The document details the design report for Team Merlin's electric motor-powered model aircraft created for the SAE Aero Design Challenge 2019. It outlines the team's objectives, design specifications, aerodynamic analysis, material selections, and manufacturing processes, emphasizing their commitment to excellence and innovation in UAV design. The report serves as documentation of their compliance with competition rules and showcases their engineering efforts over a six-month period.

1. 1
Society of Automotive Engineers
Aero Design Challenge 2019
(SAE-ADC 2019)
Design Report of an Electric Motor Powered
Radio Controlled Heavier-than-air Model Aircraft
Submitted by:
TEAM MERLIN (ADC20190152)
Nitte Meenakshi Institute of Technology, Bangalore
SAE AERO DESIGN CHALLENGE 2019

2. 2
STATEMENT OF COMPLIANCE
Certificate of Qualification
Team Name: TEAM MERLIN
Team No. ADC20190152
College: Nitte Meenakshi Institute of Technology, Bangalore
Faculty Advisor: Mr. Siddalingappa P K
Faculty Advisor’s Email: siddalingappa.pk@nmit.ac.in
Statement of Compliance
As a faculty advisor, I certify that the registered team members are enrolled in collegiate courses.
This team has designed and constructed an electric motor-powered radio-controlled airplane, for
their participation in SAE Aero Design Challenge 2019, without direct assistance from
professional engineers or RC model experts and pilots.
Mr. Siddalingappa P K,
Assistant Professor
Department of Aeronautical Engineering
Nitte Meenakshi Institute of Technology, Bangalore

3. 3
Contents
1. About TEAM MERLIN..........................................................................................................................6
1.1 Our mission....................................................................................................................................6
1.2 Our vision.......................................................................................................................................6
1.3 Current design ...............................................................................................................................6
1.4 Meet the team...............................................................................................................................7
2. Current aircraft specifications.............................................................................................................7
3. Roadmap of the design process of individual systems ............................................................................8
3.1 Aerodynamic design.......................................................................................................................8
3.1.1 Selection of airfoil for wing section................................................................................................8
3.1.2 Selection of tail-plane airfoil........................................................................................................ 10
3.1.3 Calculation of basic aircraft dimensions subject to rulebook constraints......................................... 11
3.1.4 Wing analysis in Xflr5 .................................................................................................................. 12
3.2 Fuselage Design........................................................................................................................... 14
3.3 Propulsion system........................................................................................................................ 15
3.3.1 Selection of propeller................................................................................................................... 15
3.3.2 Selection of motor and Battery..................................................................................................... 16
3.4 Avionics and control systems ....................................................................................................... 16
3.5 Materials used ................................................................................................................................... 17
3.6 Landing gear design ..........................................................................................................................17
4 Numerical Performance analysis of Systems..........................................................................................18

4. 4
4.1 Computation of total drag coefficient using finite wing theory....................................................... 18
4.2 Calculation of induced angle of attack .......................................................................................... 18
4.3 Calculation of wing loading.......................................................................................................... 19
4.4 Computation of aircraft thrust required ......................................................................................... 20
4.5 Computation of take-off performance-ground roll......................................................................... 21
4.6 Structural analysis of aircraft components................................................................................... 21
4.7 Structural analysis of thrust plate ................................................................................................. 21
5. Manufacturing process ...................................................................................................................... 23
5.1 Fuselage....................................................................................................................................... 23
5.2 Wing and control surfaces............................................................................................................ 24
5.3 Tail............................................................................................................................................... 25
6. Payload prediction graph .................................................................................................................. 27
Summary............................................................................................................................................... 28
Innovations ........................................................................................................................................... 28
Conclusion............................................................................................................................................. 28
References ............................................................................................................................................ 29

