2. 2
1. EXECUTIVE SUMMARY
This report details the conceptual and preliminary design of the High Voltage long range cargo delivery
unmanned aerial system for the Airbus Cargo Drone Challenge. This design combines simplicity and
performance in a robust package capable of the following four tasks:
1. One hundred kilometer range with three kilograms of payload
2. Sixty kilometer range with five kilograms of payload
3. Vertical takeoff and landing capability
4. Fast turnaround time between missions
1.1 Concept Motivation
High Voltage is optimized for ease of use in time critical emergency scenarios to provide medical support.
The global use of unmanned aerial systems (UASs) is ever growing, yet within the commercial and civil-
use sectors, the full potential of UASs has not been realized. The perception such systems are restricted
to only governmental or military actors is due to the inability for many modern systems to operate in complex
environments. It is important for stakeholders to have a low-cost, flexible platform capable of completing a
variety of mission with minimal risk. High Voltage seeks to overcome these challenges by designing a
vehicle with a combination of high performance, redundancy, and safety features.
1.2 Design Elements
The design is a podded fuselage with twin tailbooms extending to an inverted V-tail. This enables maximum
structural efficiency by combining load paths as much as possible. Design features include:
Redundancy from eight VTOL props and parachute recovery system
Power management system including single breaker for arming and disarming the system
Single battery compartment for faster, easier battery swapping, minimizing turnaround time
High mounted forward flight prop for minimal interference
Extensive use of composites and SLS 3D printed parts for easier manufacturing and weather
resistance
1.3 Team Introduction
The design team is composed primarily of engineering students from the Georgia Institute of Technology,
plus a faculty advisor from the research faculty. The students are part of a club that participates in design
competitions put out by professional organizations like the American Institute of Aeronautics and
Astronautics (AIAA) and the Society of Automotive Engineers (SAE). The members of the team have years
of experience with these competitions; the past two years the Georgia Tech teams have had six entries into
these competitions and won four 1st
places, one 2nd
place and one 3rd
place overall. This experience helps
guide the design and analysis process used for High Voltage.
3. 3
2. CONCEPTUAL AND PRELIMINARY DESIGN
The High Voltage design team was provided with a list of mission requirements and design features by the
organizers of the Airbus Cargo Drone Challenge. These requirements are listed below:
- The design shall be capable of vertical takeoff and landing (VTOL)
- The aircraft shall include at least one fixed wing for forward flight
- The aircraft will operate in two flight modes, a hover mode for takeoff and landing, and a forward
flight mode for efficient travel
- Maximum takeoff mass (MTOM) shall be below 25 kg
- The maximum wing span shall be below 5 meters and the maximum aircraft length shall be below
4 meters
- The aircraft shall be modular for the ease of transportation. The maximum length of the individual
parts shall not be longer than 2 meters
- The aircraft shall have a 60 km or greater range for a 5 kg payload, and shall have a 100 km or
greater range for a 3 kg payload
- A single payload bay shall be located near the aircraft center of gravity
- The cargo payload bay should be able to accommodate small payloads with a custom fixture
- The aircraft design shall avoid the use of tilting wings, tilting motors / rotors, or variable pitch
propellers.
- Minimum payload bay dimension shall be 450 x 350 x 200 mm
- The payload bay shall be located and accessible from the lower side of the aircraft and must be
interchangeable with payload bay of same size and same interface
- The cruise speed in fixed wing mode shall be at least 80 km/h
- The maximum speed shall not exceed 194 km/h
- The aircraft shall use at least 4 direct drive lift rotors/propeller but not more than 10 direct drive lift
rotors/propellers
- For energy storage, off the shelf rechargeable batteries shall be used
- Aircraft shall have the space and power for additional items outlined in the Ignition Kit
- The vehicle shall be capable of flying in all configurations while experiencing up to 10 m/s head
and cross wind
- The aircraft must be capable of 20 minute turnaround between max-distance missions
- Aircraft shall be transportable by two average size people
- Aircraft must be able to be disassembled and transportable in a variety of shipping scenarios,
including van, vessel, aircraft, truck, and ISO container
4. 4
2.1 Overview of Existing Designs
There are many existing systems that have similar mission and design requirements to those provided by
the organizers. These systems were examined to aid in the brainstorming phase of the configuration
selection. One such system is the Arcturus Jump 15, shown in Figure 1 below.
Figure 1: Arcturus Jump 15
The Arcturus Jump 15 uses a conventional wing-tail placement with a T-tail and a tractor propeller intended
for long endurance surveying and surveillance. For VTOL, the Jump 15 has four direct drive vertical lift
rotors mounted to underwing pylons arranged parallel to the freestream. When in forward flight mode, the
VTOL propellers are aligned parallel to the pylons in order to minimize drag.
