Design and Numerical Analysis of a Vertical Take-Off and Landing

(VTOL) Fixed Wing Unmanned Aerial Vehicle (UAV).

Introduction:
As the deliveries of product increasing day by day in cities, companies are facing some major
problems in delivering the products. Since the number of deliveries increasing day by day the
requirement for manpower is also increasing day by day, which increases the delivery cost.
Also, they are failing to deliver the product in time due to limitations in surface road capacity.
As well as due to constantly increasing traffic. Hence an Unmanned Areal Vehicle can be
considered a perfect solution for these problems. A Vertical Takeoff and Landing Fixed Wing
UAV can deliver the product in a very short time with high accuracy and safety. It can take off
vertically from the store and can deliver the product in the exact location as requested by the
consumer. After these considerations, a VTOL Fixed Wing completely electric Unmanned Areal
Vehicle is being proposed in this design report.
Mission profile:

Cruise Cruise
Loter Forward descent
Forward Climb Forward descent Forward Climb

VTOL Climb VTOL descent VTOL Climb VTOL descent

Ground Payload delivery

Fig.1. Mission profile

Segment details:
 Range: 6 Km
 Loitering time: 5 min
 Emergency hovering time: 2 mins.
 Max speed: 115 km/h
 Cruise speed: 108 km/h
 Cruise altitude: 250m
 Payload mass : 0.5 kg
 Vertical climb speed: 3 m/s
 Forward climb speed: 3 m/s

Constraint Analysis:
A constraint diagram was developed based on the mission profile data. The constraint graph was
developed for the cruise, climb and stall mode.

𝑇 𝑓𝑤 𝐶𝐷0 𝑊/𝑆
( ) =𝑞∗ +𝑘∗
𝑊 𝑐𝑟𝑢𝑖𝑠𝑒 𝑊/𝑆 𝑞

𝑇 𝑓𝑤 𝑉𝑣 𝐶𝐷0 𝑊/𝑆
( ) = +𝑞∗ +𝑘∗
𝑊 𝑐𝑙𝑖𝑚𝑏 𝑉 𝑊/𝑆 𝑞
𝑊 1 2
( ) = ∗ 𝜌 ∗ 𝑉𝑠𝑡𝑎𝑙𝑙 ∗ 𝐶𝐿
𝑆 𝑠𝑡𝑎𝑙𝑙 2
The diagram was plotted for thrust loading against wing loading using the above equations. From
the constraint diagram, minimum wing loading for this particular mission profile was found to be
266 N/m2. And for this wing loading, the thrust to weight ratio was found to be 0.26. Which well
below the maximum allowed thrust to weight ratio as mentioned in the problem statement.

Fig.2. Constraint Analysis

Mass Calculation:
After the design of the mission profile and constraint diagram, the approximate mass of the
Fixed-wing VTOL UAV was calculated. The total weight of the UAV was calculated as follows:
WTO = Wempty +Wenergy+Wpayload
Now the empty weight of the UAV was the total sum of the structural weight and the propulsion
weight as well as the avionics weight of the UAV.
Wempty=Wstructure+Wpropulsion,FW+Wpropulsion,VTOL+Wavionics
The energy weight refers to the weight of the battery.
Wenergy=Wbattery
Now it is not possible to calculated or estimate the weight of the battery at this stage of the
design process. For that reason, the mass fraction of the battery was calculated at this design
stage.
The mass fraction of the battery of a UAV is given by the following Equation:
tP
𝑀𝐹𝑏𝑎𝑡𝑡 = M
TO Espec ηbatt fusable

Using this generalized equation was used to find out the battery mass fraction for each segment.
The total mass fraction of battery for the entire mission was found to be approximately 0.323.
A similar approach was taken in determining the weight of the avionics and structure and
subsystem.
A reasonable estimation of the structure mass fraction is 25-35%. In this design, 35% were
selected. The avionics mass fraction was selected to be 5%. And as the proposed UAV will
contain a small camera for real-time monitoring, a subsystem mass fraction of 6% was
considered.
And as mentioned before, the payload weight is 0.5 kg.
Based on these data the total mass of the UAV was calculated through an iterative process. After
the convergence of this iterative process, the gross weight of the UAV was found to be 5.5 kg.

Propulsion selection:

At this stage, the weight of the UAV is 5.5 kg.

