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

[1] Capable Computing, MotoCalc website: www.motocalc.com, accessed March 2018.

[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.