An alternative configuration is exemplified by the XCraft X Plusone, shown below.
Figure 2: XCraft X Plusone
5. 5
The X Plusone is a tail-sitter that takes off vertically in a traditional quadrotor configuration, then transitions
into a flying wing powered by all four motors. It is intended for hobbyist film work and achieves a top speed
of approximately 60 mph. The X Plusone is also capable of extended hover on station.
In addition to the Jump 15 and X Plusone, a variety of configurations exist which do not meet the design
requirements given by the organizers but have similar mission profiles. For example, the Aerovel Flexrotor
uses a single, two-meter diameter blade with variable pitch and collective to transition from a tail-sitter
configuration into conventional forward flight. Another example is the Krossblade Sky Prowler, which uses
fold-out rotors for VTOL that then retract for forward flight while a pusher propeller accelerates the vehicle
laterally.
2.2 Creation of Figures of Merit
The design requirements explicitly describe mission requirements with measurable goals. In addition to
explicit requirements, there are a number of implicit requirements such as reducing the logistical footprint
of the vehicle to enable widespread UAS adoption. The final vehicle must balance these requirements. It is
necessary to subdivide competing requirements into figures of merit (FOMs) to select a configuration that
best satisfies both explicit and implicit requirements. The competition guidelines were analyzed to create
figures of merit as described in the following section.
Operational Flexibility: A Commercial UAS must be capable of functioning in a variety of
environments ranging from rural fields to urban rooftops to complete a variety of potential missions.
Additionally, a UAS must be able to successfully complete its mission in a multitude of conditions. Some
conditions may not have been anticipated during initial development. High operational flexibility is therefore
characterized by robustness, interchangeable parts and modularity. Vertical takeoff and landing capability
is a must, in order to eliminate the need for large, specialized facilities.
Ease of Maintenance: Maintenance costs represent one of the highest cost sectors in commercial
industries where a fleet of vehicles is used to deliver a service. A UAS must be able to minimize
maintenance costs while still achieving high operational flexibility and high performance. Ease of
maintenance is characterized by simplicity in design. The number of moving parts, as well as the absolute
number of parts, should be minimized. The system should be designed to use parts that are either available
off the shelf or fabricated with basic machining equipment.
Safety and Reliability: In all designs, safety and reliability are of paramount importance. In all
commercial applications, the ability for a platform to perform its mission again and again enables
stakeholders to have confidence in vehicle operations, minimizing investment risk. Furthermore, the safety
of human lives is of utmost importance, and the ability for a design to recover from a subsystem failure, or
at least fail in a manner that does not endanger others, is critical to this over-riding objective. Designs with
high safety and reliability are characterized by redundancy in critical subsystems, with provisions such as
emergency flight termination parachutes in order to handle catastrophic system failure.
6. 6
Turnaround Rate: Successful designs must be low cost, while still providing the necessary level of
service. The value of a platform is maximized when time spent idle on the ground is minimized. The speed
at which a vehicle can be prepared for its next mission is the turnaround rate. This turnaround rate is
constrained by the amount of service required on before the vehicle can begin its next mission. Designs
which maximize turnover rate are characterized by mechanical simplicity and ease of access to critical
components. This is especially relevant for the battery compartment and payload bay, which must be
serviced at the end of every mission.
Ease of Manufacturing: Lower cost is always important when bringing a product to market. For
UASs, it is especially critical that initial capital investment is low in order to minimize stakeholder risk. While
there are many factors that influence cost during the design process, manufacturing influenced design is
the factor which drives cost most directly during concept evaluation and selection. Designs which
maximize ease of manufacturing include a feedback loop between the manufacturing phase of product
development and both the conceptual and preliminary design stages.
Before it is possible to evaluate configurations, it is necessary to arrange these FOMs by order of priority.
This is presented below in Table 1.
Table 1: Figure of Merit Weighting Factors
Figure of Merit 1 2 3 4 5 6
Safety and Reliability 6
Ease of Maintenance 5
Turnaround Rate 4
Operational Flexibility 3
Ease of Manufacturing 3
Safety and Reliability is the highest priority, with a rating of six, since human lives are the highest priority.
Ease of Maintenance and Turnaround Rate are next in priority with a rating of 5 and 4 respectively. As
noted previously, widespread adoption relies on the long-term profitability of a system. Low maintenance
costs enable greater profit margins, making it the highest priority outside of Safety and Reliability. High
Turnaround Rate enables greater profitability by maximizing the time the system is carrying a useful
payload. Finally, Operational Flexibility and Ease of Manufacturing are both rated at three. Neither FOM is
unimportant, but they are not as critical compared to the higher rated FOMs. For example, greater
manufacturing costs are acceptable if it is possible to increase the safety and reliability of a system.