5. 5
List of Figures
Figure No. Title Page No.
1.0 Final model 6
3.1 Selig 1223 airfoil 9
3.2 Eppler 423 airfoil 9
3.3 Aerodynamic performance of selected airfoils 9
3.4 NACA 0015 airfoil 10
3.5 Coefficient of Lift vs alpha for NACA 0015 11
3.6 Streamline and Cp distribution around wing at 10 degrees
angle of attack
12
3.7 Wingtip vortices 13
3.8 Local lift distribution over finite wing 13
3.9 Xlfr5 analysis data for finite wing 14
3.10 Fuselage modeled using CATIA 15
3.11 Graph of PWM control [7]
17
3.12 Material used 17
4.1 Thrust required vs Free Stream Velocity 20
4.2 Graph of Thrust Required Vs Velocity (MATLAB) 20
4.3 Von Mises nodal stress values color map with 50 N force on
shaft axis
22
4.4 Translational displacement vector 22
5.1 Laser cut fuselage pieces 23
5.2 Completed fuselage assembly 24
5.3 Manufactured wing ribs and tail ribs using Laser cut 24
5.4 Assembly of the right Wing 25
5.5 Pair of Wings with Fuselage and Tail 25
5.6 Completed tail assembly minus vertical stabilizers 26
5.7 CG location with Payload 26
5.8 CG location without Payload 27
6.1 Payload 27

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1. About TEAM MERLIN
Team Merlin is a student team from the Department of Aeronautical Engineering, NMIT
specializing in the design, development and testing of UAVs ranging from quadcopters to fixed
wing aircraft. We also participate in UAV design challenges such as the SAE aero design challenge
2018-2019.The team was formed specifically to participate in the above event in the regular class.
Our name is a reference to both the famous Rolls Royce merlin engine as well as the famed wizard
of Arthurian legends.
1.1 Ourmission
Our current mission is to design and development of a Radio controlled aircraft for the SAE Aero
design challenge and secure a victory by dint of our Plane’s design, which has been a product of 6
months of hard work, late nights, good engineering, vigorous debate and insightful analysis.
1.2 Our vision
Our long-term goal is to popularize the hobby of RC flying and design of RC aircraft not just in
the department but in the larger student engineering community as it is the need of today's highly
sophisticated world of UAVs used for ever increasing number of tasks.
1.3 Current design
Fig 1.0: Final Model
The above image features our UAV design for the aero design challenge coming up in July. The
model is designed to be capable of carrying payloads of about 3-4.5 Kg. It features a twin boom

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configuration with rear pusher props (not visible in the current rendering) and can be adapted for
carrying medical payloads or for agricultural usages. This design is the product of maximizing the
maximum payload available given the dimensions as per the competition guidelines. Future
upgrades include but are not restricted to metal and composite construction, usage of nitro engine,
smart payloads, aerial mapping systems etc.
1.4 Meet the team
At the end of the day, any successful organization is a function of its team members who are all
striving for a shared common goal of excellence, unwavering determination and commitment to
the highest ideals of intellectual honesty and engineering ethics.
Sl. No Member Designation
1 Shaurya Gupta Team Captain and Structural Engineer
2 Guruprasad S Aerodynamicist, Report, Social Media
3 Mayur R Mahale Aircraft Performance Analyst and Designer
4 Mahashana Structural Engineer
5 Jatin N Manufacturing Engineer
6 Jnanesh Kamath Manufacturing Engineer, Report, Accounting, Social Media
7 Bhaskar Manufacturing and Logistics
Table 1.1: Team members and work distribution
2. Current aircraft specifications
Our model is a high wing twin boomed design with the propeller in the pusher configuration
between the tail booms. Materials of construction are Balsa wood and Aeroply using
cyanoacrylate, Fevicol and araldite as adhesives.
Aircraft sub-Systems Configuration
Wing  Configuration: high wing, straight, Rectangular wing
Span: 2.245m
 Chord:0.53m
 Airfoil: Eppler-423
 Root chord:0.53m
 Tip chord: 0.53m
 Aspect ratio: 4.49

