AIAA Design Build & Fly Design Report

Emirates Aviation University Page | 1

Contents
Contents ................................................................................................................................... 2
Acronyms and Nomenclature ................................................................................................. 4
1. Executive Summary .......................................................................................................... 5
1.1 Design approach 5
1.2 Performance and Capabilities 5
2. Management Summary..................................................................................................... 6
2.1 Team Organization 6
2.2 Organization chart 6
2.3 Milestone chart 7
3. Conceptual Design............................................................................................................ 7
3.1 Mission requirements 8
3.2 Scoring sensitivity analysis 9
3.3 Subsystem design requirements 10
3.4 Aircraft configurations 10
4. Preliminary Design...........................................................................................................13
4.1 Design Trades 14
4.2 Mission Model 16
4.3 Aerodynamic characteristics 18
4.4 Payload Selection 21
4.5 Propulsion 25
4.6 Stability and Control 26
4.7 Estimated Mission Performance: 29
5. Detail Design ....................................................................................................................30
5.1 Dimensions 30

5.2 Weight and Balance: 31
5.3 Structural characteristics: 33
5.4 System and subsystem design: 36
5.5 Mission performance 39
5.6 Drawing Package 39
6. Manufacturing ..................................................................................................................40
6.1 Manufacturing process investigation 40
6.2 Selection process 40
6.3 Manufacturing of Parts 42
6.4 Manufacturing Milestones 45
7. Testing plan......................................................................................................................46
7.1 Testing schedule 46
7.2 Testing objectives 46
7.3 Flight performance testing 49
7.4 Flight checklist 49
8. Performance Results .......................................................................................................51
8.1 Demonstrated performance of key subsystems 51
8.2 Demonstrated flight performance of completed aircraft: 52
8.3 Performance improvements 53

Acronyms and Nomenclature
M1 -- Mission One
M2 -- Mission Two
M3 -- Mission Three
C.G. -- Center of Gravity
RAC -- Rated Aircraft Cost
TFS -- Total Flight Score
FOM -- Figures of Merit
EW -- Empty Weight
GS -- Ground Score
TMS -- Total Mission Score
CL -- Aircraft lift coefficient
CD -- Aircraft Drag Coefficient
C 𝐷0
-- Aircraft Zero-Lift Drag Coefficient
C 𝐷 𝑖
-- Aircraft Induced Lift
Cm -- Aircraft Pitching Moment Coefficient
Cl -- Aircraft Rolling Moment Coefficient
Cn -- Aircraft Yawing Moment Coefficient
C𝑙 𝐵
-- Aircraft Lateral Stability Coefficient
C 𝑛 𝐵
-- Aircraft Directional Stability Coefficient
MTOW -- Maximum Takeoff Weight
NiCad -- Nickel-Cadmium
𝑇𝑠 -- Settling Time
𝑇𝑝 -- Peak Time
𝑇𝑑 -- Doubling time
e -- Oswald Efficiency
P -- Power
S -- Area
𝑆𝑔 -- Takeoff Roll Distance
T -- Thrust
D -- Drag
m -- Mass
V -- Velocity (ft/s)
AR -- Aspect Ratio
Re -- Reynolds number
𝜌 -- Density
𝛼 -- Angle of attack
Kv -- Motor Voltage Constant (V)
𝜂 -- Efficiency
𝜁 -- Damping Ratio
g -- Gravity
𝑉𝑇𝑂 -- Takeoff Speed
P -- Power
𝑆 𝑤 -- Wing Area
W/S -- Wing Loading
P/W -- Power Loading
T/W -- Thrust to weight ratio
𝑅 𝑇 -- Taper Ratio
𝛽 -- Sideslip Angle (degrees)

1. Executive Summary
This report comprehensively discusses and evaluates the design, testing, and manufacturing of the aircraft
designed by the students of Emirates Aviation University for the 2019-2020 Design/Build/Fly (DBF)
competition. The aircraft is designed as a banner towing aircraft able to accommodate passengers and
luggage perfectly in accordance with the mission of this year. High-wing conventional aircraft design is
preferred for a spacious configuration to carry maximum passengers and luggage. The wing is constructed
with a maximum wingspan to produce a high lifting area with minimum wing loading. It achieves short take-
offs for mission 1 and mission 3.
1.1 Design approach
The plane ‘Skywalker’ aims to leave a lasting impression of its flying capabilities in the competition. This
will be achieved by showcasing flying capabilities reliably and ensuring the required compliance with the
missions. The concept is designed by accurately following the mission requirements and availing scoring
opportunities, which is evident in the preliminary sizing of the aircraft. Trade studies were carried out, while
creating the preliminary design, to estimate the performance parameters. The design was computed into
computational fluid analysis to calculate the lift and drag. The stability of aircraft was identified by analysis
in XFLR5. The battery, motor, and propeller were thoroughly iterated in eCalc to decide the perfect
combination. After establishing the design metrics, a detailed model was created, manufactured, and then
flight tested to validate and verify the estimated performance.
1.2 Performance and Capabilities
The performance of ‘Skywalker’ is determined by the following performance capabilities shown in the flight
tests:
➢ Max Take-off weight 6.61 lbs.
➢ Short take off 20ft
➢ Benign stall characteristics
➢ Secure storage of payload in Mission 2 and 3
➢ Top speed 60+ ft/s
➢ Proven capability in windy conditions
➢ Effective deployment and release of banner
➢ Accommodates four passengers and luggage
The team is satisfied with the high wing configuration monoplane that perfectly corresponds to the required
performance for the competition. The aircraft is designed and the stability is ensured with maximum payload
carried in the fuselage. The aircraft is integrated with the right match of motors and batteries to achieve
Figure 1: Skywalker mid flight

maximum speed and laps in each mission within given flight window. The performance of the aircraft is
tested at harsh windy conditions to be prepared for the predicted flying conditions in Wichita, KS
2. Management Summary
The EAU Aviator team consisted of nine undergraduate students with two seniors, four juniors, and three
sophomores. The team followed a discipline hierarchy that establishes a strong core between seniors and
sophomores collaborating extensively to share ideas and reach project milestones. The team members
were mentored by experienced personnel, which helped to build a strong learning foundation for the team.
2.1 Team Organization
The hierarchy of the team was organized subsequently to the design phase, which is segmented into CAD,
structures, aerodynamics, propulsion, stability, and manufacturing. The manufacturing and testing of the
aircraft associated each team member to share maximum ideas and producing the best out of the design.
The senior members of the team worked as the most responsible individuals leading their team in the right
direction. The faculty advisors supported the team financially and shared their vital experience.
2.2 Organization chart
Figure 2: Team Organization Chart

2.3 Milestone chart
A milestone chart was created during the process to complete the tasks within set deadlines. The progress
throughout the process was monitored by the team leaders to maintain the discipline of the team and finish
all the milestones on time. The team manager supervised the progress of the project to ensure high
performance and smooth operations.
3. Conceptual Design
The objective of the conceptual design is to figure out the possible solutions compatible to satisfy the given
design requirements and gain high scores in the competition. Scoring analysis identifying the most sensitive
mission requirements helped the team in maximizing flight score. These scoring factors were then
translated into a conceptual study of the design. The possible design solutions were evaluated in the
deciding matrix to achieve the most effective design. The Figures of Merit (FOM) used to distinguish
amongst the three possible design configurations decided the optimum configuration.
Figure 3: Gantt chart