Alternatively, lower operational flexibility is acceptable if the ease of maintenance for a system is increased.
There are a variety of secondary considerations that are not discussed here, since they are not important
enough to warrant detailed discussion.
7. 7
2.3 Evaluation of Potential Configurations
The team further analyzed the two best existing configurations along with two new configurations derived
from existing configurations. Each of these possible configurations is given a score based on the FOMs. A
point value between one and five for each FOM is given, then multiplied by the FOM’s corresponding priority
score. This is shown in Table 2 below. For clarity, only the final multiplied point value for each FOM is
shown for each of the configurations.
Table 2: Configuration Selection
Underwing Pylon
Conventional T-
tail
Flying Quad Wing Conventional
Flying Wing
Twin-boom
Inverted V-tail
Safety and
Reliability
30 12 18 30
Ease of
Maintenance
25 20 15 25
Turnaround
Rate
12 16 16 20
Operational
Flexibility
9 9 6 15
Ease of
Manufacturing
12 9 9 15
Total 88 66 64 97
New designs include a twin-boom inverted V-tail design with eight vertical lift motors and a pusher motor
on the tail in addition to a flying wing design with underwing pylons for vertical lift. The Twin-boom Inverted
V-tail configuration, ultimately achieved 97 points, edging out the Underwing Pylon Conventional T-tail
configuration exemplified by the Jump 15. Both the Flying Quadwing and Conventional Flying Wing scored
worse than the other two designs due to safety concerns and difficult maintenance.
The Flying Quadwing and Conventional Flying Wing are particularly risky during the hover segments, as
any loss of engine power could result in immediate failure. While a flight termination parachute could
mitigate this risk, these designs receive a lower score from a reliability standpoint. Both vehicles also
received low scores in Ease of Maintenance. By its nature, a flying wing’s structure must be highly
integrated, greatly complicating any maintenance task.
Both the Twin-boom Inverted V-tail and Underwing Pylon Conventional T-tail configurations possess
superior Safety and Reliability, Ease of Maintenance, and Ease of Manufacturing. Both designs inherently
8. 8
segregate structural elements, simplifying maintenance and manufacturing tasks. Most importantly, both
possess a wing and a tail, giving superior glide characteristics and controllability in the case of unexpected
subsystem failure. In particular, the ability for either configuration to make a controlled landing in an engine-
out scenario enables the safest possible response to unexpected subsystem failure.
Ultimately, the Twin-boom Inverted V-tail surpasses the Underwing Pylon Conventional T-tail configuration
in the area of Operational Flexibility. With the twin-boom structure, the fuselage is independent of the boom
structure and can be reconfigured for different missions. The layout can be designed to be modular enabling
critical components to be swapped if damaged or unsuitable for a mission.
3. WEIGHT SIZING
3.1 Mission Profile
The basic mission profile is illustrated in the figure below, but there are additional requirements that need
to be considered in the sizing process. Both missions should be possible to complete at both sea level and
an altitude of 2000m ISA +10C. The cruise speed must be between 80 km/hr and 194 km/hr. The plane
also needs to be capable of sustained flight in 10 m/s (~36 km/hr) winds. The maximum mass of the vehicle
is 25 kg. Finally, there is a list of required fixed equipment not included in the 3 or 5 kg payload that must
be carried. This is listed in the table below, including the power draw. The optional parachute recovery
system was included to maximize safe operation of the vehicle.
Figure 3: Mission Profile
9. 9
Table 3: Fixed Equipment
Component: Mass (kg) Power (W)
Flight Computer 0.58 10
IMU 0.0575 5
ADS-B 0.155 11
Antennas 0.322 7
Camera 0.483 21
Comms 0.575 17
Parachute 1.02 10
Para Launcher 0.18055 10
Subtotal: 3.373 91
3.2 Sizing Assumptions
The sizing process required a number of assumptions before analysis could begin. These assumptions
were validated with additional analyses performed later.
Vehicle mass at the 25 kg limit
Cruise speed at the 80 km/hr minimum
Stall speed of 48 km/hr (~30 mph) at sea level to ensure gusts do not stall the airplane
Maximum lift coefficient of 1.5
Aspect ratio of 16
Oswald’s efficiency factor of 0.8
Parasitic drag coefficient of 0.02866
The combination of the vehicle mass, stall speed and maximum lift coefficient results in a required wing
area of 1.486 m2
, at a wing loading of 165.2 N/m2
.
The battery selected is the ThunderPower High Voltage TP9000-6SHV. ThunderPower is one of the most
respected names in the Li-Po battery world and the High Voltage series are specifically designed to achieve
the best possible energy density at the expense of power density. The estimated energy density is 203.2
W-hr/kg based on the reported weight of 0.984 kg per battery at the rated capacity and a voltage of 3.7
V/cell.
