing and empennage in flight ............................................................ 51

Figure 33: The NAL Grounds where the flight test was carried out ..................................................... 55

Figure 34: The aircraft taking off by hand launch ................................................................................. 55

Figure 35: The team setting up the model before flight ....................................................................... 58

Figure 36: Aircraft being launched and pilot with remote ................................................................... 58

Figure 37: Aircraft in flight .................................................................................................................... 60

Figure 38: Aircraft facing a strong crosswind on landing approach ..................................................... 62

Figure 39: the rubber band has eaten half way into the foam wing. The left rubber band has been put

back into position .................................................................................................................................. 64

Figure 40: the wing could be actuated about the carbon fibre spar. This is why the flutter was seen

from the ground .................................................................................................................................... 65

Figure 41: the missing Servo screw. ...................................................................................................... 65

Figure 42: PESIT Garden and Fountain ................................................................................................. 75

Figure 43: Professor M.R.Doreswamy Silver Jublee Block (A- Block) ................................................... 75

Figure 44: Department of Mechanical Engineering (C-Block) ............................................................... 76

Figure 45: F Block (left) and Department of Electrical and Electronics Engineering (Right) ................. 76

Figure 46: Tech Park (E-Block)............................................................................................................... 77

Figure 47: Boys Hostel blocks ............................................................................................................... 77

Figure 49: Stitched image showing the new football ground and a part of the cricket ground ........... 78

Figure 48: Stitched images showing A-block, the fountain, entry way and the Department of

Mechanical Engineering, motorcycle parking in a single frame ........................................................... 78

Figure 50: Aircraft taking off ................................................................................................................. 80

Figure 51: Aircraft on final approach .................................................................................................... 81

Figure 52: Aircraft levelling off .............................................................................................................. 81

Figure 53: Aircraft attempting Flare ..................................................................................................... 82

Figure 54: Aircraft after landing ............................................................................................................ 82

Figure 55: Autopilot chipset .................................................................................................................. 90

Figure 56: Snapshot (12.97108, 77.67567) ........................................................................................... 84

Figure 57: Snapshot (12.560668, 77.327044) ....................................................................................... 85

Figure 58: Snapshot (12.55549, 77.324981) ......................................................................................... 86

Abstract

In the present time, the need for small aircraft which can be remotely operated has

taken at most priority in the needs of society. The multiple roles that an UAV can take clearly

shows its versatility. This provided us the motivation required to design an UAV that could

be fabricated using low cost materials but not compromising on the quality of the aircraft.

We started out by defining a problem statement, “To design and fabricate an Unmanned

Aerial Vehicle capable of carrying a light weight camera and a GPS unit, capable of a

sustained flight time of 20 minutes”.

Why did we choose to carry a camera and a GPS unit as payload? The reason is simple, we

wanted our UAV to be used for aerial tracking purposes. We thought the best application for

such an aircraft would be in search and reconnaissance operations after natural calamities and

also for tracking radio collars of endangered animals.

We started off with a long drawn design process following a conceptual design approach. We

also kept looking at the UAVs present in the market and tried our best to incorporate the

positive aspects of these UAVs. Once the design was complete we entered the world of

fabrication which was completely new to us.

It was the first time we were using a myriad of tools and each day was a learning experience

right form choosing the materials to working on them. We had opportunities to visit the

industry and see how Laser cutting is carried out. We carried out a complete survey to find

out the best dealers to buy our components and raw materials from. We also built a number of

tools to fit our needs of the fabrication process. These tools were built from materials that

could be found at home and proved to be very efficient and effective in the fab process.

We had to chance to make a number of innovations and also made a number of low cost and

equally effective solutions to many problems that we faced during the course of the project.

The toughest part of the project was carrying out the flight tests. We believed in the concept

of build and fly where we build our own aircraft and fly it ourselves. A number of flights

were put in and the results were both joyous where we would celebrate a very successful

flight or sometimes would result in us heading back home to repair the aircraft. The long

endurance flights where we kept breaking the time limits that we set were the most satisfying.

The day we completed a 20 minute flight without any glitches was the most joyous moment

for our team.

During the project apart from having fun, we learnt a number of things. The first and most

important thing that we learnt is how to work as a team, then we learnt how difficult it is to

make something fly. We learnt to use a number of tools for the first time, we learnt the

challenges in fabrication, we learnt how to improve upon our mistakes.

At the end of the project we had an aircraft that could fly for the 20 minute duration specified

by the project target and had the capability to take a video on board and also had the

provision for a GPS unit to track the entire flight path, thus successfully completing our goals

and our project.

The beautiful thing about this project is that is doesn’t end here. There is always scope for

improvement. There are new areas to venture into and new things to learn.

Acknowledgements

This project would not have possible but for the opportunity that the Peoples Educational

Society Institute of Technology, Bangalore gave us.

It is a pleasure for us to thank all those people who supported and guided us to make this

project a success both directly and indirectly.

We extend our gratitude to our guide Dr. T. S. Prahlad, Chair Professor of Fluid Mechanics,

P.E.S.I.T. for allowing us to further our knowledge and gain a profound understanding of the

basic design principles involved in developing an unmanned aerial vehicle. We would like to

sincerely thank him, for offering this project to us and constantly encouraging us along the

duration of the project. Thank you sir for the patience and kindness you have given us, we

truly value it a lot.

We would also like to take this opportunity to thank Mr Satish Nair, Viable Central Asia for

his passion to help and support students. Thank you sir for your contribution to our project.

We would also like to thank Mr Prajwal of NAL, Bangalore for providing us with a number

of suggestions and ideas many of which we have incorporated into our project. We would

also like to thank him for the hardware support he provided us with and the tips he gave us on

flying. Thank you Mr Prajwal for your time and for your patience and your constant

encouragement.

We also express our gratitude to Dr. K.N.B. Murthy, Principal, P.E.S.I.T. and

Dr K. Narasimha Murthy, Head of Mechanical Department, P.E.S.I.T. for encouraging us and

partially funding our project. Thank you Sir. This project would not have been a success if

not for your generosity.

We would like to thank our parents for supporting us in our times of success and failures.

They have been our pillars of support, constantly been by our side and have helped us finish

this project successfully.

Lastly we would like to say that it was a great honour for us to be associated with all these

people and we have enjoyed each moment of the past 3 months in which we have been

learning and progressing under their guidance and support.

1) Introduction

Unmanned Aerial Vehicles (UAVs) are remotely piloted or self-piloted aircraft that can carry

cameras, sensors, communications equipment or other payloads. They have been used in a

reconnaissance and intelligence-gathering role since the 1950s, and more challenging roles

are envisioned, including combat missions. To distinguish UAVs from missiles, a UAV is

defined as “a powered, aerial vehicle that does not carry a human operator, uses

aerodynamic forces to provide vehicle lift, can fly autonomously or be piloted remotely,

can be expendable or recoverable, and can carry a lethal or nonlethal payload.”

Currently, military UAVs perform reconnaissance as well as attack missions. While many

successful drone attacks on militants have been reported, they are also prone to collateral

damage and/or erroneous targeting, as with many other weapon types.

UAVs are also used in a small but growing number of civil applications, such as fire fighting

or non-military security work, such as surveillance of pipelines. UAVs are often preferred for

missions that are too "dangerous" for manned aircraft.

Applications

1. Reconnaissance :

Reconnaissance in a purely military sense involves the constant monitoring of enemy

troop movement and formations. Reconnaissance missions require the aircraft to be

behind enemy lines for an extended period of time thus, these missions are highly

dangerous for manned aircraft. Such missions are the domain of UAVs. These aircraft

are capable of surveying enemy positions and providing real time data back to the

controllers and in case a UAV is shot down the mission commander doesn’t have to

worry about pilot casualties.

2. Remote sensing :

UAV remote sensing functions include electromagnetic spectrum sensors, biological

sensors, and chemical sensors. A UAV's electromagnetic sensors typically include

visual spectrum, infrared, or near infrared cameras as well as radar systems. Other

electromagnetic wave detectors such as microwave and ultraviolet spectrum sensors

may also be used, but are not very commonly used. Biological sensors are sensors

capable of detecting the airborne presence of various microorganisms and other

biological factors. Chemical sensors use laser spectroscopy to analyze the

concentrations of each element in the air.

3. Search and rescue:

UAVs will likely play an increased role in search and rescue in the world over the

years. They can be used to find victims and also hostages holed up.

4. Providing timely information on highway or other transportation modes on traffic

flow and incidents, and the transmission of this information to the appropriate

decision maker, are key requirements for improving traffic and incident management.

The use of Unmanned Aerial Vehicles (UAVs), equipped with video cameras and/or

other sensors, is a technically viable method of providing timely information to

support decisions regarding major traffic incidents and natural disasters, and also in

providing an improved security and safety for the public.

Steps in the Design Process

1.2.1) Literature survey:

In this step we understand what exactly a UAV is and understand its working. Here

the emphasis was on understanding the intricacies involved in the design process as

well as anticipating problems that can arise while applying the design to the actual

model. A survey was undertaken to find the best materials and places to source the

components and also the tools required to fabricate the prototype

1.2.2) Theoretical Design:

This was the first step taken into engineering design. It includes:

Initial sizing and weight estimation

Layout Drawings

Fuselage design

Wing design

Control surface design

Thrust requirements and power plant selection

1.2.3) Prototype Fabrication and flight tests:

A flight worthy prototype of the aircraft was fabricated and a number of flight tests

were conducted to check the air-worthiness of the aircraft. Each step of the test

demanded more from the aircraft in terms of performance and operation. Also we

used the flight tests to practically reduce the wing span of the aircraft. The flight tests

were also helped us gather a lot of information about parameters that could not be

tested on simulations on the computer.

2) Problem Statement

To design and fabricate an Unmanned Aerial Vehicle capable of sustained flight for 20

minutes, carrying a payload of a light weight camera and a GPS unit.

2.1) Design Considerations

1. Endurance – 20 Min

2. Payload: Camera and GPS unit

2.2) Mission

1. Hand launched Take off

2. Climb to desired altitude

3. Cruise for 19 minutes

4. Descent in 2 phases

5. Belly Landing

Two designs have been considered to achieve the goals set by the problem statement. The

first design is a delta wing model on which we had put an additional constraint of a wingspan

of 500mm. Prototype gliders of this design were fabricated and flown, but a number of

problems concerning the speeds and stability of the aircraft arose. These problems have been

discussed as a part of appendix 2. Thereafter we removed the self imposed constraint of a

500mm wingspan and designed an aircraft with a conventional rectangular wing. The initial

design phase of this planform showed us that we had overcome the problems that we faced

with the delta wing. Therefore this is the planform that we chose and further developed.

Our main aim was to try and build the aircraft using materials that are easily available in the

market, cost effective and at the same time not compromise on the quality of the aircraft.

We have tried to keep the costs low by using simple solutions to overcome a number of

hurdles that we faced in the course of completing this project.

Design Conventional Planform

20 Minute endurance

Payload of a Camera and GPS unit

3.1) Components

The aircraft is expected to be a lightweight reconnaissance aircraft and the components were selected

are listed below.

1. Global Positioning System (GPS)

2. Camera

3. Batteries

4. Motor

5. Servos

6. Electronic Speed Control

7. Transmitter and Receiver

8. Propeller

3.2) Weight of Components

Component Weight (Grams)

Servos 30

G.P.S 70

Camera 35

Receiver 28

Batteries 185

Motor 55

ESC 31

TOTAL PAYLOAD 434 Table 1:Weight Estimate

3.3) Wing Loading

The weight of the aircraft being designed has been approximated to 500 grams.

Form a group of R.C aircraft which are a powered glider design similar in weight and having a similar

conventional platform as our design we have initially assumed the Aspect Ratio as 8 and also the wing

loading to be 38.91 N/m2

Thus, A.R = 8

The wing loading is = 38.91N/m2

From the wing loading we can find the wing area; S

= 4.68918/ 38.91= 0.1205 m

2

The wing span (b) can be calculated from the Aspect Ratio and the Wing area

Wing Span = b = √

= √

= 0.9819 m

3.4) Wing Geometry

A conventional rectangular wing is chosen. The wingspan, surface area and aspect ratio are known.