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 Planform area: 1.12 sq. m
Tail plane Span:0.856m
 Chord: 0.22 m
 Planform area:0.18832m^2
Fuselage  Type: prismatic frame with square cross section
 Length:
 Cross section:
Propulsion  Power plant: Aeolian 600kv BLDC motor
 Propeller: 3 bladed 16 x 8
 Thrust coefficient: 0.98
 Power coefficient:0.86
 Battery: 22.2v ,25-30c 6S lipo battery pack(6000 Mah)
UAV Avionics  Hobby wings 100 A electronic speed controller(ESC)
 Servos (6nos)
 RC transmitter
 Receiver
Materials  Base materials: Balsa wood, Aero ply
 Adhesive : Araldite, CA glue, fevicol
Landing gear  Tricycle configuration, with steerable nose gear at the front of the
fuselage and two gears attached to the rear end of the fuselage.
Table 2.1: Current aircraft specifications
3. Roadmap of the design process of individual systems
3.1 Aerodynamicdesign
The end goal of the aerodynamic design for this competition is to generate maximum amount of
lift possible given the competition specifications in order to carry the highest possible payload that
is structurally safe and flyable. Our central focus was to select an airfoil with the highest possible
lift coefficient to enable highest lift force for a given chord and span.
3.1.1 Selection of airfoil for wing section
After using the UICC airfoil library database [15]
of airfoils, we shortlisted two airfoils i.e., Selig
1223 and the Eppler 423 based on their high values of maximum lift coefficient. These airfoils

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belong to a class of high lift airfoils producing good lift performance even at low Reynolds number.
The selected airfoil profiles are shown below.
Fig 3.1 Selig 1223 Airfoil
Fig 3.2 Eppler 423 Airfoil
We performed an analysis of these airfoils using xflr5-a popular panel method solver. The obtained
results are compared in the plots given below;
(a) (b)
Fig 3.3 Aerodynamic performance of selected airfoils: (a) Cl vs alpha and (b) Cl/Cd vs alpha for
E423 (blue) and S1223 (yellow)

10. 10
From the analysis, it is observed that E423 has a higher glide ratio than the S1223. This is preferred
as it can be utilized for the smooth landings even when the throttle is cut and the aircraft is made
to glide smoothly.
Conclusions after analysis:
 S1223 offers higher lift coefficient for the same angle of attack but stalls violently in some
Reynolds number regimes.
 The trailing edge of S1223 is thin and is at risk of breaking during maneuvers. It’s also poses
significant difficulties in manufacturing.
 Eppler 423 offers slightly lower lift coefficient but has stable value after 10 degrees. In
addition.
 E423 has a better glide ratio as compared to the S1223.
With these considerations in mind, we chose the E423 airfoil over the S1223.
3.1.2 Selection of tail-plane airfoil
It is usual practice to use symmetrical airfoils of around 12 -15% thickness in the tail section. We
decided to use NACA 0015 for its reasonable values of lift coefficient, which should be high
enough to counteract the aerodynamic moment about the main wing quarter chord.
Fig 3.4: NACA 0015 Airfoil

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2
Fig 3.5: Coefficient of lift vs alpha for NACA 0015
3.1.3 Calculation of basic aircraft dimensions subject to rulebook constraints
Based on the rulebook constraint that the sum of the wingspan, length and height from the
ground should not be beyond 170 inches, we decided the wingspan first because of the priority
being payload capacity. Through iterative calculations, we decided to use a wingspan of
2.23(87.795 inches) m and a chord of approximately 0.529 m(20.8268) which collectively give
us an area of 1.179 Sq. m due to straight non tapered wing configuration(planform area = Span ×
chord).A wing of this configuration would give us a lift force of
L =
1
⍴𝑣2
𝑆𝐶𝑙
2
Where ⍴ - density of air
V- Free stream velocity
S- Planform area
Cl- Lift coefficient
In cruise condition,
L= W, Where W is the weight of the aircraft
For our aircraft,
L=W=
1 1.225×11 1.179×1.25