The difference between the two required payloads is 2 kg and conveniently this translates into a difference
in cell count between the two missions of two battery packs. It was assumed that eight packs would be
used for the 100km mission and six packs for the 60 km mission, which is substantiated in the following
analysis.
10. 10
The battery energy is split between three different elements. The first is the required fixed equipment energy
draw; this is estimated based on the power draw multiplied by the flight time. The second is the VTOL
energy; which is split between one minute of full power for climbout plus a total of six minutes of hover
power. The final energy element is the forward flight portion of the mission. This is calculated based on the
drag of the vehicle multiplied by the distance traveled, with efficiency factors for the propulsion system taken
into account as shown below. The efficiencies include the motor and propeller efficiencies estimated using
MotoCalc as well as the power management and distribution system estimated based on the voltage drop
over the length of wire at 12 AWG with a current draw of ~30 Amps
𝑅𝑎𝑛𝑔𝑒 =
(𝐵𝑎𝑡𝑡𝑒𝑟𝑦𝐸𝑛𝑒𝑟𝑔𝑦 − 𝑉𝑇𝑂𝐿_𝑒𝑛𝑒𝑟𝑔𝑦 − 𝐹𝐸_𝑒𝑛𝑒𝑟𝑔𝑦) ∗ 𝜂 𝑚𝑜𝑡𝑜𝑟 ∗ 𝜂 𝑝𝑟𝑜𝑝 ∗ 𝜂 𝑃𝑀𝐴𝐷
𝐷𝑟𝑎𝑔
3.3 Sizing
The first element to be estimated is the VTOL energy required, which is common for both mission payloads.
This is found by using the full power level for climbout and the power required for hover. The elements of
this analysis are shown in the table below for both altitudes. The thrust and power levels came from
MotoCalc analysis using the exact motor/battery/ESC/propeller combination selected.
Table 4: VTOL Energy Required
Thrust (N) T/W Power (W) Time (min) Energy (W-hr)
Full (2000m ISA+10C) 395.4 1.62 5619 1 93.65
Hover (2000m ISA+10C) 244.7 1.0 2564 6 256.4
Total (2000m ISA+10C) 350.05
Full (SL) 403.0 1.65 5723 1 95.39
Hover (SL) 244.7 1.0 2358 6 235.84
Total (SL) 331.23
The total range of the vehicle can now be built up from the various elements as shown in the table below.
The range of the vehicle far exceeds the required 100 km for the 3 kg payload and 60 km for the 5 kg
payload. This allows plenty of margin for wind, errors in the analysis, plus any deleterious effects from hot
or cold temperature on the batteries and motors.
11. 11
Table 5: Still Air Range Analysis
3 kg Payload
2000m ISA
+10C
5 kg Payload
2000m ISA
+10C
3 kg Payload
Sea Level
5 kg Payload
Sea Level
VTOL Energy (W-hr) 350.1 350.1 331.2 331.2
Fixed Equipment Energy (W-hr) 182.0 123.8 166.5 114.7
Forward Flight Energy (W-hr) 1066.3 725.0 1100.6 752.9
Drag (N) 14.39 14.39 16.19 16.19
Propeller Efficiency (%) 73 73 73 73
Motor Efficiency (%) 91 91 91 91
PMAD Efficiency (%) 97 97 97 97
Range (km) 172.0 116.9 148.0 101.3
Flight Time (hr) 2.15 1.46 1.85 1.26
It was also desired to check the maximum sustained head wind the full mission could be completed for both
potential payloads, as well as the maximum range at the suggested wind speed of 10 m/s (36 km/hr). This
is done by using the flight time for the range analysis above, but adjusting the ground speed based on the
true air speed minus the wind speed.
Table 6: Wind Range Analysis
3 kg Payload
2000m ISA
+10C
5 kg Payload
2000m ISA
+10C
3 kg Payload
Sea Level
5 kg Payload
Sea Level
Range with 10 m/s headwind
(km) 90.8 60.5 77.6 52.0
Max headwind where required
range can be achieved (km/hr) 31.5 36 23.3 29.1
As can be seen in the table, there is only one combination where the maximum wind speed can be withstood
while still completing the required mission. All of the potential missions can be completed with a 23.3 km/hr
wind.
12. 12
3.4 Weight and Balance
The total weight and balance component breakdown for both payloads is shown in the table below. This
was used in the stability and control and is consistent with the final CAD package. The datum for all
measurements is the leading edge of the wing.
