1
Design of a Subscale Reconfigurable eVTOL Aircraft for
Transition Corridor Flight Testing
Frank K. Kozel,1 Jacqueline Q. Tu,1 and Eddie Q. Li1
Georgia Institute of Technology, Atlanta, Georgia, 30332
Matthew M. Warren2
Georgia Institute of Technology, Atlanta, Georgia, 30332
Urban Air Mobility (UAM) is an emerging class of transportation that is envisioned as a
low-cost, on-demand, point-to-point passenger air service with flights between rooftop
“vertiports” situated throughout cities. The types of aircraft being considered for UAM are
electric VTOL (eVTOL) aircraft that typically include a wing for forward flight efficiency and
distributed rotors for vertical flight and hover. Many of these aircraft also include tiltrotor
or tiltwing mechanisms to achieve "conversion" or "transition" from horizontal to vertical
flight and vice versa. The complex aerodynamics, flight dynamics, and control of transition
are among the most difficult aspects of the design of tiltwing and tiltrotor VTOL aircraft. In
this paper, we describe a scale aircraft that was designed and built to investigate the flight
performance of new configurations of eVTOL aircraft that have difficult transition
aeromechanics and flight performance. The aircraft is a modular tiltrotor that uses four rotors
for vertical lift and tilts the two rear rotors for forward thrust. The aircraft was designed and
constructed using rapid prototyping techniques with a focus on robustness and
reconfigurability. Material selection was primarily light balsa and plywood for structural
shape with metal and composite components used to transfer large structural loads. Custom
control systems based on the open source PX4 software were implemented to control the
aircraft in vertical flight and transition to forward flight. A flight test plan to measure
transition performance was developed in which the flight envelope is incrementally expanded
to manage risk. The aircraft will be used to test a variety of eVTOL configurations and to
collect representative subscale flight test data.
Nomenclature
α = angle of attack
cd = section drag coefficient
CD0 = parasite drag coefficient
cl = section lift coefficient
Clβ = roll damping coefficient
Cmα = pitch damping coefficient
Cnβ = yaw damping coefficient
CL = airplane lift coefficient
CLmax = maximum airplane lift coefficient
CD = airplane drag coefficient
Re = Reynold’s number
Wtotal = total weight
1 Aerospace Engineering Undergraduate Student, Daniel Guggenheim School of Aerospace Engineering, AIAA
Student Member. 2 Aerospace Engineering Graduate Student, Daniel Guggenheim School of Aerospace Engineering, AIAA Student
Member.
2
I. Introduction
HE primary objective for conducting this research is to advance the understanding, technology, and operating
efficiency of electric transition aircraft. A team from the Georgia Institute of Technology, consisting of one
graduate and three undergraduate aerospace researchers, worked together with the intent to develop a small prototype
aircraft with the ability to transition from vertical flight to horizontal flight. Three of the members possess expertise
in the techniques used to rapidly prototype unmanned aerial vehicles for the AIAA Design Build Fly (DBF)
competition. Using these techniques, the team fully developed the requirements for this aircraft, conducted analysis
on a design, iterated on this design, manufactured the final aircraft, and flight tested in just over four months. This
paper describes the process used by the design team to create a fully realized flying model that will provide data to
understand the dynamics of flight during transition. With this information, the team will characterize the transition
states of the model to show that rapid prototyping can yield insightful flight test data about complex and non-linear
flight mechanics for a specific eVTOL configuration before a full-scale version is ever built.
II. Project Overview
Instead of operating on a linear design process, the team used an interconnected, rigorous aircraft design cycle to
explore the design space. For example, reconfigurability of the aircraft was implemented early in the conceptual design
phase to ensure a multipurpose aircraft was produced for future research endeavors. Additionally, flight testing is
shown to not only validate the design but expose solvable defects in the aircraft conceptual design such as the stability
and control of the vehicle.
Figure 1: Aircraft Design Cycle
Each team member contributed to the manufacturing
of the vehicle which took place over several weeks.
Previous rapid prototyping experience from the team’s
senior members guided the scheduling, while knowledge
of supply chains, familiarity with construction, and
effective management techniques enabled on-time
completion. As a benefit, the modularity of design meant
no time was lost to component integration.