With these parameters we can calculate the Root Chord (Cr) of the wing

Root chord (Cr) =

( )

= 2* 0.1205/ (0.981889984* (1 + 1))

= 0.1227 m

3.5) Lift and Drag

Calculating Reynolds Number at the cruise speed of 14m/s;

Re =

= (1.225* 14* 0.1227)/ (2*10-5

)

= 105246.33

Clrequired = ( ) = (4.68918) / (0.5* 1.225* (14^2) * 0.1205)

= 0.3241

3.6) Airfoil selection

A number of low Reynolds number airfoils were analysed for their 2D chars in XFLR5. The

characteristics that we were looking for were high lift at alpha = 0°, the cruise conditions fall within

the drag bucket, and a thick airfoil for structural reasons.

The Airfoils that were shortlisted were

1) Wortmann FX – 60

2) Selig 4083

3) PT 40

4) Gemini

5) Hobie

Figure 1: Airfoil Sections

The thicker the airfoil, stronger the wing would be structurally. Therefore, structurally the Gemini,

followed by the PT 40 would produce the strongest wing.

The 2D CFD runs show that the Wortmann airfoil produces the greatest lift for a given angle of attack

followed by the Selig 4083, Hobie, PT 40 and the Gemini.

Figure 2: Combined Drag Polar

Figure 3: Combined Cl-Alpha Graph

The Hobie is eliminated because it stalls very early at an angle of 7 degrees.

The Gemini is eliminated because of its sudden stall characteristics.

The Selig 4083 was constructed out of foam, but turned out to be too flimsy, thus getting eliminated.

The Wortmann airfoil, even though it is aerodynamically the best airfoil, producing most lift and least

drag at the cruise Reynolds number, is even weaker than the selig airfoil, especially near the trailing

edge. Therefore even it is eliminated.

The airfoil thus chosen is the PT 40

We have chosen the PT40 airfoil because of its high lift and low drag value at our Cl required along with

its very gentle stalling characteristics. This implies that the operator would have a large buffer zone in

which he can recover the aircraft if it approaches stall. The airfoil is also thicker which helps make the

wing stronger.

Figure 4: PT - 40 Section

Figure 5: Cl - Alpha and Drag Polar

For this value of Clrequired, from the graph the value of alpha initial is very close to 0 Degrees

CdInduced = ( ( ) = (0.3241^2) / (3.14 * 8)

= 0.004179

Cf = 0 ( ( )) / (Log (105246.3327))2.58

= 0.007074

Cdw = 2 = 0.014149

3.7) Fuselage drag calculations

Fuselage length = 0.5m

Fuselage height = 0.056m

Fuselage width = 0.056m

Fuselage Wetted area = 0.112 m2

Cdf =

= 0.007074 * 0.112 / 0.1205

= 0.006574847

Overall Drag coefficient:

CdO = Cdw + Cdf = 0.014149 + 0.006574847 = 0.020724093

Accounting 10% more drag for interference we have;

Cd = CdO + CdInduced = 0.026976329

3.8) Velocity Correction

Our initial cruise velocity of 14m/s was assumed from an existing Powered Glider R.C model. In the

second iteration we will find the optimum velocities for maximum endurance and range.

Velocity Cl D0 Di Drag Power Sink Rate

5 2.541061 0.042068 0.474343 0.516411 2.582054 0.550641

5.5 2.100051 0.050902 0.392019 0.442921 2.436065 0.519508

6 1.764626 0.060578 0.329405 0.389982 2.339895 0.498999

6.5 1.503587 0.071095 0.280676 0.351771 2.286511 0.487614

7 1.29646 0.082453 0.242012 0.324465 2.271252 0.48436

7.5 1.129361 0.094653 0.210819 0.305472 2.291037 0.48858

8 0.992602 0.107694 0.18529 0.292984 2.34387 0.499847

8.5 0.87926 0.121576 0.164132 0.285708 2.428522 0.517899

9 0.784278 0.1363 0.146402 0.282702 2.544317 0.542593

9.5 0.703895 0.151865 0.131397 0.283262 2.690987 0.573871

10 0.635265 0.168271 0.118586 0.286857 2.86857 0.611742

10.5 0.576204 0.185519 0.107561 0.29308 3.077338 0.656264

11 0.525013 0.203608 0.098005 0.301613 3.317743 0.707532

11.5 0.480352 0.222539 0.089668 0.312207 3.590376 0.765672

12 0.441156 0.242311 0.082351 0.324662 3.895942 0.830837

12.5 0.40657 0.262924 0.075895 0.338819 4.235234 0.903193

13 0.375897 0.284378 0.070169 0.354548 4.609118 0.982926

13.5 0.348568 0.306674 0.065068 0.371742 5.018517 1.070233

14 0.324115 0.329812 0.060503 0.390315 5.464405 1.165322

14.5 0.302148 0.35379 0.056402 0.410193 5.947793 1.268408

15 0.28234 0.37861 0.052705 0.431315 6.469727 1.379714

15.5 0.264418 0.404272 0.049359 0.453631 7.031281 1.499469

16 0.248151 0.430774 0.046323 0.477097 7.633552 1.627908

Table 2: Cruise Speed Optimization

Figure 6: Graph to Obtain Optimum Cruise Velocity

From the Graph of Power Vs Velocity we see that the velocity for minimum power is approximately

8m/s.

At this velocity, the Cl value when calculated is found to be 0.9926. This is a very high value nearing

1. The Cl value can be decreased by increasing the cruise velocity of the aircraft.

We change the cruise velocity to 10m/s and recalculate the coefficients.

New Reynolds Number is calculated to be

Re =

= 1.225* 10 * 0.1227 / (2*10

-5) = 75175.95193

Recalculating the Cl value for cruise velocity of 10m/s,

Cl = 0.63520

Also recalculating the values of the Drag Coefficients:

Cdi = 0.0160572

Cdf = 0.007095308

Cdo = 0.022364601

Accounting an additional 10% of Overall Drag to account for interference, we have

Cd = 0.040658283

From this value, drag is calculated to be = 0.300117 N

Thus the power required to overcome this drag is = Power = Drag * Velocity = 0.300117 * 10

= 3.00117 W

The stall speed for the aircraft is calculated as, VStall = √

= √

= 5.940751m/s

3.9) Mission

The aircraft is hand launched from ground level by the operator. The aircraft ascends to four hundred

feet. The aircraft then loiters in the same area for nineteen to twenty minutes taking photographs. It

then descends back to the ground.

Figure 7: Mission Profile

3.10) Take-Off and Climb

Since the aircraft is hand launched, it needs to clear the ground as quickly as possible without

stalling. The initial thrust supplied to the aircraft must be enough to sustain the velocity of the aircraft

above its stall speed. The power required at take off to reach an assigned altitude of 400ft is calculated

for varying rates of climb at the take off velocity.

The equation to evaluate power consumed for varying rates of climb is

(

)

W – Weight of the aircraft

v – Velocity

D – Drag

L – Lift

θ - Angle of ascent

Figure 8: Graph to Obtain Optimum Rate of Climb

The graph also seems to almost level off at about 90 seconds. Therefore a time of 100 seconds is

taken is taken to complete the ascent phase of the flight.

3.11) Energy requirements for mission

With an engine which consumes fuel such as aviation fuel or kerosene, the range or endurance of the

aircraft can be estimated by applying the Breguet formula. Our aircraft does not use a consumable

fuel, but rather a battery to power the electric motor, which is our primary thrust producing prime

over. Thus, in this case we estimate the amount of energy (joules) required to complete the mission

and then choose a suitable battery which can provide this amount of energy.

Our mission is broken up into 3 phases.

1) Takeoff

2) Loiter

3) Descent Glide

The energy required at each stage is estimated theoretically, and then added up to obtain the energy

required for the whole mission.

3.11.1) Take-off

The power required to reach four hundred feet has been estimated to be 8 watts. The aircraft

is expected to achieve this in 100 seconds. The product of time and power gives the energy

required. This is calculated to be 800 J.

3.11.2) Loiter

The aircraft now levels off and circles the area that has to be scanned. The aircraft remains in

this phase for up to nineteen minutes. The power required for sustained flight at 10m/s has

been estimated at 3.6 watts.

3.11.3) Descent Glide

0

5

10

15

20

25

30

35

0 20 40 60 80 100 120 140

Po

we

r (W

)

Time (s)

Selection of Rate of Climb

After the aircraft has scanned a particular area, it begins an unpowered glide to the ground.

The aircraft continues in this state till it reaches an altitude of 20 feet.

The power consumed by the servos and the receiver has also been estimated.

Servos - 70mA at 11.1V for 20 minutes = 932 J

Reciever – 10mA at 11.1V for 20 minutes = 133 J

The energy required for the whole mission is estimated to be 6 kJ.

Efficiency of motor = 70%

Efficiency of propeller = 70%

Efficiency of Battery = 80%

The Energy requirement after taking into account the component efficiencies is 15.5 kJ.

Remote control aircraft pilots recommend that if an aircraft is designed to fly for certain endurance,

then, because of weather conditions, the energy requirement is taken as 1.5 times to compensate for

gusts, head winds, mid-course corrections etc.

Therefore the energy requirement = 1.5 x 15.5kJ = 23.25 kJ.

A 1000mA battery provides 36 kJ of energy. With an estimated battery efficiency of 80%, the energy

available from the battery is 28 kJ. Therefore this battery is chosen.

3.12) Stability

The aircrafts designed to be very stable longitudinally so that the on-board camera does not jerk or

shudder, thereby compromising on the quality of images taken.

The static margin is the defined as the distance between the C.G. of the aircraft and the Aerodynamic

Center of the wing, and is usually measured as a percentage of the mean aerodynamic chord (MAC).

Since our wing is rectangular, the MAC is the same as the chord.

The static margin for gliders is between five and ten percent. Greater the margin, the greater the

stability. If the margin exceeds 10%, the aircraft become ultra stable and is very difficult to manuver.

The C.G. of the entire structure including the fuselage, tail, servos, motor and esc is found to lie 16

cm from the nose of the aircraft.

The battery is the heaviest component in the aircraft at 185g and is placed in the roomiest part of the

aircraft.

The wing is then moved about till the required static margin is achieved.

Position of

Wing

Wing CG Aerodynamic Center CG of A/C Static margin

3 7.305 6.075 13.41645 49.41324047

3.5 7.805 6.575 13.44913 46.26783837

4 8.305 7.075 13.48181 43.12243627

4.5 8.805 7.575 13.51449 39.97703418

5 9.305 8.075 13.54717 36.83163208

5.5 9.805 8.575 13.57985 33.68622998

6 10.305 9.075 13.61253 30.54082789

6.5 10.805 9.575 13.64521 27.39542579

7 11.305 10.075 13.67789 24.25002369

7.5 11.805 10.575 13.71057 21.1046216

8 12.305 11.075 13.74325 17.9592195

8.5 12.805 11.575 13.77593 14.8138174

9 13.305 12.075 13.80861 11.66841531

9.5 13.805 12.575 13.84129 8.52301321

10 14.305 13.075 13.87397 5.377611114

10.5 14.805 13.575 13.90664 2.232209017

11 15.305 14.075 13.93932 -0.91319308

11.5 15.805 14.575 13.972 -4.058595176

12 16.305 15.075 14.00468 -7.203997273

12.5 16.805 15.575 14.03736 -10.34939937 13 17.305 16.075 14.07004 -13.49480147

13.5 17.805 16.575 14.10272 -16.64020356

14 18.305 17.075 14.1354 -19.78560566

14.5 18.805 17.575 14.16808 -22.93100776

15 19.305 18.075 14.20076 -26.07640985

15.5 19.805 18.575 14.23344 -29.22181195

16 20.305 19.075 14.26612 -32.36721405 Table 3: Placement of Wing to get Appropriate Static Margin

Figure 9: Variation of Static Margin with Wing Leading Edge Position

Therefore for a static margin of 10%, the leading edge of the wing is found to lie at 12.4 cm from the

nose of the aircraft.

3.13) Horizontal Tail Sizing

The tail-moment arm (TMA) is the distance between the mean aerodynamics chords of the wing and

the tail. The TMA was taken to be 2.5 times the wing’s MAC.

Therefore, TMA= 0.306841 m.

The area of the horizontal tail (HTA), is given by the formula,

HTA =

Where, WA is the Wing Area.

As the TMA was taken to be 2.5 times the MAC, HTA is 20% of the wing area.

Hence, HTA = 0.2 * 0.120513 = 0.024103 m2.

The aspect ratio of the tail was assumed to be 5.

Therefore, the span of the horizontal tail, bht = √ = 0.347151 m.

