INTRA AEROTECH UAV CHALLENGE

DESIGN REPORT

TEAM OJAS

TEAM MEMBERS

Shradha Basti
Dabulu Sahoo
Piyush Mohanty
Abhipsaa Dash
Rajiblochan Pradhan
Sai Smruti Patro
Nishant Barik
Biraj Xess
Alok Kumar Muni
Dibya Ranjan Lenka
Sujit Mohapatra
Shefali Samad
Biswarupa Biswal
Adyasha Nayak
Tarini Moharana
Satyajit Parida

VEER SURENDRA SAI UNIVERSITY OF TECHNOLOGY

BURLA, ODISHA, INDIA, PIN:768018

 CONTENTS

TABLE OF CONTENTS
SL.NO. CONTENTS PAGE NO.
1.0 INTRODUCTION 5
1.1 Aerotech Intra-club Aero Design Competition 2025 5
1.2 OBJECTIVE 5
1.3 MISSION PROFILE 5-6
2.0 EXECUTIVE SUMMARY 6
3.0 DESIGN RESEARCH 6
3.1 AIRFOIL SELECTION 6-8
3.2 PLANFORM SELECTION 8
3.2.1 WING LOADING 8
3.2.2 ASPECT RATIO 8
3.2.3 WING PLANFORM 8
3.3 FUSELAGE 9
3.4 HORIZONTAL TAIL 9
3.5 VERTICAL TAIL 9
3.6-DIMENSIONAL ANALYSIS 9
3.7 ELECTRONIC SYSTEM 10
3.7.1 MOTOR SELECTION 10
3.7.2 SERVO SIZING 11
3.7.3 THRUST ANALYSIS 11
3.8 CAD DESIGN OF AIRCRAFT 12
4.0 ANALYSIS 13
4.1 AERODYNAMICS ANALYSIS 13
4.1.1 TOTAL LIFT 13
4.1.2 TOTAL DRAG 13
4.1.3 STALLING SPEED 13
4.1.4 MOMENT FORCE 14
4.1.5 AERODYNAMIC EFFICIENCY 14
4.2 COMPUTATIONAL FLUID DYNAMICS(CFD) ANALYSIS OF
THE AIRCRAFT
14
4.3 RESULTS 14
4.3.1 OSTWALDS SPAN EFFICIENCY FACTOR 14
4.3.2 INDUCED ANGLE OF ATTACK 15
4.3.3 TOTAL DRAG 15
4.4 STATIC PERFORMANCE 15
4.4.1 THRUST AVAILABLE 15
2
 4.4.2 THRUST REQUIRED 16
4.4.3 POWER REQUIRED 16
4.4.4 POWER AVAILABLE 17
4.4.5 MAXIMUM VELOCITY 17
4.4.6 RATE OF CLIMB 18
4.4.7 GLIDING FLIGHT 18
4.4.8 TIME OF CLIMB 18
4.5 MAXIMUM ALTITUDE 18
4.5.1 ABSOLUTE CEILING AND SERVICE CEILING 19
4.6 DYNAMIC PERFORMANCE 19
4.6.1 TAKE-OFF PERFORMANCE 20
4.6.2 LANDING PERFORMANCES 20
4.6.3 V-n DIAGRAM 20
4.7 LIFTING PERFORMANCE 21
4.8 STABILITY ANALYSIS 21
4.8 STATIC STABILITY 22
4.9 DYNAMIC STABILITY 22
5.0 MANUFACTURING 23
5.1 LASER CUTTING 23
5.2 FUSELAGE 24
5.3 WING 24
5.4 TAIL 24
5.5 LANDING GEAR 25
5.6 ASSEMBLY AND SUB-ASSEMBLE 25
6.0 FLIGHT TESTING 26
7.0 CONCLUSION 26
8.0 REFERENCES 27
9.0 2D DRAWING SHEET 28
10.0 PAYLOAD PREDICTION GRAPH 29

LIST OF FIGURES
SL.NO. CONTENTS PAGE NO.
01 FIG 1.0-AIRFOIL S1223 7
02 FIG 2.0 CAD DESIGN OF AIRCRAFT 12
FIG 2.1 TOP VIEW 12
FIG 2.2 SIDE VIEW 12
FIG 2.3 ISOMETRIC VIEW 12
FIG 2.4 FRONT VIEW 12

3
03 FIG 3.0 COMPUTATINAL FLUID DYNAMIC(CFD) OF COMPLETE 14
AIRCRAFT USING ANSYS
FIG 3.1 PRESSURE CONTOUR (SIDE VIEW) 14
FIG 3.2 VELOCITY STREAMLINES (SIDE VIEW) 14
FIG 3.3 PRESSURE CONTOUR (ISOMETRIC VIEW) 14
FIG 3.4 VELOCITY STREAMLINES (REAR VIEW) 14

LIST OF TABLES
SL.NO CONTENTS PAGE NO.
01 TABLE 1.0 - AIRFOIL SELECTION 6
02 TABLE 2.0 - WING SPECIFICATIONS 8
03 TABLE 3.0 - DIMENSIONS OF DIFFERENT ASPECTS OF AIRCRAFT 9-10
04 TABLE 4.0 - MOTOR SPECIFICATION 10
05 TABLE 5.0 - THRUST TABLE 12

