FDFDF

1. KLS GOGTE INSTITUTE OF TECHNOLOGY
Belagavi, Karnataka
Department of Mechanical Engineering
&
Department of Aeronautical Engineering
SAEINDIA AEROTHON 2021
Design Report
Team Vayuputras
Faculty Advisor
Prof. Parameshwar Banakar
Department of Aeronautical Engineering
KLS Gogte Institute of Technology, Belagavi
Team Number 049
Shane Fernandes (C)
Tejas Bane
Jeetendrakumar Garag
Nupur Bagi
Rohit Raghavan
Jeet Thakkar
Tejas Takkekar

2. The V-7, concept Rendering
1

3. 2

4. TABLE OF CONTENTS
Statement Of Compliance 2
Table Of Contents 3
List Of Figures 4
List Of Symbols And Acronyms 5
1.0) Introduction
1.1) Objective
1.2) Problem Statement And Requirements
1.3) Mission Profile
6
6
6
7
2.0) Conceptual Design
2.1) Research
2.2) Design Analysis and Review Process
2.3) Design Selection Process
8
8
8
9
3.0) Preliminary And Detailed Design
3.1) Wing Selection
3.2) Fuselage Sizing
3.3) Aircraft Performance
3.4) Propulsion Selection
3.5) Material Selection
3.6) Detailed Weight Breakdown
3.7) Flight Envelope
3.8) Landing Gear
3.9) Stability and Controls
3.10) Communications Systems
3.11) Payload
3.12 ) Schrenk's curve, Spar Design, Rib
Sizing etc.
3.13) Endurance Calculation
9
9
12
13
14
16
17
17
18
20
21
22
23
23
4.0) Computer Aided Design details 25
5.0) Computational Analysis 26
6.0) Optimized Design 26
7.0) Commercially Off The Shelf Parts 27
8.0) Practical Application and Feasibility 28
9.0) Innovation 28
References 29
Appendix 30
3

5. LIST OF FIGURES:
Figure no. ………………………...……...….Figurename………..…………………………………..Page no.
1.3.1………………….Mission profile for Cargo……………………...………………………………………7
1.3.2………………….Mission profile for surveillance……………………………………………….……….7
2.3.1………………….Design methodology in aircraft design systems engineering approach…………….…9
3.1.1………………….Comparison of various fuselage airfoils……………..………………………….....…10
3.1.2………………… Fuselage airfoil - MH93……………………………..………………………………..10
3.1.3………………… Comparison of various wing airfoils………………..………………………………...10
3.1.4………………….Wing airfoil - LA2573A………………………….…………………………………...11
3.1.5………………….CL vs. ⍺………………………………………………………………………………...11
3.1.6………………….Blended wing structure……………………………….……………………………….12
3.4.1………………….Motor dimensions……………………………………………………………………...15
3.4.2………………….Thrust vs. aircraft airspeed…………………………………………………………….15
3.5.1…………………..Pultruded Carbon Fibre…………………………………………………………....….18
3.7.1………………….Vn diagram………………………………………………………………………....…17
3.7.2………………….Limit combined envelope…………………………..…………………………….…....18
3.8.1………………….Landing gear……………………………..……………………………………….…....19
3.8.2………………….Dual Parachute System…………………..……………………………………….…...19
3.9.1………………….Control system………………………..………………………………………….........20
3.10.1……………... XLRS GCSD4……………………………………………………………....…21
3.10.2………………...Communication system………………………………..………………………….…..21
3.11.1………………...Schrenk’s curve………………………………….………………………………..…...21
3.11.2………………...Spar and rib placement……………………………….………...……………………...22
4.1…………………....Frame……………………………………………………………………………….....23
4.2…………………….Component placement………………………………………………………………..23
4.3…………………….Front cruise view……………………………………………………………………..23
4.4…………………….Rear cruise view……………………………………………………………………...23
4.5…………………….Payload bay open……………………………………………………………………..24
4.6…………………….Flaps engaged………………………………………………………………………....24
5.1…………………….Static pressure………………………………………………………………………...25
5.2…………………….Turbulent pressure…………………………………………………………………….25
5.3…………………….Velocity………………………………………………………………………….……25
6.1……………………CL vs. ⍺…………………………………..…………………...…………….………....26
6.2…………………….CL/CD vs. ⍺………………………………………….………………….....……...…..26
6.3…………………….CM vs.⍺..…………………………………………………………………….………..26
6.4…………………….CD vs. ⍺………………………………………………..…………………...…………26
6.5…………………….CM vs. CL ………………………………….………………………………..…………26
6.6…………………….XFLR5 analysis of the final model…………………….……………………………..26
7.1…………………….Commercially off the shelf parts………………………...……………………………27
4

6. LIST OF SYMBOLS AND ACRONYMS
e Span Efficiency factor
5

7. 1.0) INTRODUCTION
This document details the final design made by the Vayuputras Team from KLS Gogte Institute of
Technology, with the purpose of participating in SAEINDIA AEROTHON 2021. The report explains
the methodology, overall design, analysis, performance, used to build this Unmanned aerial vehicle
named “V-7” . The main objective is to design an unmanned aerial vehicle capable of transporting
medical loads as well as having a precise telemetry system for surveillance. In addition, it must fulfill
its mission by following the SAEINDIA design requirements and be an aircraft safe enough to
operate under different mission profiles.
1.1 ) OBJECTIVE
The main objective of participating in the SAEINDIA AEROTHON is to be able to contribute
towards an Atmanirbhar Bharat and help the sector of unmanned space technology reach newer
heights with our innovative contributions. It will soon be a thing of the past when India will have
looked towards other countries to develop and procure UAV systems. Events like SAEINDIA
AEROTHON help boost our creativity as well as challenge our thinking. All this coupled with
passion and patriotism will play a huge role in endowing our nation with ground breaking new
technologies.
Through this competition we can improve our communicative and interpersonal skills which are vital
for working together in multidisciplinary teams.This along with research and implementation of our
innovations will open up new opportunities and career paths. Furthermore, we will also be able to
unleash and recognise our full potential as future engineers by working on different aspects that go
towards the successful building and completion of a project.
1.2 ) PROBLEM STATEMENT AND REQUIREMENTS
The problem statement suggests the development of an UAV which can help with transportation of
medicines and other essential commodities along with surveillance in remote areas which also fulfill
the following specifications:
S. No. Design Requirements
1 Minimum Endurance – 3 Hrs
2 Payload – 6 Kg
3 Maximum Weight – 50 Kg
4 Max Speed – 150 km/hr
5 Above Sea Level (ASL) – 6000m
6