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L=109.223 Newton
L=11.134 Kg f
This gives us a total aircraft mass of 11.134 Kg which includes the payload.
The next step was to decide the length of the fuselage based on the length of the battery,
specifications of payload bay(10 inches) .The total length from the tip of nose cone to the
trailing edge of vertical stabilizer is about 1.28m (50.44 inches).
Length + Wingspan + minimum height from the ground=170 inches
Substituting above values into the equation, we get:-
Minimum height from the ground= 31.765 inches
3.1.4 Wing analysis in Xflr5
The aerodynamic properties of finite wings differ significantly from that of its constituent airfoils
due to the effect of downwash created by wing tip vortices, which reduce the total angle of attack
(α) by an angle called induced angle of attack or αi. To investigate the effects of these aerodynamic
phenomenon on our UAV, analysis was performed on Xflr5 and the results are presented below.
Fig 3.6: Streamline and Cp distribution around wing at 10-degree angle of attack

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Fig 3.7: Wing tip vortices
Fig 3.8 Local lift distribution over finite wing

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Fig 3.9: Xlfr5 analysis data for finite wing
From the finite wing analysis, the lift coefficient is found to be around 1.25 at 7 degrees angle of
attack. This agrees with the theoretical calculation made from Aerodynamic theory in the
performance analysis section of this report.
3.2 Fuselage Design
The fuselage of the aircraft will be carrying the payload, battery, motor and other electronic
systems as well as serve as a mount for the landing gear. The fuselage has to withstand heavy
loading especially during takeoff and landing. Here the aero ply has chosen to be the fuselage
material due to its high bending stiffness and strength compared to Balsa wood.
The fundamental design of our fuselage is based on the competition specifications for the payload
bay dimensions of 4 x 4 x 10 inches. We decided to create a simple square cross section of four
aero ply plates with interlocking box joints. The plates were stuck using araldite as an adhesive.
The 3D model was created on CATIA V5.

15. 15
Fig 3.10: Fuselage modeled using CATIA
The structural analysis was performed on this model using the structural analysis toolbox in
CATIA V5 and is documented in the performance analysis section of this report.
3.3 Propulsion system
3.3.1 Selection of propeller
Based on the requirement of thrust to maintain steady flight as well as ground clearance
considerations, 16 x 8 three-blade propeller has been selected. Over the period of our literature
survey, it was noted that for the given weight category of the UAV, the propeller used happened
to be of the following configuration: Diameter = 15 to 18 Inch and Pitch = 0.5D.
In addition, the ratio of pitch to diameter obtained with the above-specified configuration, it gives
the most efficient propeller performance during steady cruise flight with a fixed pitch propeller
driven aircraft. Therefore, we considered a propeller available in the market of the nearest available
configuration.
Blades: Two with Diameter =18 inch and Pitch = 8 inch,
By applying the thrust formula with the following values from the specifications and requirement,
Diameter D= 18 inch; Pitch, p= 8inch; Propeller rotation = 6500 RPM; for static thrust, the
propeller forward air speed be V0 = 0;

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𝑑3.5
𝐹 = 4.3923 × 10−8
× 𝑅𝑃𝑀
√ 𝑝 (4.233 × 10−4
× 𝑅𝑃𝑀 × 𝑝 − 𝑉0)
We get 𝐹 = 55𝑁 = 5.5 𝐾𝑔
However, due to the lack of ground clearance in our twin boom design, a 3-blade propeller was
chosen. For the same power consumption by a 3-blade propeller as that of a 2-blade propeller and
the pitch was assumed same.
The relation is found to be: 𝐷 = 𝐷 (
𝐵1
)
𝐵2
1⁄4
Where D2 is the diameter of 3 bladed propeller
B2 is the number of blades = 3
D1 is the diameter of 2 blade propeller
B1 is number of blades = 2
3.3.2 Selection of motor and Battery
The selection of motor is based on the torque required to use the 16 x 8 inch 3 bladed propeller,
which in turn was selected based on thrust required to lift our payload. From the available motor
in the market, we used an Aeolian 600kv BLDC motor for the aircraft,
22.2v ,25-30c 6S lipo battery pack for powering the motor and the electronic speed controller is
used, which will give us a flying time of approximately 5 minutes without any failures and for the
electronics and servos, a separate battery of 11.1 volts 1100mAh was chosen which will
also function without failures.
3.4 Avionics and control systems
Electronic speed controller or ESC is the control system used to control the BLDC via pulse width
modulation or PWM .We selected the Hobby wings 100A ESC because of the maximum current
draw of the motor which is in the range of 100 A.
2 1