LIST OF GRAPHS
SL.NO CONTENTS PAGE NO.
01 GRAPH 1.0 - AIRFOIL SELECTION USING XFLR 5
GRAPH 1.1 (Cl vs Cd) 7
GRAPH 1.1 (Cl vs α) 7
GRAPH 1.3 (Cd vs α) 7
GRAPH 1.4 (Cl/Cd vs α) 8
GRAPH 1.5 (Cm vs α) 8
02 GRAPH 2.0 - DRAG POLAR 15
04 GRAPH 3.0 - THRUST REQUIRED 16
05 GRAPH 4.0 - POWER REQUIRED 17
06 GRAPH 5.0 - POWER Vs. AIRCRAFT SPEED 17
07 GRAPH 6.0 - TIME OF CLIMB 18
08 GRAPH 7.0 - Absolute and Service Ceilings 19
09 GRAPH 8.0 – v-n GRAPH 20
10 GRAPH 9.1 ROLLING MOMENT CURVE 22
GRAPH 9.2 PITCHING MOMENT CURVE 22
11 GRAPH 10.0 PAYLOAD PREDICTION GRAPH 29

4
1.INTRODUCTION
1.1 Aerotech Intra-club Aero Design Competition 2025
AeroTech Intra-Club Design Challenge 2025 is a dynamic initiative hosted by the senior
members of the AeroTech Club, aimed at immersing newly joined members in the
exciting world of aircraft design. This hands-on event not only fosters collaboration
among teammates but also offers an invaluable opportunity for inductees to identify
their strengths and areas for improvement. With guidance from experienced mentors,
participants gain practical exposure and build foundational skills crucial for future
aerospace endeavours.

1.2 OBJECTIVE

To engineer an aircraft that achieves maximum lift capacity while staying within strict
spatial and weight limitations. Designs should emphasize a low empty weight and a high
payload fraction to enhance overall efficiency.

Key Constraints & Specifications

• Dimensional Limit:
The combined sum of the aircraft’s length (nose to tail), width (wingtip to
wingtip), and height (ground to highest point) must not exceed 180 cm.

• Weight Requirement:
The aircraft’s total weight must be 1.5 kg or less, excluding payload.

1.3 MISSION PROFILE

The design process starts with a literature review on conceptual and preliminary design
considerations. After data acquisition, the parameters were iterated for the detailed
design of the aircraft, and performance criteria were optimized.

Conceptual Design: After reading several design reports, we have come to the point to
decide upon the aircraft design & configuration.

➢ Rectangular wing configuration - for better aerodynamic efficiency.

5
 ➢ T-wing Configuration - reducing the turbulence effect by fuselage at a higher
angle of attack; improved lateral flight control; decrease in overall height to
increase aspect ratio and wingspan to improve lift.
➢ Flaps used as a high lift device.
➢ Tapered fuselage - to reduce skin/parasitic drag.

2.EXECUTIVE SUMMARY

The primary objective of this initiative is to develop essential engineering skills through
the innovative design and construction of a radio-controlled aircraft. This project
encourages the application of advanced techniques and strategic thinking to deliver a
design that excels in both performance and efficiency. A critical goal is to create an
aircraft capable of earning points across multiple phases of a competitive event—
demonstrating superior lift capabilities while maintaining a low structural weight. This
delicate balance between payload optimization and minimal self-weight reflects
technical proficiency and thoughtful design.

3.DESIGN RESEARCH
3.1 AIRFOIL SELECTION
At 25 m/s, the Reynolds number is calculated as:

For stall conditions in our mission profile, the average Reynolds number is 184,634. We
conducted thorough aerofoil research to achieve high lift and gentle flight stability
under these conditions. Selection was based on key aerodynamic and practical factors,
including:

Sl.No Air-foil Specifications

Max thickness 9.1% at chord
1 (e387-il) E387
Max camber 3.2% at 44.8% chord
Max thickness 8.7% at 27.9% chord
2 (ag35-il) AG35
Max camber 2.3% at 37% chord
Max thickness 14.7% at 29.7% chord
3 (mh113-il) MH 113
Max camber 6.4% at 47.7% chord
Max thickness 12.1% at 19.8% chord
4 Selig S1223 (s1223-il)
Max camber 8.1% at 49% chord
TABLE 1.0 AIRFOIL SELECTION

6
- Lift and drag performance

- Stall behaviour and pitching moment

- Structural and manufacturing feasibility

- Aerodynamic centre location

- Flutter, vibration, and operational environment

- Compatibility with the target Reynolds range

The analysis conducted using XFLR5 software identified the Selig S1223 (s1223-il)
airfoil as the most favourable.