8. 1.3 ) MISSION PROFILE
UAVs are used for a multitude of operations right from cargo, surveillance,reconnaissance to
research.The need arises for Indian made UAVs which are robust and versatile and can be used for a
range of missions within and outside our borders. Keeping in line with the problem statement for
SAEINDIA AEROTHON we have decided to design our UAV exclusively for cargo delivery. The
cargo delivered will primarily be Medical supplies ranging from tablets, IV’s, vaccines to small
organ transplant containers.
Below is highlighted the typical mission profile for our UAV. On loading the cargo in the payload
bay and giving fly commands the UAV will steadily climb after taking off. Following this it will
cruise out to reach the desired location within a ceiling altitude of 3000m.Once the location is
reached, an onboard camera will be activated which can be used to monitor the payload drop and the
UAV will descend to a lower altitude from where the payload drop of the medicines, essentials will
take place. Later, the UAV will climb back to a higher altitude and cruise back only to descend back
at the initial or reprogrammed position.
Fig. 1.3.1 - Mission profile for cargo
The designed UAV can also be equipped with a modular camera which will allow for exclusive
surveillance missions in the same profile. It has been designed such that an exclusive image capture
system in the form of a high end surveillance camera can also be added in the payload bay as an
extension.This makes the UAV highly versatile to different missions that are arising in our domestic
market.
Fig. 1.3.2 - Mission profile for surveillance
7

9. 2.0 ) CONCEPTUAL DESIGN
2.1 ) RESEARCH
The blended wing aircraft was our primary option, so we dove deeper into the various aspects
surrounding it through comparative studies on conventional aircraft. After careful consideration,
discussion and evaluation of the various kinds of aerial vehicles which can be suitably used to
successfully complete the expected tasks while meeting all the required design specifications, a
blended wing UAV is going to be our approach.
Initially, the team was divided into two divisions, namely Design and R&D and were tasked with
scouring resources for extensive study of current unmanned aircraft as well as future developments
from around the world. Through the use of standard, internationally prescribed aircraft design,
control, stability, innovation and analysis books and resources the BWB was finalised and verified
through the subsequent engineering procedure.
2.2 ) DESIGN ANALYSIS AND REVIEW PROCESS
Due to the fact that BWB UAV are relatively uncommon, commercial aircraft such as the Airbus
MAVERIC and Boeing 797 BWB were carefully studied along with military BWB such as the
Northrop Grumman X-47B and the B2 to gauge the fundamental differences in flight
construction,control and design. It is a known fact that the blended wing craft has unstable pitch
owing to the absence of the tail section. To make the craft statically as well as dynamically stable, it
is very important to set the CG (center of gravity) of the craft ahead of its neutral point. This gives
the craft strength to continue its stable state after encountering any gust or turbulence. Using XFLR5
and SOLIDWORKS it was ensured that the CG was placed accordingly, we were able to analyse this
neutral point after multiple iterations.
Along the same lines, the lateral stability or the yaw of the UAV can be improved by giving it a
sweep. When the UAV encounters turbulence, the drag forces which are acting on the wings change
their moment arm. As a result of this the craft is able to regain its initial stability. In addition to this,
the sweep also increases the control surface areas as well as the moment arm from the CG.
The BWB has flat surfaces and sharp edges,reducing its Radar Cross-Section (RCS)and making it a
Stealth aircraft by virtue of its design enabling it to undertake surveillance missions without the fear
of getting spotted on radar. When radar signals hit the surface, they are reflected obliquely, away
from the radar antenna.
Once the CAD models were prepared, they were tested through CFD and static analysis so as to
confirm the designs ability to fly and withstand the environmental factors. The flight envelope was
also plotted to ascertain the vehicle's safety whilst flying.
The following table specifies the overall objectives of the team ahead of the development of the UAV
which were also extensively scrutinised during the final review process:
1 Innovation
2 Maximum payload possible
3 Good lift performance
4 Good structural design
5 Improved endurance and aircraft performance
8

10. 2.3 ) DESIGN SELECTION PROCESS
To establish the best results and prepare a holistic report the engineering process adopted was based
on Mohammad H. Sadraey’s methodology and adapted by the team, as shown in the following
diagram:
Fig. 2.3.1 Design Methodology in Aircraft Design Systems Engineering Approach
Through the establishment of multiple preliminary designs,study and analysis of the resources
gathered it was decided after multiple iterations, to proceed with the blended wing configuration set
at the particular values for angle of sweep,angle of attack ,etc. these arguments put forth the most
aerodynamically efficient and stable model compared to conventional configurations,setting up the
detailed design. Key details such as the effect of wetted area reduction due to absence of
empennage, reduction in the skin friction drag were evaluated. At the same time, the wing loading as
well as the spanwise lift distribution of the all-lifting design, interference drag reduction because of
the smooth blended wing centerbody intersection were all fitted around the preliminary design to
gradually reach the final detailed design.
3.0) PRELIMINARY AND DETAILED DESIGN
3.1) WING SELECTION
Airfoil Structure:
The most important components for flying an aircraft are its wings and there are various different
types namely : Rectangular, elliptical, tapered, delta, trapezoidal, ogive, swept-back, forward-swept,
and variable sweep.
As stated by NASA: “The BWB airframe merges efficient high-lift wings with a wide airfoil-shaped
body, allowing the entire aircraft to generate lift and minimize drag. This shape helps to increase
fuel economy and creates larger payload (cargo or passenger) areas in the central body portion of
the aircraft.”
The basic idea was to come up with a design in which the fuselage would facilitate the lift while also
maintaining the nose down moment, hence reflex airfoils were considered.
9