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Fig 3.11: Graph of PWM control[7]
An example of PWM in an idealized inductor driven by a voltage source modulated as a series of
pulses, resulting in a sine-like current in the inductor.
3.5 Materials used
The fabrication of Designed RC airplane is done using balsa wood and aeroply together. Aeroply
is used for the fuselage and wing spars, while the ribs will be cut from balsa wood. This gives us
the right mixture of strength and weight saving which is necessary for the competition as our
structural weight should not exceed 5kg.
Fig 3.12: Material used (a) Balsa wood of thickness (3mm and 6mm), (b) Aero ply of thickness
(3mm and 2mm)
3.6 Landing gear design
The tricycle arrangement has a single nose wheel in the front and two or more main wheels slightly
aft of the center of gravity. Tricycle gear aircraft are the easiest to take-off, land and taxi, and

18. 18
𝐿
a
consequently the configuration is the most widely used on aircraft. Our model will be having on
steerable nose wheel on the fuselage and two more on the rear end of the fuselage.
4 Numerical Performance analysis of Systems
4.1 Computation of total drag coefficient using finite wing theory
There is a difference in aerodynamic performance of airfoils (infinite wing) and real world finite
wing. This is due to various factors such as downwash, aspect ratio, ground effects, wingtip
vortices etc. The total drag coefficient (not including effect of tail) is computed from the
aerodynamic theory of finite wings, which are presented below [1]
.
𝐶 = 𝐶 𝐷,0 + 𝐶 𝐷,𝑖
Where CD - Total drag coefficient
CD, 0 - Drag coefficient at zero lift = 0.02
CD, i - Induced drag coefficient and 𝐶 𝐷,= 𝐾𝐶2
CL - Lift Coefficient = 1.2447
K =
1
𝜋𝑒𝐴𝑅
We get CD = 0.1598
4.2 Calculation of induced angle of attack
The presence of downwash over a finite wing creates downwash, which reduces the effective angle
of attack of the wing and creates a component of induced drag. This can be visualized as a tilting
back of the lift vector. This results in the formation of wingtip vortices.
The Induced AOA is given by 𝛼 = 𝐿
𝜋𝐴𝑅
Induced Lift coefficient: 𝐶 = (1+ ak)
𝑖

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Where, k =
1
(e×A×R×π)
a
𝐶 =
(1 + a/(e × AR × π))
e = 0.75 to 0.8 (Ostwald's efficiency factor)
Aspect Ratio (AR) = 4.41
0.5
a =
0.08726
= 5.73
a 5.73
𝐶 =
(1 + ak)
=
(1 + (5.73 × 1/(0.75 × 4.41 × π)
= 3.694
At maximum Angle of Attack 27°
𝐶 𝐿,𝑚𝑎𝑥 =
(3.964 × 27 × π)
180
= 1.74
=
1.74
=0.12259 rad or 7.2°
4.41×π
4.3 Calculation of wing loading
Wing loading is the total weight of an aircraft divided by the area of its wing. The stalling speed
of an aircraft in straight, level flight is partly determined by its wing loading. An aircraft with a
low wing loading has a larger wing area relative to its mass, as compared to an aircraft with a high
wing loading. The faster an aircraft flies, the more lift can be produced by each unit of wing area,
so a smaller wing can carry the same mass in level flight. Consequently, faster aircraft generally
have higher wing loadings than slower aircraft. This increased wing loading also increases takeoff
and landing distances. A higher wing loading also decreases maneuverability,
Wing loading factor: n =
w
… … … . . kg/m^2
A
(10 × 9.81)
n =
1.12
n=87.589 N/m2
or 8.928 kg/m2
𝑖