Fig 1.0 Selig S1223 (s1223-il) airfoil

GRAPH 1.1 (Cl/Cd vs α) GRAPH 1.2 (Cl vs α) GRAPH 1.3(Cd vs α)

7
 GRAPH 1.4(Cl vs Cd) GRAPH 1.5(Cm vs α)
GRAPH 1.0 - AIRFOIL SELECTION USING XFLR 5

3.2 PLANFORM SELECTION

A rectangular wing has a constant chord from root to tip, cost-effective, and offer safe,
predictable stall behaviour—ideal for training and model aircraft.
It provides uniform lift distribution across the span and supports stable flight
characteristics—ideal for low-speed handling. Based on XFLR5 analysis, the 2D lift
coefficient was found to be 2.25, leading to a required wing area of 0.225 m², as
calculated using the standard lift equation.
3.2.1 WING LOADING

1.1 𝑘𝑔
𝑊𝑖𝑛𝑔 𝑙𝑜𝑎𝑑𝑖𝑛𝑔 = = 4.8 𝑘𝑔/𝑚2
0.225 𝑚2

3.2.2 ASPECT RATIO

1.05 𝑚
𝐴𝑠𝑝𝑒𝑐𝑡 𝑅𝑎𝑡𝑖𝑜 (𝐴𝑅) = =5
0.21 𝑚
3.2.3 WING PLANFORM

SL.NO. WING SPECIFICATIONS MEASUREMENTS

01 Wingspan 1.05m
02 Wing Area 0.225m2
03 Wing Load 4.8kg/m2
04 Aspect Ratio 5
05 Mean Aerodynamic Chord 0.21m
06 Taper Ratio 1
TABLE 2.0 WING SPECIFICATIONS

8
3.3 FUSELAGE
The fuselage serves as the aircraft's central body, housing all on board components.

𝑇𝑜𝑡𝑎𝑙 𝑙𝑒𝑛𝑔𝑡ℎ = 0.7 × 1.05 𝑚 = 0.74 𝑚

3.4 HORIZONTAL TAIL

Maintains aircraft’s longitudinal stability and pitch control.
0.5 × 𝑀𝐴𝐶 × 𝑊𝐴 𝑂. 5 × 0.2𝐼 × 0.225
𝐻𝑜𝑟𝑖𝑧𝑜𝑛𝑡𝑎𝑙 𝑇𝑎𝑖𝑙 𝐴𝑟𝑒𝑎 (𝐻𝑇𝐴) = = = 0.053𝑚2
𝑇𝑀𝐴 0.44
The dimensions of elevator area, it controls the lateral movement of the aircraft.
𝐸𝑙𝑒𝑣𝑎𝑡𝑜𝑟 𝑎𝑟𝑒𝑎 = 40% × 𝐻𝑜𝑟𝑖𝑧𝑜𝑛𝑡𝑎𝑙 𝑇𝑎𝑖𝑙 𝐴𝑟𝑒𝑎 (𝐻𝑇𝐴) = 0.021𝑚2
Where, HTA=Horizontal Tail Area
MAC=Wing’s Mean Aerodynamic Chord (meters)
WA=Wing Area (in sq. Meters)
TMA=Tail Moment Arm (in meters)
3.5 VERTICAL TAIL
Directional stability, achieved through a vertical tail.
𝑉𝑒𝑟𝑡𝑖𝑐𝑎𝑙 𝑇𝑎𝑖𝑙 𝐴𝑟𝑒𝑎(𝑉𝑇𝐴) = 35% × 𝐻𝑇𝐴 = 0.018𝑚2
Rudders help in control of rotation of aircraft in the vertical axis.
1
𝑅𝑢𝑑𝑑𝑒𝑟 𝐴𝑟𝑒𝑎 = × 𝑉𝑒𝑟𝑡𝑖𝑐𝑎𝑙 𝑇𝑎𝑖𝑙 𝐴𝑟𝑒𝑎 = 0.062𝑚2
3
3.6 DIMENSIONAL ANALYSIS
S.NO. STRUCTURAL COMPONENTS DIMENSIONS
01 Total Length 1.17m
02 Fuselage Length 0.74m
03 Wingspan 1.05m
04 Wing
Average chord Length 0.21m
05 Wing Area 0.225m2
06 Aspect Ratio 5
07 Flaperon 0.21m
08 Horizontal Tail Area 0.532m2
09 Horizontal Stabilizer Area 0.319m2
10 Horizontal Stabilizer Aspect Ratio 4
11 Horizontal Stabilizer Span 0.35m
12. Horizontal Stabilizer
Average Chord Length 0.08m
Root Chord 0.114m

9
 Tip Chord 0.057m
13 Elevator Area 0.0212m2
14 Elevator Width 0.059m
15 Vertical Tail Area 0.0186m2
16 Vertical Stabilizer Area 0.0124m2
17 Vertical Stabilizer Span 0.14m
18 Vertical Stabilizer
Root Chord 0.1m
Tip Chord 0.05m
Taper Ratio 0.5
19 Rudder Area 0.0062m2
20 Rudder Aspect Ratio 3.5
21 Rudder Span 0.147m
22 Rudder Chord 0.042m
TABLE 3.0 DIMENSIONS OF DIFFERENT ASPECTS OF AIRCRAFT

3.7 ELECTRONIC SYSTEM

The propulsion system was carefully engineered to deliver sufficient thrust to counteract drag
during cruise while maintaining a lightweight profile. A 1250 KV brushless motor, rated for 3S
operation, was selected for its peak current of 34 A and a maximum power output of 500 W.
This motor drives a 10-inch diameter propeller with a 5-inch pitch, powered by a 3S Li-Po
battery with a capacity of 2200 mAh and a voltage of 11.1 V.