11. For the fuselage,Various airfoils like MH93, NACA25112, MH78, S5010, etc. were considered and
analysed using xflr5 with set parameters like a ceiling altitude of 6000m, Mach number of 0.063 and
Reynolds number 1,23,543, Density and Viscosity. MH 93 was best suited for our design due to its
high Cl value without compromising due to an increased thickness. It had a max thickness of
15.99% to its chord length and a Max Cl of 1.13 at 12⁰ angle of attack.
Fig. 3.1.1 - Comparison of various Fuselage Airfoils
Fig. 3.1.2 - Chosen Fuselage Airfoil- MH93
For the wings, various airfoils of lesser thickness were compared to the selected MH 93 so that the
transition between the fuselage and wing airfoils wouldn't lead to any hindrance in airflow. Airfoils
like fx60126_3, MH42, LA2573A, etc. were compared. Upon conducting various iterations on an
xflr5 based prototype, LA2573A proved to be the perfect pair for MH93.
Fig. - 3.1.3 Comparison of various Wing Airfoils
10

12. Fig. 3.1.4 - Chosen Wing Airfoil- LA2573A
The Cl vs alpha graphs of the selected airfoils:
Fig. 3.1.5 - Cl vs alpha
Wing structure:
Using the selected airfoils, several prototypes were created within xflr5(Further mentioned in
Sec.[6.0]). It is known for conventional planes with empennage that the tail is the factor that
maintains the pitching stability.
In case of a Blended wing body, the absence of tail is compromised by positioning the Components
in the aircraft in such a way that the Centre of Gravity(CG) always lies nearer to the leading edge as
compared to the Neutral Point. This was ensured by studying the CM vs 𝛼 graphs until a satisfactory
condition was reached.
The value of Neutral Point was found with iterations in CG value in xflr5 until a line parallel to
X-axis in CM vs 𝛼 graph was observed. The Neutral Point (XNP) is at a distance of 0.958 m from
the Leading Edge.
Also, the value of CG was derived using SolidWorks, where all components, strength members and
skin materials were added along with their mass in the designed prototype.
The ‘fuselage’ contributes to the total lift hence is considered as a wing while reading the dimensions
and other parameters. UAV configuration according to best values:
11

13. Specifications Values
Wing Span 4 m
Wing Area 3.23 m2
Aspect Ratio 5
Root to Tip Sweep 30.43o
Mean Aero Chord
(MAC)
1.16 m
Mean Geo Chord 0.81 m
CG from leading
edge (XCG)
0.83 m
Fig. 3.1.6 - Blended Wing Structure (xflr5)
Internal Wing Structure:
Aside from the shape and configuration of the wing, the structure inside needs to be taken into
account as well when designing wings. The purpose of the wing structure is to increase the solid
mechanics of the wing and make the wing more rigid. The structure inside the wing will help to
distribute the load evenly, so the stresses will not peak in some areas.
Wing skin:
The forces that act on the wing act on the wing skin first. The wing skin is a crucial component of the
whole wing structure. Lightweight composite materials were considered so as to conform with our
objective of improved endurance whilst maintaining robustness. More mentioned in section [3.5].
3.2) FUSELAGE SIZING
Along with increased lift and better fuel efficiency, another major advantage of a blended wing is its
increased fuselage carrying volume. Our selected airfoil for the fuselage has a 2m chord length
varying to 0.7 metres at the sides (fuselage-wing root intersection). The below values were directly
obtained from CAD designing.
12

14. Max fuselage thickness (along the centre chord) = 0.319 m
Max fuselage thickness(wing root chords)= 0.112 m
Max carrying= 0.3 m3
Payload bay Capacity = (0.6✕0.32✕0.24)m3
= 0.046 m3
3.3) AIRCRAFT PERFORMANCE
Theoretical Development :
The drag polar for an aircraft with is found out using [11],
CD = CDo + kCL
2
(1)
Where CDo is the zero lift drag and viscous effects dominate these as well as shear and also the
approximated flow.
kCL
2
talks about the lift-dependent drag made up of vortex drag and viscous pressure.
This above equation can also be dimensionalized by multiplying q (dynamic pressure) and S
(reference area), giving rise to
D = qS(CDo + kCL
2
) (2)
Upon xflr5 analysis, the value of CDo was found to be 0.00010 from the Cd vs Alpha graph and the
value of CD was 0.000469043 and 0.00925 for angle of attack of 1° (for cruising) and 6° (for
climbing) respectively. The CD values calculated using the formula were compared with the XFLR5
values. A minor error of 6% was observed using equation 1.
Drag value was found to be 0.27552 N and 12.2625 N for angle of attack of 1° (for cruising) and 6°
(for climbing) respectively.
Equations for an aircraft in steady and level flight,
Preq = D x U (3)
Where Preq is the power required to overcome the drag D of our UAV at the given velocity U.
Now assuming that the steady level flight requires the equality of the aircraft’s lift L and weight W.
Substituting
L = W = 0.5⍴U2
SCL , the above equation can be converted into:
Preq = 0.5⍴U3
SCDo + (2W2
k / ⍴US)
Preq for the aircraft in steady and level flight was estimated to be 554.23W.
Equations of motion in a steady climb [8]:
The center of gravity of the airplane moves at a constant velocity along a straight line inclined to the
horizontal at an angle Ɣ. Since the flight is steady, the acceleration is zero and the equations of
motion in climb can be obtained by resolving the forces along and perpendicular to the flight path
and equating their sum to zero.
T - D - WsinƔ = 0 (4)
13

