2 Abstract generation of Mas rover will be li.docx

2
Abstract
generation of Mas rover will be limited by constraints of
Martian topography,
and thereby would benefit from knowing viable routes prior to
dispatch. In addition to the rover,
an aircraft is being sent with the Mars 2020 mission to prove
the merit of flight on Mars and
to attempt scouting traversable paths. A coaxial helicopter
design has been selected to meet the
needs of the mission without exceeding volumetric limitations
of interplanetary
transport. By opting for a rotocopter design, the scout can only
generate lift as a function of
propeller thrust and is unable to traverse more than six-hundred
meters due to the power necessary
to maintain lift in the low-density Martian atmosphere. Route
information recovered
from missions would be limited by such short-range flight. To
optimize mission range, a tilt-
rotor aircraft with collapsible fixed-wings has been devised.
Uniting short
takeoff/landing rotorcraft capabilities with the increased range
of fixed-wing airplanes creates an
optimal performance profile for an exploratory scout. Rather
than relying on propeller-thrust to
generate lift after takeoff, a tiltrotor aircraft will achieve high

forward velocity by rotating
the direction of thrust from vertical to forward oriented. As
forward velocity increases, a fixed
wing, designed for ultra-low Reynolds number conditions,
generates the lift required for cruise
flight. To combat volumetric constraints, the wing is
collapsible, reducing its footprint in
transit. the combination of lift generation methods has the
potential to advance the extent of
exploratory missions beyond what is currently possible,
accelerating timelines and saving money
simultaneously.
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Table of Contents
1 Executive Summary
...............................................................................................
............................. 6
2 Introduction
...............................................................................................
.......................................... 6
2.1 Motivation
...............................................................................................
...................................... 6
2.2 Problem Statement

...............................................................................................
......................... 7
2.3 Stakeholder needs
...............................................................................................
.......................... 7
2.4 State of the Art
...............................................................................................
............................... 8
2.5 Approach
................................................................................ ...............
.............................................. 8
3. Materials and Methods
...............................................................................................
.................... 9
3.1 Materials
...............................................................................................
.............................................. 9
3.1.1 Foam board insulation
...............................................................................................
................... 9
3.1.2 Poplar, Bass, and Balsa Woods
...............................................................................................
..... 9
3.1.3 Graphene Tape
......................................................................................... ......
............................ 10
3.1.4 Polylactic acid (PLA) 3d printing filament

................................................................................ 10
3.2 Equipment
...............................................................................................
.......................................... 10
3.2.1 Arduino Micro
...............................................................................................
............................ 10
3.2.2 Motors
...............................................................................................
......................................... 10
3.2.3 Servos
...............................................................................................
.......................................... 10
3.2.4 RC Controller
...............................................................................................
.............................. 10
3.2.5 Flight Controller
...............................................................................................
.......................... 11
3.3 Software
...............................................................................................
............................................. 11
3.3.1 Fusion 360
...............................................................................................
................................... 11
3.3.2 Autodesk Flow Design
...............................................................................................

................ 11
3.4 Manufacturing methods
...............................................................................................
..................... 11
3.4.1 Airfoil creation
...............................................................................................
............................ 11
3.4.2 Tilt rotor assembly
...............................................................................................
...................... 12
3.4.3 Collapsible wing
...............................................................................................
......................... 12
4 Results
...............................................................................................
...................................................... 12
4.1 Specifications, Constraints, Standards
...............................................................................................
... 12
4.2 Concepts
....................................................................................... ........
................................................. 13
4.2.1 Decision Matrices
...............................................................................................
............................... 14
4.3 Detailed Designs
...............................................................................................