3.1 Mission requirements
The final score of the AIAA design-build and fly competition is calculated by multiplying the written report
score with the mission score. The total mission score can be calculated by the addition of the scores
attained in each mission. Mathematically they can be represented as
Score= written report score*total mission score (1)
Total mission score=M1+M2+M3+GM (2)
Mission 1 – Test flight
There will be no payload on the flight for this mission. The aircraft will be required to be present in the
staging box. There is a limited length in which the aircraft has to take off. The length will be 20 feet. It is
mandatory for the teams to complete 3 laps within the given 5-minute flight window. The timer begins when
the throttle is given for the first time. In order to receive a score, a successful landing should be achieved.
M1 = 1.0 for successful mission
Mission 2 – Charter flight
This mission has passengers and luggage as payload. The number of passengers that are being flown can
be selected by the teams but cannot exceed the maximum limit selected by the organizers, which will be
revealed at the time of the tech inspection process. This mission has no specified take-off length. A 5-
minute window is given for this mission, and all the teams should complete 3 laps in the given time frame.
The scoring is directly proportional to the number of passengers flown and inversely proportional to the time
taken. The formula used to calculate score attained in this mission is
M2=1+
𝑓𝑙𝑜𝑤𝑛
𝑛𝑜.𝑜𝑓 𝑝𝑎𝑠𝑠𝑒𝑛𝑔𝑒𝑟𝑠
𝑡𝑖𝑚𝑒
𝑚𝑎𝑥𝑖𝑚𝑢𝑚
𝑛𝑜,𝑜𝑓 𝑝𝑎𝑠𝑠𝑒𝑛𝑔𝑒𝑟𝑠
𝑡𝑖𝑚𝑒
Mission 3 – Banner flight
The only payload for Mission 3 is the banner. Similarly, to Mission 1, aircraft should enter the staging box
in the flight configuration. The maximum take-off length is 20 feet. After the first upwind turn, the banner
has to be deployed using a command switch present on the remote. A 10-minute window is provided for
this mission. The timer starts when the throttle is advanced for the first time. The score received in this
mission is calculated based on the following formula
M3=2+
𝑓𝑙𝑜𝑤𝑛 (𝑛𝑜.𝑜𝑓 𝑙𝑎𝑝𝑠∗𝑏𝑎𝑛𝑛𝑒𝑟 𝑙𝑒𝑛𝑔𝑡ℎ)
𝑚𝑎𝑥𝑖𝑚𝑢𝑚 (𝑛𝑜,𝑜𝑓 𝑙𝑎𝑝𝑠∗𝑏𝑎𝑛𝑛𝑒𝑟 𝑙𝑒𝑛𝑔𝑡ℎ)

Ground mission
This mission is used for demonstration of mission 2 and 3. Initially, the aircraft will be kept in a 10ft by 10ft
with the uninstalled maximum allowed number passengers, luggage, and the uninstalled banner. The
assembly crew and a pilot may perform in this mission with only assembly crew team members allowed to
touch the aircraft. The ground mission starts as the official says, “GO.” The assembly crew member will
then load the passengers and luggage and will run back to the start-finish line. The pilot will then verify the
activeness of the flight controls. The assemble crew member will start again behind the start/finish line and
will start as soon as the official says, “GO.” This time the assembly crew member will unload the passenger
and luggage and will install the banner in the stowed configuration on to the aircraft. The pilot will again
verify the activeness of the flight controls and will hold the aircraft in a vertical position with the tail down to
demonstrate the deployment and release of the banner. The score is given by the following formula:
GM=
min 𝑡𝑖𝑚𝑒
𝑡𝑖𝑚𝑒
where min time is the fastest time for all teams.
3.2 Scoring sensitivity analysis
Scoring sensitivity analyzed explicated the important performance factors in all missions. This year’s
mission 2 and 3 requires to carry passengers and luggage along with the banner towed externally to the
aircraft. According to sensitivity analysis, maximizing the number of passengers and luggage and reducing
the time to complete 3 minimum laps will help to score high in mission 2. The aircraft with the capability to
carry a banner with maximum length in a 10-minute flight window and completing maximum laps will achieve
the best scores. These sensitive parameters are plotted in MATLAB to evaluate the effect on the total score
with a change in performance variables. The figure below depicts the sensitivity analysis.
Figure 4: Sensitivity Analysis graph

3.3 Subsystem design requirements
The mission requirements translated into subsystem design requirements design helps in improving the
overall score. After understanding the mission requirements, the team revealed the following subsystem
design requirements:
Take-off
The aircraft needs to take-off within a short take-off distance of 20ft in mission 1 and mission 3. This design
requirement affects the selection of the wing area and the propulsion selection of the aircraft.
L/D ratio
The aircraft must have a high lift to drag ratio to carry a maximum payload for a long duration. The aircraft
needs to have a minimum of drag to carry passengers and luggage and complete a minimum of 3 laps in a
short time. High L/D ratio will help in short take-offs, better climb performance, and glide ratio.
Restraint system
The restraint system is required to secure the passenger and luggage in the aircraft. The restraint system
must be designed expertly to quickly unload the passengers and luggage in a short time to maximum ground
mission score. The banner mount must be kept simple for easy installing of the banner
Banner sizing
The banner must be sized light and easy to control in flight. The banner deployment in midflight increases
the drag and disrupts aerodynamics of the aircraft. The design must consider the weight and size of the
banner as windy conditions are forecasted in Wichita, Ks.
3.4 Aircraft configurations
The team evaluated various aircraft configurations to decide the successful aircraft design for the
competition. The configuration which yielded the highest score was selected as the final configuration of
the aircraft. These possible configurations were compared using the figure of merit (FOM) translated by the
mission requirements given in the rules.
Fuselage configuration:
The fuselage was analyzed for different configurations: airfoil, conventional, and pod. Weight was
determined as the critical deciding factor as the aircraft needs to carry passengers, luggage, and the banner
towing mechanism in mission 2 and mission 3. The conventional fuselage was concluded as the best choice
because it is simple and easy to size compared to other configurations and offers better space to
accommodate a higher number of passengers and luggage. The conventional fuselage helps in attaching
the banner towing mechanism at the rear end of the fuselage, unlike the other configurations.

Wing configuration
High wing, low wing, and mid-wing configurations were evaluated for the best wing configuration. Each of
these configurations was evaluated based on the crucial deciding factors, weight, simplicity, payload
capacity, and stability. Low wing configuration has the worst stability when compared to the other
configurations and does not offer significant advantages when it comes to the other deciding factors. A mid-
wing configuration is a poor option when it comes to weight because of its wing root intersection of the
fuselage. It provides less space for the fuselage to accommodate passengers and luggage. High wing
configuration relatively offers better characteristics as it has the best lateral stability and is relatively simple
and easy to control. Hence, most importantly allows plenty of space for the fuselage to carry a maximum
number of passenger and luggage and weights the least.
Table 1: Aircraft configuration table
Table 2: Wing Configuration table

Tail configuration:
Three different configurations: Conventional, T-tail, and V-tail, are evaluated above for weight, simplicity,
and stability. The aircraft needs to carry a high payload in a windy environment and requires larger control
surfaces to cope with the resistance of wind and drag created by the banner. Therefore V-tail is a relatively
complex configuration that has the least stability and is not a favorable option. T-tail configuration is
unsuitable as it is prone to suffer from a dangerous deep stall and may lose pitch control when the wing is
stalled at high angles of attack. Hence, the conventional tail is the best configuration. It offers better stability
and good stall characteristics with its large control surfaces.
Propulsion:
A single tractor motor is selected from the above configurations evaluated as it gives the highest efficiency.
Furthermore, it reduces weight and is simple to assemble. Multi pusher configuration is the least favorable
option because of its poor efficiency and difficult manufacturability. A multi tractor is a suitable option, but
the aircraft does not require high power requirements. A single tractor motor is more than enough to produce
the required thrust for the mission with a lighter weight structure.
Table 4: Propulsion configuration
Table 3: Tail configuration table

Landing gear configuration:
Tricycle configuration consists of two main wheels beneath the wing and a nose wheel for steering located
aft of the aircraft CG. The disadvantage so of this configuration is that it causes drag due to the wake
generated by the propeller. Though it has reasonable control relatively lacks behind tail dragger. Bicycle
configuration comprises of two main center wheels and two wingtip wheels. This type of configuration adds
excessive weight and drag with added wheels and struts. The favorable configuration is the tail dragger as
it creates less drag and allows much easy rotational clearance during short take-offs.
4. Preliminary Design
After the completion of the conceptual design, preliminary sizing was done to determine the size of different
components of the aircraft. Preliminary sizing is a critical phase in designing aircraft and influences aircraft
performance. Trade studies were considered to evaluate the best possible sizing of the aircraft.
Design Methodology
The EAU Aviators’ team conducted a preliminary study to ensure aircraft sizing complies with the mission
requirements. The team worked on different design aspects of the aircraft, including aerodynamics,
propulsion, stability, and control. These essential design requirements were computed for a thorough
analysis. The aerodynamics of the aircraft was analyzed using Computation fluid analysis in Solid works.
The aircraft was estimated for the zero-lift drag condition and the best L/D max to evaluate the minimum
power, thrust, and speed required for optimum flight conditions. The propulsion system was iterated with
different combinations of battery and motor to find a system providing flight time required in the missions.
The aircraft stability was evaluated for knowing its ability to respond to aerodynamic forces in flight. The
XFLR5 software was used in this process to help to size the aircraft with consideration to the stability
needed in the design. The figure below depicts iteration process.
Table 5: Landing gear configuration

Figure 5: Iteration process overview
4.1 Design Trades
The preliminary sizing was focused on designing aircraft with the lowest engine power and wing loading.
The matching plot was plotted to determine the required wing reference area and the subsequent sizing of
the overall aircraft. The matching graph plotted with Wing loading vs. power loading based on key
performance parameters, which helped the team to conduct parametric studies of the aircraft.
The essential parameters used by the team in matching graph determine the wing and engine sizing are
the following:
Stall speed: The stall speed is one of the leading design parameters when it comes to design aircraft with
the least possible take-off run. The aircraft with the lowest stall speed results in lower wing loading design,
hence increasing the wing reference area and ensuring safe and short take-off.
(
𝑊
𝑆
) =
1
2
𝜌𝑉𝑠2
𝐶𝑙 𝑚𝑎𝑥
Maximum speed: Another excellent performance for design the aircraft is maximum speed. The maximum
speed required for an aircraft design depends on its weight, wing area, and engine power. The acceptable
region of the matching plot below, which satisfies the maximum speed requirement, is determined through
the following equation below. The Vmax is inversely proportional to the value of power loading, and thus
with better aerodynamic characteristics, aircraft require the least power to fly at Vmax. The region below
the graph is acceptable for the design.