15. L - WcosƔ = 0 (5)
It was seen above, that the Clmax value of the primary fuselage was 12o
. However, due to lack of an
empennage, the wing and fuselage is alone responsible for its pitching stability. This comes at an
expense of only being able to operate the blended wing aircraft at an angle far less than the Clmax
value. Ideally this value lies between 5o
to 7o
. Keeping in mind the change of 𝛼 with a gust of wind,
an 𝛼 of 6o
was chosen.
The thrust required for ascend was estimated using equation (4) and it was found out to be 63.53N.
The vertical component of the flight velocity (VC) is given by,
VC = VsinƔ = (T - D / W)V (6)
The vertical component of the velocity (VC ) is called rate of climb and also denoted by R/C. It is
also the rate of change of height and denoted by (dh/dt). Hence,
VC = R/C = dh/dt = VsinƔ = (T - D / W)V (7)
Using the thrust and drag values estimated above, the rate of climb VC was estimated to be 3.39 m/s
or 203.85 m/min.
Rearranging equation (4), we get,
T = D + WsinƔ (8)
Multiplying equation (8) by velocity V on both sides we get,
TV = DV + WVsinƔ (9)
TV = DV + WVC (since VC = VsinƔ) (10)
Where TV is the Power Available, DV is the Power dissipated in overcoming drag, and WVC is the
rate of increase in Potential Energy.
Thus, the power available was estimated to be 2065.03W, Power dissipated in overcoming drag was
estimated to be 368.56W and the rate of increase in Potential Energy was estimated to be 1664.76 W
3.4) PROPULSION SELECTION
When considering the various options available for propulsion and energy systems, an emphasis was
laid on sustainable, clean energy.The impact of the fuel on the environment defines its sustainability.
Our target of maximum efficiency and endurance was met by the use of Hydrogen fuel cells coupled
with electric engines due to their light weight, easy availability and low acoustic footprint. Fuel cells
have a higher specific energy/Kg and are compact. This is a highly desired property when
considering energy systems for UAVs due to the space constraint.
Hydrogen Fuel Cell:It uses a polymer electrolyte membrane fuel cell (PEMFC) that vaporizes
oxygen and hydrogen. Even though they have lower specific power output when put up against
batteries on one hand, they have higher specific energy on the other. This contributes heavily towards
a craft’s endurance as it gives an efficiency of about 60-70%. Intelligent Energy’s IE-Soar 2.4kW
fuel cell was chosen for the energy system.
This cell functions utilising a constant intake of hydrogen from the tank and combines it with the
oxygen available in the atmosphere. The by-product is water which escapes into the atmosphere in
the form of vapour.
Another advantage of using this cell is the fact that it does not need cooling and is capable of
operating at subzero temperatures.
Hybridisation: Like the name suggests, the idea is to create a propulsion system by combining
hydrogen with the conventional batteries (Li-ion) to make the best use of both their advantages. The
basic idea is that the fuel cells will power the UAV when in cruise condition and in times when high
power consumption is required, the Li-ion will fulfill the additional power needs.
14

16. Electric Engine: To stay true the objective of sustainable transportation, our team opted for an
electrical propulsion system. The Turnigy RotoMax 1.60 Brushless Outrunner Motor was selected
due to its exceptional qualities.
Specifications:
➔ RPM: 231kV
➔ Max current: 80A
➔ Watts: 2960kV
The motor has the following dimensions:
➔ Shaft (A):10mm
➔ Length (B):58mm
➔ Total Length (E):128mm
Fig. 3.4.1- Motor dimensions
Propeller: While selecting the propellers we had to carefully consider the number as well as the
shape of the blades that suit our design.Rigidity of the material is a very important factor taken into
consideration mainly because it increases the propeller's efficiency as well as reduces its acoustic
footprint.Lightweight Carbon 3010 material was chosen for the propeller.
For the propellor, the non-dimensional coefficients can be defined accordingly :
Propellor diameter= 21in
Pitch = 13 in
RPM (for ascending) =6930
The thrust was calculated using the formula :
(11)
Fig. 3.4.2 - Thrust vs. Aircraft Airspeed
The maximum thrust obtained (Static Thrust) for the given Propeller parameters, during take-off is
13.92903 kg. This gives the Thrust to Weight (T/W) ratio of 0.279 during take-off, while the
propulsion system is capable of producing the maximum static thrust of 21.18761 kg, which gives
the maximum T/W ratio equal to 0.424. The propeller parameters were carefully iterated to get the
15

17. desired value of T/W ratio.The thrust vs aircraft speed graph shows the linear variation of thrust with
respect to speed upto 150km/h.
3.5) MATERIAL SELECTION
Materials for the UAV were chosen such that the most advanced composite materials available for
aerospace applications were used. Keeping in mind that a light weight aircraft translates into
increased flight time, materials with low density but very high strength have been chosen. The
materials chosen are also easily manufacturable and can employ latest manufacturing processes in
additive manufacturing such as pultrusion,SLS,Injection and laser molding can be other methods.
Materials used are discussed below:
Pultruded carbon rods and Balsa Wood(Ribs & Spars): Pultruded carbon fiber rods offer
extremely high strength and stiffness because they contain 95% unidirectional carbon fiber running
longitudinally. Carbon Fiber Solid Rods are made from continuous carbon fibers in a resin matrix of
either vinyl-ester or epoxy. Balsa wood has also been used in certain shorter members to further
reduce the weight due to its low density of 0.04 g/cm3
.
Properties:
➔ Density: 1.8 g/cm3
.
➔ Tensile Strength: 2.34 GPa
➔ Glass Transition Temperature: 100° C
➔ Fiber Volume: 62%
➔ Matrix Material Bisphenol Epoxy Vinyl Ester
Fig. 3.5.1 Pultruded Carbon fibre
Pultrusion is the method of manufacturing continuous lengths of a fiber-reinforced section by pulling
continuous strands of carbon fiber through resin and a former before curing the resin all in one
process.
Features:
➔ Very Strong and Light
➔ Very high Rigidity
➔ Very low coefficient of expansion
PEEK (Skin): PEEK (polyetheretherketone) is a high-performance engineering plastic with
outstanding resistance to harsh chemicals, and excellent mechanical strength and dimensional
stability. It offers hydrolysis resistance to steam, water, and sea water.PEEK polymers are obtained
by step-growth polymerization by the dialkylation of bisphenolate salts. It has the ability to maintain
stiffness at high temperatures and is suitable for continuous use at temperatures up to 170°C. Due to
its exceptional dielectric properties it has good radar transmission and transmittance, further
improving our objective of surveillance capabilities.
Properties -
➔ Density: 1.2 g/cm3
.
➔ Young’s modulus: 3.9 GPa
➔ Elongation at break: 150%
➔ Thermal conductivity: 0.25W/mK
16