Material and component costs were documented
thouroughly to demonstrate the affordability of a small
scale research vehicle. Figure 2 shows a visualization of
the main component groupings and their contributions to
cost. As the team expected, the largest contributor to cost
were the electronics. The flight control system and the
propulsion system electronics made up 83% of the budget.
III. Design Process
A. Requirements
The success of the research depends upon the aircraft’s ability to satisfy the desired mission requirements in a
reliable and effective manner, namely providing flight test data during transition between horizontal and vertical flight
for a given configuration. Stability in both vertical and horizontal flight modes and robustness for many flight tests
factored heavily into the design. Additionally, electric propulsion shall be used for its ease of use in model remote
control (RC) flying, reliability as a system, and in the spirit of a future full scale eVTOL aircraft. Multiple power
T
Adhesives
4% Structural
Materials
13%
Propulsion
System
37%
Flight
Control
Electronics
46%
Adhesives
Structural
Materials
Propulsion
System
Flight Control
ElectronicsTotal Cost: $2,838.16
Figure 2: Cost Analysis
3
systems shall be used in a multirotor fashion due to the availability and low cost of hobby grade electric motors. The
aircraft shall also be a Group 2 UAV, a vehicle with a maximum takeoff weight (MTOW) between 20 and 55 lbs, that
will provide flight test data with the intent to pass or fail a potentially feasible urban air mobility (UAM) eVTOL
configuration. Lastly, modular components shall be used for ease of transportation, replacing damaged pieces, and for
the ability to test multiple aircraft configurations.
B. Risk Assessment
The primary risk is complete loss of control of the aircraft, leading to a crash. This risk is mitigated by reducing
both the probability of occurrence and severity of consequence. Temporary loss of control is probable by nature of
this research vehicle for transition dynamics. They key is to ensure a high probability of control recovery. This is
achieved by creating a simple and stable fixed-wing platform with predictable gliding characteristics, and testing
transition maneuvers at sufficient altitude for a power-off recovery. In the event of temporary loss of control, the
vehicle is designed to allow immediate manual override by the pilot into “airplane mode” or, if appropriate, reversion
to the onboard flight controller for “multirotor mode”. The severity of consequence is measured in injury first and cost
second. The potential for injury to any people, including flight test crew and pilot, is low due to the small size of the
vehicle and remote operation over open fields. Small size also reduces the cost of the project and the financial burden
incurred for any crashes or failed components.
Secondary risks include structural failure, especially when landing. To reduce this potential, the aircraft is made
of components that can be easily replaced. Furthermore, additional plywood structure was added around areas known
to experience high stresses such as the attachment points for the landing gear and motors. The same modular system
used for testing different aircraft configurations is also used to replace broken or failed subsystems.
Other risks from a broader, systems level include susceptibly to manufacturing errors by engaged engineering
students and engineering complexity. Strategies used to combat this were jigging the wing structures for more precise
assembly, instituting a team hierarchy to enforce standardized construction procedures, and routinely performing
inspections and tests leading up to final assembly.
C. Configuration Selection
Figures of Merit (FOMs) were used to analyze different airplane configurations, as seen in Table 1, with higher
weights corresponding to more important characteristics. Considerations for the merit values were determined from
prior DBF experience with similar unmanned aerial vehicles (UAV) of the same general size. The vertical flight mode
would be performed with four motors in a quadrotor arrangement around the center of gravity, and transition would
occur by rotating motors to a new position. The quadrotor approach is a simple solution that is well understood and
that allows for a simple transition mode. This design also allows for modifications to utilize an octo-rotor configuration
or other combinations of tilting and stationary motors. Another consideration was each configuration’s modularity to
enable new rotor arrangements and transition strategies.