(
W
P
) =
σnp
1
2
𝜌𝑠𝑙 V3
max
CD0
1
(
𝑊
𝑆
)
2k
ρVmax
W
S
Take-off run: The acceptable region in the matching plot satisfying the take-off run requirements is
analyzed from the equation below. The take-off run increases with power loading (W/P) values increasing.
Hence, take-off runs more significant than 1 does not meet the take-off run requirements, and the region
below the graph is accepted as take-off run requiring least power loading (W/P).
(
𝑊
𝑃
) =
1 − 𝑒𝑥𝑝 (0.6 𝜌 𝐶 𝐷𝐺 ST0
1
(
𝑊
𝑆
)
)
𝜇 − (𝜇 +
𝐶 𝐷𝐺
𝐶𝐿𝑅
) [𝑒𝑥𝑝 (0.6 𝜌 𝐶 𝐷𝐺 ST0
1
(
𝑊
𝑆
)
)]
.
𝑛 𝑝
𝑉𝑇𝑂
Rate of climb: The rate climb is mainly dependent upon the engine power and the maximum lift-to-drag
ratio (L/D) max. Thus, the Rate of climb value, as analyzed by the equation below, with ROC value in the
denominator of the equation, is inversely proportional to the value of power loading. The aircraft with better
(L/D) max will consume less thrust of the engine, helping aircraft to fly for a more extended period of time
and attain maximum laps in the missions.
(
𝑊
𝑃
) =
1
ROC
𝑛 𝑝
+
√
2
𝜌 × √3 × CD0
𝐾
× (
𝑊
𝑆
) × (
1.155
(𝐿 𝐷⁄ ) 𝑚𝑎𝑥 × 𝑛 𝑝
)
The figure 6 below illustrates the design constraint graph.

The design point in the matching graph above is the intersection of stall speed and maximum speed. This
point shows the lowest engine power and satisfies all the performance parameters discussed above. The
team evaluated the following design constraints from the matching graph.
Table 6: Values obtained from design constraint graph
4.2 Mission Model
The competition follows the flight course shown below for all the flight missions. The flight course is roughly
estimated to be 2500 ft per lap. There is 180 degree upwind and downwind turns and a 360-degree turn.
The take-off and landing segments of the flight depend on the aircraft performance. The aircraft is required
to have short take-off and land safely within the runway. The aircraft achieving a higher number of laps
along this flight path will attain higher scores. The orientation of the flight course will be adjusted according
to the winds in Wichita, Ks, at the time of the competition. The figure below illustrates the flight path of the
AIAA DBF competition 2020.
Figure 6: Design constraint graph

The design requirements at different flight segments were evaluated to keep them in consideration when
working on the design constraints of the aircraft. The design requirements included the study of each flight
segment and uncertainties expected in the fly-off. The table 7 below depicts the design requirements of
each flight segment.
Uncertainties
Gust loads:
The weather in Wichita, Ks will be windy at the time of the competition. This makes flight difficult for the
aircraft as gust loads from winds will be disrupting the performance of the aircraft. The aircraft at normal
flight conditions with 1 g at steady flight and 2g at turns will be actually facing more g’s in windy conditions.
The aircraft needs to be aware of the accelerated stall while performing 180 degree and 360 turns. As the
stall speed increases with the square root of the load factor, the aircraft must fly within safe and stall-free
banking angles to avoid loss of stability and control. As shown below in the graph, at 60 degrees of banking
angle, a 41 % increase in stall is observed. Hence, it was decided by the team to bank aircraft at below 60
Figure 7: Flight path of AIAA flyoff
Table 7: Design requirements and constraints of each flight segment

degrees to ensure the minimal increase in the stall speed. The figure below illustrates the change in stall
speed percentage with banking angle.
Interference drag:
The interference drags caused by the shape and body of the aircraft are uncertain and ambiguous to
evaluate primarily because of the uncertain windspeed flowing around. The aircraft will be expected to
produce more drag and fly at lower velocities due to the interference drag. Hence, it is vital to design aircraft
with maximum performance to compensate for any uncertainty present in the competition.
4.3 Aerodynamic characteristics
Airfoil Selection: Selecting the appropriate airfoil is very crucial for a stable aircraft. Airfoil selection aims
to select an airfoil that would provide high lift coefficient, low drag coefficient, high efficiency, and low
pitching moment characteristics. The airfoil data were looked up in the airfoil tools website. After iteration,
3 airfoils were chosen to be compared NACA 2412, Clark Y SM, and S9000. The Reynolds number was
approximated to be around 200,000. Our aircraft mission requirements include a short takeoff at 20ft and a
stable aircraft at high winds. Hence, a higher 𝐶𝑙 𝑚𝑎𝑥
with low 𝐶 𝐷 is crucial for our aircraft. The table below
illustrates various parameters of all the three chosen airfoils. Data for all the airfoils is obtained by
Airfoiltools.com.
Figure 8: Banking angle vs Stall speed graph

From the airfoil database above in table 1, Clark Y SM is chosen to be the most suitable airfoil. The decision
was influenced based on 𝐶𝑙 𝑚𝑎𝑥
of 1.3584 at takeoff with a relatively low 𝐶 𝐷 of 0.06036 at a higher AOA of
15.75°. A higher Cl/Cd ensures that flight is efficient, and turning maneuvers are stable. Clark Y smoothed
airfoil is shown below in figure 9.
Lift and drag analysis:
The aerodynamics team conducted a computational fluid analysis to estimate the aircraft’s performance in
each mission. The aerodynamics characteristics of Mission 3 are different from the Mission 1 and Mission
2, giving 28.52 % increase in the overall drag of the aircraft. The banner stowed externally to the aircraft
in Mission 3 helps to increase the overall drag of Mission 1 and Mission 2. The aerodynamics helped the
team to evaluate the shape of the designed aircraft and calculate the power required from the propulsion
system The figure 10 below depicts the drag contribution of each component in different missions of the
competition.
Figure 9: Airfoil Cl vs Cd graph
Table 8: Airfoil comparison table

Drag power
The drag of each mission was utilized to calculate the drag power, which is the power required to overcome
the drag of the aircraft. The Mission 1 and Mission 2 producing a similar amount of drag was compared
with Mission 3 with the help of the equation below:
𝐹 = 𝐹𝐷 ∗ 𝑣 =
1
2
𝐶 𝐷 𝜌𝑣3
𝐴
The figure 11 below depicts that the M3 requires higher power compared to rest of the missions. Hence,
M3 is the most important mission when selecting the avionics of our aircraft design.
Figure 10: Drag analysis of each mission

4.4 Payload Selection
Banner Material Selection:
The team conducted a wind tunnel test to select the best suited banner material. The wind tunnel tests were
conducted at various velocities of different fabrics. Mission 3 score was taken into account when testing,
hence each fabric being tested had different surface roughness and stiffness. The table 9 below depicts
the different fabrics used for wind tunnel testing and their properties.
The properties taken into account when testing each fabric are:
❖ Fabric properties effect: The drag decreases with the smoothness of the surface. However, drag
increases with a flexible fabric as it will produce more oscillations.
❖ Wind Speed: The drag increases with the increasing wind speed.
❖ Reynolds number: A greater Reynold’s number decreases the overall banner drag.
Figure 11: Drag power vs Velocity (ft/s) graph
Table 9: Material properties table