The chord of the horizontal tail, Crht =

= 0.06943 m.

y = -6.2908x + 68.286

-40

-30

-20

-10

0

10

20

30

40

50

60

0 5 10 15 20Stat

ic M

argi

n

Dist. From L.E. of Wing to Nose

Longitudinal Stability

3.14) Elevator Sizing

The larger the elevator area, in proportion to the horizontal tail’s total area, the more effective the

elevator, as shown by the graph below.

Figure 10: Elevator Sizing

From the graph above, we can see that a value of Se/St lying between 0.3 to 0.35 gives an elevator

effectiveness of 70%. Assuming the value of Se/St as 0.35, we get;

Se = 0.35 * 0.024103 = 0.008436 m2.

3.15) Vertical Tail Sizing

Since the vertical tail is placed on the horizontal tail, the root chord and the tail-moment arm of the

vertical tail is same as that of the horizontal tail.

Crvt= 0.06943 m

TMA = 0.306841 m

Area of vertical tail is given by, Svt=

Where, Cvt is the tail volume coefficient, taken from the table below.

Taking the value of Cvt for a sailplane, i.e. Cvt= 0.02.

Thus, Svt =

= 0.007713 m

2

Taking a taper ratio (k) for the vertical tail as 0.6, we get the value of tip chord as,

Ctvt = 0.6 * 0.06943 = 0.041658 m.

The span of the vertical tail is given by, b =

( ( ))

Thus, b is calculated to be 0.13886 m.

3.16) Rudder Sizing

A rudder area of 30% of the vertical tail area is found to optimum for a model this size.

Therefore, Sr = 0.3 * Svt = 0.3 * 0.007713 = 0.0023139 m2.

3.17) Aileron Sizing

The Aileron was sized from a group of RC Powered Glider having similar characteristics as our

aircraft;

The Aileron Dimensions are 20cm X 5cm

Figure 11: Tail Volume Ratios

3.18) Components Selected

The aircraft is expected to be a lightweight reconnaissance aircraft and the components were selected

are listed below.

1. Global Positioning System (GPS)

2. Camera

3. Batteries

4. Motor

5. Servos

6. Electronic Speed Control

7. Transmitter and Receiver

8. Propeller

An exhaustive survey of components that meet the design requirements was carried over the internet.

3.18.1) Global Positioning System (GPS)

The GPS unit chosen is the 65CBTOOTH GPS Receiver. The specifications of the

component are as shown in the table below.

Specifications

Dimensions 75 x 42 x 16 (mm)

Weight 70 g

Battery Life 18.5 hours (full charge)

Reacquisition Time 0.1 seconds

Time log 300 hours

GPS data stored 90,000

Recharge Time 2.5 hours

Data Recorded Latitude, Longitude and Speed

Cost Rs. 2395

The 65C BTOOTH unit was chosen as it is a lightweight unit, refreshes its location every 0.1

seconds and has sufficient memory for storing position data. The data recorded can be

transferred to a computer via a USB cable and can be used with applications such as Google

Maps to trace the flight path.

3.18.2) Camera

Two types of cameras were considered the first was the Spy Hidden DVR Micro Camera

DV and the second a 3.2MP Spy Pen Camera.

Specifications DVR Micro Camera Spy Pen Camera

Camera Resolution 720 x 480 at 30 fps 640 x 480 at 30 fps

Photograph Resolution 1600 x 1200 pixels 1024 x 768 pixels

Battery Life 60-80 minutes 120 minutes

Recording Time 1 hour Memory Dependent

Data Storage 2-16gb 2gb External Card

Cost Rs. 1200 Rs. 1250

Although the two cameras are almost identical in specifications, a pen camera is preferred as

it smaller in size, lighter and also has external storage thus making it user dependent.

3.18.3) Motor

A brushless DC motor was selected to act as the power plant of the aircraft. A brushless

motor was selected as it is more efficient, produces less heat and is more reliable than a

brushed DC motor. The motor selected was the Emax 1270/13.

Specifications

Recommended Model Weight 1000 grams

Shaft Diameter 4 mm

Weight with cables 55 grams

Dimensions 28.5 x 28.5 mm

Continuous Current 25 Amps

Maximum Thrust 800 grams

Cost Rs. 726

3.18.4) Servo Motors

Two servo motors are required to control the control surfaces namely one for the elevator and

one servo for the aileron. The servo chosen was the Power HD Servo. This servo was chosen

because of its light weight and small size.

Specifications 2 plastic gear servos and 1 metal gear servo

Weight 9 grams

Torque 1.8 kg-cm

Speed 0.12 sec/60

Size 22.5 x 11.5 x 24.6 mm

Cost Plastic Gear: Rs 275

Meatl Gear: Rs 480

3.18.5) Propeller

The diameter of the propeller was calculated to be 9 inches. Since the aircraft requires a high

torque during takeoff, a propeller of diameter 8 and a pitch of 4 taken, also this specification

of propeller makes the motor most efficient i.e. a 8X4 propeller. APC 8X4 Propeller is

chosen

Features

High Thrust

Low Noise

Gas filled nylon for strength and durability

Rs. 225

3.18.6) Receiver

Futaba PCM FP-R138 DP

Specifications

Number of channels 8

Dimensions 64 x 35.3 x 20.8 mm

Frequency 72 MHz

Crystal 16

Weight 28 gms

3.18.7) Electronic Speed Control

E-Max ESC

Specifications Programmable ESC

Continuous Current 25 Amps

Dimensions 50x28x12 mm

Surge Current 30 Amps

Weight 31 grams

Cost Rs 1000

FABRICATION

4.1) Fuselage Fabrication

The fuselage is the part of the aircraft that will hold the payload. We chose to go ahead with a box

type fuselage in which all the components could be easily fit in.

We chose to build two fuselages:

1. The first one is made of Balsa

2. The second one is our lower cost fuselage made of Coroplast

4.1.1) Balsa Fuselage

Steps in Fabrication

1. The side panels were drawn on the Balsa sheet and were cut out.

2. Three plywood pieces measuring 5cmX5cm and 3mm thickness were cut out. These pieces

were used as the Motor Mount and as supports along the length of the fuselage. The two

pieces that would be places along the length of the fuselage needed to have a provision for the

wires to be threaded through. This was achieved by cutting out holes in the plywood pieces.

3. The squares were places at predetermined points and glued to the balsa cut outs with fevicol

and placed in a vice to dry overnight. Additional supports were given to the motor mount

using beading pieces. A third piece was glued in at the point where the fuselage starts to curve

upwards.

4. Two additional pieces of plywood measuring 3cmX3cm and 1cmX1cm were cut out. These

pieces were placed to help hold the gentle taper of the fuselage. The 1cmX1cm piece was

placed at the end of the fuselage thus completing the sides of the fuselage.

5. The bottom of the rectangular section is a single piece of balsa glued onto the sides. The rear

of the bottom portion is tapped with brown tape. This helps in saving weight and also helps in

accessibility of components in this section.

6. The cover of the fuselage is a flap. The reason for choosing a flap is that it provides for

accessibility to the components in the fuselage. The flap is a rectangular piece of Balsa that

has been attached to the body with the help of ribbons. The ribbons make it possible for the

flap to be opened and closed. For the rear we traced out the shape and cut out a piece of

Balsa and glued it onto the fuselage.

Building the empennage:

1. The empennage was fabricated using Balsa. The vertical tail and the horizontal tail were

drawn on the balsa sheets and were cut out.

2. The pieces were then sanded such that the leading edges were rounded and the trailing edges

were sharp.

3. The elevator was cut out from the horizontal tail piece and was attached back to the horizontal

tail with the help of ribbons. The ribbons allow for the upward and downwards deflection of

the elevator.

4. The vertical tail and the horizontal were attached using hot glue. This assembly in turn was

placed on the fuselage with a bead of hot glue.

Figure 12: The Fuselage Sections Being Cut Out from the Balsa Piece

Figure 13:The Plywood Pieces Being Positioned Along With the Balsa Side Panels in the Vice and Being Glued Using Fevicol

Figure 14: The curve of the aft section of the fuselage being done

Figure 15: The fuselage during fabrication placed in the vice (Top View)

Figure 16: Leading edge of the empennage being sanded into shape

Figure 17: Fuselage after completion

Figure 18: Fuselage with flap open

Figure 19: Fuselage with flap open

Placing of Elevator Servo:

The servo was placed along the centreline of the fuselage on the top surface. We had to make sure that

there was enough movement of the servo arm to facilitate the corresponding movement of the

elevator.

The control line from the servo horn to the elevator horn was then bent and fitted making sure there

was enough clearance for the horn to move freely and that the control line did not interfere with this

movement.

4.1.2) Coroplast Fuselage:

Balsa as a raw material is quite expensive. As our main aim was to keep the costs for materials low

we chose to build a second fuselage using coroplast. Coroplast (Corrugated Plastic) is basically two

thin sheets of plastic with ridges in between. It is a very strong material and is quite economical to

use. It is slightly heavier than Balsa but the difference is not much to cause a major difference in the

flight performance.

Fabrication of the coroplast fuselage:

1. Three sections i.e the sides and the bottom of the fuselage were drawn onto the sheet of

coroplast as a development. This was then cut out and the places where a fold would be done

were cut such that only the top plastic layer was cut but the bottom layer remained.

2. Three plywood pieces measuring 5cmX5cm and 3mm thickness were cut out. These pieces

were used as the Motor Mount and as supports along the length of the fuselage. The two

pieces that would be places along the length of the fuselage needed to have a provision for the

wires to be threaded through. This was achieved by cutting out holes in the plywood pieces.

3. The plywood pieces were placed in similar positions as the Balsa fuselage. The only

difference was the use of hot glue to stick the plywood to the coroplast. Also we did not use

the 3cmX3cm and 1cmX1cm pieces at the tapered region.

4. The tail region was built by hot gluing the 3 pieces such that the correct taper was achieved.

5. To cover the tail region, a coroplast sheet was cut into the correct shape and tapped to the

bottom using duct tape.

6. The flap in this model is different from the Balsa model. We cut the single flap into 3

different flaps thus providing us access to the individual bays if required. Also we found that

it was necessary to remove and replace the battery connections a number of times. With a

single flap it was necessary to remove the wing every time that had to be done.

We chose to have a single small flap at the front which would provide easy access to the

battery leads without the need to remove the wing.

Building the empennage:

1. The empennage was fabricated using Balsa. The vertical tail and the horizontal tail were

drawn on the balsa sheets and were cut out.

2. The pieces were then sanded such that the leading edges were rounded and the trailing edges

were sharp.

3. The elevator was cut out from the horizontal tail piece and was attached back to the horizontal

tail with the help of ribbons. The ribbons allow for the upward and downwards deflection of

the elevator.

4. The vertical tail and the horizontal were attached using hot glue. This assembly in turn was

placed on the fuselage with a bead of hot glue.

Placing of Elevator Servo:

Our aim was to try and make the elevator servo sit a little more flush with the fuselage surface

as compared to the Balsa fuselage. We achieved this by cutting out a piece of plywood and

cutting out a slot on this in which the servo would fit into. This piece in turn was hot glued to

the inside of the tail coroplast section.

The servo sat much more flush to the surface and at the same time with enough clearance to

prevent any interference from the elevator control line-servo horn assembly.

4.2) Wing Fabrication

4.2.1) Material

The material chosen to build the wing was High Density Foam. The reasons for choosing this material

to build the wing are:

1. Foam as a construction material is very light

2. It is comparatively easier to work into the complex air foil shapes than balsa

3. It is highly durable once given the necessary treatments which will be explained later

4. It is an easily available material and is very economical

We had to find a way to cut the air foil shape from the foam blocks. On doing some research we

found that the most effective method of cutting foam is using a Hot Wire Cutter.

4.2.2) Hot Wire Cutter

A Hot Wire Cutter basically uses an electric current to heat a wire under tension. When this wire is

passed through the foam the wire cuts the foam.

Fabrication of Hot Wire Cutter

We chose to build the Hot Wire Cutter using cost effective and easily available materials.

Materials Used:

Plywood (5mm thickness)

Compression Springs

Guitar String

Nails and Screws

Steps in Fabrication:

We first cut three pieces of plywood.

o Two pieces measuring 30cm X 5cm

o One Piece measuring 100cm X 5cm

These were arranged in the shape of a U such that the longer piece forms the bridge and the

two smaller pieces form the legs. They were fastened in such a way that the hinge provided

for motion

A string was drawn between the free ends of the smaller pieces. Provision was given for wire

leads to be attached to the string.