Table 1: Configuration Evaluation
Configuration
FOM Weight Flying Wing Blended Wing Conventional Canard
Control
Effectiveness
5 3 5 5 3
Internal
Volume
4 2 2 5 5
Simplicity 3 4 3 3 3
Modularity 2 1 2 5 4
Total 14 37 46 64 52
4
D. Sizing for Fixed-Wing Operation
The team determined that testing the model at high wing loadings for RC aircraft, about 4 lb/ft2, would simulate
more difficult transition dynamics indicative of full-scale aircraft. However, the team also determined that flying at
higher wing loadings would involve more risk, so the design incorporated the ability to fly at a low wing loading of 2
lb/ft2 for initial flight testing verification. The energy-based constraint diagram in Figure 3 was then used to determine
minimum power to weight for fixed-wing operation. Solid lines represent the minimum performance metrics that the
vehicle would need to surpass. Changing between 2 and 4 lb/ft2 is simply achieved by adding larger batteries to the
aircraft, hence the design point with higher wing loading in Figure 3 also has lower power to weight given the same
motors. The restriction of takeoff was added to enable fixed-wing verification testing.
Figure 3: Sizing Constraint Diagram (note: Power/Weight is indicative of fixed-wing operation)
E. Propulsion System Selection
The required propulsion system performance for fixed-wing operation was defined in the sections above. In hover,
the aircraft must have the ability to climb with margin for variable RPM control. For this reason, the team chose
maximum thrust equal to 1.6 times the MTOW of the aircraft where motor would provide 8 lbs of static thrust. During
forward flight, the aircraft thrust required is not nearly as substantial, so only two of the motors would be tilted. The
thrust required at high speed cruise speed was determined as 3.8 lbs total, 1.9 lbs per motor. No single fixed pitch
propeller is well suited for both cruise and hover requiring a compromise between the two.
The cruise and hover thrust both provide constraints on the motor selection. The motor selection criteria were to
meet the thrust requirements while minimizing the weight of the motor and power consumption. A 6-cell series 22.2V
lithium polymer (LiPo) battery was chosen to narrow the field for motor selection. This voltage was selected for its
compatibility with many off-the-shelf speed controllers and battery chargers. Component selection at higher voltages
becomes more limited and subsequently more expensive.
Many motor companies provide static thrust data allowing
for an initial screening. The most promising motors were then
analyzed with a variety of propellers using MotoCalc[1] which
is a commercial motor and propeller analysis tool that the
research team has already validated with wind tunnel
testing[2]. The results from MotoCalc were used to determine
the thrust and power statically and at the cruise airspeed.
Larger diameter propellers with a lower pitch performed
better statically but did not have enough pitch to generate the
required thrust at the cruise speed as seen in Figure 4.
The final propulsion system is a Cobra 3525-12 motor
with an APC 12x6 propeller. This motor should produce 8 lbs
of static thrust and 1.9 lbs of thrust at the cruise speed. Each
motor requires 55 A of current at full power. To reduce the
current load on the batteries and wiring, the airplane was
wired in two halves. Each battery connects to two electronic
speed controllers (ESC) and two motors and must supply 110 A of current for short bursts. The wires between the
battery and two ESCs is 8 AWG to handle the maximum burst current. In the light configuration, two 3.45 Ah batteries
0
5
10
15
20
25
30
0 1 2 3 4 5
Po
wer
/Wei
gh
t (W
att
s/lb
s)
Wing Loading (lbs/ft2)
TO-100 ft
TO-250ft
Cruise-90ft/s
Cruise-70ft/s
Climb-10ft/s
Climb-3ft/s
Lower Bound
Design PointUpper Bound
Design Point
Figure 4: Thrust Curves
5
rated for a 75C discharge were used to reduce the aircraft weight. The vehicle will have a hover time of around 6 min
with these small batteries at a continuous discharge of 10C. Once the weight is increased to the heavy configuration,
two 10 Ah batteries will be used to increase the flight time and reduce the discharge rate on the battery.
A Castle Creations Phoenix 60 ESC was chosen to match the motor, with a 60 A and 22.2 V rating. The ESC also
features data logging at 10 Hz which will be used during flight testing as described in Section V. The ESC can record
current, voltage, power, motor RPM, and commanded throttle.
F. Airfoil Selection
A Reynolds number (Re) of 300,000 was calculated using sea level cruise conditions for this aircraft and is
referenced for airfoil selection analysis. Four airfoils with favorable low Reynolds number characteristics were
selected from the UIUC airfoil database[3] and analyzed in XFOIL[4]. Figure 5 shows the chosen airfoils, their drag
polars, and their cl vs AoA plots. The figure also includes wind tunnel data[3] used to validate XFOIL for the SD7062.