The available wind tunnel had a test section size of 40 x 30 cm. Hence each fabric was cut into dimensions
of 35 x 20 cm to safely fit in the test section. The fabrics above were tested at velocities ranging from 5 m/s
to 25 m/s. At each velocity, the drag which the material creates is estimated. The figure 12 below illustrates
the drag created by each fabric due to its characteristics.
It can be seen from the figure 12 above that double-thick cotton produces the most drag. However, Satin
material having the smoothest surface with medium stiffness produces the least drag. Hence satin material
being most suited for mission 3 is selected for our banner. The figure 13 below depicts the wind tunnel test
of the selected fabric, satin.
Figure 12: Wind tunnel test of banner graph

Banner Size Selection:
The banner size was selected on the basis of aspect ratio and the best-suited banner length for maximum
mission 3 score. The team investigated the effects of various aspect ratios on banner drag in the wind
tunnel. The figure 14 below further illustrates the relationship of aspect ratio with the drag coefficient.
Figure 13: Wind tunnel test of satin material
Figure 14: Banner Aspect ratio vs Cd

Banner Length
The selection of the ideal banner length required the best combination of banner drag and mission 3 score.
This method of evaluation required a giant wind tunnel. Several lengths of the banner had to be tested to
obtain precise values of drag. The team couldn’t perform this method due to limited resources. Hence an
alternative method was chosen to evaluate the closest value of banner drag at various lengths to select the
best length. Due to a more significant wind tunnel being unavailable, the material oscillations were not taken
into account. Hence, the banner drag was estimated using the turbulent drag equation across a flat plate.
The aircraft was assumed to be flying at 50 ft. Hence the values obtained for these conditions are mentioned
below.
❖ Density (𝜌): 0.0023746
❖ Dynamic Viscosity (𝜇): 3.73 ∗ 10−7
❖ Velocity: 65 ft/s
The table 10 below illustrates the formulae used to estimate the banner drag.
Using table 10, several values of lengths were used to obtain drag created at each length. The team first
estimated the minimum drag a banner can create by using the minimum dimensions mentioned in the rules.
The drag estimated at 10 x 2 inches was very minimal and would provide optimal performance in mission
3. However, considering the mission 3 score depends on both the banner length and the number of laps,
this banner length of 10 inches could not be used as it will only yield a higher number of laps performed.
Hence iterations on the banner size were performed to estimate a suitable banner size, which will yield both
the parameters, higher number of laps, and more significant banner length. The table 11 below depicts the
iterations performed to evaluate the drag each length of the banner creates.
Table 11: Results of Banner drag at various lengths
Table 10: Key formulae for Banner analysis

Number of passenger’s selection
The team decided to use the maximum banner length of 23.62 inches, as it will yield an excellent banner
length score. However, to maximize the number of laps, the team came up with a strategy to fly the aircraft
at less velocity to minimize the banner drag.
The number of passengers were selected in a way that the aircraft consists of least wing loading and yields
highest score according to the M2 formula, No. Of pass/time. The table 12 below consists of wing loading
with different number of passengers.
The team decided to choose 4 as the maximum number of passengers and luggage as it ensures aircraft
with low wing loading which helps in reducing the weight of propulsion system as less power required to
complete the Mission. The number of passengers selected comfortably fits within the fuselage dimensions
of the aircraft which was selected to be minimal to reduce the structural weight of the aircraft.
4.5 Propulsion
The preliminary sizing of the propulsion system focus, providing the necessary power for the short take-off
and carrying a payload with ease. Along with these requirements, the propulsion must be able to provide
enough power to comprehensively complete the maximum number of laps in each mission. The team used
ecalc to analyze and evaluate the minimum power requirement at each mission, depending on the take-off
weight of the aircraft and the maximum flight endurance required.
The three main components which make up the propulsion system include the battery, motor, and propeller.
The selection of these components relied heavily on their abilities to provide efficiency in each mission
❖ Battery selection - The battery is what provides electricity within the propulsion system. After
reviewing the mission requirements, the amount of power needed to complete each mission, and
take-off power was calculated. Taking into account the required power, weight contribution, and
discharge rate, a battery was selected.
❖ Propeller selection - The size of the propeller chosen affects the cruise speed and the runway
distance needed for take-off.
Weight (lbs.) Wing loading lb/ft^2
0.638 0.167454068
1.276 0.334908136
1.914 0.502362205
2.552 0.669816273
3.19 0.837270341
Table 12: Selection of Passengers

❖ Motor Selection – The motor selection is based on the efficiency and power output it provides.
The avionics team used Propcalc software for selecting the best combination of avionics. Several
combinations of components were compared. The combination yielding the highest flight time, and the least
power consumption was chosen for our aircraft. The table 13 below shows the comparison of three different
motors.
It is evident from the table () above that Rimfire 0.55 (480) Kv is the most efficient motor. It provides a
maximum flight time of 15 minutes and consumes the least current. The motor was selected to meet the
maximum flight time of 10 mins in mission 3 in the competition. The team decided to choose rimfire because
it is an out runner brushless motor, and it provides high torque. This allows the aircraft to accelerate quickly,
especially when short-take offs are required.
4.6 Stability and Control
The stability of an aircraft is its ability to respond to aerodynamic forces and flight inputs. In order to fly an
RC aircraft, it needs to be statically and dynamically stable. The static stability is the corrections an aircraft
makes when disturbed in pitch mode. A positively static aircraft will return back to its original altitude when
perturbed. Dynamic stability is the aircraft’s response over time when disturbed. The aircraft will have
oscillations, but they should always dampen out so that the aircraft can achieve trim condition.
Static Stability
It is impossible to fly a statically unstable aircraft. Hence the static stability was deeply examined. The static
stability relies on the aircraft's static margin. It is the distance from the neutral position of the aircraft to its
center of gravity. For an aircraft to be statically stable, its center of gravity should be forward of the neutral
point. This will create a positive aerodynamic moment to restore the aircraft's original position after
perturbation in flight. An aircraft with negative static stability will mean the center of gravity is aft of the
neutral point. This will create a negative moment that will continue increasing the aircraft's angle of attack
until it stalls. A neutrally stable aircraft has the center of gravity placed at the same point as the neutral
point.
The theoretical range of the static margin is set to be 5-15 %. Static margin in this range would mean the
aircraft has positive static stability, it requires less effort by the pilot and will be easy to maneuver. The static
Table 13: Motor selection table

margin of less than 5% means the aircraft will be overly responsive to any control surface inputs. Meanwhile,
the static margin of above 15% means the aircraft will require excessive pilot's effort to handle it.
The challenge of this year was to place the payload and banner towing mechanism accordingly so that the
center of gravity remains forward of the neutral point and keeping the center of gravity within acceptable
limits. Another challenge was to release the banner in such a way that the center of gravity doesn't change.
To encounter this, the payload and the banner towing mechanism was placed right under the quarter chord
of the wing. This will eventually result in a positively static aircraft. The table 14 below illustrates the static
stability results for each mission.
The tail sizing is done in such a way that the stability derivatives satisfy Level 1 flying conditions so that
less effort is required by the pilot to handle the aircraft. The table 15 below illustrates the key input
parameters used to obtain the required stability derivatives.
Dynamic Stability
Dynamic stability was obtained by creating a model in XFLR. The dimensions used were obtained in sizing
analysis. Static margin analysis aided in precisely placing each component in place and getting the moment
of inertia required. Root locus was obtained for each mission, which depicted eigenvalues for longitudinal
Table 14: Static stability results of each mission
Table 15: Input Parameters

and lateral modes. The figure 15 below illustrates the root locus of the aircraft with payload and banner
fitted within.
The root locus depicts the spiral pole has a very slight positive real component, which illustrates that if the
aircraft is pushed laterally by a strong gust, then it will begin to roll and eventually dive if there is no pilot
input. However, due to the spiral pole being very slowly convergent, the pilot can still maintain control of
the aircraft by using appropriate inputs. Moreover, the Phugoid, Short period, Dutch roll, and Roll modes
are all dynamically stable. This means the aircraft, when disturbed, will eventually return to its steady level
flight condition without any pilot input. The table 16 below illustrated the dynamic characteristics of all the
missions of the aircraft
Figure 15: Root locus

4.7 Estimated Mission Performance:
The performance of the aircraft was estimated for different missions in the competition. The aircraft
performance was evaluated with the help of the aerodynamic results of CFD simulation, XFLR5, and the
propulsion characteristics of the aircraft from the eCalc calculator. The mission 2 requires a longer runway
distance to take off compared to mission 1 and mission 3 as it has a higher maximum take-off weight with
4 passengers, luggage, and the banner mount. In mission 3, it is comparatively easier to take off within a
short distance than mission 2, but the windy conditions efforts to disrupt the aircraft performance as it
deploys the banner. Different materials for the banner were hence tested during the flight test to select the
lightest and airworthy banner cloth for the mission 3. The table 17 below depicts the estimated mission
performance at each mission.
Table 16: Results of dynamic stability analysis