.................................... 15
4.3.1 Prototype 1
...............................................................................................
...................................... 19
4
4.3.1.1 Description
...............................................................................................
............................... 19
4.3.1.2 Results
...............................................................................................
...................................... 20
4.3.1.3 Lessons Learned
...............................................................................................
....................... 20
4.3.2 Prototype 2
...............................................................................................
...................................... 20
4.3.2.1 Description
...............................................................................................
............................... 20
4.3.2.2 Results
...............................................................................................
...................................... 20

...............................................................................................
....................... 20
4.3.3 Prototype 3
...............................................................................................
...................................... 21
4.3.3.1 Description
...............................................................................................
............................... 21
4.3.3.2 Results
...............................................................................................
...................................... 21
...............................................................................................
....................... 21
4.3.4 Prototype 4
...............................................................................................
...................................... 21
4.3.4.1 Description
...............................................................................................
............................... 21
4.3.4.2 Results
...............................................................................................
...................................... 22
...............................................................................................
....................... 22

4.3.5 Prototype 5
...............................................................................................
...................................... 22
4.3.5.1 Description
...............................................................................................
............................... 22
4.3.5.2 Results
...............................................................................................
...................................... 23
...............................................................................................
....................... 23
4.4 Additional Analysis
...............................................................................................
............................... 23
5
Discussion...............................................................................
................................................................. 26
6 Context and Impact
...............................................................................................
................................ 27
6.1 Economic Analysis
...............................................................................................
............................ 27
6.2 Environmental Impact Analysis
...............................................................................................
......... 27

6.3 Social Impact Analysis
...............................................................................................
...................... 28
6.4 Ethical Analysis
...............................................................................................
................................. 28
7 Project Management Update
...............................................................................................
................. 28
7.1 Team organization
...............................................................................................
............................. 28
7.2 Schedule and milestones
...............................................................................................
.................... 28
7.3 Project Budget
...............................................................................................
.................................... 28
8. Summary and Conclusions
...............................................................................................
.................... 29
8.1 Project Reflection
..................................................................................... ..........
............................... 29
8.2 Senior Design During Covid-19
...............................................................................................
........ 29

8.3 Insights Gained During Project
...............................................................................................
.......... 29
8.4 Project Conclusions
...............................................................................................
........................... 29
5
9 Future
Work......................................................................................
.................................................. 30
10 References
...............................................................................................
............................................ 30
11 Appendices
...............................................................................................
...................................... 31
11.A Detailed Project Management
...............................................................................................
......... 31
List of Figures
Figure 1: Shear stress of Poplar wood vs Bass
wood.................................................................................. 10
Figure 2: Wing Design Decision Matrix

...............................................................................................
...... 15
Figure 3 Raf 6 Airfoil
...............................................................................................
.................................. 16
Figure 4 CL vs Alpha
...............................................................................................
................................... 16
Figure 5 CD vs Alpha
...............................................................................................
.................................. 17
Figure 6 Cl/CD vs Alpha
...............................................................................................
............................. 18
Figure 7 Power Required Vs Velocity for known flight
conditions ........................................................... 19
Figure 8 : Physical Weight 2-D Schematic
...............................................................................................
.. 19
Figure 9 : CAD Model of Printed Wind Tunnel Testing Airfoil
................................................................ 20
Figure 10 : Rotor Thrust Test Prototype
......................................................................................... ......
...... 21
Figure 11 : Motor Housing Original Prototype
...........................................................................................
22

Figure 12 : Rotor-Motor in Wing Mounting Assembly
.............................................................................. 22
Figure 13 : Wing Prototype Assembly
...............................................................................................
......... 23
Figure 14 Vertical Take-off ground effects
...............................................................................................
. 24
Figure 15 Fixe Wing Flow Conditions
...............................................................................................
........ 25
Figure 16 Fixed Wing Flow (2).
...............................................................................................
.................. 25
Figure 17 Propeller Downwash Test Assembly
.......................................................................................... 27
Figure 18 Fall Quarter Gannt Chart
...............................................................................................
............. 31
Figure 19 : Updated Gantt chart depicting winter and spring
term progress .............................................. 31
Figure 20 : Budget for the overall project
...............................................................................................
... 32