18. Features -
➔ Superior tensile strength
➔ High operating temperature
➔ Excellent creep resistance
➔ Low friction
➔ Exception dielectric properties
3.6) DETAILED WEIGHT BREAKDOWN
A summary of the weight breakdown is mentioned below:
No. Component Weight (kg) No. Component Weight (kg)
1 Hydrogen fuel cell 4.4 9 Pitot tube 0.041
2 Hydrogen tank 4.8 10 Valves for hydrogen tank 0.05
3 Camera 1.2 11 Wires 0.10
4 Servo 0.06*6=0.36 12 Receiver 0.03
5 Motor 0.84 13 Analog Video transmitter 0.072
6 ESC 0.165 14 Ribs and Spars 9.869
7 GPS 0.13 15 Outer skin 15.612
8 Parachute 1.3 16 Propeller 0.30
Total 39.269
3.7) FLIGHT ENVELOPE
The V – n diagram depicts the aircraft limit load factor as a function of airspeed. One of the primary
reasons for this diagram is that the maximum load factor; that is inferred from this plot is a reference
number in aircraft structural design.
Fig. 3.7.1 - Vn Diagram
17

19. From the above graph we can infer that: Results:
For, Mass=40 kg
Wing area= 1.42 mᒾ Min load factor(-)= -4
Cruising speed= 20 m/s Dive speed= 30 m/s
Max load factor(+)= 10
The general condition, 𝒏𝒏𝒆𝒈 ≥ 𝟎. 𝟒 𝒏𝒑𝒐𝒔 for normal and utility aircraft is satisfied by our designed
UAV.
At the speed above dive speed, destructive phenomena such as flutter, aileron reversal, and wing
divergence can occur that lead to structural damage, failure or disintegration. This speed limit is a
red-line speed for the aircraft and must never be exceeded.
Load Gust Diagram:
The atmosphere is a dynamic system that puts forth a variety of phenomena. Some of these include
turbulence, gust, wind shear, jet stream, mountain wave and thermal flow. We look at only gust, as it
is not predictable, but occurs during most high altitude flights. When an aircraft experiences a gust,
the immediate effect is an increase or decrease in the angle of attack. Below is plotted the combined
gust envelope for the V-7:
Fig. 3.7.2 - Limit Combined Envelope
Considering:
MAC= 1.16 Sweep angle (deg)= 30.43
AR= 5 Cruise Mach Number= 0.05882
The blue dashed line represents the limit of vertical gust load during cruise and the red dashed line
indicates the limit of vertical gust at dive speed.
18

20. 3.8) LANDING GEAR
The typical configuration of a landing gear are tricycle, bicycle, tailwheel or unconventional gear and
its complexity is highly affected by the need or not of a retraction system. Through careful
consideration the slidable skids were chosen for take off and Parachute system for landing. The skid
is attached to the under belly of the fuselage and is wide enough for opening the payload bay and can
also be mounted on a Catapult Rail, this results in a lighter and simpler solution.
The landing gear is sized such that a reasonable clearance is given between the Rails and all other
parts of the aircraft in its compressed position.
For landing we describe the development of a Dual Parachute system with fall detection and quick
release based on an accelerometer-gyroscope installed in the aircraft that will help in Safe landing.
According to the below formula:
(12)
The Radius of Parachute deployed, r = 1.88m
The optimum descent speed achieved will be 7.5m/s preventing any damage to the aircraft.
Fig. 3.8.1 - Landing Gear Fig. 3.8.2 - Dual Parachute System
3.9) STABILITY AND CONTROLS
Stability:
The Static Margin is a percentage value which tells us about the aircraft’s static stability. For
Unmanned Aircrafts this value should ideally lie between 5% to 40% of MAC.
Static Margin SM, is numerically calculated by:
(13)
𝑆𝑀 = −
(𝑋𝐶𝐺 − 𝑋𝐴𝐶)
𝑀𝐴𝐶
Here, XAC is the distance of the Aerodynamic Centre from the leading edge.
It is to be noted that the Aerodynamic Centre(XAC) and the Neutral Point(XNP) are the same point for
a tailless design (Source: Wikipedia). Hence,
= 0.1103 of MAC (ie. SM = 0.1279 m)
𝑆𝑀 = −
(0.83 − 0.958)
1.16
Static margin for the V-7 was analysed using xflr5(Section [6.0]).
Numerical Value of SM = 11.03% of MAC
Analysed xflr value of SM = 10.86% of MAC
Percentage error = 1.541%
19

21. Controls:
Control surfaces of tailless aircraft are an interesting part of design due to the absence of
conventional tail. The control surfaces for pitch and yaw control for these aircraft are totally different
from conventional aircraft. The absence of tail rudder could be substituted by other control surfaces
such as split drag flaps, inboard and outboard ailerons, winglets, rudders or Thrust Vectoring. The
problem of absence of the elevator can be solved by substituting it with elevons. The elevons are
aircraft control surfaces that serve the functions of both the elevators and the ailerons. They are
installed on each side of the aircraft at the trailing edge of the wing. If the elevons on both sides are
moved in the same direction they will cause a pitching moment. If moved in the opposite direction
(one up, one down) they will cause a rolling moment.
Elevon Sizing: Chordelevon = 30% * Chordwing tip region = 0.12 m
Lengthelevon = 0.589m
Airfoilelevon = Flat Plate
For yaw control there are two possible designs for BWB aircraft,first one is placing the vertical tail at
the tips of wings rather than the aft of tail like conventional aircraft. Second one is by using split drag
flaps (rudders) as yaw control surfaces.
Split drag flaps consist of upper and lower flaps that will be deflected oppositely. This device works
as a drag producer in order to generate yawing moments. Deflection of the flaps on one side of the
wing produces asymmetric drag force and, as consequences, a yawing moment is produced that
rotates the nose of the aircraft toward the deflected flaps. To improve the effectiveness of split flaps
they are located near to the wing tips. This provides a long moment arm and will give greater yawing
moment for the BWB aircraft.
However the effectiveness of the control surface is directly proportional to the distance of rudders
from the CG and also the size of the control surface. For a blended wing, the wing tips lie nearer to
the CG, hence for the rudders to be effective, it would take up huge surfaces hence increased drag
and weight.
Therefore for our model, we chose the split flap concept.
Flaps also have a major effect on lift even at a low angle of attack. The dimensions of the flaps
designed were: Chordflap = 30% * Chordwing root region = 0.22 m
Lengthflap = 0.577 m
Airfoilflap = Flat Plate
ESC
The following ESC was selected for use, HOBBYWING Platinum HV 120A V4
Specifications:
➔ Input Voltage: 6S-14S Lipo
➔ Cont./Peak Current:130A/160A
Control system:
Control system includes flight controllers, sensors and all the servos which control the control
surfaces of the V-7.
Flight controllers:Pixhawk 2.1
Servos: 2000 Series Dual Mode
Specification:
Gear Ratio: 300:1
Voltage Range: 4.8V - 7.4V
Stall Current (7.4V): 3,000mA
Max PWM Range: 500-2500μsec
Fig. 3.9.1 - Control System
20