5. Detail Design
The detail design section is the final step of the design. It aims to further refine the conceptual and
preliminary design sections. This section will focus on attaining the specific goals the aircraft requires to
meet the specifications.
5.1 Dimensions
The table 18 below depicts the overall dimensions of the aircraft. It also specifies the propulsion system the
aircraft will use.
Table 17: Estimated performance of sky walker in each mission.
Table 18: Dimensions of Skywalker

5.2 Weight and Balance:
The weight and balance of an aircraft is the critical aspect of its performance and stability. Skywalker weighs
about 6.17 lbs. When fully loaded. The balance of the aircraft varies based on the position of the
components within the aircraft. If the mass is too far forward, the elevator won’t have adequate control
power to control the aircraft in pitch. If the mass within the aircraft is too far aft, then the aircraft will not have
an adequate static margin to restore unstable pitching moment. The goal was to distribute the masses in
such a way that the center of gravity remains between 25 – 30% of the wing quarter chord, and the static
margin remains in the range of 5 – 15%. The detailed model was made in Solid Works. Each component
was applied with its material, which accounted for its density and other structural properties. The reference
point was taken as the nose of the aircraft. The center of gravity was found using the center of mass in
Solid works. The masses and center of gravity of each component were multiplied to attain the moment
arm of each component. The figure 16 below illustrates the CG location of the aircraft.
Figure 16: Center of gravity of Skywalker

Table 19: Weight and balance table of each mission.

5.3 Structural characteristics:
An aircraft is exposed to various loads and stresses during flight. These loads are distributed among each
component of the aircraft. Skywalker is designed to withstand all types of loads. The structure of the aircraft
was designed such that it can withstand a load of 3.4g at a maximum gross weight of 6.16 lbs. The aircraft
structure is completely made of balsa wood. Additional strength is added to the aircraft by using plywood
in high-stress areas like fuselage bulkheads. The aircraft consists of a fuselage initially constructed by using
balsa wood. Plywood is added as a second layer to add rigidity. A removable high wing is made entirely by
balsa wood. The empennage is also made entirely by using balsa wood. The structural concepts and flight
loads were deeply investigated to achieve the optimal structure that can overcome these challenges. The
figure 17 below illustrates the structural characteristics of the aircraft at maximum weight.
To analyze the structural integrity of some components where the maximum stress is being applied, FEA
in solid works was conducted. FEA is a built-in feature in solid works. It virtually simulates a component to
applied forces and loads. Although FEA is a powerful tool for CAD simulation, it also has its limitations. The
FEA tool assumes that the material being used exhibits uniform behavior. In reality, these structures are
being laser cut and glued together. FEA also assumes the structure is a solid member which is attached
ideally. However, the pros outweigh the cons of the FEA, making it a very efficient and powerful tool to
overcome design restraints.
During the design stage, each component was modeled in solid works. The components which receive
maximum stress were the wing and fuselage. FEA was conducted on these components to refine their
design and achieve acceptable strength without extra weight.
Figure 17: Structural V-n diagram

Fuselage
The fuselage structure was very critical as this year’s challenge was to accommodate payload and attach
an external banner mechanism. The added weight in the aircraft meant the structure had to be extremely
rigid yet lightweight. The fuselage structure was mainly made of plywood. It comprises 7 bulkheads: one at
the front to mount the motor, one at the leading edge of the wing, one at the trailing edge of the wing, which
will withstand the loads exerted by the wing onto the fuselage. The rest of four bulkheads are attached at
equal distances in the rear fuselage to support the incoming loads and empennage. The aircraft consists
of three frames each made by plywood to support loads. The overall fuselage is kept lightweight by having
large lightening holes at the sides where the shear flow is least. To prevent buckling and deformation of the
fuselage sides, stringers are attached to the corners of the fuselage. The lower frame of the fuselage, which
is attached to the landing gear and banner mechanism, is made by a thick sheet of plywood to increase
overall rigidity. The passengers and luggage compartment inside the fuselage are made separate but under
the wing. The compartments are made using balsa sheets of 3/32” thickness.
Wing
The wing is required to support the whole aircraft during flight and turns. It should ideally be lightweight and
rigid. Hence, a combination of balsa and plywood is used to make the wing structure. Balsa wood being
among the lightest wood with high rigidity allows it to be extensively used in the aircraft. The wing structure
is deeply examined as it supports half the weight of the aircraft in steady level flight and twice the weight
during turns. The wing structure includes 20 ribs made by balsa and placed equally at 3.75 inches each.
The first rib, the last rib and the two inner ribs which connect the wing sections are reinforced with an added
plywood rib. The wing has two spars made of plywood of 3/32" are added at the quarter chord along with a
shear web. The ribs withstand most of the compressive bending loads. The spars at the quarter chord have
horizontal grains that prevent bending stress in flight. Shear webs have vertical grains that take most of the
shear stress. Leading-edge and trailing edge balsa wood sheet provides the wing torsional rigidity and
maintains the airfoil shape. Monokote is used to cover the wing to provide added strength and provide a
smooth surface.
Figure 18: Fuselage structure

FEA is conducted by fixing one end of the semi wing and evenly distributing 3.08 lbs half of the aircraft
weight load to all the ribs. Considering the yield strength of the Balsa wood to be 20.7 MPa. The maximum
Von misses stress achieved is 15.41 MPa, and maximum displacement of the wing is 0.303 inches. The
factor of safety of 0.74 indicates the wing has structural integrity, and it won't break when maximum loads
are applied. The figures below illustrate Von misses stress and maximum deflection of the wing at half the
weight of the aircraft.
Empennage
The vertical and horizontal stabilizers are constructed using sheets of balsa wood. Thin sheets of 3/32” are
used to shape both the tails. The grains of each of the tail was placed such that the tails can withstand
maximum bending stresses. Lightening holes applied as the tail needed to have some weight to stabilize
the aircraft. The elevator and rudder are made using separate sheets of balsa and attached to the tail using
hinges. Attachment of the control surfaces with the empennage isn’t kept very tight, so the control surfaces
need less hinge moments to move.
Figure 19: Wing structure
Figure 20: Finite Element Analysis of wing structure

Landing Gear
The landing gear is constructed using aluminum. Thin sheet of aluminum is molded into appropriate shape
so the landing can be formed. Aluminum is considered instead of other materials due to its added strength
and ability to not deform during severe impacts. The aircraft has tail landing gear configuration, hence the
front landing gear included two tires and the rear included one. Both the landing gears were attached to the
fuselage using screws.
5.4 System and subsystem design:
Propulsion
The final propulsion system of the aircraft comprises of Rimfire 0.55 (480 kv) motor, 5s (5000mah) battery
and 14 x 7 APC propeller for each mission. The propulsion system helps the aircraft to complete every
mission of the competition convincingly, with enough battery percentage remaining. The team planned to
choose propulsion with maximum power settings to accommodate the payload and yield higher number of
laps in the charter and banner flight. The team devised this strategy to increase the overall score of the
team.
Figure 21: Empennage structure
Figure 22: Landing gear structure

Controls
In this section, a Servo PDI - 6215MG is used, these servos position the control surfaces such as the
horizontal stabilizers, elevators, rudders, and flaps. They deliver torque depending on their voltage input,
for 4.8 V (197.95 oz/in) and for 6V (212.79/in). This Servo weighs 62 grams and can lift to a maximum of
15 kg, which is lighter than DS3225MG Servo that weighs 75g that also lifts 15 kg as well. Concluding that
the PDI-6215MG better than the DS3225MG Servo. The Servo PDI-621MG is capable of delivering the
necessary torque in order to operate at a minimum and maximum speed of an aircraft.
Radio Control
The radio control used is a Futaba T8J 2.4Ghz S-FHSS 8-Channel Airplane Radio system along with an
R2008SB Receiver. In case of failure, a fail-safe option is available on all 8 channels, which means that it
will pass through a signal to the flight control surfaces in this occurrence to avoid failure. A Futaba T8J
requires four AA batteries in order to use the transmitter. This radio controller is programmable and
eliminates other aircraft to operate on the same channel. Aircraft features include: 6 programmable mixes,
flaperons differential rate, flap trim, elevator, Elevon, V-Tail mixing, Differential Ailerons, Airbrake/Landing,
(Flap/Elevator, Elevator/Flap, Aileron Rudder) Mixing, Gyro Sensitivity, Pitch Curve, throttle Delay,
Throttle/Needle Mixing and Idle down.
Restraint system
Restraint systems were manufactured with plywood from
a laser-cutting machine to restrain the passengers and
luggage into it. The idea of designing the restraint system
was kept to be simple and easy to load and unload,
especially taken into consideration the ground mission of
the competition. The space in the restraints below was
designed to be a tight fit for the passengers and luggage
to avoid movement midflight.
Banner Mechanism
The banner mechanism has great significance in this year’s competition. The significant constraints when
designing the banner mechanism were the placement of the banner, safe, and quick deployment. The team
came up with the idea to place the banner mechanism under the wing of the aircraft so that the cg doesn’t
change much. The banner mechanism consists of a two switch servo motor, a thick balsa plate, balsa rod
attached at the leading edge of the banner, two plastic rings attached at each end of the balsa rod, a ribbon
hooked to each end of the balsa rod and a plastic ring attached to the end of the ribbon. The figure 24
below depict the components used to assemble the banner mechanism.
Figure 23: Restraint system for passengers and luggage’s