Figure 5: Section Lift and Drag Characteristics for Selected Low Reynolds Number Airfoils
The team chose to use the SD7062 airfoil for the main wing. Among its competitors, the SD7062 has lower lifting
capability, however, it has the most favorable geometry for internal structure to support the motors and transfer load.
The stall characteristics are also favorable with stall occurring at higher angles of attack and occurring slowly rather
than abruptly. Additionally, the team has validated the SD7062’s performance through flight testing on similar aircraft
DBF projects.
G. Drag Analysis
The Hoerner drag buildup method was used to approximate
the parasite drag, CD0, found to be 0.0403. This method is only
a rough approximation of the actual CD0 of the aircraft. The
Vehicle Sketch Pad (VSP)[5] parasite drag function provided
a drag estimate for the wing surfaces and fuselage. VSP is not
suitable for motor pods and landing gear, for which drag was
estimated through a table of similar component shapes using
Hoerner’s Fluid Dynamic Drag[6]. The total parasite drag
buildup of the components is shown in Figure 6. It is intended
that initial testing of the vehicle will produce data to calculate
the actual CD0 to validate the drag estimation method.
H. Wing Configuration and Tail Sizing
Athena Vortex Lattice (AVL)[7] is a vortex lattice code that was used to perform stability and aerodynamic
analysis. The wing area was chosen based on the wing loading in Section D. The wing span was chosen to
accommodate two sets of propulsion pods per side, should octo-rotor or other configurations be investigated in the
0.25
0.5
0.75
1
1.25
1.5
1.75
0 5 10 15 20
c l
AoA (deg)
DAE 31
SD 7062
E 203
NACA 6414
SD7062 Wind Tunnel
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
0 0.02 0.04 0.06 0.08 0.1
c l
cd
DAE 31
SD 7062
E 203
NACA 6414
0.000920.00227
0.0082
0.0045
0.00450.0095
0.01043 Vertical Tail
Horizontal Tail
Landing Gear
Power Pod 1
Power Pod 2
Fuselage
Wing
Figure 6: Zero Lift Drag Coefficient Build up
6
future. The wing portion that supports propulsion pods is straight and untapered for simplicity. Swept polyhedral tips
were added for lower induced drag and improved spiral characteristics.
The center of gravity (CG) was placed at the wing quarter chord to reduce the wing’s nose down pitching
contribution, and reduce the tail down load. The horizontal tail placement and size targeted a static margin between
8% and 12% and an elevator deflection smaller than 15 degrees to trim at a nose-heavy CG at CLmax. The vertical tail
was sized for a Cnβ greater than 0.06 rad-1. Tail sizing was performed for 4 different tail configurations: T-tail, V-tail,
conventional tail, and cruciform tail. The T-tail was chosen to minimize interaction with the tilting propeller wash and
reduce stability changes through transition. Figure 7 shows the loading and lift distribution at an AoA of 12 degrees.
Figure 7: Aircraft AVL Virtual Model and Predicted Trefftz Plot
I. Wing Structural Analysis
The wing was chosen to be removable for transportation, maintenance, or configuration changes. The wing utilizes
a carbon fiber tube as the main spar to transfer the bending and torsional loads from the motors and aerodynamic
forces. Carbon fiber was chosen for its high stiffness to weight ratio, high strength, and off-the-shelf availability. Balsa
and plywood ribs comprise the rest of the structure, where balsa wood is used as ribs to hold the shape of the airfoil
and plywood is used for the motor and fuselage attachment hardpoints as well as the polyhedral joint. Spar sizing on
an eVTOL aircraft with podded motors is dictated by conventional fixed-wing loads and the ability to retain control
authority in VTOL operation. The tube spar’s VTOL sizing criterion was an angular deflection of less than 5 degrees
relative to the fuselage during maximum yaw conditions in hover.