We then hammered nails near the hinge on which the compression spring would be attached.

The spring keeps the string in tension and also provides strength to the entire assembly.

We also gave a provision for changing the tension in the string by providing two different

positions for the spring. This helped us change the tension if required.

The leads from the string are connected to an adaptor which in turn are connected to a wall

socket

Figure 20: Hot Wire Cutter

4.2.3) Steps in Fabrication of Foam Wing:

1. The airfoil coordinates were taken from the UIUC online database. These coordinates were in

turn fed into Solid Works after multiplying the values with the predetermined chord length

and a curve was drawn. The curve was then extruded to get a solid. On converting the Solid

from 3D to 2D, one of the views gives the airfoil profile. A print out of the air foil was taken.

2. The airfoil print out was then stuck onto a sheet of 3mm plywood and the sheet was sanded

down to give the shape as close to the airfoil outline as possible. Two of these were made.

3. A number of markings were made on the templates. These markings help provide a reference

during the cutting process making sure that the two persons operating the hot wire are at the

same point during the cutting process.

4. The two templates are placed on either side of the foam block and stuck to it using fevicol.

The Hot Wire is then inserted and made to run along the templates from the leading edge to

the trailing edge and is removed at quarter chord.

5. The excess foam is broken off and the wing is removed. The wing now has the two plywood

templates stuck on either side. These are then removed using a blade.

6. A groove was to be cut into which the spar would sit. This is done again using the hot wire.

We mark off the outline of the spar on either side of the cut foam wing. The hot wire is

lowered and follows the outline drawn thus creating the groove for the spar to sit in.

7. The carbon fibre spar is then taken and fevicol is applied on three sides of the spar. This is

then slid into the groove and a slight pressure is applied to keep it in place. The groove is then

covered with a layer of paper to help give a neat finish to the top surface of the wing.

8. The section where the fuselage is to seat is marked off. This is to place the servo along the

centre line of the wing. A groove of 4.5cmX3.5cm is shaved off the wing. Into this a plywood

piece in which a cut out for the servo has been done is placed and fixed in place using fevicol.

9. Once the aileron position is chosen, a slot is cut in the wing for the aileron to sit.

10. Application of MonoKote:

MonoKote is commercially available light weight plastic shrink wrap film available

in various colour schemes with an adhesive on one side, used to cover and form the

surfaces of a model aircraft.

The MonoKote is cut into shape such that is forms a wrap around the wing and is then

applied with the help of a hot iron. The adhesive on one side is heat activated. Once

the layer sticks to the surface a different temperature is set on the iron, this is to

shrink to layer onto the wing.

11. The aileron is made of balsa wood. The aileron is placed on the foam wing and taped to it in

such a manner that it is free to move up and down.

12. The last step in the fabrication of the wing is placing the servo and attaching the control rods

from the servo to the ailerons. We chose a simple system in which a single servo is used to

control both the ailerons. When one aileron is pushed down the other is pulled up. The control

rods are simple cycle spokes that have been bent in such a manner to achieve the movement

necessary.

What we learnt using the Hot Wire Cutter?

It was a first time experience for the team building and using the hot wire cutter and we learnt a

number of things while using it. As the process was carried out we found a number of ways to make it

faster and make the finish better.

A number of challenges and how we overcame them have been listed below:

During the course of fabrication we first used balsa as the wood to make the templates. On

carrying out the process we found out that the hot wire started to eat into the balsa thus

getting stuck in the template itself. The wing that we got first had a very bad finish and was

discarded. We then tried out the same using plywood and found that the wire glided smoothly

on it and thus we choose to go with plywood to make the templates.

In the first few wings we faced the problem of a bow along the trailing edge. After a number

of wings were cut we found that the cause for the problem was that the middle portion of the

wire tends to move slower than the extremes thus causing the bow. The problem was rectified

by moving beyond the trailing edge of the template and staying in the foam for 5 seconds thus

giving enough time for the wire to take up a straight shape.

While cutting the wing we faced the problem of uneven heating along the length of the wire.

The cause of the problem was found to be in the method we had connected the leads to the

wire. We changed this and found that the wire heated evenly.

While cutting the groove for the spar we found that at the centre portion of the wing the

groove was very shallow but at the ends it was very deep. We tried raising the middle portion

and cutting the groove. This proved to work but we required a number of passes and had the

problem of the wire cutting all the way through the foam.

We found out that the best solution was to move the hot wire very slowly. This once again

helps the natural bow that is formed in the wire to come back to the straight line

configuration. Once the first cut is made, the cutter is kept at the bottom of the groove for 5

seconds and then moved up slowly.

With every wing that we made we were able to get a much better finish and were able to get a finished

wing much faster.

Another major problem that we had to overcome was how to prevent the rubber bands which hold the

wing in place on the fuselage from eating into the foam?

The solution that we arrived at is a simple one. We used a slightly think plastic sheet and wrapped it

around the central portion of the wing where the rubber bands would be placed. The sheet was hot

glued into place. The hot glue provided additional strength to this sheet in turn preventing the rubber

bands from eating into the foam wing.

The first wing that we made took not less than three days but as we kept making wings we learnt a

number of things and also were able to implement a number of methods to help increase our

efficiency of production. At the end we were able to make a whole wing right from printing out the

shape to attaching the servos in One day.

Figure 21: The plywood template being placed on the foam block

Figure 22: The foam cutter being run on the template. Cutting the leading edge of the wing

Figure 23: The foam cutter being run on the template. Cutting the trailing edge

Figure 24: Breaking away the excess foam around the cut wing

Figure 25: the wing being removed from the block

Figure 26: The cut wing removed and placed on a table

Figure 27: Applying the Monokote layer using the hot iron

Figure 28: A number of attempts to get the correct linkage between the Aileron Servo and the Ailerons

Figure 29: Placing the Plastic protective sheath around the wing to hot glue into place

Figure 30: Completed Wing (Only MonoKote application left)

5) Glider Tests

We carried out a number of glider tests in order to determine and check a number of parameters:

Whether the tail size was adequate enough to provide stability to the aircraft

To check if the wing could sustain the impact of a belly landing

To check if the foam was a good material to absorb the impact of a belly landing

We built a 1:1 scaled wing using the foam and a 1:1 scaled empennage using coroplast. These were

hot glued onto a stick which represented the fuselage.

5.1) Glider test without a spar

The first test was carried out without a spar in the wing. On impact the wing broke and this suggested

that the placement of a spar in the wing was compulsory.

Figure 31: Broken wing without spar

5.2) Glider Tests with Balsa Spar

The second set of tests was carried out with a new wing having twin Balsa spars and wrapped with

brown tape. We found that the wing with the twin spars absorbs the impact very well and the only

damage that resulted from the landings was a small chipping of the foam on the trailing edge.

5.3) Glider tests with Carbon Fibre spar

The third set of tests was carried out with a single carbon fibre spar placed at quarter chord. The spar

gave immense strength to the wing and was clearly better than having a double balsa spar. Once again

the only problem was the slight chipping of the foam on landing impact. This could be easily

prevented by giving a layer of protective coating on the surface of the foam.

The final set of tests was carried out with a wing which had a carbon fibre spar and a layer of

MonoKote applied on its surface. This solved the problem of the trailing edge chipping on landing.

We decided to go ahead with this form of the wing.

Figure 32: Glider with full scale wing and empennage in flight

FLIGHT TESTS

6) Maiden Flight

Date: 3rd

May 2011

Time: 07:30

Location: NAL grounds

Weather Conditions: Moderate Winds, Overcast conditions

6.1) Setup

1. Wing Span= 95cm

2. 1300 mah Battery

3. Coroplast Fuselage

4. Wing fixed by Rubber Bands being wrapped around the Fuselage

6.2) Summary

This was the first time we were flying our aircraft. Our main aim was to take-off and land.

We wanted to check if the aircraft was air worthy and if the entire setup would work.

It was also the first ever flight for our pilot

6.3) Flight Path

1. To Take-off at full throttle

2. Turn

3. Land without damaging the model

6.4) Objectives Achieved

1. The take-off was achieved successfully

2. Turn and glide were achieved successfully

3. Landing was perfect

6.5) Duration of Flight - 21 seconds

6.6) Comments

Form our first flight we were able to infer the following:

The aircraft took off at a very steep R/C. Thus the motor was switched off at the peak of

ascent.

The aircraft began an unpowered glide covering 2 circuits of the ground showing that the

aircraft has very good glider characteristics.

The aircraft landed smoothly without any damage. The landing was not entirely a smooth

touch down but we managed to land it successfully. The foam layer that we had added for

protection at the nose of the aircraft had done its job well. In fact on inspection we saw that

the foam layer had taken the impact and that there was no damage on the balsa fuselage.

6.7) Damage Reported

There was no damage as such to the aircraft. Only the foam layer that had been added for

protection had taken sum mud on its surface which was a clear indication that the aircraft

landed on the layer.

Figure 33: The NAL Grounds where the flight test was carried out

Figure 34: The aircraft taking off by hand launch

7) Flight Test Number 1: First Flight with Coroplast

Fuselage

Date: 5th May 2011

Time: 16:30

Location: Disused Test Track

Weather Conditions: Moderate Winds

7.1) Setup

1. Wing Span= 95cm

2. 1300 mah Battery



7.2) Summary

This was the first flight of the coroplast Fuselage model. Our main aim was to check if there

were any changes brought about due to change in the Fuselage material from traditional Balsa

to coroplast. The coroplast fuselage is slightly heavier than the Balsa model. Also this was the

first flight in moderate wind conditions in an open field.

7.3) Flight Path

1. To Take-off at a medium rate of climb

2. Carry out a Left Aileron turn

3. Trim the control surfaces in flight

4. Carry out a complete circle

5. Carry out a Straight in approach

6. Try a smooth landing



2. Left Aileron Turn

3. Trimming was carried out during the flight

4. The circle was successfully achieved

5. Lining up with the runway was done successfully

7.5) Duration of Flight - 1 minute

7.6) Comments

From this first flight we learnt the following:

During launch the aircraft must be held such that the orientation of the wings is parallel to the

ground. As we near the launch velocity the aircraft starts to lift off and the person launching

must simply let go of the aircraft and MUST NOT try and push it into the air

The control surfaces are adequate and provide a good amount of control for roll and pitch

movements of the aircraft.

The control surfaces need to be trimmed in the first flight. This is because of the removable

wing. Every time the wing is mounted and dismounted (for transportation), there is a need to

remove the servo horn. This in turn causes a change in the Neutral Position of the control

surfaces. A certain amount of setting can be done on the ground by visual reference but the

actual feel for the controls can be got only in flight, thus the trimming has to be carried out in

flight.

Once trimmed the aircraft is very stable. We were able to achieve a complete 360 degree

aileron turn. The turn was a large radius turn.

The aircraft had to be given additional throttle during the turn as it tends to loose altitude as it

carries out the maneuver.

Lining up for landing proved to be the most challenging part of the flight as this was carried

out for the first time. The aircraft is made to gain altitude and made to point towards the

runway. The throttle is brought to Zero

As the aircraft comes closer to the ground the attitude control was given through inputs to the

elevator causing the aircraft to flare.

The landing was not entirely a smooth touch down but we managed to land it successfully.


There was no damage as such to the aircraft; there were slight bruises that were purely

superficial. The following was noted:

The wing had displaced slightly forward

No damage to the Empennage

No damage to the Fuselage

A complete check of all the servos and the motor was carried out and No damage was

noted.

Figure 35: The team setting up the model before flight

Figure 36: Aircraft being launched and pilot with remote

8) Flight Test Number 2: Acrobatics and Manoeuvrability

Test

Date: 5th May 2011

Time: 17:00


Weather Conditions: Moderate Winds

8.1) Setup

1. Wing Span= 95cm

2. 1300 mah Battery



8.2) Summary

This was the second flight of the coroplast Fuselage model. Our aim in this flight was to

check the durability of the aircraft by putting it through high G turns and through few

acrobatic manoeuvres. We also wanted to check how effective the rubber bands were in

holding the wing in position by putting the wing through this highly demanding test.