Banner Mechanism Assembly
The banner is folded with the balsa rod and attached to the mechanism in a horizontal orientation. Most of
the ribbon is kept inside the banner, however some of the ribbon is used to fold the banner so it doesn’t
touch the ground. One end of the banner is attached to the balsa plate, and the other end is attached to the
servo mechanism. The Figures 25 and 26 below illustrate the assembly of the overall banner mechanism.
Figure 25: Front banner mechanism assembly
Figure 24: Banner mechanism components
Figure 26: Rear balsa plate

Deployment
The deployment system operates by placing the ring attached with the ribbon close to the servo motor and
the balsa rod ring next to it. During take-off, both the rings are locked using the servo rod. Deployment
occurs when the servo rod moves back one switch. The figures 27 and 28 below further illustrates the
deployment of the banner.
Release
The banner is released before landing the aircraft. During the deployment of the banner, the ring attached
to the ribbon is locked in between the servo rod. Before landing, the servo rod is moved further back, and
the ribbon is released. The figure 29 and 30 below illustrates the banner release after deployment.
5.5 Mission performance
The corresponding mission results were calculated below, depending on our aircraft design performance
discussed previously in the sections. The final score of each mission was calculated by the formulae given
in the rules. The flight mission scores in the table 20 below are compared with the best-assumed score
expected in this year’s competition.
5.6 Drawing Package
Figure 27: Banner and rope fixed
Deployment
Figure 28: Banner deployed
Figure 28: Ribbon end locked
Release
Figure 30: Ribbon end released
Table 20: Mission performance table

10.00
1.20
3.24
55.09
10.27
20.00
1.27
1.84
32.56
27.55
1.16
14.46
20.00
43.97
7.00
6.50
1.50
10.94
14.65
A A
B B
C C
D D
4
4
3
3
2
2
1
1
AIAA DESIGN BUILD AND FLY
2019/2020
3 VIEWS DRAWING
EMIRATES AVIATION UNIVERSITY
SIZE
C
SCALE 1:12 SHEET 1 OF 1DIMENSIONS ARE
IN INCHES
DATE:
DRAWN BY:
CHECKED
BY:
22/2/2020
MUHAMMED AHNUF
DR. HICHAM
MUHAMMED NAZIM
SKYWALKER

3
18
19
15
17
20
4
6
5
11
12
7
2
8
9
14
16
10
1
13
ITEM NUMBER COMPONENT MATERIAL QTY.
1 FUSELAGE FRONT COVER BALSA 1
2 BULKHEAD 2 PLYWOOD 1
3 SIDE FRONT FUSELAGE PLYWOOD 1
4 MOTOR MOUNT PLYWOOD 1
5 BOTTOM FUSELAGE PLYWOOD & BALSA 1
6 UPPER FUSELAGE PLYWOOD 1
10 REAR FUSELAGE COVER BALSA 1
11 ELEVATOR BALSA 1
12 TAIL BALSA 1
13 LANDING GEAR ALUMINIUM 1
14 REAR LANDING GEAR ALUMINIUM 1
15 RIBS BALSA 18
16 SHEAR WEB BALSA 2
17 RIBS COVER BALSA 2
18 FRONT RIB COVER BALSA 2
19 AILERON BALSA 2
20 WING ATTACHMENT ROD CARBON FIBER 1
A A
B B
C C
D D
4
4
3
3
2
2
1
1
2019/2020
STRUCTURAL ARRANGEMENT
SIZE
C
SCALE 1:12 SHEET 1 OF 1WEIGHT: 2.64 LBS.
DATE:
DRAWN BY:
CHECKED
BY:
20/2/2020
MUHAMMED AHNUF
DR. HICHAM
MUHAMMED NAZIM
SKYWALKER

A
B
C
DETAIL A
SCALE 1 : 1.75
1
DETAIL B
SCALE 1 : 1.75
2
DETAIL C
SCALE 1 : 2
3
D
EG
DETAIL D
SCALE 1 : 0.95
7
DETAIL E
SCALE 1 : 2.25
6
DETAIL G
SCALE 1 : 2.25
4
5
H
K
DETAIL H
SCALE 1 : 1.5
8
DETAIL K
SCALE 1 : 1.4
9
ITEM NO. COMPONENT DESCRIPTION QTY.
1 TAIL SERVO
RUDDER AND HORIZONTAL
TAIL SERVOS
2
2 BOTTOM SERVO
BANNER MECHANISM
CONTROL
1
3 ESC 100 A 1
4 PROPELLER 14 X 7 INCHES 1
5 MOTOR RIM FIRE 0.55 480 KV 1
6 BATTERY 5000 mAH Electrifly 1
7 RECIEVER R2008SB
8 LEFT WING SERVO LEFT AILERON SERVO 1
9 RIGHT WING SERVO RIGHT AILERON SERVO 1
A A
B B
C C
D D
4
4
3
3
2
2
1
1
2019/2020
SYSTEM LAYOUT
SIZE
C
SCALE 1:12 SHEET 1 OF 1ALL DIMENSIONS
ARE IN INCHES
DATE:
DRAWN BY:
CHECKED
BY:
22/2/2020
MUHAMMED AHNUF
DR. HICHAM
MUHAMMED NAZIM
SKYWALKER

F
H
BANNER TOWING
MECHANISM
B
C
BANNER MECHANISM TABLE
PAYLOAD TABLE
THE RING ATTACHED
TO THE BALSA PLATE
HOLDS ONE END OF
THE BANNER.
8
4
5
7
BANNER LEADIND EDGE
ATTACHED WITH BALSA ROD
& RING LOCKED BETWEEN
SERVO ROD. TWO SWITCH
SERVO ALLOWS DEPLOYMENT
& RELEASE OF BANNER.
1
4
5
3
2
6
DETAIL F
SCALE 1 : 1.5
3
1
DETAIL H
SCALE 1 : 1
4
2
RESTRAINT SYSTEM
ITEM
NO. COMPONENT MATERIAL QTY.
1 PASSENGER
MOLD PLYWOOD 1
2 LUGGAGE MOLD PLYWOOD 1
3 PASSENGER BALSA WOOD 4
4 LUGGAGE BALSA WOOD 4
ITEM
NO. COMPONENT DESCRIPTION QTY.
1 SERVO MOTOR RELEASE AND
DEPLOY 1
2 BALSA PLATE ATTACHED WITH
SERVO 2
3 PLATE HOLES FOR ROD 1
4 BANNER SATIN CLOTH 1
5 ROD FOR
BANNER
ROD WITH RING TO
ATTACH BANNER 1
6 SERVO ROD ROD TO LOCK THE
BANNER 1
7 RING PLACE REAR
BANNER 1
8 REAR BALSA
PLATE BALSA WOOD 1
A A
B B
C C
D D
4
4
3
3
2
2
1
1
2019/2020
PAYLOAD ARRANGEMENT
DRAWING
SIZE
C
SCALE 1:12 SHEET 1 OF 1ALL DIMENSIOS
ARE IN INCHES
DATE:
DRAWN BY:
CHECKED
BY:
22/2/2020
MUHAMMED AHNUF
DR. HICHAM
MUHAMMED NAZIM
SKYWALKER