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4.4 Computation of aircraft thrust required
The thrust required is an important parameter in aircraft performance and has to be matched by the
thrust available produced by the engine. The lowest point of the thrust required curve translates to
the highest glide ratio (Cl/Cd) value attainable [1]
.
Fig 4.1: Thrust required vs Free Stream Velocity
Using the drag values we computed in the previous section, we plotted the graph of thrust required
vs free stream velocity using MATLAB. The results are given below,
Fig 4.2: Graph of Thrust Required Vs Velocity (MATLAB)

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The data follows the trend described by the theory and validates our choice of propeller and motor
to provide the necessary thrust available to sustain steady un-accelerated flight.
4.5 Computation of take-off performance-ground roll
The distance between the starting point of an aircraft and the point at which its wheels completely
leave the ground is called ground roll[1]
. Takeoff distance is the sum of ground roll and the distance
taken to cover an obstacle of standard size designated separately for civilian and military aircraft.
For The competition however, a calculation of ground roll is sufficient. The ground roll we have
calculated is less than the maximum runway length available thus validating our designs capability
to take off within given constraints.
Ground roll is given by Sg =
1.21(W/S)
ρ∞×g×(T/W)×𝐶L max
Sg =
1.21(9.81×10/1.12)
1.225×9.81×(4.5/10)×1.74
Sg = 36.95 ft
This is well within the prescribed runway length of 200 feet for the regular class and is thus
acceptable.
4.6 Structural analysis of aircraft components
We used FEM software to do our structural analysis. These software have stood the test of time
and are used everywhere in the industry today as an aid to the design process along with other
technologies like CFD(computational fluid dynamics) and CIM(computer integrated
manufacturing).Our analysis is performed on the structural analysis toolbox of CATIA V5 which
is widely used in the aerospace industry.
4.7 Structural analysis of thrust plate
Structural analysis of thrust plate with 50 N thrust force acting on the motor shaft axis has been
done.

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As we can see from the results of the FEM analysis, the peak displacement value is of the order of
a hundredth of an mm which a negligible is considering the excellent bending properties of
Aeroply. We can conclude that the thrust plate can safely withstand the thrust forces.
5. Manufacturing process
There is a world of difference between the processes of the computer aided design phase of any
project and the manufacturing phase. Manufacturability was always a key priority for us during
designing the aircraft, it was the primary reason we opted out of the S1223 airfoil in the first place.
We have currently fabricated all our sub-assemblies and are waiting for electronic components to
arrive. We have made extensive use of laser cutting technology to give us the accurate profiles
designed in CAD as well as obtain the components with the desired tolerances so that the fits are
neither too loose nor too tight.
5.1 Fuselage
The fuselage is a 4 piece square assembly. We obtained aeroply sheets and got them laser cut as
per our design.
Fig 5.1: Laser cut fuselage pieces
We performed the assembly by gluing the box joints with araldite and after it had dried, we used
CA glue to seep into the cracks and provide additional strength, The glue is several time stronger
than the material itself giving us great confidence in the joint, we then treated the wood with a

24. 24
locally sourced wood finisher and left it overnight and found that its strength had improved
considerably.
Fig 5.2: Completed fuselage assembly
5.2 Wing and control surfaces
Fig 5.3: Manufactured wing ribs and tail ribs using Laser cut
The wing and control surfaces were by far the most laborious and exacting tasks in the entire
manufacturing process and involved all the team members. The process was similar to the tail
assembly, using spacer jigs to align the ribs and finally glued with CA glue. The completed
assemblies of left and right wings were treated with a wood treatment compound and left to cure.