8.3) Flight Path

1. To Take-off at a High rate of climb

2. Carry out a sharp Left Aileron turn and level off

3. Carry out a complete 360 degree turn

4. Carry out a tighter 360 degree turn

5. Level off

6. Clock wise direction loop

7. Sharp 180 degree turn

8. Anti-clock wise loop

9. Line up and perform a Straight in approach Landing



2. Sharp Left Aileron Turn

3. 360 degree turn

4. Tighter 360 degree turn

5. CW loop

6. 180 degree turn

7. Anti-clock wise loop

8. Straight in approach landing partially successful

8.5) Duration of Flight: 1.10 minute

8.6) Comments

The aircraft is very manoeuvrable and is able to take sharp turns quite easily. The loss in

altitude during the sharp turns is kept under control by increasing the throttle.

The aircraft was able to pull both directional loops with ease and it was easy to get it back in

control.

The rubber bands held the wing in position during the flight. There was no shifting of the

wing during the flight.


There was no damage as such to the aircraft; there were slight bruises that were purely

superficial. The following was noted:

The wing had not displaced on this landing

Slight superficial damage to the Empennage



noted.

Figure 37: Aircraft in flight

9) Flight Test Number 3: Heavy Cross Winds with

Aborted Landing

Date: 5th May 2011

Time: 17:15


Weather Conditions: High Cross winds

9.1) Setup

1. Wing Span= 95cm

2. 1300 mah Battery



9.2) Summary

This was the third flight of the coroplast Fuselage model. Our aim in this flight was to fly the

sorties that we would be doing during an aerial tracking mission. The flight would test the

ability of the aircraft to fly straight line paths at a level attitude and at the end of the run carry

out a turn and repeat the flight in the opposite direction a slight distance offset from the first

run.

9.3) Flight Path

1. To Take-off at a moderate rate of climb

2. Carry out a Left aileron turn

3. Fly a straight path

4. Carry out a tight 180 degree turn


6. Carry out a tight 180 degree turn


8. Line up for a landing

9. Attempt a flared, kiss landing



2. Left Aileron turn

3. Sortie of straight line fly-by’s followed by the turns

4. Line up

5. Successful touch down

9.5) Duration of Flight: 2.30 minutes

9.6) Comments

The aircraft is very manoeuvrable and was able to achieve the sortie easily.

On landing approach the aircraft faced a strong crosswind. The effect of the gust was to roll

the aircraft CW direction. The pilot tried to put the nose into the incoming wind in order to

perform a Crab Landing but the gust was too strong and blew the aircraft off course. He

managed to line it back up but just 10 feet above the ground the aircraft faced yet another

crosswind gust throwing the aircraft out of the line up once again.

We were able to power up and gain altitude, finish a circuit and line up again and land

successfully. This clearly proved that the motor produces enough lift to help the aircraft climb

even at very low airspeeds.


There was no damage as such to the aircraft

The wing had not displaced on this landing




noted.

Figure 38: Aircraft facing a strong crosswind on landing approach

10) Flight Test Number 4: Wing Failure and First

Crash

Date: 5th May 2011

Time: 17:30


Weather Conditions: Heavy Wind Conditions

10.1) Setup

1. Wing Span= 95cm

2. 1300 mah Battery



10.2) Summary

This was the fourth flight of the coroplast fuselage model in one day. This was to subject the

aircraft to rugged use by continuously taking off and landing in short duration flights.

10.3) Flight Path

1. To Take-off at a moderate rate of climb

2. Carry out a Left aileron turn

3. Carry out a number of circular sorties

4. Line up with the runway

5. Land



2. Left Aileron turn


10.6) Comments

The weather conditions got worse after take-off. There were strong guts that were blowing the

aircraft making it extremely difficult to control. The pilot tried to bring the aircraft back in for

a landing but wasn’t successful in doing so

While flying into the wind the aircraft tended to pitch up and stall in spite of throttling up to

100%. The reason was the strong head winds that were hitting the aircraft.

While attempting the final sortie, the aircraft stalled at a high AoA. From the ground a flutter

was visible on the right wing. We immediately realized that the wing had failed. The aircraft

crashed and the chase crew was sent out to recover the aircraft.


The following was noted on recovery:

The rubber bands that hold the wing had slipped from its support and had eaten into

the foam all the way till the carbon fibre spar. This had caused the wing to turn

around the spar. This was the flutter that was visible from the ground.

The Aileron servo had lost one of the mounting screws.

Empennage had superficial damage.


Figure 39: the rubber band has eaten half way into the foam wing. The left rubber band has been put back into position

Figure 40: the wing could be actuated about the carbon fibre spar. This is why the flutter was seen from the ground

Figure 41: the missing Servo screw.

10.8) Analysis of the Crash

We carried out a detailed examination of the crash in order to find the reason why it had occurred and

find a solution to prevent it from occurring in subsequent flights

From our study we found two main reasons as to how the crash had been caused:

1. The Aileron servo had lost one of its two mounting screws. This could have caused an

unbalanced force on the ailerons which in turn would have caused excess forces on the wing

and displaced it from its position. This in turn would have caused the rubber bands to slip and

eat into the foam causing failure of the wing

2. The second reason was that the rubber band had slipped from its position and eaten into the

foam wing. This in turn would have caused the flutter of the wing. The flutter in turn would

have pulled on the control rods of the aileron servo and caused a pulling and pushing force on

it. This in turn caused the servo to move and pulled out the servo mounting screw.

Of the two reasons the second one seems more likely.

10.9) Changes in Design

After the crash we made a few changes in the design to prevent the same problems from occurring

again:

1. The size plastic that we wrap around the central portion of the wing was increased from a

span of 10cm to 20cm in order to make sure that there is no way the rubber bands can slip.

2. The Aileron servo mount material was changed from Indian Balsa to 3mm plywood. Also a

layer of hot glue is applied on the surface of the plywood. The two layers ensure that the

servo is held tight and secure.

11) Flight Test Number 5: Gusty Weather Flight

Date: 15th May 2011

Time: 15:45


Weather Conditions: Overcast Skies, High Velocity Gusty Winds

11.1) Setup

1. Wing Span= 85cm, High Aspect Ratio

2. 2300 mah Battery



11.2) Summary

The main aim of the test was to check if the aircraft structure can withstand the extremely

gusty winds and also to evaluate the flight dynamics of the aircraft in these conditions

11.3) Flight Path


2. Carry out a Left turn

3. Take a circuit above the runway

4. Land



11.5) Duration of Flight: 3 seconds

11.6) Comments

1. The aircraft levelled off, elevator control was lost abruptly and the aircraft went into a

shallow dive.

2. Crash landed on the runway

11.7) Crash Analysis

1. As soon as the launcher released the aircraft, a side gust shifted the wing, making it skew

with respect to the fuselage.

2. The trailing edge of the wing jammed the elevator servo, thus fixing the elevator in one

place.

3. This caused the aircraft to go into a shallow dive


1. The wing had displaced from its position, but was structurally alright

2. The coroplast fuselage absorbed the shock of impact well and did not show any sign of failure

3. The propeller shattered into 2 fragments

11.9) Changes to Design: None

12) Flight Test Number 6: First Long Endurance

Flight

Date: 13th May 2011

Time: 12:00


Weather Conditions: Clear Skies, Low Winds

12.1) Setup


2. 2300 mah Battery



12.2) Summary

This was the first flight with the high aspect ratio 85 cm wing. The main aim of the flight was

to fly for 5 minutes, the longest endurance flight till date.

We wanted to check how much the battery would drain by during the flight, check by how

much the components inside the aircraft heat up during a long flight and also check how the

aircraft performs during long duration flights.

12.3) Flight Path




4. Carry out circuits for the 5 minute duration


6. Smooth Landing


1. Flight duration of 5 minutes was achieved




5. A number of circuits were carried out


12.5) Duration of Flight: 6 minutes

12.6) Comments


The aircraft performs well on long duration flights. The pilot was comfortable flying the

aircraft for the long duration. He did not feel the need to continuously give input to the

aircraft thus proving it to be stable which is an advantage for long duration flights as it helps

prevent pilot fatigue from setting in.

For most of the flight we were able to fly at 75% throttle.

The battery and the other electronic components heated up normally. No overheated

components were noted.

The lining up and the landing were carried out successfully.

While landing the aircraft hit a low bush. The impact was on the right wing. When checked it

was noted that no damage had been sustained. The foam wing had absorbed the impact very

well.


There was no damage as such to the aircraft in spite of the wing hitting the bush. The

following was noted:

The wing had displaced slightly at an angle. This was a result of the impact into the

bush. The wing was put back into its correct position immediately with no effort at

all.

No damage to the leading edge which clearly shows that the foam wing is ideal for

impact absorption



A complete check of all the servos and the motor was carried out and no damage was

noted.

13) Flight Test Number 7: Flight Test for

Manoeuvrability

Date: 13th May 2011

Time: 12:20


Weather Conditions: Clear Skies, Low Winds

13.1) Setup


2. 2300 mah Battery



13.2) Summary

The main aim of the test was to check the aircrafts manoeuvrability by making it to carry out

a “Figure of 8” turn followed by a number of circles each lesser in radius. These tests help us

determine the shortest turn radius of the aircraft and help the pilot gain knowledge on how the

aircraft performs and responds while carrying out the manoeuvres.

13.3) Flight Path



3. Carry out a Figure of 8 Turn

4. Try a Zero Radius Turn

5. Carry out a number of circular circuits each with a smaller radius

6. Smooth Landing




3. Two Figure of 8 turns were carried out

4. Zero radius turn was attempted

5. Short radius circuits were carried out


13.6) Comments


The aircraft carried out the Figure if 8 turn with ease.

A zero radius turn was attempted. In order to carry out the turn, the pilot needs to cut back on

the throttle, give a nose up moment through the up elevator and give a full directional aileron.

The resulting forces cause the aircraft to pitch up and roll.

The problem with this maneuver is that the aircraft loses altitude. The turn was attempted and

the aircraft did respond to the inputs and the turn was partially carried out but the aircraft

started to lose altitude so preventive action was taken, the turn was aborted and a circuit was

carried out.

A number of turns were carried out and each of the turns had a smaller radius as compared to

the previous turn. The aircraft handled and responded well.

The lining up and the landing were carried out successfully. The throttle was cut off after

lining up and the aircraft flared about 15 feet above the ground and landed in a perfect

landing.


There was no damage to the aircraft.

14) Flight Test Number 8: First Flight in PESIT with

On-board Camera

Date: 16th May 2011

Time: 13:00

Location: PESIT

Weather Conditions: Clear Skies, moderate Winds

14.1) Setup


2. 2300 mah Battery



5. On board camera

14.2) Summary

The main aim of the test was to fly our first flight over PESIT. The additional payload on this

flight was the pen camera. We wanted to try and take a video from the aircraft of the college

campus in order to retrieve few aerial pictures

14.3) Flight Path


2. Carry out 2 circuits of college

3. Line up and carry out a Smooth Landing



2. The circuits were successful

3. A smooth landing was achieved

4. On board camera successfully recorded a video and images were retrieved.


14.6) Comments

Through this flight we were able to conclude the following:

1. That the pen camera is good enough for the aerial photography that we wanted to achieve.

The video quality was good.

2. The camera placement provided a clear view of the ground and there were no hindrances

during the flight.

We noted that the camera tends to shake when the aircraft is throttled up.


There was no damage to the aircraft or the camera.

Pictures of the college obtained from the on-board camera after processing:

Figure 42: PESIT Garden and Fountain

Figure 43: Professor M.R.Doreswamy Silver Jublee Block (A- Block)

Figure 44: Department of Mechanical Engineering (C-Block)

Figure 45: F Block (left) and Department of Electrical and Electronics Engineering (Right)

Figure 46: Tech Park (E-Block)

Figure 47: Boys Hostel blocks

Pic:

Figure 49: Stitched image showing the new football ground and a part of the cricket ground

Figure 48: Stitched images showing A-block, the fountain, entry way and the Department of Mechanical Engineering, motorcycle parking in a single frame

15) Flight test number 9: 20 Minute Flight

Date: 17th May 2011

Time: 11:00


Weather Conditions: Clear Skies, No Winds

15.1) Setup


2. 2300 mah Battery



5. Camera on Board

15.2) Summary

The objective of this test flight was to see whether the aircraft with this configuration can match the

project requirement of 20 minutes flight time. The aircraft was to be flown just like it would on a real

mission in very large oval shaped patterns and with very gentle turns.