As the thrust to weight ratio of the UAV is 0.26 the Thrust required for the forward flight mode
of the UAV is 14.029 N.
Now the required motor power for the cruise was calculated through the formula:
T∗V
𝑃𝑚𝑜𝑡𝑜𝑟 =
ηpropeller

Taking the propeller efficiency 65% the power required for the cruise is calculated to be 650 W.
A brushless DC motor will be used that can supply a maximum power of 1 hp.
For the VTOL mode, the thrust to weight ratio was calculated as follows.
𝜌 ∗ (𝑉𝑟𝑜𝑐 )^2 ∗ (𝑆𝑡𝑜𝑡𝑎𝑙 /𝑆𝑤𝑖𝑛𝑔 )
T/W = 1 + 𝑇𝑚𝑓
(𝑊/𝑆)

From this, the thrust to weight ratio was calculated as 1.165.

So the total power required for vertical climbing was 570 W. As 4 motor will be used, the power
required per motor is 142 W. So 4 brushless DC motor will be used. Whose maximum power
supply is 200 W.

Airfoil Selection:
From the constraint diagram the wing loading, (W/S) was selected as 266 N/m2. So from this, the
required lift coefficient during the cruise segment can be obtained.
W
2∗(S)
𝐶𝑙,𝑎𝑖𝑟𝑓𝑜𝑖𝑙 = 2
0.9 ∗ 0.95 ∗ ρ ∗ V𝑐𝑟𝑢𝑖𝑠𝑒
From the above equation, the required airfoil was calculated to be 0.568.
The NACA 4415 airfoil was selected as wing airfoil. It is very suitable for this type of low-speed
UAVs. And it meets the required lift criteria. It has a thickness ratio of 15%. Which will help to
delay the stall. Fig.3 shows the NACA 4415 airfoil geometry. Fig.4 shows the airfoil lift at
different angles of attack. It can be seen that it has a stall angle of 12 degrees. And the angle of
zero lift is approximately -4 degrees. Fig.5 Shows the Cl/CD at different angles of attacks.

Fig.3 NACA 4415 airfoil

 Fig.4 Cl vs Alpha graph of NACA 4415 airfoil

Fig.5 Cl/Cd vs Alpha graph ofNACA 4415 airfoil

Wing Design:
As mentioned above the NACA 4415 airfoil is used in the main wing geometry. The area of the
wing was calculated as follows:
WTo 5.5 ∗ 9.81
𝑆𝑤 = = = 0.19919 𝑚2
W 266
(S)

Since the wingspan was limited to 120 cm in the problem statement, the span of the designed
UAV is selected to be 120 cm. The dimensions of the main wing are given below.
 Wingspan: 120 cm
 Area: 2000 cm2
 Taper ratio: 0.8
 Aspect Ratio: 7.2
 Root chord: 18.45 cm
 Tip chord: 14.75 cm
 Mean aerodynamic chord: 16.67 cm
 Wing sweep: 0 deg.
 Wing dihedral: 0 deg.
 Wing twist: 0 deg.
One of the most important phenomena in wing design is the wing angle of incidence. It is
defined as:
𝐶𝐿,𝑐𝑟𝑢𝑖𝑠𝑒
𝑖𝑤 = ( ) + 𝛼0𝑙𝑟 − 0.4 ∗ 𝜀𝑡
𝐶𝐿𝛼

Again CLα is defined as :

2∗𝜋∗𝐴
𝐶𝐿𝛼 =
2 + √4 + (𝐴2 (tan𝛬21/2 +1 − 𝑀2 )

Using these equations the iw was calculated as 2.48.

Horizontal Tail design:

The horizontal tail was calculated using the Volume coefficient method. According to this
method:
𝑆𝐻𝑇 ∗ 𝑙𝐻𝑇
∀𝐻𝑇 =
𝑆𝑤 ∗ 𝐶𝑤
Here 𝑙𝐻𝑇 is the distance between the mean aerodynamic center of the wing to the mean
aerodynamic center of the tail.
For this current design, the horizontal tail volume coefficient was taken as 0.72. And the ratio of
the main wing area to the horizontal tail area was taken as 0.2. So using these values the 𝑙𝐻𝑇 was
calculated as 59.99 cm. The overall dimensions of the horizontal are given below.
 Area: 398.34 cm2.
 Span: 55.00 cm
 Taper ratio: 1
 Aspect ratio: 7
 Chord: 7.28 cm
 𝑙𝐻𝑇 : 59.99 cm
 Airfoil: NACA 0012
The horizontal wing tilt angle was later determined by numerical simulation in xflr5 software.