Table 7: Weight and Balance for Both Missions
Name Quantity Unit Mass
(kg)
Net Mass
(kg)
X distance
(mm)
Moment
(mm-kg)
Payload Mechanism 1 1.35 1.35 165.10 223.17
Tail 1 0.78 0.78 1136.65 885.50
VTOL Motor Assemblies 4 0.57 2.27 127.00 288.03
Forward Flight Motor 1 0.19 0.19 1008.38 193.23
Wing Outboard Sections 2 0.41 0.82 165.10 134.80
Landing Gear 4 0.23 0.93 127.00 117.52
Wing Inboard Sections 1 0.59 0.59 165.10 96.61
Fairings 2 0.11 0.22 127.00 27.65
Tailbooms 2 0.04 0.07 355.60 26.18
Extensions 2 0.03 0.06 -635.00 -40.32
Negative Bus 1 0.22 0.22 -186.94 -41.55
Positive Bus 2 0.15 0.30 -186.94 -55.97
Flight Termination Chute 1 1.02 1.02 127.00 129.54
Flight Termination Launcher 1 0.18 0.18 499.62 90.21
Launcher Cradle 1 0.15 0.15 499.62 72.52
Antennas and Externally
Mounted Systems
1 0.38 0.38 -295.28 -112.50
ADS-B 1 0.11 0.11 550.42 62.42
IMU 1 0.06 0.06 127.00 7.49
Breaker Switch 1 0.03 0.03 -186.94 -5.09
FCC 1 0.61 0.61 550.42 337.05
Coms 1 0.58 0.58 550.42 317.07
Camera Dome Holder 1 0.61 0.61 -406.40 -248.86
Camera System 1 0.48 0.48 -486.72 -234.02
Fuselage Structure 1 1.50 1.50 127.00 190.68
Empty Total 13.50 182.28
Battery One 1 7.73 7.73 -95.67 -739.43
Payload One 1 3.00 3.00 177.80 533.39
Mission One Total 24.23 93.07
Battery Two 1 5.91 5.91 -95.67 -564.98
Payload 2 1 5.00 5.00 177.80 888.75
Mission Two Total 24.41 105.67
13. 13
4. PROPULSION SELECTION
Motor and propeller databases were generated using the power requirements dictated by the Weight Sizing
process. Propulsion selection was done in two stages for the VTOL and Forward Flight stages of the mission
profile. In order to simplify the design, the propulsion system for both stages were sized to use the same
battery pack. Both stages were sized using fixed pitch propellers, as stipulated by the requirements.
MotoCalc, a commercially available motor analysis tool, was then used to estimate the motor efficiency,
static thrust, and thrust at specified flight conditions for a variety of propeller and motor combinations. The
combinations which equaled or surpassed the power requirements discussed above were analyzed further.
4.1 VTOL Stage
The goal for the VTOL propulsion system was to weigh as little as possible while being capable of
generating a thrust-to-weight ratio of 1.6 at 2000 m ISA +10C. This level of thrust would mean that we could
lose two motors and still have 20% more thrust than required to maintain hover.
Direct-drive brushless out-runner motors with low motor constants (Kv) were preferred for the VTOL Stage
of the mission. The selection from the database was narrowed to the Tiger Motor (T-Motor) U5, U7, U8,
U10, U11, Antigravity 4004, and Antigravity 4006 motors on account of their low Kv ratings and proven high
reliability. The propellers tested were T-Motor 16x5.4CF, 20x6CF, 26x8.5CF, 28x9.2CF and 30x10.5CF
propellers. Of this selection space, the T-Motor U10 paired with the T-Motor 26x8.5CF propeller, the T-
Motor U11 paired with the T-Motor 30x10.5CF propeller, and the T-Motor U8 paired with the T-Motor
28x9.2CF propeller met the power and thrust requirements. However, the T-Motor U11 paired with the
30x10.5CF propeller was deemed too heavy, and the T-Motor U10 paired with the 26x8.5CF propeller drew
more power than the T-Motor U8 with the 28x9.2 propeller. The lower weight and higher efficiency of the
T-Motor U8 with the T-Motor 28x9.2CF propeller drove its selection as the VTOL propulsion system.
4.2 Forward Flight Stage
The goal for the forward flight stage was to create the thrust required to balance the drag as estimate in the
range sizing while maximizing propeller and motor efficiency. Direct-drive brushless out-runner motors with
moderate motor constants (Kv) were preferred for the Forward Flight Stage of the mission. The selection
from the database was narrowed to the Hacker A30-10XL, A30-12XL, A40-10L, and A40-14L motors on
account of their moderate Kv ratings and proven high reliability. The propellers tested were APC 14x10E,
16x10E, and 17x12E propellers. Of this selection space, only the Hacker A40-14L paired with the APC
17x12E propeller met the power and thrust requirements. This combination was selected as the forward
flight propulsion system.