15.3) Flight Path




4. Carry out circuits for the 20 minute duration


6. Smooth Landing





4. A number of circuits were carried out

5. Flight duration of 20 minutes was achieved


7. Very smooth soft belly landing

15.5) Duration of Flight: 20 minutes and 36 seconds

Figure 50: Aircraft taking off

Figure 51: Aircraft on final approach

Figure 52: Aircraft levelling off

Figure 53: Aircraft attempting Flare

Figure 54: Aircraft after landing

15.6) Comments

1. The aircraft matched the project requirement of 20 minute flight time.

2. The flight was carried out with a camera on board.

3. The aircraft was successfully trimmed to fly stick free

4. The aircraft can sustain itself at about 52-53% thrust

5. There was no loss in power for the duration of the flight, i.e. the throttle setting did not

need to be changed once the aircraft was trimmed.

6. The aircraft is very stable when flying in a straight line.

15.7) Damage Reported: None

The team with the aircraft after the successful 20 minute flight

16) Comparison of Google Earth Snapshots with Snapshots Taken From the P❺

The photographs retrieved from the on board camera were compared to

images from Google Earth. We carried out one trial flight with a GPS

system on board. The photographs below show a part of the flight path and

where a snapshot was taken by the UAV, on Google Earth.

Figure 55: Snapshot (12.97108, 77.67567)

Figure 56: Snapshot (12.560668, 77.327044)

Figure 57: Snapshot (12.55549, 77.324981)

17) What Makes the P❺ Different?

The USP of our model is its base cost. Since the aircraft is either made of coroplast or balsa,

and not an expensive, exotic material such as Kevlar or carbon fibre, the base cost of the

model as a whole comes down drastically. A lot of the materials used to build the aircraft

were not specialized aeromodellers’ materials, but things that we adapted such that they

served the purpose. For example, the wing coating where the rubber bands sit is usually

coated with a layer of fibre glass. This gives a very strong smooth finish, but is very

expensive and also has a very large weight penalty. We used a stick file to do the same job,

thus essentially reducing the cost of coating from about 400 rupees to 10 rupees, while having

a lower weight penalty. Of course, the strength provided by the stick file is much lower than

that of the glass fibre, but the strength provided was enough. We have used this same

approach a number of times in numerous places, thus effectively lowering the cost of the

aircraft.

As part of our literature survey, we looked at the cost of similar sized UAVs in the market.

One of the aircraft we found was a low cost mini UAV called Featherlite, manufactured by

Aeroart. This low cost UAV has a wingspan of 1.9m and has an endurance of 1 hour. It costs

5,03,700 Indian rupees (7900 Euros).

The cost break up of our aircraft with an autopilot (Future Work) integrated is as given below

The following is the cost estimate for this aircraft if it is to be sold as a kit:

The aircraft is available in two specifications:

o Balsa fuselage

o Coroplast fuselage

Both the kits will come with the same equipment but would differ only in the fuselage

material and hence their costs

17.1) Cost Split Up:

1. Electronics:

Component Quantity Cost (Rupees)

Tipple 20C 2300 mah 3S 1 1450

EMAX ESC 25A 1 1000

EMAX Grand Turbo

Motor GT2210/13

1 725

Futaba Tx/Rx 4Ch 1 6000

Servo Plastic Gear 2 550

Servo Metal Gear 1 480

3.5 mm Gold Connectors 2 90

TOTAL 10,295

2. Raw material:

o Balsa Fuselage:

The Balsa fuselage requires two sheets of Balsa measuring 1m X 10cm.

Each sheet costs Rs 350. In addition, half a sheet of Balsa is required to build

the Empennage.

Plywood reinforcement at 3 sections along with the motor mount = Rs 50.

Thus material cost for fuselage = Rs 925

o Coroplast Fuselage:

The coroplast fuselage requires sheet of coroplast measuring 60cm X 50cm.

This sheet would approximately cost Rs 50.

Plywood reinforcement at 3 sections along with the motor mount = Rs 50.

The coroplast model also would require half a sheet of Blasa costing Rs 175,

Thus the total cost for the Coroplast fuselage = Rs 275

Foam:

Both the models require foam for building the wing. The approximate cost for foam to

build a wing is around Rs 50

3. Additional Components:

Components Cost (Rupees)

Carbon Fibre Rod 200

Horns and Clevis 25

MonoKote, 2 colours 200

Propeller (2 nos) 180

Cycle Spokes (4 nos) 10

TOTAL 615

4. Miscellaneous Items:

Item Cost (Rupees)

Glue 100

Double Sided Tape 10

Cutter 25

Plywood Sheet (3mm) 50

TOTAL 185

Fuselage Type Balsa Coroplast

Cost (Rupees) 12070 11420

The cost of the aircraft if it is to be sold as a Ready to Fly (RTF) kit is approximately 11000

rupees. This places our coroplast model very competitively in the trainer aircraft class of

remote controlled aircraft.

Another aircraft with specifications similar to ours can be found at the following address:

http://www.jackshobbies.in/products_big.aspx?imgid=1061

The cost breakup of the components if the P❺ was to be converted into a fully-fledged UAV

is as given below

Cost of Camera = Rs 1250

Cost of GPS = Rs 2900

Autopilot = Rs 18,900 (400 USD)

UAV Specific PDA = Rs 50,000

Fuselage Type Balsa Coroplast

Cost (Rupees) 85,120 84,470

As mentioned above, The Featherlite manufactured by Aeroart costs 5,03,700 Indian rupees

(7900 Euros).

The above estimate shows that the cost of our UAV would be less than a fifth that of the

Featherlite. Inspite of having a lower endurance, the fact that our UAV is more compact and

cheaper would place it above the Featherlite when it comes to non-military operations such as

disaster management, vehicular movement tracking updates and animal radio collar tracking.

http://www.jackshobbies.in/products_big.aspx?imgid=1061

18) Future Work

18.1) Autopilot

Every UAV is armed with an autopilot. The autopilot enables the operator to fly the aircraft by only

inputting target coordinates on a computer. The autopilot also assists in stabilizing the aircraft and

optimizing the various systems onboard the UAV. The cheapest quality UAV autopilot available

today is an open source autopilot whose code is available online called ARDUPILOT from

www.diydrones.com. This autopilot costs roughly 18000 rupees and provides both in flight

stabilization and GPS navigation.

Figure 58: Autopilot chipset

18.2) Live telemetry and Video feed

Our aircraft records video and GPS tracking data onboard and this data is recovered when the aircraft

lands. A UAV’s current position needs to be known to the operator while it is in flight, along with a

live video feed of the area it is flying over. The telemetry (tracking) part is taken care of by the

autopilot. The video feed is an optional upgrade with the ARDUPILOT, but costs approximately 5000

rupees more.

http://www.diydrones.com/

18.3) Extension of Flight time

Improvements in the aerodynamics of the aircraft by making it more streamlined, adding appropriate

wingtips, designing a specialized airfoil for our application.

Use of exotic materials such as composites to lower the weight of the structure as well as strengthen it

more.

On board system managers that optimize the energy usage.

18.4) Portability

The whole aircraft along with its controller station needs to be made portable since the range of the

aircraft is limited. Due to this short range the aircraft needs to be taken close to the area that needs to

be scanned and then launched. This means that the operator would have to carry it to such a location.

Thus the design, lightening and strengthening of all the controller and its subsystems needs to be done

keeping this in mind.

19) Conclusion

The challenge of designing, fabricating and flying a mini unmanned aerial vehicle for at least

20 minutes has been achieved.

Over the course of 2 months, 2 separate designs were created on paper and their gliders were

evaluated. The aircraft chosen was then fabricated over the period of a month and the aircraft

first flew on the 10th

of May, 2011 for 18 seconds.

Subsequently a number of flight tests were carried with various configurations to study their

effects on flight dynamics while the aircraft was flying. Each flight pushed the known limit of

the aircraft a little more, whether it was the flight time or the structural strength.

The design objectives of the project were achieved on the 17th

of May, 2011 when the aircraft

flew non-stop for more than 20 minutes.

During the project apart from having fun, we learnt a number of things.

The first and most important thing that we learnt is how to work as a team, then we

learnt how difficult it is to make something fly.

We learnt to use a number of tools for the first time, we learnt the challenges in

fabrication

We learnt how to improve upon our mistakes and we learnt the hard way the truth

behind the adage, “learn from your mistakes”.

Each day was a new challenge and a learning curve and we overcame these challenges

through hard work, team spirit and sheer determination.

This is not the end of this project, but just the beginning. The aircraft is yet to be pushed to its

utter limits, optimized to fly better, and with auto-pilot integration, our aircraft has the

potential to be the finest and cheapest unmanned flight surveillance solution to a number of

demanding situations.

Appendix A: Software Used

SolidWorks 2010

SolidWorks is a 3D mechanical CAD program and developed by Dassault Systems. SolidWorks is

currently used by over 1.3 million designers and engineers in 130,000 companies worldwide.

SolidWorks was the CAD tool of choice for this project. All CAD models and drawings were done

using SolidWorks. SolidWorks was preferred as it has an easy to understand interface that can easily

be used by both amateurs and professionals to create CAD models of objects very easily and with a

very high degree of realism and accuracy in a very short span of time.

The SolidWorks curve wizard is unique to SolidWorks in terms of the ease with which files having

coordinates of the curves can be imported. This curve wizard was especially helpful for us, as we had

to make drawings and models of different airfoils that we were going to use. Many other CAD

software such as SolidEdge were tried but the ease and the functionality of SolidWorks scored over

all other software. The ability of the software to store the models and drawings in different formats

was also of great benefit to us.

SolidWorks is more than just a CAD tool. It can be used to carry out structural and flow anlaysis as

well as computational fluid dynamic (CFD) simulations. Real life conditions and constraints can be

applied to the model and simulated to understand the behaviour of the model.

XFLR5

XFLR5 is an analysis tool for airfoils, wing and planes operating at low Reynolds Numbers. It

includes:

XFoil’s Direct and Inverse analysis capabilities.

Wing design and analysis capabilities based on the lifting line theory, on the vortex lattice method,

and on a 3D plane method.

XFLR5 is capable of plotting an airfoil in 2D by importing the coordinates from a notepad file. This

feature of XFLR was not only made evaluation of airfoil parameters easy but also saved us valuable

time.

XFLR5 was used to evaluate airfoil parameters such as the lift vs angle of attack, the coefficient of lift

vs coefficient of drag etc. The program evaluated these characteristics at different angles of attack,

Reynolds Number and mach number. The results were represented in a graphical form, which made it

easy for us to interpret and understand the airfoil performance.

XFLR5 is available as freeware and is used by students and professionals around the globe to

understand and study the performance parameters of airfoils.

MotoCalc

MotoCalc is a program for predicting the performance of an electric model aircraft power system,

based on the characteristics of the motor, battery, gearbox, propeller or ducted fan, and speed control.

By specifying a range for the number of cells, gear ratio, propeller diameter, and propeller pitch,

MotoCalc can produce a table of predictions for each combination.

MotoCalc can predict weight, current, voltage at the motor terminals, input power, output power,

power loss, motor efficiency, motor RPM, power-loading, electrical efficiency, motor RPM, propeller

or fan RPM, static thrust, pitch speed, and run time. By producing a table of predictions, MotoCalc

helps in determining the optimum propeller size and/or gear ratio for the aircraft.

MotoCalc can also carry out in-flight analysis for a particular combination of components, predicting

lift, drag, current, voltage, power, motor and electrical efficiency, RPM, thrust, pitch speed, propeller

and overall efficiency, and run time at various flight speeds. It can also predict stall speed, hands-off

level flight speed, throttle, and motor temperature, optimal level flight speed, throttle, maximum level

flight speed, rate of climb, and power-off rate of sink.

MotoCalc's graphing facility can plot any two parameters against any other (for example, lift and drag

vs. airspeed), making it easier to interpret performance of the aircraft.

For particular requirements, such as a minimum run time, maximum current, or maximum power loss

(which is dissipated as heat), MotoCalc's filter facility can be used to filter out the unacceptable

combinations.

As we were beginners to electric flight, MotoCalc’s MotoWizard was able to guide us and give

suggestions regarding the ideal power system for our requirements by asking a few questions about

the model and our preferences. The results of the analysis were very detailed and explained in simple

language how the aircraft and the power system would perform under different conditions.

Appendix B: Original Design Delta Wing Plan form

500 mm Wing span

20 Minute Endurance

Wing Loading:

The relationship between wing size, weight, and speed is embodied in the "Great Flight Diagram",

which plots weight against cruising speed shown below.