25. 25
Fig 5.4: Assembly of the right Wing
Fig 5.5: Pair of Wings with Fuselage and Tail
With all these various sub-assemblies completed, only the complete assembly remains and we
remain quietly confident in the next phase of our efforts i.e., testing and analysis of stability and
control.
5.3 Tail
After the fuselage, we proceeded to assemble our tail by lining up our balsa tail ribs with spacer
jigs and then passing balsa spars through them. We used CA glue to secure the spars and then
inserted soft balsa stringers and began sanding the protruding portion of the stringers to follow the
airfoil geometry.

26. 26
Fig 5.6: Completed tail assembly minus vertical stabilizers
Vertical stabilizers were then inserted on the left and right side of the tail and glued with CA
adhesive.
The location of center of Gravity with and without payload is as shown below:-
Fig 5.7: CG location with Payload

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Fig 5.8: CG location without Payload
6 Payload prediction graph
Fig 6.1: Payload Fraction Vs Density Altitude

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Summary
This section succinctly summarizes our effort to produce a payload lifting RC aircraft under the
given design constraints of the competition that stands apart from other teams in terms of design,
build quality, innovations and most importantly: payload carrying capacity.
Innovations
Innovative features of our model are:
1. Simplest possible 4 piece fuselage assembly with excellent strength which facilitates rapid
loading and unloading of payload through the front portion which has a detachable nose.
2. Usage of high lift airfoil (E423) in the wing to give us good payload carrying capacity.
3. Steerable landing gear in the nose which gives us excellent control during takeoff, landing
as well as taxiing.
4. Increased ground clearance by using 3 blade propeller of smaller diameter instead of using
a larger equivalent two bladed propeller for a given thrust.
5. Usage of pusher configuration: The main advantage of pusher airplane is that the position
of the propeller, right behind the fuselage, increases the overall efficiency of the plane by
reducing the profile drag.
6. Highly optimized design of ribs and fuselage which save a lot of weight and increase
payload fraction by strategically creating pockets in non-stressed regions while
maintaining safe amount of strength
7. Reduced fuselage length and thus weight savings are obtained by using twin boom
configuration.

29. 30
Conclusion
To conclude. We have successfully managed to design and fabricate a radio controlled aircraft in
the regular class having a total weight of about 10 kg and a payload carrying capacity upwards of
4.5 kg. We hope to see the aircraft perform well in the upcoming competition in July. This has
been a learning experience for us as aeronautical engineering students and we are immensely proud
of our work and confident in the capability of our UAV.
References
[1].”Aircraft performance and design” -John D Anderson Jr, Tata McGraw Hill Publications, 2010
[2].”Fundamentals of Aerodynamics”- John D Anderson Jr, Tata McGraw Hill Publications, 2010
[3].”Mechanics of flight”- AC Kermode -Himalayan books,2004 edition
[4].Getting started with MATLAB-quick introduction for scientists and engineers -Rudra pratap
[5].Oxford University press
[6].“THE WRIGHT STUFF” final report for SAE West by the Department of Mechanical
Engineering, Northern Arizona University
[7].https://www.ecalc.ch/ -eCalc online RC calculator for electrical systems
[8].https://aerotoolbox.net
[9].SAE Aero Design Challenge 2017 - Design Report from NIT Calicut
[10]. ”Goat works” Worcester polytechnic Institute (WPI) -SAE Aero Design East 2012 Micro
Class Design Report.
[11]. http://www.xflr5.com/xflr5.htm
[12]. https://www.rcbazaar.com

30. 31
[13]. Reza, Mirza Md Symon & Mahmood, Samsul Arfin & Iqbal, Asif. (2016). Performance
Analysis and Comparison of High Lift Airfoil for Low-Speed Unmanned Aerial Vehicle.
10.5281/zenodo.1468120.
[14]. https://m-selig.ae.illinois.edu/ads/coord_database.html UICC airfoil database
[15]. http://airfoiltools.com/
[16]. Lakshmi GS, Balmuralidharan P, Sankar G, K Selvaraj, N Balachandran, “High Lift Two-
Element Airfoil Design for MALE UAV Using CFD”
[17]. https://forum.flitetest.com/index.php