14. 14
4.3 Power Management and Distribution (PMAD)
The propulsion system sizing determined that a total of nine motors and either six or eight batteries will be
used, depending on the mission. All the batteries and all the loads will be connected in parallel to maximize
the amount of power available to the aircraft in hover and minimize the effects of battery resistance during
cruise. In addition, the number of electrical junctions throughout the aircraft were minimized to simplify
assembly, with independent wires running from each motor’s speed controller to the electrical junctions
mounted in the front of the fuselage near the batteries. The resulting circuit is shown below. Note the
inclusion of a break switch that enables the airplane to be quickly made safe for payload retrieval and
battery swapping.
Figure 5: PMAD Block Diagram
The negative leads of all the batteries and motors are connected to a common negative bus, while the
positive leads are connected into two separate busses for motors and batteries. These two busses are
connected using higher gauge wire through a breaker/power switch rated for 300A, which protects the
system in the event of a short but serves a more important purpose as a safety switch for the entire aircraft.
Mounted on the exterior of the front of the fuselage, opening the switch as soon as the aircraft has landed
disables the propulsion system, allowing cargo to be loaded or unloaded without any fear of the propellers.
The switch is closed when the aircraft is prepared for takeoff.
Battery Battery Battery Battery Battery Battery Battery Battery
Motor Motor Motor Motor Motor Motor Motor Motor Motor
Negative Bus
Battery
Positive Bus
Motor
Positive Bus
300 A
Breaker
/ Switch
15. 15
The wire gauge for the distributor wires was selected based on the tradeoff between resistive losses and
wire weight inherent to the problem. This design requires 78 feet of wire to run to all motors, including the
longer run to the tail motor for forward flight. The efficiency can be computed from the formula below:
𝜂 = 1 −
𝐼𝑠𝑢𝑠𝑡𝑎𝑖𝑛𝑒𝑑 𝑅 𝑤𝑖𝑟𝑒
𝑉𝑛𝑜𝑚𝑖𝑛𝑎𝑙
The choice of 14-gauge wire resulted in a transmission efficiency of 98% during hover and 97% during
cruise, due to the longer wires to the cruise motor. The total wire weight is 0.98 lbs, plus about 0.22 lbs for
the switch.
In a production aircraft the connections to the battery packs would be made with stiff connectors which
connect as the batteries are slid into place, enabling a rapid turnaround time. The six battery pack would
also include plastic covers to slide over the connectors intended for the seventh and eighth batteries to
prevent short circuits.
5. AERODYNAMICS
5.1 Airfoil Selection
Flight conditions were estimated at a Reynolds number (Re) of approximately 450,000. Hundreds of airfoils
from the UIUC database were analyzed through a MATLAB script at the estimated Re and filtered based
on their thickness and complexity. Complex airfoil geometry, such as high camber, can result in
manufacturing errors that will negatively affect vehicle performance. Low thickness airfoils typically have a
small leading edge radius, resulting in more abrupt stalls at lower angles of attack. Increased airfoil
thickness allows for increased geometric stiffness while reducing structural weight. After filtering, the
resulting airfoils were further analyzed on maximum section lift coefficient and lift to drag ratio. The four
airfoils chosen were the SD7062, the Falcon, Clark Y, and the S822. Drag and lift coefficient curves for the
four airfoils were constructed using the 2-D panel code XFOIL and are shown below.
Figure 6: Experimental Lift and Drag Characteristics for Selected Airfoils from UIUC
0
0.5
1
1.5
2
0 10 20
SectionLiftCoefficient
Angle of Attack, Degrees
Clark Y
Falcon
S822
SD7062
0
0.5
1
1.5
0 0.01 0.02 0.03
SectionLiftCoefficient
Section Drag Coefficient
Clark Y
Falcon
S822
SD7062
Re = 450,000 Re = 450,000
16. 16
Examination of the drag polar shows that the SD7062 and S822 airfoils have a more stable sectional drag
coefficient over longer ranges of lift coefficient than the other airfoils. This indicates that for a given range
of lift coefficients around 0.6 (the approximate cruise CL), the drag remains relatively low and constant. This
is important as lift will vary across the wings due to downwash from the wingtip vortices and environmental
variables such as local wind. The SD7062 was chosen as the main wing airfoil because it has high thickness
of 14% and superior maximum lift and lift-to-drag ratio.
5.2 Planform Selection and Lift Distribution
The wing planform needs to have an aspect ratio of 16 based on the range analysis, but needs to be better
defined. The wing was modeled using AVL, a freely available vortex lattice analytical tool developed by Dr.
Drela of MIT. The wing was modeled with the SD7062 airfoil, and is shown below including the tail which
will be discussed later.