An electronic speed controller (ESC) rated for a burst current of 60 A was integrated, ensuring
seamless compatibility with the power system. The design prioritizes endurance, guaranteeing
a minimum operational time of three minutes to support multiple high-speed circuits with an
adequate safety buffer.

3.7.1 MOTOR SELECTION

S.no. Motor Propeller Size Thrust(G) Power(W)

01 T-Motor AT4125 540 KV 10×5 5200 1123
02 DYS D3536-6 1250 KV 10×5 2470 166.5
03 Sunnysky X2216 880 KV 10×5 1850 127.7
TABLE 4.0 MOTOR SELECTION

The DYS D3536-6 (1250KV) brushless motor was selected due to its high thrust output (~2.47
kgf) when paired with a 10×5” propeller on a 3S battery, which comfortably exceeds the total
aircraft weight of 1.5 kg. It offers a safe thrust-to-weight ratio (~1.64), supports efficient cruise

10
performance with manageable current draw (~15A), and is compatible with a wide range of
ESCs and propellers. Its proven reliability and availability make it an ideal choice for
competition UAVs requiring both power and stability.

3.7.2 SERVO SIZING

Maximum torque required for our constants is given by the following equation: - Torque
required = 8.5×10-6[C2 V2 L sin(S1)tan(S1)/tan(S2)] C = Control surface chord in cm, L =
Control surface length in cm, V = Speed in MPH, S1 = Max control surface deflection in
degrees, S2 = Max servo deflection in degrees For optimal performance and effective
maneuvering, the control surface chords are set to one-third of the main stabilizer's
chords, resulting in the following dimensions. Elevator as 8.12 cm × 60.96 cm and
Rudder as 6.09 cm × 22.60 cm. Performance analysis indicates a minimum take off
speed of 25 m/s, which equals approximately 55.9 MPH. Using the equation above, the
calculated maximum torque required at take-off conditions is:

Elevator: 0.67 kg.cm and Rudder: 0.14 kg.cm

Both values fall well below the selected servo's maximum rated torque of 2.2 kg·cm,
providing a safe operational margin and ensuring reliable surface actuation during flight.

3.7.3 THRUST ANALYSIS

To meet the competition's performance and stability requirements, an in-depth thrust

analysis was conducted based on motor-propeller characteristics and aerodynamic

considerations. A DYS D3536-6 1250KV brushless outrunner motor was selected for its

reliable performance in fixed-wing UAV applications. The motor is powered by a 3S

(11.1V) LiPo battery and paired with a 10×5 inch propeller, chosen after evaluating

thrust, current draw, and efficiency across multiple configurations.

Static Thrust Evaluation:

Experimental data and community benchmark tests for the DYS D3536-6 motor indicate

that with a 10×5” propeller at full throttle, the system produces approximately:

11
Thrust: ~1.9–2.0 kgf (18.6–19.6 N)

Current Draw: ~15A

Power Consumption: ~165W

T = TCρV2D2

TABLE 5.0 THRUST TABLE

RPM/V 1250

Diameter×Pitch 10× 5

Thrust 312gf

3.8 CAD DESIGN OF AIRCRAFT

FIG 2.1:Top View FIG 2.2:Side View

FIG 2.3:Isometric View FIG 2.4:Front View

12
4.0 ANALYSIS

4.1 AERODYNAMICS ANALYSIS

Aerodynamics analysis was employed to systematically evaluate and refine the aircraft

design, ensuring optimal performance and stability.

4.1.1 TOTAL LIFT

1
𝑇𝑜𝑡𝑎𝑙 𝑙𝑖𝑓𝑡 = Ρ𝐴𝑉 2 𝐶𝑑 = 96.04𝑁 (𝑤ℎ𝑒𝑟𝑒, 𝐶𝑙 = 1.75)
2

4.1.2 TOTAL DRAG

1
𝑇𝑜𝑡𝑎𝑙 𝐷𝑟𝑎𝑔 = 𝜌𝐴𝑉 2 𝐶𝑑 = 1.09 𝑁 (𝑤ℎ𝑒𝑟𝑒, 𝐶𝑑 = 0.02)
2

4.1.3 STALLING SPEED

It is the minimum velocity required by an aircraft to maintain its levelled flight.

2×𝑊
𝑆𝑡𝑎𝑙𝑙𝑖𝑛𝑔 𝑆𝑝𝑒𝑒𝑑 (𝑉𝑠𝑡𝑎𝑙𝑙𝑖𝑛𝑔 ) = √ = 1.88 𝑚𝑠 −1 (𝑤ℎ𝑒𝑟𝑒, 𝐶𝑙𝑚𝑎𝑥 = 2.25)
𝜌 × 𝐴 × 𝐶𝑙𝑚𝑎𝑥