In fixed-wing operation, the wingtip deflection was calculated with simple Euler-Bernoulli beam theory equations
to ensure it did not exceed 5% of the span under the maximum loading conditions. A root bending moment of 290 lbs-
in generates a maximum stress of 10 ksi which is only 3.7% of the ultimate tensile stress. A maximum shear force of
18 lbs is at the root. The 2 lbs power pods mounted under the wings decrease the root shear stress and bending moment.
Figure 8 displays the displacement, shear force, and bending moment as a function of semi-span.
Figure 8: Shear, Bending Moment, and Deflection Diagram Along Semi-Span
0
50
100
150
200
250
300
350
0
4
8
12
16
20
0 5 10 15 20 25 30 35
Mo
men
t (l
b-i
n)
Shea
r (l
bs)
Span Location (in)
Shear
Bending Moment
0
0.1
0.2
0.3
0.4
0.5
0 5 10 15 20 25 30 35
Def
lect
ion (
in)
Span Location (in)
7
J. Propulsion Attachment Structure Design
The propulsion mounting structure was designed to be
simple, rigid, and modular. Each motor interfaces with a carbon
fiber tube that mounts directly to the wing. Euler-Bernoulli
beam bending theory was used to determine the size of carbon
fiber rod for a deflection of less than 5% of the protruding length
of the carbon rod. Mounting holes were drilled to allow for
quick and easy attachment to the hardpoints in the wing.
The tilt mechanism design was kept simple and the servo
was oversized to account for the dynamics of rotating a spinning
motor with a distributed mass on the end. The motor assemblies
were encased in custom 3D printed fairings to reduce drag.
Figure 9 shows an image of the tilt mechanism and fairing housings. Figure 9: Modular Propulsion Assembly
K. Software Utilization
A variety of software was used within the conceptual design block including Open VSP, AVL, XFOIL, Excel,
AutoCAD, and SolidWorks. Furthermore, data and design experience was brought from similar aircraft made for
AIAA DBF projects. This knowledge, when integrated with the tools mentioned previously, allows for a rapid
conceptual design phase.
SolidWorks is utilized to create a high-fidelity model of the entire aircraft following the conceptual design phase.
A CAD view of the model is shown in the Appendix. From this 3D model, 2D CAD drawings are created using
AutoCAD for precision laser cutting of the components used to build the aircraft. The specific applications for all the
design and manufacturing tools are shown in Table 2.
Table 2: List of Computer Aided Design Tools
Program Application
Dassault Systèmes SolidWorks 3D Modeling & Finite Element Analysis
Autodesk AutoCAD 2D Modeling & Laser Cutting Vectors
Microsoft Excel Aircraft Sizing, Structural Analysis, & Bookkeeping
Vehicle Sketch Pad (Open VSP) 3D Modeling & Aircraft Visualization
Athena Vortex Lattice (AVL) Aircraft Stability Analysis
XFOIL Airfoil Analysis & Selection
From the AutoCAD drawings, an Epilog Fusion laser cutter is used to cut balsa and plywood providing an easily
repeatable, cheap, precise, and quick way of creating complicated parts. The application allows for tight tolerances on
components and precision in construction. The wing ribs were aligned and positioned correctly using laser cut jigging
structures. The jigs were built into the parts and discarded after completion of the assembly. Laser cutting also
eliminates the need for fasteners which are replaced with finger joints.
IV. Stability and Control
The aircraft flight dynamics were characterized separately in the three main phases of flight (quad-copter, airplane,
and transition) using linear time invariant models. The control system is based on the concept of control scheduling.
Separate control laws will be running on the flight control system, and one will be selected to send outputs to the
actuators. The controller is selected based on the angle of the tilting rotors. This tilt angle determines the level of
coupling between motor thrust changes and yawing, rolling, and pitching moments. The steady vehicle forward flight
speed will also be highly dependent on tilt angle because it dictates the amount of thrust that is projected along the
horizontal and vertical directions.
The control system will be implemented on the Pixhawk flight control hardware using custom flight control
software based on the open source PX4 code. The Pixhawk includes 3-axis gyros and accelerometers, pitot airspeed
sensor, GPS, and magnetometer. While highly developed, the PX4 software is both more complex and less flexible
than custom control systems. Some portions of the PX4 software will be kept, such as the Extended Kalman Filter
used to stabilize sensor outputs and telemetry system, while the actual vehicle control code will be customized.