Fig: The Great Flight Diagram

Weight Estimation by Component Breakup:

Components Weight (grams) Servos 30

GPS 70

Camera 35

Receiver 28

Batteries 100

Motor 145

ESC 25

Structure

Fuselage 72

Wing 73.575

Tail 20

Carbon fibre

2m 10

TOTAL WEIGHT 608.575

This weight is a very rough approximation of the all up weight of the aircraft. Therefore the weight

considered for design was assumed to be double this, due to factors such as reinforcement of the

structure, glue, additional components, wiring etc.

The weight of the aircraft being designed has been approximated to 1.25 Kg

From the Great Flight Diagram, we see that the weight of our aircraft comes in the range of the

Snowy Owl and the Osprey.

Considering the Osprey’s characteristic values from the graph we have:

Cruising speed ≈ 15 m/s

Wing Loading = W/S ≈ 100 N/m2

= 10.1936 Kg/m2

From the wing loading we can calculate the Wing Area (S)

= 1.250 / 10.1936 = 0.12262 m

2

Wing Span (b) has been predefined in the design criteria as 500mm,

Aspect ratio = A.R =

= 0.5

2 / 0.12262 = 2.038736

http://cgi.ebay.in/CORONA-CS-928BB-Analog-Servo-/150550079927?pt=LH_DefaultDomain_203&hash=item230d7be9b7#ht_2981wt_754

http://cgi.ebay.in/USB-GPS-Receiver-Data-Logger-Vehicle-Tracker-BluetootH-/270698136998?pt=LH_DefaultDomain_203&hash=item3f06ddc5a6#ht_7300wt_924

Wing Geometry:

Now the wingspan, surface area and aspect ratio are known.

Taper ratio is the ratio of tip chord to root chord.

We assume the entire wing to be swept with a taper ratio (k) of 0.45;

This is because of structural reasons. If the taper ratio is one, the most amount of lift is produced but

due to a uniformly distributed load the wing becomes weak structurally. If the taper ratio is zero the

wing is structurally very strong but is aerodynamically inefficient since a lot of lift producing span is

lost. A taper ratio of 0.45 is the optimum value taking structural safety and lift producing ability into

account.

Root chord (Cr) =

( )

= 2* 0.12262 / (0.5 * (1 + 0.45))

= 0.338276 m

Tip chord = k * (Cr)

= 0.45 * (0.338276)

= 0.152224 m

Mean Aerodynamic Chord = MAC = (

)

= 0.667 * (0.45 + 1/(1 + 0.45)) * (0.338276))

= 0.2570 m

Lift and Drag:

Calculating Reynolds Number at the cruise speed of 15m/s ;

Re =

= (1.225* 15* 0.2570)/ (2*10-5

)

= 236129.7

Clrequired = ( ) = (1.250 * 9.81) / (0.5* 1.225* (15^2) * 0.12262)

= 0.725624

Airfoil selection

For aircraft that operate in the region of low Reynolds number and have a requirement of high lift,

there is a series of specialized airfoils created by Michael S. Selig. These airfoils are now standard on

this size of aircraft. We have chosen the selig1210 because of its high lift and moderate drag along

with its very gentle stalling characteristics. This implies that the operator would have a large buffer

zone in which he can recover the aircraft if it approaches stall.

Fig: Cl vs Alpha Fig Cl vs Cd

For this value of Clrequired, from the graph the value of alpha initial = -3.5 Degrees

CdInduced = ( ( ) = (0.725624^2) / (3.14* 2.038736)

= 0.082249

Cf = 0 ( ( )) / (Log (2255325.43))2.58

= 0.006002

Cdw = 2

= 0.012003

Fuselage drag calculations:

Fuselage length = 0.5m

Fuselage height = 0.06m

Fuselage width = 0.06m

Fuselage wetted area (A) = 0.075m2

Cdf =

= 0.005943*0.075 / 0.122625

= 0.003671

Overall Drag coefficient:

CdO = Cdw + Cdf = 0.011886 + 0.003671 = 0.015674

Accounting 10% more drag for interference we have;

Cd = CdO + CdInduced = 0.015674 + 0.082249 = 0.099491

Velocity Correction:

Our initial cruise velocity of 15m/s was assumed from the great flight diagram. In the second iteration

we will find the optimum velocities for maximum endurance and range.

Velocity Cl D0 DL D Cd VSink

5 6.530612 0.028234 12.50957 12.5378 6.677229 -5.11225

6 4.535147 0.040658 8.687199 8.727857 3.227899 -4.27051

7 3.331945 0.055339 6.382432 6.437772 1.74926 -3.67498

8 2.55102 0.07228 4.88655 4.95883 1.031607 -3.23512

9 2.015621 0.091479 3.860978 3.952457 0.649676 -2.90089

10 1.632653 0.112938 3.127392 3.240329 0.431424 -2.64247

11 1.3493 0.136654 2.584621 2.721276 0.299435 -2.4411

12 1.133787 0.16263 2.1718 2.33443 0.215841 -2.28446

13 0.966067 0.190864 1.850528 2.041392 0.160825 -2.16417

14 0.832986 0.221358 1.595608 1.816966 0.123426 -2.07442

15 0.725624 0.25411 1.389952 1.644061 0.097286 -2.01108

16 0.637755 0.28912 1.221637 1.510758 0.078572 -1.97122

17 0.564932 0.32639 1.082142 1.408532 0.064891 -1.9527

18 0.503905 0.365918 0.965244 1.331162 0.054702 -1.954

19 0.452258 0.407705 0.866314 1.274018 0.046988 -1.97401

20 0.408163 0.45175 0.781848 1.233598 0.041061 -2.01198

21 0.370216 0.498055 0.709159 1.207214 0.036447 -2.0674

22 0.337325 0.546618 0.646155 1.192773 0.032812 -2.13994

23 0.30863 0.59744 0.591189 1.188629 0.029916 -2.22944

24 0.283447 0.65052 0.54295 1.19347 0.027587 -2.33584

25 0.261224 0.70586 0.500383 1.206242 0.025696 -2.45921

26 0.241517 0.763458 0.462632 1.22609 0.024149 -2.59966

27 0.223958 0.823315 0.428998 1.252312 0.022872 -2.75739

28 0.208247 0.885431 0.398902 1.284333 0.021811 -2.93262

29 0.194132 0.949805 0.371866 1.321671 0.020924 -3.12566

30 0.181406 1.016438 0.347488 1.363926 0.020177 -3.33682

31 0.169891 1.08533 0.325431 1.410761 0.019545 -3.56645

32 0.159439 1.156481 0.305409 1.46189 0.019008 -3.81492

33 0.149922 1.22989 0.28718 1.51707 0.018548 -4.08264

34 0.141233 1.305558 0.270536 1.576094 0.018153 -4.37001

35 0.133278 1.383485 0.255297 1.638782 0.017811 -4.67746

36 0.125976 1.463671 0.241311 1.704982 0.017516 -5.00545

37 0.119259 1.546115 0.228444 1.774559 0.017258 -5.35443

38 0.113065 1.630818 0.216578 1.847397 0.017034 -5.72486

39 0.107341 1.71778 0.205614 1.923395 0.016837 -6.11722

40 0.102041 1.807001 0.195462 2.002463 0.016663 -6.53199

41 0.097124 1.898481 0.186044 2.084524 0.01651 -6.96966

42 0.092554 1.992219 0.17729 2.169508 0.016375 -7.43073

43 0.088299 2.088216 0.16914 2.257355 0.016255 -7.9157

44 0.084331 2.186471 0.161539 2.34801 0.016148 -8.42507

The induced drag is minimum at high velocity and very high at low velocity. The parasite drag is

minimum at low velocity and maximum at high velocity. The total drag caused due to these two

components is minimum at the point of intersection of these two drags. The velocity at the minimum

drag is used to get a rough estimation of the cruise velocity of the aircraft.

The velocity at which the sink rate is minimum is the velocity for maximum endurance, from the

graph this is approximately 18.5 m/s.

0

0.5

1

1.5

2

2.5

3

3.5

0 10 20 30 40 50

D

r

a

g

Velocity

Drag vs Velocity

D0

DL

D

The point of intersection of the curve and the tangent drawn from the origin to the curve gives the

velocity for maximum range. This is approximately 22 m/s.

Taking the velocity value for maximum range and recalculating the Reynolds Number,

Re =

= 1.225* 22*0.2570 / (2*10

-5) = 291226.6

Also recalculating the Cl value for maximum endurance velocity,

Cl = ( ) = 12.2625 / (0.5*1.225*18.52*0.122625)

= 0.477035

Fig: Clvs Alpha

For the new value of Cl, it can be seen that the initial alpha value shifts to -4 degrees (Show by the

orange line)

New value of Drag = 1.3189 N

Power required for cruising = = 1.3189 * 18.5

= 24.4 W

Mission

The aircraft is hand launched from ground level by the operator. The aircraft ascends to four hundred

feet. The aircraft then loiters in the same area for nineteen to twenty minutes taking photographs. It

then descends back to the ground.

Takeoff and climb

The takeoff velocity of the aircraft is approximately 12m/s, 2 cases have been considered for the

climb analysis.

In the first case, the aircraft immediately begins to climb at 12 m/s. In the second case, the aircraft

accelerates to 18.5 m/s(cruise speed) and then begins to climb. The power required to climb to 400

feet is plotted against the time it takes the aircraft to attain this altitude and a suitable time is chosen

such that the power consumption is not very high.

The equation to evaluate power consumed for varying rates of climb is

(

)

W – Weight of the aircraft

v – Velocity

D – Drag

L – Lift

θ - Angle of ascent

The graph clearly shows that the power required to climb at 18.5m/s is about 10 to 20 watts lower

than if the aircraft begins to climb immediately at 12m/s. The graph also seems to almost level off at

about 50 seconds. Therefore a time of 60 seconds is taken is taken to complete the ascent phase of the

flight.

Energy requirements for mission

With an engine which consumes fuel such as aviation fuel or kerosene, the range or endurance of the

aircraft can be estimated by applying the Breguet formula. Our aircraft does not use a consumable

fuel, but rather a battery to power the electric motor, which is our primary thrust producing prime

over. Thus, in this case we estimate the amount of energy (joules) required to complete the mission

and then choose a suitable battery which can provide this amount of energy.

Our mission is broken up into 4 phases.

1) Takeoff

2) Loiter

3) Descent Glide

4) Flare

0

20

40

60

80

100

120

140

160

180

200

0 20 40 60 80 100 120 140

PO

WER

(w

atts

)

TIME (s)

Power v/s Time

18.5 m/s

12 m/s

The energy required at each stage is estimated theoretically, and then added up to obtain the energy

required for the whole mission.

1) Take-off

The power required to reach four hundred feet has been estimated to be 49 watts. The aircraft

is expected to achieve this in 60 seconds. The product of time and power gives the energy

required. This is calculated to be 2925 J.

2) Loiter The aircraft now levels off and circles the area that has to be scanned. The aircraft remains in

this phase for up to nineteen minutes. The power required for sustained flight at 18.5 m/s has

been estimated at 24 watts.

3) Descent Glide

After the aircraft has scanned a particular area, it begins an unpowered glide to the ground.

The aircraft continues in this state till it reaches an altitude of 20 feet.

4) Flare

The aircraft now slows down and increases its angle of attack to 5 degrees. This is a

controlled stall phase where the aircraft nose points up, but the aircraft descends to the ground

very slowly. It is also called feathering. This phase requires a power of 30 watts.

The energy required for the whole mission is estimated to be 31171.87 joules. The battery selection is

done based on this requirement.

Battery Type Rating

(mah)

Battery

Energy(J)

Battery

Efficiency

90%(J)

Propeller

Efficiency

85%(J)

Motor

Efficiency

85%(J)

No. of

Batteries

GB/T118287 SAMSUNG BATTERY 1500 19980 17982 15284.7 12992 3

Sony Erricson BST 38 970 12920.4 11628.36 9884.106 8401.49 4

Nokia bl 5c 1020 13586.4 12227.76 10393.6 8834.557 4

Rcforall lipobattery 1300 51948 46753.2 39740.22 33779.19 1

Nokia BP 4L 1500 19980 17982 15284.7 12992 3

The efficiency of the battery, propeller and the motor are taken into account and 5 lithium polymer

batteries are considered for our energy requirements. The Rcforall battery turns out to be the lightest

and cheapest option. Therefore it is the battery we have chosen.

Stability

The aircraft s designed to be very stable so that the onboard camera does not jerk or shudder, thereby

not compromising on the quality of images taken.