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1 Executive Summary
Objective
Develop a tilt rotor drone with foldable fixed wing in
order to optimize range of flight of the Martian air
reconnaissance vehicle.
Performance goals are achieving flight under the
conditions of the Martian surface, providing a greater
range of flight than that of the current solution,
Observing the terrain of mars from an aerial
perspective at high resolution and Vertical Takeoff
and Landing on uneven surfaces of Mars.
Technologies include high torque rotor; foldable
wing mechanism, Thin Raf-6 airfoil.
Approach
Maintaining flight at higher speeds and for longer
durations than purely thrust driven prototypes by
1. Achieving lift in ultra-low Reynolds number
conditions using Raf-6 airfoil
2. Tilting powerful multi-rotors in order
to change their direction of thrust capability
3. Creating a collapsible fixed wing, thereby
meeting the mandated footprint considered
feasible for transport by NASA.
Key Milestones

3/15 Wind Tunnel Testing
4/23 Truss Model Testing
4/27 Prototype Ready for Test Flight
4/30 Prototype Presentation
2 Introduction
2.1 Motivation
The quest to observe and document the red planet has been an
ongoing mission for the better part
of the past five decades. Beginning with the myth of lost
civilizations on its surface that fed popular culture
and science fiction, Mars has been focus of public interest. This
interest has persisted as researchers made
discoveries such as indications of water having once existed in
a liquid state on the planet at some point in
its history. The existence of liquid water on the Martian surface
is among one of the many aspects of Mars
that makes it a treasure trove of scientific insight. A planet
which once was able to sustain liquid water may
have also been able to sustain life in some form. Understanding
what has led to the change that Mars has
undertaken from its days of liquid water to its current state of
frigid barrenness has the potential to answer
questions about the formative years of our solar system as well
as the formative years of life on Earth.
While several potentially habitable Earth analogs have been
identified and numerous are hypothesized to
exist, the fact that Mars is s own solar system has made it the
most accessible specimen that we
believe may have had earth like environment. This makes
geological samples and high-quality data
recovered from the planet very valuable to for the progress of

scientific efforts. Martian exploration is also
a necessary steppingstone in human space travel, similarly to
the way in which the in which the lunar
landing was a milestone for human exploratory capability in the
. Mars is a probable destination
for a manned mission in the future and fully understanding the
planets resources and hazards is critical to
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the success of such an endeavor.
are going to shape exploratory missions of the future,
presumably to even more distant parts of our galaxy.
These are all the reasons for which billions of dollars have been
spent in the effort to reach Mars
and retrieve meaningful data on its environment and conditions.
Devices including satellites, landers and
rovers have been deployed in the past to image, collect
atmospheric and topographical data, and samples.
The Spirit and Curiosity rovers, delivered in 2004, being the
most recent of these deployments were sent to
collect high resolution images from the surface of the planet
and conduct field studies on surface samples
utilizing onboard geological laboratories. The next generation
of rover is planned to land on Mars in
February of 2021. Mars 2020 mission rover is the first to be
equipped with a drill to probe beneath
the surface of Mars in search of signs of life supporting
conditions. It is also the first rover to be
accompanied by an aerial scout. This scout is a solar powered,
coaxial helicopter and is slated to be the first

aircraft to fly on another planet.
The mobility of land rovers, like the 2020 rover, is fundamental
to exploratory efforts. Although
they are designed with considerable regard to this need, there
are still complications created by the highly
varied and ultimately unknown geography of the Martian
surface. Land surveyance vehicles are subjected
to movement and directional constraints due to topography,
thereby necessitating a means to assess viable
routes prior to rover dispatch to optimize time spent moving
between sites of interest. Unmanned
rotocopters such as the aforementioned helicopter, have been
devised for scouting operations to allow for
aerial assessment of potential routs for the rover. Aerial
vehicles provide similar resolution images to those
collected by rovers but can more easily access remote areas and
can traverse distances much more quickly
than a land rover. The intention of including this helicopter to
the Mars 2020 mission is to prove that flight
in a Martian atmosphere is possible and a viable option for
future missions all while the aircraft fills a
valuable scouting role in the 2020 mission.
2.2 Problem Statement
to achieve, is not without its own limitations. Achieving flight
on Mars presents complications due to the
atmospheric density relative to that on Earth.
Reduced atmospheric conditions require significantly greater
power to produce sufficient thrust for the
generation of lift than comparable terrestrial rotorcrafts. Low
density atmosphere contributes to ultra-low
Reynolds numbers, compared to Reynolds numbers experienced
by conventional aircrafts flying in standard
Earth altitudes. Early attempts to combat atmospheric