8
A. Aircraft
The simplest flight condition to consider
is in standard forward flight. AVL was used
to determine static stability derivatives. This
information was combined with the mass
moments of inertia found in CAD to
determine dynamic stability using the full 6
degree of freedom linearized equations found
in Philips’ Mechanics of Flight[8]. The
aircraft control system will include closed-
loop stabilization, but the airplane was
designed to be stable in cruise with an open
loop. This allows the pilot to take manual
control of the aircraft while testing
experimental control architectures. The
aircraft stability coefficients were
determined at a steady level cruise of 50 ft/s.
The static stability analysis verified that the
aircraft is statically stable with 9.1% 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 as
shown in Table 3. The three oscillatory dynamic modes are shown. The roll mode is highly overdamped and the spiral
mode is unstable with a 15 second time to double that the pilot can easily control.
B. Quad-copter
In hover, the flight loads will be evenly split between the four motors. The airplane’s aerodynamic surfaces are
modeled as flat plates that slew through the air and damp motion as a quad-copter. The rotors also provide some
damping, but rigid rotors in hover are statically unstable. Control will be provided by differential thrust since the
airplane control surfaces are ineffective in hover. The change in thrust from any one motor will result in roll, pitch
and yaw moments a quad-copter uses for control. These highly coupled controls are handled well by a computer
system for mixing. A simple proportional, integral, derivative (PID) controller is added to the flight computer to
artificially stabilize the quad-copter. PID control is well understood and the team has experience tuning PID controllers
for quad-copters.
C. Transition Mode
The vehicle dynamics of the aircraft with tilted rotors is more complex than either the hovering flight or cruising
flight because there are air loads and the dynamics from the four quad-copter motors. The motor tilt will not be used
for stability control due to the slow actuation time. Effectively, the model in transitioning flight is the combination of
the existing dynamical models with additional consideration for the roll-yaw coupling that occurs with partially tilted
motors. The transition flight mode will use a Linear Quadratic Regulator. Tuning a PID controller is less practical in
transition because of the changing effectiveness of the control surfaces and the motor variable RPM control. The
accuracy of the aircraft model can be improved from the previous testing of the airplane mode and quad-copter mode.
The aerodynamic parameters can be updated from the forward flight testing, and the thrust parameters can be updated
from the quad-copter testing. The controller will be designed to fly continuously at any motor tilt angle.
V. Testing
A. Component Testing and Verification
To ensure expected operational characteristics, testing was performed on each component and subsystem before
installation on the aircraft. All control surface and tilt mechanism servo actuation was verified by a servo tester as
well as the transmitter/receiver system. The electric propulsion system involved the most extensive subsystem tests.
Static thrust test data was collected from each of the motors to verify their performance. The motors produced 10%
less thrust in static testing than predicted which may result in a lower MTOW. Next, the structure of the propulsion
system was tested. The entire assembly was bolted, as it would be on the wing, onto a secure test stand upside down
and the motors were run at full RPM for one minute. The structure passed the stress test with motors in the vertical
flight position, so the rear motor was rotated to its horizontal flight position and run at full RPM for one minute.
Static Stability
Inputs
Wtotal(lbs) 10
CL 0.68
α (deg) 4.7
Stiffness Coefficients
Cmα (rad-1) -0.437
Clβ (rad-1) -0.140
Cnβ (rad-1) 0.061
Static Margin % Chord 9.1
Dynamic Stability
Mode Short-Period Phugoid Dutch Roll
Time to Half (s) 0.136 13.31 3.696
Damping Ratio 0.618 0.071 0.043
Undamped Freq. (s-1) 8.252 0.735 4.327
Damped Freq. (s-1) 6.485 0.733 4.323
Table 3: Static and dynamic stability
9
Figure 11: Completed Aircraft
Having passed this test, the tilt mechanism was tested through its range of motion with the rear motor spinning at full
RPM. All components passed their verification tests and were installed on the aircraft.