Figure 7: AVL Model
The wing has a small amount of taper on the outboard sections, which also have 5 degrees of dihedral.
The taper modifies the lift distribution to make it more elliptical and reduce the induced drag to further
improve performance. The dihedral improves roll stability, makes it less likely for a wingtip to impact the
ground, and makes it possible to control the airplane with only rudder-elevator. The wing is also mounted
at an incidence angle of 2.2 degrees, which enables the wing to generate sufficient lift for cruise while
allowing the fuselage and tailbooms to be in their minimum drag condition. The resulting lift distribution is
shown below. The estimated Oswald’s efficiency factor is 0.97 not including the effect of the fuselage and
boom, such that the sizing assumption of 0.8 is clearly conservative.
17. 17
Figure 8: Lift Distribution
5.3 Parasitic Drag Estimation
The parasitic drag coefficient was estimated based on a form factor component drag buildup. The equations
used were taken from Roskam’s Airplane Aerodynamics and Performance book, with the wing, tail and
landing gear modeled as lifting surfaces and the fuselage and tailboom modeled as fuselages. The wing is
the dominant factor in the drag, based on its large contribution to the total wetted area. The drag is
consistent with the estimation made as part of the range sizing described earlier.
Component CD0
Wing 0.01323
Fuselage 0.00845
Tailboom 0.00610
Landing Gear 0.00055
V-Tail 0.00033
Total 0.02866
Figure 9: Parasitic Drag Estimates
Wing
Fuselage
Tailboom
Landing Gear
V-Tail
18. 18
6. STABILITY AND CONTROL
Static and dynamic stability characteristics were computed to ensure that the aircraft would be able to
successfully complete the design mission. AVL was used to determine static stability derivatives. This
information was combined with the principal moments of inertia found in CAD to determine dynamic stability
behavior using the full six degree of freedom linearized, coupled differential equations found in Philips’
Mechanics of Flight. The most important static derivatives, deflections, and the static margin are seen on
the left side of the table below, while the most important high-frequency dynamic mode behavior is seen on
the right. The static stability analysis confirmed that the design is statically stable with 14.1% margin. The
dynamic analysis indicated that the aircraft is stable in all high-frequency modes, with damping ratios and
frequencies within the expected ranges for small unmanned vehicles. This analysis was completed based
on the 3 kg payload.
Table 8: Relevant Stability Parameters for Elevator-Balanced Trim Conditions
Static Stability Dynamic Stability
Inputs
mtotal(kg) 25 Mode
Short-
Period
Dutch
Roll
Roll
V(km/hr) 80 Damping Rate (s-1
) 4.17 0.58 14.64
Aerodynamic
Parameters
CL 0.6 Time to Half (s) 0.17 1.20 0.047
α (deg) 0.0 Damping Ratio 0.58 0.15
β (deg) 0.0 Damped Freq. (s-1
) 5.91 3.85
Stiffness
Coefficients
Cm,α (rad-1
) -0.820 Undamped Freq. (s-1
) 7.24 3.90
Cl,β (rad-1
) -0.113 Control
Cn,β (rad-1
) 0.070 Cn,δr (deg-1
) 0.0015 δa (deg) 0
Static Margin % Chord 14.1 Cm,δe (deg-1
) -0.0366 δe (deg) 1.43
In addition to the stability, the trim needed to be analyzed in a range of conditions. The center-of-gravity
shifts slightly between the different battery arrangements required for the 3 kg and 5 kg payload missions,
and both needed to be checked. In addition, the cruise speed and stall speed need to be trimmable. Finally,
the ruddervators are split to allow for redundancy and it was checked if either the upper or lower
ruddervators failed that the airplane could still be trimmed. As can be seen in the table below the maximum
required trim angle is 14.6 degrees, enabling plenty of margin for the control system and any required
maneuvering. Note that this includes the inherent nose down pitching moment generated due to the thrust
line of the forward flight motor being well above the vertical cg position. The thrust generates a pitching
moment of coefficient of 0.085 that must be trimmed out in the worst case scenario.
19. 19
Table 9: Longitudinal Trim Analysis
Required Trim Angle Forward CG
Cruise (deg)
Aft CG
Cruise (deg)
Forward CG
Stall (deg)
Aft CG
Stall (deg)
All Ruddervators -0.9 -0.1 -6.1 -4.1
Top Ruddervator Only -2.2 -0.3 -14.6 -9.6
Bottom Ruddervator Only -1.6 -0.2 -10.3 -6.9
7. STRUCTURAL SIZING AND MATERIAL SELECTION
7.1 Structural Layout
The primary goal for the structural layout was to ensure that all loads were accounted for and have an
adequate load path to the major load-bearing components. A diagram of these loads and their relevant load
paths can be found below in Figure 10. Each of the tailbooms must transfer the motor, ground, and
aerodynamic tail loads longitudinally along the tailboom spar into the wing spar. In turn, the wing spar must
transfer the loads from the tailbooms, along with the aerodynamic wing loads, into the fuselage. The use
of composite tubes throughout ensures a simple, stiff, strong structure. Many of the brackets connecting
the load bearing pieces are made from SLS 3D printed nylon.