22. GPS:The Here 2 GPS was selected as it has a Concurrent reception of upto 3 GNSS (GPS, Galileo,
GLONASS, BeiDou) coupled with industry leading 167 dBm navigation sensitivity, all whilst
maintaining high security.
Sensors include:Compass, Gyro, Accelerometer: ICM20948 and a barometer.
3.10) COMMUNICATIONS SYSTEM
The selected ground control system is the XLRS GCSD4 V2
Ground station Telemetry system.
Professional Portable Ground Control StationTelemetry with a long
range of 200Km.
It has an embedded PC running on windows 10, Wifi and
Bluetooth.Internal battery specifications with power back up of 8-12
hours mounted with biquad antenna of 12dBi (BQ89) that significantly
improves performance, safety and range in GCSD4-V2
Fig. 3.10.1- XLRS GCSD4
Telemetry:
The flowchart below explains the entire telemetry system and its connections:
Fig. 3.10.2 - Communication system
Radio control receiver:The RXLRS professional radio control receiver has long-range capabilities
up to 200 km,with Mavlink telemetry and radio modem transparent of 38.4kb to 100kb.
Video transmitter: XOSD3B OSD + Analog video transmitter with 8 video channels and inputs for
2 video cameras with PAL format.
Pitot tube: Air speed meter MS4525DO with pitot tube.
3.11) PAYLOAD
The V-7 has been designed to carry a maximum payload of 10 Kg.
Aircraft weight= 40Kg MOTW= 50Kg
Payload Bay Volume= 0.046 ㎥
21

23. 3.12 ) SCHRENK'S CURVE, SPAR DESIGN, RIB SIZING
Fig. 3.11.1 - Schrenk’s curve (local lift vs half span distribution) (XFLR5 generated curve)
The Schrenk’s Curve tells us about the variation in generated lift in the wing along the span. Through
the analytical method Local Lift vs Span graph was studied. Ideally, for conventional aircraft, the
curve is the mean curve between an Ellipse and a linear line of corresponding Lift values. Here, the
graph is studied from the tip towards the root. It can be observed that the curve is ideal, but at 0.575
m from the centre the lift suddenly increases. This indicates that the fuselage also contributes to the
major portion of the lift.
Lightweight materials were used and a minimalist approach was taken up by the team in developing
the ribs and spars to make the V-7 as light as possible.
Fig. 3.11.2 - Spar And Rib Placement
➔ Drawing added in Appendix
3.13 ) ENDURANCE CALCULATION
Average power consumption calculations : The power required for the UAV to ascend and descend
was calculated to be 2065.03W (Section 3.2). At a Rate of Climb of 203.85 m/min (Section 3.2), the
UAV will take 14.7167 min ≈ 15 min to reach an altitude of 3000m. After which, it can cruise for
210 min on a power saving mode, and consume 554.23W of power. The power that the electronics
(except motor and ESC) will consume will be ≈ 175W. Hence, the average power required for the
complete flight mission will be ≈ 970W..
Fuel Consumption calculations:The mass of hydrogen stored in a cylinder is proportional to the
pressure and the efficiency of the fuel cell varies with load.
22

24. Average Power Consumption by the UAV= 970W
Hydrogen Tank Capacity=210g
𝑓𝑢𝑒𝑙 𝑐𝑜𝑛𝑠𝑢𝑚𝑝𝑡𝑖𝑜𝑛(𝑔/ℎ) = 𝑃𝑜𝑤𝑒𝑟 (𝑊) ÷ (𝐸𝑛𝑒𝑟𝑔𝑦 𝑐𝑜𝑛𝑡𝑒𝑛𝑡 𝑜𝑓 𝐻𝑦𝑑𝑟𝑜𝑔𝑒𝑛 × 𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑐𝑦)
(14)
Energy content of Hydrogen= 33.3Wh/g
efficiency= 0.56
Therefore, the fuel consumption = 52g/h of Hydrogen
𝐸𝑛𝑑𝑢𝑟𝑎𝑛𝑐𝑒 = 𝑇𝑎𝑛𝑘 𝑐𝑎𝑝𝑎𝑐𝑖𝑡𝑦(𝑔) ÷ 𝑓𝑢𝑒𝑙 𝑐𝑜𝑛𝑠𝑢𝑚𝑝𝑡𝑖𝑜𝑛 𝑝𝑒𝑟 ℎ𝑜𝑢𝑟(𝑔/ℎ)
From the above calculations, an endurance of 4 hours or 240 minutes (varies with payload) can be
achieved, fulfilling the objective of improved endurance.
4.0) COMPUTER AIDED DESIGN DETAILS
The team used SOLIDWORKS and Fusion 360 for designing the V-7. Below attached are the
2D and 3D technical drawings, helping the reader visualise the UAV.
➔ All Dimensions are in m.
Fig. 4.1 - Frame Fig. 4.2 - Component placement
Fig. 4.3 - Front cruise view Fig. 4.4 - Read cruise view
23

25. Fig. 4.5 - Payload bay open Fig. 4.6 - Flap engaged
24

26. 5.0) COMPUTATIONALANALYSIS
The team used ANSYS Workbench to perform the V-7’s simulation studies (FEA and CFD)
CFD modeling is an inevitable part of multiphase flow investigation. It is used to describe the flow
characteristics and get insights into the flow pattern beforehand. First the model was imported into
the workbench with a mesh sensitivity of 50mm. Later, the time step and a maximum number of
iterations per time-step were selected to be 3 seconds and 100 iteration cycles, respectively.
The Velocity of air constrained at 35 m/s.
Coming to the post processed results from Static Pressure the maximum pressure is experienced by
the nose and the wing tips which is 360 Pascals.
In the turbulence graph, we can see that the effect of wind turbulence at the nose and tail of the plane
is low and maximum at the evelons with a maximum value of 20.7 m2
/s2
.
From the velocity graph we can see that there is minimal change in overall velocity of air around the
aircraft and is maximum at the upper edge of the nose i.e. 44.8 m/s. This justified that the plane
travels with uniform velocity around the surface and is safe.
Fig. 5.1 - Static pressure Fig, 5.2 - Turbulent pressure
Fig. 5.3 - Velocity
6.0) OPTIMIZED DESIGN
XFLR5
XFLR5 was used to initiate and optimise the V-7’s design. Several iterations were made until the
most desirable values were obtained. A thorough review of each component through a series of loops
was conducted until all the parameters matched ideal conditions.
The models were tested under fixed parameters, such as,
Velocity = 20 ms-1
, altitude = 6000m , temperature = - 23 o
C
and a varying range of angle of attack(𝛼).
The value of CG was derived using SolidWorks, where all components, strength members and skin
materials were added along with their mass.
The graphs obtained for our final design are given below:
25