B. Flight Testing
Initial flight testing commenced after inspection of the structure, power systems, and electronic connections and
completion of subsystem demonstration in an operational capacity on the airframe. At this point, the research is
concurrent with the report. The first phase of flight testing involves flying the aircraft only as a quadrotor. During this
phase, the controller gains will be tuned until the aircraft hover and input control responses are crisp and controllable.
Flight data of current draw of the motors and RPM values will be recorded by the electronic speed controllers and
collected for model verification.
The next phase of flight testing will involve placing the aircraft in fixed wing mode and taking off from a runway.
The aircraft will be trimmed and data will be collected concerning the current draw and rpm as well as the range of
horizontal flight speeds from stall to maximum steady level speed. Air speeds will be measured with a pitot static
probe during the maneuvers. This data will be processed and used to assist with the control of the next flight test
maneuver, transition from vertical to horizontal flight. Transition will be performed at a high altitude and will be
controlled with an onboard flight control computer programmed to perform the maneuver. The pilot will maintain a
manual override ability to switch the vehicle into vertical flight mode for safe landing in the event of a failed transition.
The control system will be adjusted until a reliable transition from vertical flight to horizontal flight is achieved.
Following this, the same procedure will be performed for transitioning from horizontal flight to vertical flight and
tuning of the system will occur. The entire process is summarized in Figure 10.
Figure 10: Flight testing flow chart
C. Envelope Expansion and Data Collection
Once the aircraft has exhibited it can reliably transition back and forth between the different flight modes, various
data collection experiments will be conducted. One of the experiments includes visualizing the flow around lifting
surfaces, such as the wing and tail, by attaching yarn tufts. Other tests will determine the flight envelope by measuring
control inputs and the resulting vehicle response in the form of power consumption and vehicle motion.
D. Configurations
This research aircraft can be reconfigured with different tail, wing, and fuselage configurations. Initially, different
tail configurations will be tested, and the same procedures as outlined in the flight testing section will be followed for
trimming the aircraft and tuning the flight control system. Eventually, new power systems will be implemented and
tested using all the available attachment hardpoints in the wing.
VI. Conclusion
The team completed the design, manufacture, and flight test of the
vehicle in just over four months as a testament to the ability to rapidly
iterate and prototype a design. The goal of this research project was to
provide a vehicle through which the aerodynamic phenomenon of
transition can be characterized through flight testing. The process by
which this vehicle was created can be applied to any small-scale
configuration for initial testing before a full-scale prototype is built. Risks
are mitigated by the low cost of components for a small representative
vehicle, and modularity allows replacement or modification of
components. Additional transition corridor characterization flight testing
of other configurations can be performed through simple replacement of modular components. With this vehicle,
shown in Figure 11, the team hopes to collect flight test data to develop the actual operational understanding of this
transition aircraft in different flight conditions with various environmental factors. Further research into different
configurations using this strategy can save time and cost into the development of an urban air mobility (UAM) vehicle.
10
Appendix
11
Acknowledgments
The authors thank Dr. Brian German for his support as the faculty advisor for this research project.
References
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[2] 2015 Georgia Tech DBF Team, Buzz Killington, AIAA DBF website: www.aiaadbf.org/2015TopReports.aspx,
pp 54-55, accessed March 2018.
[3] Selig, M.S., R.W. Deters, G.K. Ananda, and J.B. Brandt, “UIUC Propeller Database”, University of Illinois at
Urbana-Champaign [online database], accessed 10/2015, UIUC website:
http://mselig.ae.illinois.edu/props/propDB.html.
[4] Drela, Mark and Harold Youngren, XFOIL website: web.mit.edu/drela/public/web/xfoil, accessed March 2018.
[5] McDonald, Rob et al., Open VSP website: www.openvsp.org, accessed March 2018.
[6] Hoerner, Dr. Sighard F., Fluid-Dynamic Drag, Bricktown, New Jersey, 1965.
[7] Drela, Mark and Harold Youngren, AVL website: web.mit.edu/drela/public/web/avl, accessed March 2018.
[8] Phillips, Warren F., Mechanics of Flight, 1st. ed., Hoboken, New Jersey, 2004, pp 647-674.