Reynolds Number at stalling speed=1,50,000

4.1.4 MOMENT FORCE

Co-efficient of moment (𝐶𝑚𝑎𝑐 ) = is 𝐶𝑚𝑐𝑔 at 0 lift 𝐶𝑚𝑎𝑐 = −0.12

1
So, 𝑀𝑎𝑐 𝜌 × 𝐴 × 𝑉 2 × 𝐶𝑚𝑎𝑥 × 𝑐 = 1.38 𝑁𝑚 (𝑖𝑛 𝑛𝑒𝑔𝑎𝑡𝑖𝑣𝑒)
2

4.1.5 AERODYNAMIC EFFICIENCY

13
 𝐿𝑖𝑓𝑡 𝐹𝑜𝑟𝑐𝑒
𝐴𝑒𝑟𝑜𝑑𝑦𝑛𝑎𝑚𝑖𝑐 𝐸𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑐𝑦 = = 16.5
𝐷𝑟𝑎𝑔 𝐹𝑜𝑟𝑐𝑒

4.2 COMPUTATIONAL FLUID DYNAMICS(CFD) ANALYSIS OF THE

AIRCRAFT

FIG 3.1: Pressure (Isometric View) FIG 3.2: Velocity (Isometric View)

FIG 3.3: Pressure (Side View) FIG 3.4: Velocity (Side View)

4.3 RESULTS
4.3.1 OSTWALDS SPAN EFFICIENCY FACTOR
𝐶𝑙2
= 0.09
𝜋×𝑒×𝐴×𝑅
⇒ 𝑒 = 1.1𝜋 × 5 × 0.09 = 0.85

𝐻𝑒𝑟𝑒, 𝑒 = 𝑂𝑠𝑡𝑤𝑎𝑙𝑑’𝑠 𝑆𝑝𝑎𝑛 𝐸𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑐𝑦 𝐹𝑎𝑐𝑡𝑜𝑟.

4.3.2 INDUCED ANGLE OF ATTACK

𝐶𝑙2 𝐶𝑙2 𝐶𝑙2
𝐷𝑖𝑛𝑑𝑢𝑐𝑒𝑑 = ⇒ (𝐶𝑙 )(𝛼𝑖 ) = ⇒ (𝛼𝑖 ) = = 0.07 𝑟𝑎𝑑𝑖𝑎𝑛𝑠
𝜋×𝐴×𝑅 𝜋×𝐴×𝑅 𝜋×𝐴×𝑅
= 4.01°

4.3.3 TOTAL DRAG

14
𝑇𝑜𝑡𝑎𝑙 𝐷𝑟𝑎𝑔 (𝐷) = ½ 𝜌𝐴𝑉 2 𝐶𝑑 (𝑤ℎ𝑒𝑟𝑒, 𝐶𝑑 = 𝐶𝑑𝑜 + 𝐶𝑑𝑖 )
1 1 𝐶2
= 2 𝜌𝐴𝑉 2 𝐶𝑑𝑜 + 2 𝜌𝐴𝑉 2 𝜋×𝑒×𝐴×𝑅
𝑙

= 8.38 N (𝐶𝑑𝑜 = Profile Drag) + 10.48 N (𝐶𝑑𝑖 = Induced Drag) = 5.8 N (Total Drag)

GRAPH 2.0:Polar Drag

4.4 STATIC PERFORMANCE

4.4.1 THRUST AVAILABLE
𝑇ℎ𝑟𝑢𝑠𝑡
Propellor efficiency of 100% throttle = = 6.12
𝑀𝑒𝑐ℎ𝑎𝑛𝑖𝑐𝑎𝑙 𝑃𝑜𝑤𝑒𝑟

𝑃𝑎𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒 = 𝜂 × 𝑃𝑠ℎ𝑎𝑓𝑡 = 6.12 × 500 = 3060 𝑊

(From Motor Specification Table)

𝑃𝑎𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒 3060
So, 𝑇ℎ𝑟𝑢𝑠𝑡 𝐴𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒 = = = 122.4 𝑁
𝑉∞ 25

𝑑 3.5
𝐹 = 4.392399 × 10−8 . 𝑅𝑃𝑀 (4.23333 × 10−4 . 𝑅𝑃𝑀. 𝑝𝑖𝑡𝑐ℎ − 𝑉0 )
√𝑝𝑖𝑡𝑐ℎ

4.4.2 THRUST REQUIRED

Thrust Required (to overcome drag)

2
𝐶𝑙2
𝑇 𝑟𝑒𝑞. = ½ 𝜌 𝐴 𝑉 (𝐶𝑑𝑜 + )
𝜋×𝑒×𝐴×𝑅

15
 2
2𝑊 2
= ½ 𝜌 𝐴 𝑉 𝐶𝑑𝑜 + )
𝜌 × 𝐴 × 𝑉2 × 𝜋 × 𝑒 × 𝐴 × 𝑅
63.43
= 0.002(𝑉 2 ) +
𝑉2

GRAPH 3.0:Thrust Required

4.4.3 POWER REQUIRED

𝐶𝑙2
(where, 𝐶𝑑𝑖= )
𝜋×𝑒×𝐴×𝑅

= 31.25 watt (zero lift power required) + 2.53 watt (lift induced power required)

1
So, Minimum Power Required, 𝐶𝑑𝑜 = 𝐶
3 𝑑𝑖
1
(Power Required)min = 𝜌𝐴𝑉 2 (𝐶𝑑𝑜 + 𝐶𝑑𝑖 )
2
= 31.25 (0.12) = 3.75 watt

GRAPH 4.0: Power Required

4.4.4 POWER AVAILABLE

16
 𝑃𝑜𝑤𝑒𝑟 𝐴𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒 = 𝑇ℎ𝑟𝑢𝑠𝑡 × 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦

𝑦 = 1.05 × 10−5 − 1.412 × 10−5 𝑥

GRAPH 5.0: Power Vs. Aircraft Speed

4.4.5 MAXIMUM VELOCITY

1
2 2
𝑇𝐴 𝑊 𝑊 𝑇 4×𝐶
(𝑊 ) max × ( 𝑆 ) + ( 𝑆 ) √(𝑊𝐴 ) 𝑚𝑎𝑥 − 𝜋 × 𝑒 × 𝐴𝑑𝑜× 𝑅
𝑉𝑚𝑎𝑥 = = 233.69 𝑚𝑠 −1
𝜌∞𝐶𝑑𝑜
[ ]

4.4.6 RATE OF CLIMB

𝑅 𝑀𝑎𝑥. 𝑃𝑜𝑤𝑒𝑟 𝐴𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒−𝑀𝑎𝑥.𝑃𝑜𝑤𝑒𝑟 𝑅𝑒𝑞𝑢𝑖𝑟𝑒𝑑 𝑡𝑜 𝑂𝑣𝑒𝑟𝑐𝑜𝑚𝑒 𝐷𝑟𝑎𝑔
Rate of Climb (𝐶 ) = 𝑊𝑒𝑖𝑔ℎ𝑡

490.25−33.78
= = 42.34𝑚𝑠 −1
1.1×9.8

4.4.7 GLIDING FLIGHT

𝑊𝑒 𝑘𝑛𝑜𝑤 𝑡ℎ𝑎𝑡, 𝐷 = 𝑊 (𝑠𝑖𝑛 Ѳ), 𝐿 = 𝑊 (𝑐𝑜𝑠 Ѳ)
1 1
So, 𝑇𝑎𝑛(Ѳ) = 𝐿 => (Ѳ)𝑚𝑖𝑛 = 𝑡𝑎𝑛−1 ( 𝐿 )𝑚𝑎𝑥 = 5.37°
𝐷 𝐷

In a Gliding Flight, lift can be written as

1
𝑊𝑐𝑜𝑠𝜃 = 2 𝜌𝐴𝑉 2 𝐶𝑙

17
 2 × 𝑊𝑐𝑜𝑠𝜃
𝑉=√ = 0.68°
𝜌 × 𝐴 × 𝐶𝑙

4.4.8 TIME OF CLIMB

1
It is the area between altitude and 𝑅 .
( ⁄𝐶 )
𝑚𝑎𝑥

From the graph below we found it to be 6.9 seconds.

GRAPH 6.0:Time of Climb

4.5 MAXIMUM ALTITUDE

4.5.1 ABSOLUTE CEILING AND SERVICE CEILING
For Absolute Ceiling, R/C = 0
𝑀𝑎𝑥.𝑃𝑜𝑤𝑒𝑟 𝐴𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒−𝑀𝑖𝑛.𝑃𝑜𝑤𝑒𝑟 𝑡𝑜 𝑜𝑣𝑒𝑟𝑐𝑜𝑚𝑒 𝑑𝑟𝑎𝑔
= 0 ms-1
𝑊𝑒𝑖𝑔ℎ𝑡

P Ract = P Ro (ρo / ρ)½ => ρ = 0.006 kgm-3

(at 34,000m)

For Service Ceiling, R ⁄ C = 100 ft/min = 0.508 m/sec

𝑀𝑎𝑥.𝑃𝑜𝑤𝑒𝑟 𝐴𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒−𝑀𝑖𝑛.𝑃𝑜𝑤𝑒𝑟 𝑡𝑜 𝑜𝑣𝑒𝑟𝑐𝑜𝑚𝑒 𝑑𝑟𝑎𝑔
= 0.508ms-1
𝑊𝑒𝑖𝑔ℎ𝑡

P req. act = P req. o (ρo / ρ) ½ => ρ = 0.44 kgm-3

(at 32,000m)

18
 GRAPH 7.0:Absolute and Service
Ceilings

4.6 DYNAMIC PERFORMANCE

Aircraft performance during take-off, ascent, and landing was assessed, focusing on
propulsion efficiency, aerodynamics, and structural integrity. The landing gear is
optimized for safe, efficient runway operations, with key emphasis on power-to-weight
ratio and aerodynamic lift.

4.6.1 TAKE-OFF PERFORMANCE

16ℎ 2
( )
Ground Effect Factor (𝜙) = 𝑏
16ℎ 2
= 0.98
(1+ )
𝑏

h - height of the wing above the ground

b – wingspan (Given by McCormick Theory)

1 2
1 2
𝐶𝑙2
𝐷 = 𝜌𝐴𝑉 𝐶𝑑𝑜 + 𝜌𝐴𝑉 Ø = 5.26𝑁
2 2 𝜋×𝑒×𝐴×𝑅
V Take-off = 1.2VStall = 1.2 × 6.35 =7.62 𝑚𝑠 −1

The take-off velocity is generally 20% higher than stalling velocity for safety margin.