27. Fig. 6.1 - CL vs 𝛼 Fig. 6.2 - CL/CD vs 𝛼
Fig. 6.3 - CM vs 𝛼 Fig. 6.4 - CD vs 𝛼
It can be seen from the graphs that our model
satisfies all the required parameters.
In CL vs 𝛼 a steep positive slope is achieved. The
values of CL Increase largely without a change in
the slope as a result of the flaps(Section[3.9]). This
condition is ideal for Climbing,
The CM vs 𝛼 graph gives us the “trim angle”. A
lower value is desired in case of blended wing
so as to avoid pitching instability.
The CL/CD vs 𝛼 graph tells us about the aircraft’s
efficiency at different angle of attack.
A peak value of 160 achieved at just 1o
. Hence, it
it was decided to cruise our aircraft at 1o
to
maximize fuel efficiency. Fig. 6.5 - CM vs CL graph
The CD vs 𝛼 shows us the very low value of drag for lower angle of attacks in which our aircraft will
typically operate.
26

28. The analytical Static Margin k,
SM = - (slope of CM vs CL graph) (15)
∴ SM = - = 0.1086 = 0.1086
(−0.02)
(0.184)
An image of detailed xflr5 analysis in 3D veiw is provided.
The yellow lines signify drag, green lines the lift, pink the downwash while the purple lines show us
the stream.
Fig. 6.6 - XFLR5 analysis of the final model
7.0) COMMERCIALLY OFF THE SHELF PARTS
COTS were chosen based on factors such as affordability, compatibility, ease of installation, sizes,
tolerances,safety and most importantly quality. Below is a list of the same:
Components Selections Components Selections
Motors Turnigy-Rotomax 100cc
brushless motor
Basic Camera Octopus Epsilon 135 Day
payload
Electronic Speed Controller HOBBYWING Platinum HV
200A V4.1
Surveillance Camera Octopus Epsilon 180
Multi-mission EO/IR
surveillance system
Flight Controller PixHawk 2 Orange cube Hydrogen Cell Intelligent Energy
GPS Module Here2 Telemetry Receiver system RXLRS
Pitot Tube Pitot-Static Tube for UAVs Analog Video Transmitter XOSD 3B
Servo Spectrum A6320 Torque High
speed Metal BL HV servo
-- --
Table 7.1 - Commercially Off The Shelf Parts
27

29. 8.0) PRACTICALAPPLICATION AND FEASIBILITY
The V-7 was designed and developed to undertake a variety of missions from any given
location,under any conditions. Today, India imports drones and UAVs from foreign manufacturers to
meet its domestic needs. It becomes very important to develop indegenous UAVs to save on import
costs and to establish oneself as a global player in developing advanced unmanned systems. Our
team designed a revolutionary design in the blended wing, something that is uncommon in the Indian
aero domain. The V-7 uses the latest hydrogen cell technology through the employment of Intelligent
Energy’s 2.4kw fuel cell and the highly advanced RXLRS telemetry system. Special emphasis was
laid on selecting the most sustainable materials which also reduce weight whilst maintaining high
strength.
The V-7 is designed to undertake medical deliveries within a range of 150Km and can carry large
medical payloads in its insulated payload bay. It is also equipped with a standard Epsilon camera
which can relay video to the ground station during flight. We take pride in the conception of a
modular camera system which allows the V-7 to be converted into a dedicated surveillance UAV just
by the addition of the modular Octopus Epsilon 180 camera in the payload bay. This functionality
makes the V-7 equipped to take on a variety of missions with ease.
The UAV uses a easily swappable or refillable hydrogen tank coupled to the fuel cell, which uses
ambient atmospheric oxygen and gives out water vapour as a by-product, making the V-7 non
polluting.
9.0) INNOVATION
Whilst developing the V-7, innovation was taken as a key objective and subsequently the following
systems were conceptualised:
● The conceptualisation of a BWB for cargo delivery in itself is an innovation as the wider
fuselage of the BWB allows for larger payloads to be carried.
● The V-7 can be used as both a surveillance as well as a cargo vehicle,by just the addition of a
camera.The flight controller supports 2 cameras, a fixed camera and a detachable modular
camera. The modular Epsilon 180 camera is easily mounted in the payload bay when the need
be.
● Application of pultruded carbon fibre for the development of ribs and spars due to its
exceptional strength to weight ratio.
● Innovative landing system which combines simple aircraft skids with a parachute for safe and
energy efficient landing.
● Fuel Cell Hybridisation with Li-ion batteries to improve peak power output.
28