6. Manufacturing
Manufacturing aims to construct a lightweight and robust design. The constructed design should have
enough structural integrity to withstand all flight loads. The team considered several manufacturing
processes. The manufacturing process chosen for each component represented the ideal combination of
ease of manufacturing, material weight, experience, and cost.
6.1 Manufacturing process investigation
Composite material construction
Composite material construction provides high precision and relatively better strength to weight ratio when
compared to the other manufacturing processes. The composite construction is very durable, but the
process is expensive and time-consuming. It involves complexities in the manufacturing as with many
subsequent procedures to be followed during the process.
3D-Printing
3D- printing saves time in designing complex mechanisms and structures which are relatively time-
consuming. This process helps in creating molds and other parts involving the least strength requirements.
The main concern of this manufacturing technique is its comparatively least strength to weight ratio.
Balsa / Plywood construction
Building with balsa/plywood offers to be effective than the other manufacturing processes. It can be altered
into different thicknesses, providing lower density than a composite structure and helping in manufacturing
a lightweight construction with desired strength characteristics. The process involves parts to be designed
in a CAD program and then precisely cutting them with a CNC laser cutting machine.
Foam construction
Foam helps in manufacturing lighter structures and are easily shaped in different shapes. Foam is a
cheaper manufacturing technique but are less useful when making structures with a large volume. Foam is
less durable compared to other manufacturing techniques.
6.2 Selection process
The techniques chosen were deeply investigated to optimize the manufacturing of all the components. The
final technique was chosen according to the system of Figure of Merits (FOM), rating each one with
importance factors.
Table 21: Manufacturing process selection
table

Cost: The limited manufacturing budget was allocated, so the team decided to reduce the cost with
consideration of not comprising with the quality of the construction.
Experience: The techniques were selected based on the team’s prior experience and their capabilities with
the manufacturing techniques.
Ease of Manufacture: The team was manufacturing aircraft for the first time, so it is essential to follow an
easier manufacturing procedure that can provide high-quality parts within a satisfactory timeline.
Strength: The structural integrity of the aircraft depends on the selected manufacturing technique. Hence,
the technique was selected depending on its contribution of high strength to weight ratio to the aircraft’s
robustness.
Weight: The weight of the aircraft shares a vital role in the selection of technique as with this year’s mission
requirement of short take-off in 20ft. The aircraft needs to be manufactured with a technique adding the
least weight to the structure.
Comparison of Materials and Manufacturing techniques:
The materials and manufacturing techniques were compared in the Figure of Merit (FOM) system. Both
tables below were rated from 1 (lowest) to 5 (highest) and then multiplied with FOM importance factors to
distinguish amongst them the best material and manufacturing technique. The final material selected is
Balsa and plywood, which will be manufactured by Laser cutting technique. Hence, the team evaluated it
as a cost-saving and less time-consuming decision. The tables 22 and 23 depicts the materials considered
for manufacturing.
Table 22: Material comparison table
Table 23: Material selection table 2

6.3 Manufacturing of Parts
Wing
Ribs and spars:
The team decided to use a mix of plywood and balsa for the construction of ribs. The areas of the wing
exposed to the maximum stress were constructed with plywood to enhance the structural strength under
flight loads. The leading-edge rib of each semi-wing was made up of plywood, whereas the rest of them
were of balsa. The leading-edge of each rib was design as a flat surface to ease the process of attaching
it with the spar. The figure 31 below shows the attachment of balsa and plywood ribs attached to the spar.
Reinforcement:
The team then manufactured shear webs with high precision with laser cut and glued them to the quarter
chord of the robs. The shear web was supported additionally with plywood doubler to strengthen the root
of the wing. The figure 32 below shows the arrangement of reinforcement added to the ribs.
Wing cover and coating:
The wing was covered with a sheet of balsa by gluing it to the top of the ribs. The cover was designed in
CAD and then precisely cut in laser to ensure the cover fits the right size of the wing. The section of it were
kept hollow to reduce the weight of the overall wing structure without effecting the strength of the structure.
Figure 31: Wing ribs and spars
Figure 32: Reinforcements used for the wing structure

The wing was then finally coated with a heat shrink adhesive infused plastic covering called Monokote. It
protects the wing from significant damage and is very durable.
Fuselage
The fuselage structure comprises of mainly plywood material. After an in-depth investigation, it is found that
plywood has greater structural integrity as compared to balsa wood. The construction of fuselage
components is mainly composed of CNC laser cutting machine. The fuselage structure consists of frames,
bulkheads, outer cover, and reinforcements.
Frames: The fuselage structure comprises of 4 frames. Two side frames and the top frame are made using
a thick sheet of plywood. The bottom frame is reinforced with added plywood sheet as it supports most of
the loads on the fuselage. The required shape is achieved using a laser cutting machine.
Bulkheads: The fuselage comprises of 6 bulkheads. They are all made using plywood. Bulkheads are
shaped by laser cutting thick rectangular sheets of plywood. The bulkheads closer to the nose 1 and 2
provide the primary structure to hold the aircraft wing with the fuselage. The other bulkheads hold the
primary structure of the aircraft. They are attached to the fuselage frames using epoxy adhesive.
Motor mount: The motor mount comprises of three thick sheets of plywood. Rectangular plywood sheets
are used to drill holes for the motor attachment. The sheets are attached together using epoxy adhesive.
Lightening holes: The overall weight of the fuselage structure is reduced by removing the unwanted
material without any loss of overall stiffness and strength of the aircraft.
Surface covering: Monokote is used to cover the fuselage structure. It provides adequate surface rigidity
and smoothness.
Empennage
The horizontal and vertical tailpieces were made from plywood sheets. The sheets shape was formed from
the CNC laser cut machine. The elevator and rudder are formed using separate sheets. They are connected
Figure 33: Wing cover

to the tail using hinges. The empennage is covered using Monokote, which increases its overall strength
for bending and torsion.
Landing gear
The undercarriage was made using a cut out sheet of 2024-T6 aluminum. The sheet was appropriately
bent using a press brake. It was attached to the fuselage structure using screws.
Passengers and Luggage compartments
The team decided to initially 3D print the passenger and luggage. The passenger was printed from an ABS
plastic material, and luggage was printed from a carbon fiber material. The team discarded the option of
using 3d printed passengers and luggage as it was challenging to add the required weight in the 3D printed
passengers then compared to the wooden passengers. The wooden material provided a much better
surface finish, which eased the process of painting and decoration. The table 24 below depicts the material
selected to manufacture the payload.
Weight addition
Passengers are made up of wood, which initially weighed 1 ounce had to be added with weight to meet the
required minimum weight of 4 ounces in the competition. The team came up with a solution to make the
Table 24: Payload material selection table

passenger hollow from inside and add steal chunks in a metal scrap recycling. The hole made to add weight
was filled with adhesive to permanently close it.
Luggage was also manufactured from wood and was added with additional weight to meet the requirement
of 1 ounce. The luggage originally weighed 0.42 ounce, which was then added weight by drilling a hole and
adding metal bolt nuts into it. The final weight of the luggage after adding weight came approximately to 1
ounce.
6.4 Manufacturing Milestones
The team divided the manufacturing of the aircraft above into different milestones to complete the work
accordingly to the set deadlines. The manufacturing was planned in a manner that the team saves enough
time to perform rigorous testing on the final prototype. The figure 35 below shows the planned and actual
manufacturing plan for the competition.
Figure 294: Weight addition in the passengers
Figure 35: Manufacturing plan of EAU Aviators

7. Testing plan
The airworthiness of the aircraft was verified by various testing and checks that must be conducted to
validate the reliability, structure, and performance of the aircraft, its components, and its various systems.
Data obtained from these tests depict the errors and drawbacks of the current design and improve any
future designs of this aircraft. This testing plan includes both ground tests and flight tests.
7.1 Testing schedule
Test methods are divided into ground tests and flight tests. The testing schedule constructed below helped
the team to achieve a proper sequence of the testing phase. This will also help to obtain appropriate data
and feedback regarding the current design. Thus, displayed below is a Gantt chart used to monitor the
progress of all the tests such that it is conducted accordingly.
7.2 Testing objectives
The main objective of testing aircraft is to match the requirements of the competition. The testing phases
are divided into subsequent sections as Propulsion, Structure, stability/control, and flight performance.
Propulsion testing
Propulsion testing was conducted to analyze the capability and the performance of the propulsion system
individually and the combination of various subsystems. Propulsion testing narrows down the motor
selection to the best-suited motor. The objective of this test type is to verify the speed and range related
characteristics of the aircraft. This is achieved using thrust tests and test flights to record these parameters.
Figure 36: Testing plan of EAU Aviators