complications resulted in fixed wing drone prototypes
but were ultimately discarded in favor of the more
maneuverable, transportable co-axial helicopter design
This increased power requirement results in a diminished
battery life of mere minutes per reconnaissance mission. While
this is sufficient for the current scope of the
need to be efficient utilizing power
to produce lift in order to extend their range. Increasing the
range of unmanned aerial vehicles on
Mars while maintaining maneuverability and
takeoff/landing/hover capabilities of rotorcraft has the
potential to advance extent and capabilities of exploratory
missions beyond what is currently possible with
landbound rovers, accelerating timelines and saving money
simultaneously.
2.3 Stakeholder needs
To produce the optimal aircraft for reconnaissance flights on
Mars would be very valuable to the
scientific community. Organizations at the forefront of
extraterrestrial exploration such as NASA, ESA and
ISRO, and companies with vested interest in the exploration of
space such as Blue Origin, SpaceX and
Boeing would all be stakeholders in the advancement of
extraterrestrial aerial surveying vehicles. In order
8
to successfully fulfil the roll of an exploratory scout, a product
needs to meet both the performance profile

required of scouting activities and the physical/spatial demands
of a deployment mission. To fulfill the role
of scout on a Martian mission a product must be capable of the
following functions:
1. Achieving flight under the conditions of the Martian surface
2. Providing a greater range of flight than that of the current
solution
3. Observing the terrain of mars from an aerial perspective at
high resolution
4. Maneuvering capabilities sufficient for observation of
specified areas
5. Vertical Takeoff and Landing on uneven surfaces of Mars
To meet the physical/spatial demands of transportation and
deployment the following needs must be met
by the product:
6. Occupying a similar spatial envelope to that of the current
solution in transit
7. Having similar mass to that of the current solution
A product would need to fulfill all the requirements listed above
in order to be considered successful.
2.4 State of the Art
The state of the art for this project has been based on Mars
2020 scout helicopter. Current
mission specifications have been used to demonstrate the
acceptable metrics for the Mars Extended Range
Scout.
Chart 1: Mars Helicopter Up to Date Specifications
Mass 1.8 kilograms

Weight 4 pounds on Earth; 1.5 pounds on Mars
Width Total length of rotors: ~4 feet (~1.2 meters) tip to
tip
Power Solar panel charges Lithium-ion batteries,
providing enough energy for one 90-second flight
per Martian day (~350 Watts of average power
during flight)
Blade span Just under 4 feet (1.2 meters)
Flight range Up to 980 feet (300 meters)
Flight altitude Up to 15 feet (5 meters)
Flight environment Thin atmosphere, less than 1% as dense as
Earth's
2.5 Approach
To optimize range of flight of the Martian air reconnaissance
vehicle in the future, a tilt rotor drone
with collapsible fixed wing has been devised. The tiltrotor
aircraft originally was created from the demand
for short takeoff and landing / maneuverability of helicopters
coupled with the increased range of fixed
wing airplanes. Taking advantage of these principles when
applied to an unmanned aerial vehicle creates
an optimal performance profile for an exploratory scout. This is
made possible by making use of a fixed
wing designed to achieve lift in ultra-low Reynolds number
conditions, coupled with powerful multi-rotors
capable of changing direction of thrust. This combination of lift
generation methods allows the tilt rotor
design to be capable of maintaining flight at higher speeds and
for longer durations than the purely rotor

wing design has been created to be collapsible, thereby meeting
the mandated footprint considered feasible
for transport by NASA. Increasing the range of a UAV on Mars
while maintaining maneuverability and
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takeoff/landing/hover capabilities has the potential to advance
extent and capabilities of exploratory
missions beyond what is currently possible with landbound
rovers, accelerating timelines and saving money
simultaneously.
3. Materials and Methods
To accomplish the goal of creating the Martian VTOL
prototype, the engineering team first
identified necessary traits for flight as a function of physical
design. Though contradictory, it was
paramount that the body and wings be both light weight and
strong, thereby allowing airfoil generated lift
at the lowest velocity possible while still maintaining aircraft
control during transformation. Further, it was
necessary that the tilt rotor component of the prototype was
capable of actuating at the midpoint of each
wing, thereby physically allowing the combination of an
expandable wing and tilt rotor design to work in
tandem. Since it was necessary to have actuation at the mid-
section of each wing, the structural integrity of
the wing had to be increased, thereby necessitating an analysis
into lightweight high tensile strength wooden
materials. Further, since the components used in design were so