30. REFERENCES
1.Unmanned Aircraft Systems: UAVS Design, Development and Deployment by Reg Austin.
2.Small Unmanned Fixed-wing Aircraft Design: A Practical Approach by Andrew J. Keane,
András Sóbester, James P. Scanlan.
3.Modelling and Control for a Blended Wing Body Aircraft by Martin Kozek and Alexander
Schirrer.
4.Aircraft Design: A Systems Engineering Approach by Mohammad H. Sadraey.
5.Mechanics of Flight by A.C Kermode.
6.Aircraft design: A conceptual Approach by Daniel P. Raymer.
7.Theory, Design and applications of UAV by A.R. Jha.
8.Flight Dynamics-I by Prof. E.G. Tulapurkara, NPTEL.
9.Design of Blended Wing Body Aircraft thesis paper by Randhir Brar, San Jośe State
University,USA.
10. Propulsion System for a small Unmanned Aerial Vehicle by Oscar Andersson and
Dennis Wilkman, KTH Royal Institute of Technology,Sweden.
11. Range and Endurance Estimates for Battery Powered Aircraft by Lance W. Traub.
12. Design, Manufacturing and Flight Testing of an Experimental Flying Wing UAV by
Pei-Hsiang Chung, Der-Ming Ma and Jaw-Kuen Shiau, Tamkang University,Taiwan.
13. Aerodynamic design Optimization Studies of a Blended Wing Body Aircraft by
Zhoujie Lyu and Joaquim R.R.A. Martins, University of Michigan,Ann Arbor,Michigan.
14. Stability Study and Flight Study Simulation of a Blended Wing Body UAV by Thomas
Dimopoulos,Pericles Panagiotou and Kyros Yakinthos, Aristotle University of Thessaloniki,
Greece.
15. Robert C. Nelson, Aircraft Stability and Automatic Control, McGraw-Hill, Second
edition, 1998.
29

31. APPENDIX
Additional data, drawings and pictures for components used onboard the V-7.
Ribs and Spars technical Drawing:
➔ All dimensions are in mm.
30

32. I. Motor:The Turnigy RotoMax 1.60 Brushless Outrunner
➔ RPM: 231kv
➔ Max current: 80A
➔ Watts: 2960w
➔ No load current: 37V/1.47A
➔ Internal resistance: 0.028 ohm
➔ Pole Count: 24
➔ Weight: 849g
➔ Diameter of shaft: 10mm
II. Electronic speed control: HOBBYWING
Platinum HV 120A V4
III. Flight controller: PixHawk 2 Orange
cube:32bit STM32F427 Cortex-M4F core with
FPU.
• 168 MHz / 252 MIPS
• 256 KB RAM
• 2 MB Flash (fully accessible)
• 32 bit STM32F103 failsafe co-processor
• 14 PWM / Servo outputs
• Abundant connectivity options for additional peripherals (UART, 12C, CAN).
• Integrated backup system for in-flight recovery and manual override with dedicated processor and
stand-alone power supply (fixed-wing use).
• Backup system integrates mixing, providing consistent autopilot and manual override mixing
modes.
• Redundant power supply inputs and automatic failover
• External safety switch
• Multicolor LED main visual indicator
• High-power, multi-tone piezo audio indicator
• microSD card for high-rate logging over extended periods of time.
IV. GPS:
Interfaces
• 5x UART (serial ports), one high-power capable, 2x with HW flow control
• 2x CAN (one with internal 3.3V transceiver, one on expansion connector)
• Spektrum DSM/ DSM2 / DSM-X® Satellite compatible input
31

33. • Futaba S.BUS® compatible input and output
• PPM sum signal input
• RSSI (PWM or voltage) input
• 120
• SPI
• 3.3v ADC input
• Internal microUSB port and external microUSB port extension
Power System and Protection
• Ideal diode controller with automatic failover
• Servo rail high-power (max. 10V) and high-current (10A+) ready
• All peripheral outputs over-current protected, all inputs ESD protected
Voltage Ratings
Pixhawk can be triple-redundant on the power supply if three power sources are supplied. The three
rails are: Power module input, servo rail input, USB input.
Normal Operation Maximum Ratings
Under these conditions all power sources will be used in this order to power the system
• Power module input (4.8V to 5.4V)
• Servo rail input (4.8V to 5.4V)
• USB power input (4.8V to 5.4V)
GPS System Used
32

34. V. Communications System:The block diagram explains the entire ground to air connection system
for the V-7 based on the selected telemetry systems;
Communication System
Landing system:
YANGDA Saver drone parachute system is specially designed to prevent multirotor and fixed-wing
planes from crashing in the air when they have problems like low voltage, mechanical issues,etc.
The Saver UAV parachute system can trigger the parachute ejection in just 0.1 seconds to save the
UAV.
Two methods to trigger the parachute ejection:
Method one: the Saver UAV parachute system has a built-in sensor and electrical core to detect the
drone attitude 100 times per second, and can eject the parachute out in the shortest time once the
drone is detected out of control.
Method two: trigger the ejection through the RC controller, and the propeller rotation will stop
accordingly.
Advanced ejection technology:
The first generation ejection technology is using mechanical springs, which can only support a max
weight of 20Kg and the second ejection technology is using compressed carbon dioxide, which can
support a max weight of 30Kg.
The Saver drone parachute system uses third-generation ejection technology: low-temperature
propellants, which is triggered by the high-voltage voltaic arc. Its weight is less than 10g, but
supports a max weight of 100kg.
33

35. The Ejection barrel can be reused with new propellants, the ejection barrel can be reused without any
limit.
Specifications:
Dimension: Φ105mm / H165 mm
Installation size: 116mm*116mm*M4*4
Connector: 5PIN waterproof connector
Voltage:5V
Payload: ≤ 30KG
Parachute diameter: 3.5m
Battery: 380mAh
Endurance: 8H
Trigger method: Attitude / Zero gravity / PWM / Serial Port
Attitude detection range: ±90°
Trigger angle: ±80°
Weight: 650g
Zero gravity trigger:0.5g / 1.6s
Data communication: Two-way
Power on wake-up: Yes
Power off time:10s
Two canisters each capable of carrying 30 kgs are used.
Parachute Canister
34

36. Stealth Technology in Blended Wing Body UAV:
Most conventional aircraft have a rounded shape and this makes them aerodynamic, but it also
creates a very efficient and effective radar reflector. The round shape means that the radar signal has
a larger area to hit the plane, and at least some of the signal gets reflected back.
Stealth can be achieved by two methods:
➔ Aircraft shape
➔ Absorbent materials
The V-7 has a BWB configuration which means the flat surfaces coupled by its sharp edges reflect
the radar signal away from the antenna. This gives it a radar signature smaller than a bird, making it
virtually impossible to trace. The use of an advanced composite material like PEEK which has
excellent dielectric properties further deflects radar signals. The two primary methods of achieving
stealth are confirmed by the V-7, giving it exceptional abilities to run surveillance missions within
and outside Indian borders. The payload is carried inside the body and the engine is electric, hence
reducing its thermal and acoustic footprint.
Working of Stealth Technology
35