S Take-off = = 8.07 ms

4.6.2 LANDING PERFORMANCES

V Landinig = V Stall *1.3 = 1.3*1.88 = 2.44 ms-1
19
V Landing avg = 0.7 *V landing = 1.708ms-1
L = ½ × 1.225 × 0.5475 (6.48)2 × 0.3 = 0.08 N
D = ½ ρ A V2 C do + ½ ρ A V2Ø (Cl2 /π × e × A × R) =0.04 N
4.6.3 V-n DIAGRAM
Load Factor (n) = Lift/Weight

GRAPH 8.0: V-n Graph

4.7 LIFTING PERFORMANCE

In our pursuit of heavy-lift aircraft capabilities, we are targeting a payload

fraction of 80% and a maximum cubic loading of 3.0. Following an
evaluation of an aircraft that reached these targets with a modest cubic
loading of 2.76, we refined the design to enlarge the wing area as much as
possible while adhering to defined limitations, including container size
requirements.

4.8 STABILITY ANALYSIS

The stability of an aircraft refers to its ability to maintain or return to a state

of equilibrium after a disturbance. This characteristic is analysed from two
perspectives: Static Stability, which concerns the initial tendency of the
aircraft to return to equilibrium, and Dynamic Stability, which involves the
motion over time as it attempts to return to that state.

XFLR5 is a powerful tool designed for analysing aerofoils, wings, and

complete aircraft configurations operating at low Reynolds numbers. It
employs methods such as the Lifting Line Theory (LLT) and the Vortex

20
Lattice Method (VLM) to evaluate aerodynamic performance with a focus
on low-speed flight.

Using XFLR5, one can comprehensively assess an aircraft's stability in all

three principal axes:

• Longitudinal Stability (pitching),

• Lateral Stability (rolling),
• Directional Stability (yawing).

This enables the design and optimization of aircraft to ensure balanced and
controllable flight characteristics.

4.8 STATIC STABILITY

Static stability refers to a UAV’s inherent tendency to maintain or return to a

stable flight position following a small disturbance, without requiring
continuous corrective control inputs. It is a fundamental characteristic
that influences how naturally an unmanned aerial vehicle (UAV) holds its
intended flight attitude.

Several key factors affect static stability, including:

• Wing Configuration and Airfoil Design – influencing lift distribution

and aerodynamic balance.
• Centre of Gravity (CG) – a forward CG generally enhances stability,
while an aft CG can reduce it.
• Stabilizing Features – such as horizontal and vertical tail surfaces,
which generate restoring moments to counteract disturbances.

21
 GRAPH 9.1 Rolling Moment Curve GRAPH 9.2: Pitching Moment Curve

Proper consideration of these design aspects ensures that a UAV exhibits

predictable and controlled behaviour in flight, especially during low-speed
or autonomous operations.

4.9 DYNAMIC STABILITY

Dynamic stability examines the time-dependent behaviour of an aircraft

after it has been disturbed from its equilibrium position. Unlike static
stability, which only indicates the initial tendency to return to equilibrium,
dynamic stability determines how the aircraft responds over time—
whether it returns smoothly, oscillates, or diverges from its original state.

This analysis involves solving the equations of motion derived from the
aircraft's aerodynamic and inertial properties. By calculating the roots of
the characteristic equation, one can determine the nature of the
response:

• Stable (damped oscillations or exponential decay),

• Marginally stable (sustained oscillations),
• Unstable (growing oscillations or divergence).

Understanding dynamic stability is crucial for ensuring that a UAV or

aircraft maintains controlled flight over time, especially in response to
atmospheric disturbances or control input changes.

22
5.0 MANUFACTURING
The manufacturing phase marked the transition from design to reality, emphasizing
precision, structural integrity, and functional performance. Advanced fabrication
techniques and high-quality materials were employed to ensure the construction of a
durable and efficient aircraft.

5.1 LASER CUTTING

Aircraft components were meticulously modelled using SOLIDWORKS CAD software.

Precision cutting of 3mm aero-ply sheets was carried out using a laser cutter, ensuring
high dimensional accuracy and clean edges. One of the key challenges addressed
during this stage was identifying optimal laser settings for different wood thicknesses
while incorporating appropriate tolerances for larger structural cutouts. Despite these
complexities, the laser cutting process—focusing primarily on the fuselage and wing
ribs—was successfully completed within a two-day timeframe.

5.2 FUSELAGE

The fuselage construction began with reinforcing the central circular frame using four
primary hollow aluminium rods, which provided structural support and rigidity. These
were supplemented with 40 cm birch wood rods to enhance the framework. A total of
17 custom-cut birch wood plates were integrated, supported by secondary ribs
designed to reduce overall weight without compromising structural strength. The design
also accounted for internal routing of wiring and optimal placement of electronic
components, contributing to improved aerodynamic stability, functional efficiency,
and enhanced durability during landings or minor impacts.

5.3 WING

Spanning 1.05 meters, the wing was fabricated using 3mm aero-ply, with ribs precisely
laser-cut to achieve aerodynamic accuracy. These ribs were spaced assembled using
lightweight hollow aluminium rods for added stiffness and alignment. Both right and
left wings were joined securely with nuts and bolts, ensuring structural integrity and
ease of assembly. Additionally, thin carbon fibre rods were employed to mount the
flaps and ailerons, offering high strength-to-weight ratios and responsive control
surface movement.