niche to the Martian VTOL prototype, it
was necessary to 3d print physical housings at the expense of
weight to ensure a streamline functional
design. Having assessed that the prototype must be as light
weight as possible while still structurally sound,
the following materials have been used:
3.1 Materials
3.1.1 Foam board insulation
Used in creation of both inboard and outboard section of airfoil,
as well as the creation of the
fuselage. Insulation board demonstrated a good candidate due to
its ability to maintain semblance of
structural rigidity in the face of expected loading conditions
generated in flight while simultaneously being
light weight. Further, foam board insulation proved fairly easy
to physically manipulate, thus allowing
multiple rapid prototypes of disparate airfoil concepts.
3.1.2 Poplar, Bass, and Balsa Woods
To reinforce both inboard and outboard sections of the
collapsible wing, it was necessary to embed
wooden dowels into the under-side leading edge and trailing
edge of each wing section. The embedded
dowels acted in two phase support, both increasing structural
rigidity, and allowing fixation points from
inboard wing sections to fuselage. In determining which wood
to use, tensile and shear stress graphs were
used to determine average elastic failure as a function of
loading condition. It was determined through
research that while poplar wood is heavier than both bass and
balsa wood, the increase in mechanical
properties warranted use in reinforcing the leading section of
each wing. Bass wood was used to reinforce

the trailing wing length, as it acted as a middle ground for both
weight and tensile strength between poplar
wood and balsa wood. Balsa wood, the lightest and weakest of
the structural woods used, was intermittently
embedded into airfoil sections to reinforce foam areas heavily
bored out to accommodate electronic
components.
10
Figure 1: Shear stress of Poplar wood vs Bass wood
3.1.3 Graphene Tape
To further assuage the application of shear stress away from the
foam board and into the structural
dowels, anisotropic graphene was applied to the embedded
wood to react shear more favorably.
3.1.4 Polylactic acid (PLA) 3d printing filament
Since many of the housing components for motor rotation had to
be made custom, lightweight PLA
filament was used as an easily replicable light weight testing
means. PLA levers were created to interface
with connecting surfaces of both the motor actuator on each
wing, and the motors themselves. Further, PLA
printed prototypes of airfoils and rotors were created for
experimental testing involving dimensional
analysis.
3.2 Equipment

3.2.1 Arduino Micro
The Arduino was used to aid in the prototyping and building
process. Due to the nature of rapid
code prototyping made possible through use of Arduino, the
engineering team was able to test the various
electronics to make sure they were functioning and interacting
as designed. Code was made to test the
motors with the props connected in order to get an accurate
reading and the ability to control the rotational
speed. Further, the same methodology was used to test servo
functionality, ensuring the servos were able
to articulate both smoothly and at an acceptable speed.
3.2.2 Motors
Two DYS D3542 1450KV motors were selected to generate the
thrust needed for the aircraft. These
motors are a bit large and this is due to the heavy weight of the
prototype and the fact only 2 motors are
used to lift the entire aircraft. These motors are easy to hookup
to the electronics since all the wires come
with connections and this eased in the construction.
3.2.3 Servos
Two different types of servos were selected to be used in the
aircraft. Two larger DS3225 25kg
servos are used to rotate the motors. These larger servos are
needed since the motors are large and heavy
and with the prop rotation a lot of force is needed cause rotation
during flight. The other type of servo used
was a 9g micro servo. These small servos were used for
articulating control surfaces since the specified
function did not require large torque.