Vertical Tail design:

The vertical tail was designed by the same approach. The volume coefficient is expressed as:
𝑆𝑉𝑇 ∗ 𝑙𝑉𝑇
∀𝑉𝑇 =
𝑆𝑤 ∗ 𝑏𝑤
Where 𝑙𝑉𝑇 is the distance between the mean aerodynamic center to the vertical tail aerodynamic
center.
Here 𝑆𝑊 /𝑆𝑉𝑇 was taken as 0.1. The value of 𝑙𝑉𝑇 was calculated as 58 cm. For this particular type
of UAV
The overall dimension of the vertical tail is given below:
 Total area: 199.17 cm2.
 Span: 11.8 cm
 Taper ratio: 0.8
 Aspect ratio: 1.4
 Root chord: 9.37 cm
 Tip chord 7.49 cm
 Mean aerodynamic chord: 8.46 cm
 Sweep: 20 deg.
 𝑙𝐻𝑇 : 59.99 cm
 Airfoil: NACA 0012
A 3D view of the designed UAV is given in fig.6. The xflr5 software was used to design this
UAV. An overall sketch of the entire UAV is given in Fig.7.
 Fig.6 3D view of the UAV

Fig.6 2D sketch of the UAV

Center of Gravity:
All the mass was inputted in the xflr5 software. And the location of the C.G was found as,
X= 5.43 cm
Y= 0.00
Z= 0.074 cm

Numerical Analysis:
The flight characteristic and stability of the plane was analyzed using the xflr5 software.
Initially, when the tilt angle of the horizontal tail was at zero degree, the UAV had a negative
pitching moment at zero degree angle of attack. This was due to fact that the main wing was at
an incidence angle of 2.5 deg.
After some iterative simulation in the xflr5, it was found that, if the horizontal tail is slightly
angled upward (0.2 deg) the pitching moment of the aircraft becomes zero. Fig.7 shows the
UAV’s pitching moment against the angle of attack. Here we can see that at zero degree angle of
attack the pitching moment is zero. So it is clear that the UAV will be stable on cruise mode.
Fig.8 Shows the total lift of the UAV against the pitching moment. From this graph, it is clear
that at the zero pitching moment the UAV will produce sufficient lift to fly. Fig.9 is the Cl/Cd vs
Alpha graph of the aircraft

Fig.7 picthing moment vs Alpha graph

 Fig.8 Cm vs Cl graph

Fig.9 Cl/Cd vs Alpha graph

Stability Analysis:
After designing the entire aircraft and analyzing it at the cruise mode a stability analysis was
conducted on the UAV in the xflr5 software.
The stability was calculated for zero degree angle of attack. It was observed that at zero degree
angle of attack if the nose is pitched upward due to some reason, the UAV came back to its
original stable position in less than one second. Which indicates that the UAV has longitudinal
static stability. Fig.10 explains the longitudinal static stability of the UAV.
At the same time, another analysis was conducted on the UAV for lateral stability. It was
observed that if the UAV rolled in any direction, it came back to its original stable position in
less than one second. Which indicates that the UAV has lateral static stability. Fig.11 explains
the lateral static stability of the UAV.

t = 0 sec t = 1 sec

Fig.10 Longitudinal stability analysis

t = 0 sec t = 1 sec

Fig.10 Lateral stability analysis

Design file drive link:
https://drive.google.com/drive/folders/1QTDHqJs8sl3rFkRg6JdB7heaKGz2h9da?usp=sharing

Conclusion:
A VTOL Fixed Wing UAV was proposed in this design report. The UAV has a range of 6 km. It
can deliver any payload of 0.5 kg within that range. The wingspan of this UAV is only 120 cm.
Making it capable of landing in a very short area. Also, it has a very moderate empty weight
which increases the efficiency of the UAV. Also, it has zero carbon emissions. It is completely
electrical and doest not need any fossil fuel. Overall it is a very efficient zero-carbon emissions
Unmanned Areal Vehicle.