Figure 10: Load Paths of Major Forces
20. 20
7.2 Material Selection
The materials used throughout the design were selected for the best combination of manufacturability,
weight, durability, and cost. Carbon fiber was selected for the spars due to its superior strength-to-weight
ratio and the availability of off-the-shelf parts for these applications. Square tubes were selected for the
tailboom spar to make interfacing with the spar easier, while round tubes were selected for the wing and
tail spars for their better torsional stiffness for a given cross sectional area. Plywood was selected for the
motor mounts, bulkheads, and the majority of the ribs of the aircraft for its cost and ease of manufacturing.
Fiberglass was selected for the skin of the aircraft for its high strength-to-weight ratio and the ability to
create a smooth, durable semi-monocoque structure. SLS 3D printed nylon was selected for the payload
retention mechanism, connections between carbon fiber tubes, and landing legs for its durability and the
ability to manufacture parts with complex geometries. A breakdown of the material for each of the major
structural components is shown in Table 10.
Table 10: Major Structural Component Material Breakdown
Aircraft Structure Material
Skin 3oz Fiberglass
Wing & Tail Spars Carbon Fiber Twill Wrapped Round Tubes
Tailboom Spar Carbon Fiber Twill Wrapped Square Tubes
Motor Mounts, Ribs, and Bulkheads Plywood
Wing-Fuselage & Wing-Tailboom Ribs SLS 3D Printed Nylon
Tail-Tailboom & Tail Apex Brackets SLS 3D Printed Nylon
Payload Retention Mechanism & Landing Legs SLS 3D Printed Nylon
7.3 Wing and Tailboom Spar Sizing Analysis
A sizing analysis was conducted for the two structural members seeing the most load, the wing and tailboom
spars, in order to minimize weight. A preliminary Euler-Bernoulli beam bending analysis was first conducted
to approximate what spar cross section would be acceptable. The tailboom spar study was conducted by
placing a point load equal to the thrust generated by the pairs of VTOL motors at the correct distance from
the location where the wing spar crosses the tailboom spar and neglecting the tail force for a more
conservative estimate. The Euler-Bernoulli beam bending analysis for the selected tailboom spar, shown
in Figure 11, resulted in a maximum deflection fore and aft of 0.81 mm and 2.41 mm, respectively.
21. 21
Figure 11: Euler-Bernoulli Tailboom Spar Bending Analysis
The wing spar study was conducted by examining a half span under a 2g loading. A uniform distributed
load equal to takeoff weight was placed along the length of the half span and a point load equal to the thrust
generated by both pairs of VTOL motors were placed at the wing station where the tailboom spar crosses
the wing spar. The Euler-Bernoulli beam bending analysis for the selected wing spars, shown in Figure 12,
resulted in a maximum tip deflection of 121.7 mm.
Figure 12: Euler-Bernoulli Wing Spar Bending Analysis
Before final selection of the wing and tailboom spars, finite element analysis was used to verify the results
from the Euler-Bernoulli beam analysis. The results of the tailboom spar analysis of the more demanding
load case aft of the wing spar are shown in Figure 13, while the results of the wing spar analysis are shown
in Figure 14. In both cases, the finite element analysis deflections are slightly greater than those from the
Euler-Bernoulli Beam Analysis, with a 2.59 mm maximum deflection for the aft portion of the tailboom and
a 126.5 mm tip deflection of the wing spar. These deflections are still within tolerable limits.
0.00
0.50
1.00
1.50
2.00
2.50
3.00
0 0.2 0.4 0.6 0.8 1 1.2 1.4
Deflection(mm)
Distance from Wing Spar (m)
Fore
Aft
0.0
5.0
10.0
15.0
20.0
25.0
30.0
35.0
40.0
45.0
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8
Deflection(mm)
Distance from Aircraft Centerline (m)
22. 22
Figure 13: Tailboom Aft Spar Finite Element Analysis
Figure 14: Wing Spar Finite Element Analysis
The final result is a simple, modular structure capable of handling the required loads with minimal deflection.
In all cases, the chosen limiting factors was deflection based and not stress based. In the team’s
experience, carbon fiber is capable of handling much higher deflections than the substructure, and therefore
if it is designed too close to the stress limit there will be substructural failure. This is mitigated by the minimal
deflections results indicated in the above analysis.