5.4 TAIL

An Inverted T-tail configuration was selected to optimize lift and maintain clean airflow
over the horizontal stabilizer. The tail assembly was crafted as a unified structure from

23
3mm aero ply, ensuring both durability and aerodynamic precision. The horizontal and
vertical stabilizers were connected through a robust extended aluminium rod
extending from the fuselage. Elevators and rudders were then mounted using carbon
fibre rods, allowing for precise control inputs and enhanced manoeuvrability.

5.5 LANDING GEAR

Landing gear configurations generally fall into two primary categories: the tricycle
model and the taildragger model. After conducting extensive structural analysis and
simulations on both types, the team opted for the tricycle landing gear, owing to its
superior stability during taxiing, take off, and landing. The landing gear assembly was
fabricated by bending a durable flat aluminium sheet using a precision bending
machine. The final gear height was set at 17.5 cm from the ground, incorporating
rubber wheels with a diameter of 10 cm to provide effective shock absorption and
ground clearance.

5.6 ASSEMBLY AND SUB-ASSEMBLY

The aircraft assembly followed a three-plane fixture methodology, ensuring alignment

accuracy and structural precision. Construction began with the integration of major
sub-assemblies: the fuselage, wing, and tail sections. Laser-cut 3 mm aero-ply
panels were securely joined using a combination of birch wood rods and hollow
aluminium rods, which formed the skeleton of the fuselage.

The electronic systems were systematically embedded within the fuselage, with
careful planning for cable routing and component access. For the wing, aerofoil-
shaped ribs were reinforced and positioned with exact spacing, while the horizontal
and vertical stabilizers were assembled using an extended aluminium rod for central
alignment and rigidity.

Once sub-assemblies were complete, the wing and tail sections were mounted onto
the fuselage. Wiring for servos and power systems was finalized, followed by
verification of control surface movements. The landing gear was installed, and all
major electronic placements were optimized relative to the aircraft’s Center of
Gravity (CG).

To further enhance the aircraft's surface finish and strength, the entire frame was
wrapped with a 1 mm balsa sheet. A final inspection was carried out to ensure that all
electronic systems were operational, and the aircraft exhibited balanced lateral and
longitudinal stability, confirming its readiness for the SAE DDC competition.

24
6.0 FLIGHT TESTING

The testing phase commenced with a full-scale prototype to evaluate airworthiness

and identify potential design adjustments. Following initial assessments, a powered
version of the aircraft underwent multiple flight trials under varying payload
conditions.

Early tests highlighted areas for improvement, particularly in aerodynamic stability,

control responsiveness, and landing gear resilience. In response, iterative
modifications were made, resulting in enhanced flight performance during
subsequent trials.

A master flight test schedule was developed and followed rigorously, ensuring
structured data collection, consistent analysis, and effective performance
benchmarking. These efforts culminated in a well-balanced, stable, and fully functional
aircraft, meeting the project’s goals and competition requirements.

7.0 CONCLUSION
After numerous design iterations, extensive simulations, and rigorous testing, Team

AEROTECH is confident in the aircraft’s capability to meet and exceed the objectives of

the SAE DDC competition. The integration of a high-lift aerofoil, lightweight materials,

and robust structural components has resulted in a well-engineered platform

optimized for performance, stability, and reliability.

Throughout the development process, the team encountered and overcame several

technical and logistical challenges. These were addressed through collaborative

problem-solving, external mentorship, and the application of core engineering

principles gained at Veer Surendra Sai University of Technology. This journey not only

strengthened the team's technical skills but also fostered resilience, adaptability, and

teamwork.

In conclusion, the final aircraft stands as a testament to the team's dedication,

innovation, and countless hours of hard work. It reflects not only their technical
25
proficiency but also serves as a strong representation of the engineering talent and

academic rigor of their university.

26
27
10.0 REFERENCE
➢ INTRODUCTION TO FLIGHT 8TH EDITION by J.N. Anderson

➢ by Martin Simons

➢ by J.N. Anderson

➢ by Andy Lennon

➢ : LINK-> http://www.airfoiltools.com/

➢ ->
• https://youtube.com/playlist?ist=PLFW6IRTa1g83B1HdU2mece6QLeBrts
pL7&si=4Rt0WzCDpXOTSIEX
• https://youtube.com/playlist?list=PLOzRYVm0a65ey1nPnbhfrz59Hvu-
NVob7&si=zVELbc9VvCnIUMi
• https://youtube.com/playlist?list=PLwr4eb5N5Dr5dKUghcHOuRxCkTMs
MFSsM&feature=shared

➢ ->
https://youtube.com/playlist?list=PLtl5ylS6jdP6uOxzSJKPnUsvMbkmalfKg&si=RIvD

28
Appendix A – Technical Data Sheet

Team Name: OJAS

College Name: Veer Surendra Sai University of Technology, Burla

Shown below is the the plot of density altitude vs weight of the airplane, depicting the

predicted payload capacity of the airplane. The steps taken to create the plot are

discussed.

GRAPH 10.0: PAYLOAD PREDICTION GRAPH

29
30