The stability along the lateral axis of the aircraft has been carried out, with a static margin of 10%,

which is ideal for delta wing aircraft. The components have been placed as per space constraints, and

then a dead weight is placed at the aft most position of the aircraft to shift the center of gravity

appropriately.

Component Length

(cm)

Weight

(Grams)

Moment

(gm-cm)

Motor 1.75 145 253.75

Rudder servo+ESC 4.625 20 92.5

Reciever + Camera 9.35 63 589.05

Wing 38.72414 73.575 2849.128

Battery 46.5 93 4324.5

GPS 46.25 75 3468.75

Elevon Servos 23.125 18 416.25

Fuselage 25 72 1800

The above table places the C.G. of the aircraft without the dead weight at 24.65cm from the nose of

the aircraft.

From the position of the wing, the expected center of gravity with a 10% static margin is calculated to

be at 28.15cm from the nose.

Therefore a dead weight of lead is added to the aft most position of the aircraft to pull the C.G. aft so

that the aircraft balances at 28.15cm from the nose. This dead weight has been calculated to be

94gms.

Control Surface sizing:

Aileron Sizing:

The aileron span is taken as 85% of the wetted span.

Therefore Aileron span = ba = = 0.85*(50-6) = 37.4 cm

Taking the ratio of Aileron span to Wing span we get 0.748

Reference: Aircraft Design: A conceptual approach by Daniel.P.Raymer

From the graph for a value of 0.748 we get the ratio of Aileron chord to Wing chord as 0.13

Therefore Aileron chord = 0.13* 0.257012 = 3.341155 cm

Tail Sizing:


From the table above we assume the Tail volume coefficient for a Military cargo/ bomber as 0.08

For delta wings the general practice is to consider the tail area as 25% of the wing area

Therefore, St = 0.25 * 0.122625 = 0.03065625 m2

Calculating Lvt which is the distance between the Aerodynamic centers of the tail and wing.


Lvt = ( ) = 0.08 * 0.5* .122625 / 0.03065625

= 0.16 m

Assuming a taper ratio for the vertical tail as 0.8 and the span to be 15cm we get;

Cr = 0.227083333m

Ct = 0.181666667m

Also, calculating the value for MAC; MAC = 0.205216049m

Considering AC to be 25% of MAC, AC= 0.051304012m

Leading edge of the root chord of the vertical tail is therefore calculated to be at 39.399cm from the

nose.

Glider Fabrication

The main aim of the glider was to ascertain whether the delta wing design was feasible and if the

aircraft showed good gliding characteristics. The glider was a scaled down version of the original

design having a wingspan of 300mm with the other dimensions geometrically scaled down

accordingly.

Material Selection

Since the glider was a scale model, the material that is chosen should be the best trade-off between

weight of the structure of the aircraft and cost required to make the glider.

Based on this selection criterion, three materials were considered for the fabrication of the glider.

They were:

I. Glass Reinforced Plastic (GRP) (Density- 2000 Kg/m3)

II. Balsa Wood (Density – 200 Kg/m3)

III. Corrugated Plastic (CoroPlast) (Density – 1000 Kg/m2)

GRP was the densest and most expensive material, hence was rejected. Balsa Wood, though the least

dense among the three was twice as expensive as a corrugated plastic sheet. Taking these factors into

account, corrugated plastic was the preferred material for the glider.

Wing fabrication

1. The location of the ribs in the wing was taken into account and the chord of each rib was

calculated.

2. A template of each rib was then made using SolidWorks.

3. Using this template, the ribs were laser cut on a 5mm thick sheet of corrugated plastic. The

laser used was a carbon dioxide laser running at 15W.

4. The central spar, leading edge and trailing edge of the wing were made of balsa wood in order

to reduce the structural weight of the wing.

5. To accommodate the central spar, the ribs were arranged up with their trailing edges

positioned along a straight line; the position of the central spar on each rib was marked and a

hole was punched.

6. For the leading edge, a square cross-section balsa rod was taken and sanded into shape. To

accommodate the leading edge, an L shaped cut was made in the front portion of the ribs.

7. The ribs were positioned along the trailing edge and stuck to the balsa using hot glue. The

central spar was then passed through the ribs and glued into place with hot glue. Similarly, the

leading edge was glued onto the ribs, making sure that the leading edge was sitting snugly in

the cuts made in the ribs.

8. Kite paper was used for the outer skin of the wing. Kite paper was used as it is cheap, easily

available and gives a very smooth surface finish.

9. The kite paper was first cut according the size of the wing and stuck onto the ribs using

fevicol. Kite paper once placed over the ribs was completely soaked in water. Kite paper has

the ability to shrink as it dries giving a very smooth and taut outer skin.

10. To facilitate the shrinking, the wing was placed in the sun and left to dry.

Fig: Wing with wet kite paper wrap being left to dry in the sun

Fuselage Fabrication

1. The four side panels of the fuselage along with front and rear pieces to close the fuselage

were drawn on a sheet of corrugated plastic as a development. Each piece was then cut

out from the sheet.

2. The left side panel of the fuselage and the base piece were then placed at a 90-degree

angle and hot glued in place. Similarly, the right panel of the fuselage was placed and hot

glued.

3. To support the side panels of the fuselage and prevent them from caving in, small right-

angled triangular pieces of balsa were placed at regular intervals along the edge where the

panel was glued onto the base.

4. The top of the fuselage was then hot glued in place. To close the fuselage the front and

the rear pieces were hot glued at the front and the rear.

5. To accommodate the wing, holes were made in the side panels at the locations where the

leading edge, the central spar and the trailing edge met the fuselage.

6. Duct tape was used to close the openings formed because of cutting the corrugated plastic

sheet.

Empennage Fabrication

1. The empennage was fabricated using corrugated plastic. The vertical and horizontal tails

were drawn on a sheet of corrugated plastic as a development. These pieces were then cut

out carefully from the sheet.

2. The midpoint of the horizontal tail was marked and the vertical tail was hot glued in

place.

3. To support the vertical tail and provide a larger area to stick the vertical tail to the

horizontal tail, two rectangular pieces of balsa wood were stuck using hot glue at the

junction where the vertical tail joined with the horizontal tail.

4. This assembly was the placed on the fuselage and using a bead of hot glue was stuck

firmly to the fuselage.

Fig: Laser cutting of the ribs

Fig: Kite paper attached to the Rib and spar structure

Fig: Fuselage being drawn on the coroplast sheet

Fig: The panels hot glued with wood supports

Problems faced during glider fabrication

1. The biggest problem we faced was regarding the laser cutting of the ribs from corrugated

plastic. Very few places in Bangalore have machines that are capable of laser cutting

corrugated plastic. After a weeklong search, we were able to locate a company in Whitefield

by the name of ViableAsia, which could help us out.

2. During the initial fabrication phase of the fuselage, the fuselage was cut such that the ridges

were in a direction perpendicular to the ground. With such an arrangement, the ability of the

corrugated plastic to absorb heavy impact was reduced considerably. To overcome this, the

fuselage was cut keeping the ridges parallel to the ground.

3. When the wings were attached to the fuselage the trailing edge of the wing was not snug fit

and had a tendency to slip out even after a bead of hot glue was applied. To prevent this, a

rectangular piece of balsa was inserted in the hole creating a snug fit for the wing.

4. In order to get a smooth surface after shrinking, the kite paper must be stuck very carefully

over the ribs to ensure that it is smooth and there are no ridges or contusions.

Glider Tests

Glider tests were carried out to evaluate the following parameters:

To evaluate the aircraft’s roll and lateral stability.

To evaluate the aircraft glide performance.

To evaluate the ability of the material to absorb the impact of a belly landing.

The glider tests were carried out on a 1:0.6 scale model of the aircraft. The glider tests were carried

out both indoors and outdoors.

Glider Test – I

This test was carried out outdoors. The aircraft’s centre of gravity was found to be lying closer to the

aft of the aircraft making the aircraft tail heavy. To pull the centre of gravity forward, ballast in the

form of coins were added to the nose of the aircraft.

The glider was hand launched from a height of 6 feet.

Evaluation of glider performance

During the initial flight, the glider had a propensity to turn left soon after launch. This was overcome

by trimming the ailerons and by altering the throwing action. After trimming, the glider was launched

again and the glider displayed good gliding performance as well as good roll and lateral stability and

glided a distance of over 12 feet. The glider belly landed on gravel and showed no signs of damage.

Fig 1: The Correct throwing action

Fig 2: Glider after launch

Fig 3: Glider landing in gravel

Glider Test – II

This test was carried out indoors. The aircraft’s centre of gravity was found to be lying closer to the

aft of the aircraft making the aircraft tail heavy. To pull the centre of gravity forward, ballast in the

form of coins were added to the nose of the aircraft.

The glider was hand launched from a height of 10 feet.

Evaluation of glider performance

The glider initially showed good gliding performance but a few seconds into flight, the aircraft stalled

resulting in a spin and a nosedive. The nosedive resulted in a nose first crash into the ground. The stall

was caused because the angle at which the glider was launched was incorrect. No signs of damage

were observed on the glider.

Fig 1: Incorrect angle of launch

Fig 2: Glider stalls in mid flight

Fig 3: Nosedive because of stall

Fig 4: Resulting Crash

The angle at which the glider was launched was corrected and the glider was launched again. This

time the glider showed good gliding performance and glided a distance of 10 feet. The glider landed

smoothly and no signs of damage were observed.

Fig 1: Corrected angle of launch

Fig 2: Glider after launch

Fig 3: Smooth landing

The glider tests conclusively proved that the aircraft exhibits good gliding performance and lateral

and roll stability. The tests also proved that corrugated plastic was capable of withstanding a belly

landing under different ground conditions.

Problems with the Delta Wing configuration:

1. Sweep

There were 2 basic problems with the swept planform

Construction of the internal frame of the swept wing was difficult and time

consuming because if there was any inaccuracy, the pieces would not fit together.

Stall characteristics – Though the swept wing can go to higher angles of attack

without stalling than a conventional planform, the aircraft finds it very difficult to

recover once it has stalled and begins to spin.

Fig :Dh 108 Swallow – Delta Wing Notorious For Spin Instability

2. No weight tolerance

There was no tolerance added for any kind of extra weight in the form of glue, wires,

packaging. Cfd runs on Solidworks showed us that we could expect a lift of 500gms at cruise

condition, whereas our all up weight was 503gms.

3. No Construction tolerance

A small warp or structural defect causes instability in flight. The model does not have

forgiving flight characteristics, in that a small mistake from the pilot puts the aircraft in an

unrecoverable state.

4. Difficulty faced during hand launch:

The high wing loading causes the aircraft to have a very high stall speed of 9 m/s (32.4

kmph), which in turn means that the take-off velocity of the aircraft is about 12 m/s (43

kmph). Since our aircraft is a hand launched aircraft, this take off velocity becomes

impractical.

In order to overcome the above mentioned difficulties:

1) We removed the 500 mm constraint so that the wing loading remains low enough.

2) We changed the planform from a delta to a conventional planform. The conventional

planform includes a wing +horizontal stabilizer + vertical stabilizer. Also the wing would be a

rectangular wing.

3) We started with a 1m wing span, and then practically shortened the span while increasing the

chord length, keeping the wing loading a constant and seeing how short a wingspan we could

achieve so that the aircraft’s flight characteristics suit our requirements.

References

1. Aircraft Design: A Conceptual Approach; Daniel P. Raymer

American Institute of Aeronautics and Astronautics Inc Publication

2. Model Aircraft Aerodynamics; Martin Simons

A Special Interest Model Books Publications

3. Basics Of R/C Model Aircraft Design; Andy Lennon

AirAge Media Publications

4. Aircraft Performance and Design: Dr J.D.Anderson Jr

A WCB/McGraw-Hill Publication

5. Beginners Guide To Radio Control Airplanes, Nickademuss

www.instructables.com

Bibliography

Apart from the above sources, these websites were helpful in understanding the practicalities of

flying and fabricating remote controlled aircraft.

1. www.rcgroups.com

2. www.ebay.in

3. www.rcindia.com

4. www.rcthrustcalc.com

5. www.rcdhamaka.com

6. www.rcforall.com

7. www.bananahobby.com

8. www.instructables.com

http://www.instructables.com/

http://www.rcgroups.com/

http://www.ebay.in/

http://www.rcindia.com/

http://www.rcthrustcalc.com/

http://www.rcdhamaka.com/

http://www.rcforall.com/

http://www.bananahobby.com/

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