Propulsion testing will be conducted by constructing a thrust stand. The thrust stand helps determine data
such as:
❖ Propeller efficiency values
❖ Thrust power (Hp)
❖ Current Amp drawn
Using this data and reference data regarding the performance of the motor from MotoCalc, the team was
able to determine the performance difference and motor capabilities. Using such a data, it will be easier to
further optimize the thrust characteristic of the aircraft such that better scores can be achieved.
Thrust stand:
The avionics team decided to make a thrust stand to compare the actual propulsion system performance
to the theoretical predictions from Motocalc software. The design of the thrust stands circuit measures
thrust, rpm, voltage, and current. The main objective of the avionics team was to compare the static thrust
of the motor with different propeller combinations and find the right match of propulsion components. The
circuit below was printed on to a PCB board and then integrated with LCD to display measurements.
Figure 30: Block diagram of thrust stand
Figure 38: Thrust stand setup

Structural testing
The Structural testing focuses on the wing loading capacity and the overall structural integrity of the aircraft.
This type of testing is considered as a necessity as the aircraft is meant to:
❖ Able to fly in windy conditions
❖ Built to carry payload and the banner towing mechanism
Thus, it seems reasonable to ensure the wing and the airframe, in general, can withstand bending and
torsion loads. One of the essential types of testing in this subcategory is the wingtip load test. This was
conducted by simulating the loads that this aircraft would typically experience on the wing of the aircraft.
The performance of the wings was then analyzed. The maximum wing deflection was evaluated with the
payload and other forces on the wing using Finite element analysis. The data proved the wing to be strong
enough to withstand mission conditions.
Wing Tip Test
A wing tip test was conducted to calculate the maximum wing tip loads. The test helped in the evaluation
of the loads while in cruise with a maximum weight.
Figure 39: Avionics team working on the thrust
stand Figure 31: Final thrust stand setup
Figure 321: Wing tip test performed on sky walker

Stability and Control testing
The objection of this stability and control testing category is to confirm the aircraft's CG location, range of
the flight control surface movements, and knowing the optimal flying condition to help achieve the best
performance. This can be achieved by conducting the tip tests as well as flight tests such as takeoff,
landings, and mock runs of the missions. The aircraft was tested in windy weather conditions to determine
the aircraft stability and control in a similar weather environment in Wichita, Ks. The pilot gave feedback
regarding the control and flying characteristics of the aircraft in such an environment. It was recorded that
the aircraft can be flown headwind with a wind speed of 16-18 ft/s.
7.3 Flight performance testing
In this section, the testing focuses on improving flight performance during mission runs. The aircraft
performed a routine flight of each mission. The pilot tested the handling capabilities and characteristics of
flight for the aircraft and suggest changes that can be made to improve control aircraft. This testing phase
began with a maiden flight of the first prototype and then followed by the second one.
7.4 Flight checklist
The flight checklists displayed below are the checklist for the propulsion test and the checklist for the flight
test. The checklists are making it easier to achieve the intended specification and flight characteristics to
score well.
Table 25: Flight test performance

Propulsion test (pre-test) checklist
The checklist listed used before the start of the propulsion test ensures the safe integration of the propulsion
components. The team rigorously performed the propulsion pre-flight check as propulsion serves as the
most sensitive section of the aircraft. Hence, the safety of the aircraft was not compromised.
Pre-Flight Checklist
The following pre-flight checklist planned ensured the aircraft undergoes the testing phases without any
complications to the system and the frame. The checklist eliminates the risk of failure and ticks all the
required boxes for successful missions.
Date: Prototype:
Time: Battery type:
Area Task Condition Check
Propeller installement…......................…......................Yes/No
Motor Installement…......................…......................Secure/Detached
Connections…......................…...................... Connected/Detached
Power source (Batteries)…...................... Connected/Detached
Throttle control…......................…...................... Up/Down
Data Systems…......................…...................... On/Off
Custom code…......................…......................Running/Inoperative
Testing Rig…......................…...................... Safe/Unsafe
Propulsion Testing
Additional information:
Pre-Test Checklist
Figure 33: Pretest checklist
Date: Prototype:
Time: Battery type:
Location: Weather condition:
Area Task Condition Check
Battery connected and fastened…......................…......................Established
Servo motors functional…......................…......................Operational
Motor functional…......................…...................... Operational
Power source (Batteries) charge…...................... Max charge
Throttle control…......................…......................Operational and Down
Reciever and RC connection…......................…......................Established
Passenger/Luggage position…......................…......................Fastened
Banner mechanism status…......................…......................Secured and stable
Pins position…......................…...................... Locked
Position of the COG…......................…...................... Ideal
Connection to flight controls…......................…......................Established
Rudder function…......................…...................... Operational
Elevator function…......................…...................... Operational
Flap function…......................…...................... Operational
Closed position…......................…...................... Yes
Attached…......................…...................... Yes
Functionality…......................…...................... Operational
Propulsion and Avionics
Control surfaces
Payload
Banner Mechanism
Additional information:
Pre-Flight Checklist
Figure 34: Preflight checklist

8. Performance Results
8.1 Demonstrated performance of key subsystems
Propulsion performance
Propeller efficiency can be estimated using the thrust stand. The propellers sizes used for the test are 14 x
7, 12 x 8, and 15 x 8. The 14 x 7 propeller diameter is the most optimal among the three based on its
efficiency at flight condition. Thus, this type of propeller is selected for the Rimfire 0.55. The system would
be inefficient if the propeller and the motor are not matching together as the combination will be unable to
produce the necessary power in every mission. The figures 43 and 44 below illustrate comparison of
different propeller sizes with thrust and power produced.
Banner flight analysis
Mission 3 results are vital for the team’s chances of succeeding in the competition. The maximum banner
length of 60 inches being selected, a strategy had to be devised to achieve a higher M3 score. The team
decided to perform several flight tests at various throttle percentages. This was done to evaluate the battery
consumption at each throttle percentage. The aircraft was tested in between safe throttle percentages from
50% - 75%. The figure 45 below depicts the results of the mission 3 flight tests.
Figure 43: Comparison of predicted and actual values of thrust Figure 35: Comparison of predicted and actual thrust

The results illustrate that the aircraft can perform mission 3 at high throttle percentages. However,
considering the weather uncertainties and aircraft safety, the team decided to fly the aircraft at 65% throttle
for the first half of the flight. This will save the maximum battery percentage for the latter half of the flight.
Hence according to the battery percentage remaining, the pilot will fly the aircraft faster during the last few
minutes. This will not only ensure aircraft safety but also allow us to obtain a more significant number of
laps and higher overall M3 score.
8.2 Demonstrated flight performance of completed aircraft:
The team tested “Skywalker” for multiple flight tests to evaluate and validate the predicted performance of
the aircraft. Mock flight tests were conducted to check the aircraft’s performance in each mission. The
aircraft was tested within weather conditions similar to Wichita, Ks. The aircraft performed each mission on
a windy day, hence helping the team to analyze its performance in most difficult flying conditions. The flight
test helped to determine different required performance parameters in different missions. The table below
illustrates the results of each flight test.
Figure 36: Throttle percentage vs battery percentage left

The flight test was also compared to performance predictions. The data is tabulated below in table 27.
8.3 Performance improvements
The testing results of the aircraft made it evident that an increase in battery capacity is required. The team
decided to increase the battery from 4s cell to 6s cell. The significant increase in battery capacity ensured
that the aircraft could produce a higher amount of power. Mission 3, producing the most drag, needed the
extra power to achieve maximum results. Moreover, flight tests of missions 1 and 2 illustrated the vast
capabilities of the aircraft. The aircraft speed, endurance, and stability were optimal in the first two missions.
Whereas aircraft in mission 3 was unable to deploy the banner successfully in the first attempt. The team
observed from the unsuccessful flight, the banner needed to have a weight on the bottom-leading edge for
it to be deployed, and the banner must not be excessively folded. This helped the banner to deploy much
faster and eliminate the risk of rope tangling in the midflight.
Table 26: Flight tests of each mission
Table 27: Predicted and actual performance of sky walker

After further improving the aircraft. All the missions were performed again with optimal results. The final
flight tests of the aircraft are shown below
Figure 37: Real flight path of sky walker
Figure 38: Mission 1 and 2 flight tests
Figure 39: Successful flight test post banner deployment

Bibliography
Airfoil tools . (2020). Retrieved from Airfoil tools: http://airfoiltools.com/userairfoil/index
Jr., J. D. (2010). Fundamentals of aerodynamics. New York: McGraw-Hill .
Katz, J. (2016). Automotive Aerodynamics . West Sussex: Wiley and Sons .
Kurowski, P. M. (2018). Engineering Analysis with SOLIDWORKS Simulation . Mission: Kansas.
Lent, C. S. (2013). Learning to Program with MATLAB. Chennai: Wiley.
Nelson, D. R. (1998). Flight Stability and Automatic Control. San Francisco: McGraw-Hill.

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