3.2.4 RC Controller
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The controller used is a FS-i6. This controller was obtained
through Dr. Yousuff and it has all of
the required functions to operate the prototype aircraft. The
controller has multiple channels available for
input designation, thereby ensuring all various motors and
servos deemed necessary to achieve flight are
controllable.
3.2.5 Flight Controller
The flight controller used is the Matek F405. This controller
board is the brains of the aircraft,
so it had to have all the features abilities to control the aircraft
in both flight modes. This controller can run
open source code, thereby allowing the engineering team the
ability to design coding solutions that
smoothly articulate multiple control surfaces with just a single
movement of the RC controller.
3.3 Software
3.3.1 Fusion 360
CAD modeling software was used primarily to assess adjoining
fixtures of prototypes that scaled
considerably in size during prototyping. By determining fixed
area ratios as a function of aircraft stability
margins, it became possible to scale CAD models to better
visualize prototype designs prior to manufacture.

Assessing flaws in design and creating tools to better articulate
design concepts to staff and corollary
advisors was paramount in identifying flawed logic in
mechanical component design. Further, the created
CAD models were used as visual templates to create physical
models against after design had been finalized
3.3.2 Autodesk Flow Design
Autodesk Flow Design was used to assess laminar flow over the
entirety of the prototype to estimate
generated lift instead of conducting physical experiments. By
assessing lift as a function of simulation and
then validating the results through experimentation, the
simulation model was validated for accuracy,
thereby lending credence to the validity of using flow
simulation software to analyze lift generated on the
Martian surface.
3.4 Manufacturing methods
3.4.1 Airfoil creation
To create physical wing sections for use in the prototype, the
senior design team explored a variety
of methodologies that made use of foam board insulation. Early
attempts at wing creation resulted in folding
foamboard over two-dimensional airfoil cross sections that were
strategically placed along the length of
wing section for increased rigidity. To the benefit of this
method, the end result was indeed lighter in weight
than that of a solid foam core, but the structural integrity of the
wing was deemed incapable of maintaining
shape in flight at the velocity necessary to generate lift. The
second and most promising airfoil prototype
was created by meticulously sanding 1 x 12-inch foam
insulation sections into continuous airfoil shapes.

This process was achieved by 3d printing an airfoil template of
the desired length and tracing the shape on
both sides of the 12-inch foam section. The foam board was
then shaped and smoothed with high grit
sandpaper to create a uniform shape capable of generating a
laminar flow. Finer grit sandpaper was then
used to further diminish any asperities in material, and to shape
the wing section into the desired airfoil.
The end result of this method, while higher in weight than the
initial prototype, proved capable of
maintaining shape at the velocity necessary to generate lift
purely as a function of airfoil shape. Further, by
using a solid foam section, there was ample room to embed
actuators and necessary wiring into the wing
section by creating form fitting indentations with a Dremel.
Structural integrity to both inboard and
outboard airfoil sections were increased through imbedding
balsa, poplar, and bass wood strategically
throughout the underside of each wing section. Embedded wood
sections also acted as points of affixation
to the fuselage. To reduce the effects of shear stress on the wing
sections, anisotropic graphene tape was
adhered to embedded wooden sections.
12
3.4.2 Tilt rotor assembly
In order to create a prototype that was capable of both a
horizontal take off and fixed wing
transformation, it was necessary to strategically place both
motors at a maximum distance away from the

left and right of the aircraft butt line respectively, at point of
collapsible wing expansion. To accomplish
this task, actuators were embedded into the foamboard airfoils,
and reinforced with strategically placed
balsa wood to ensure that the counter moment generated through
actuation was properly resolved and not
traveling throughout the entire system. A 3d printed l-bracket
with connection points to the actuator was
then affixed to the assembly, followed by attaching the motor in
similar design. The finalized prototype
manages to rotate from zero degrees to ninety degrees without
structurally compromising the wing or
generating unresolved reactionary moments.
3.4.3 Collapsible wing
The collapsible wing was created by adjoining two shaped and
reinforced airfoil sections at their
respective ends through use of a small hinge. Embedded in the
onboard section of the airfoil lies a smaller
actuator with a lever arm capable of generating enough torque
to lift the collapsible airfoil section without
causing an unresolved reactionary moment. Embedded in the aft
(collapsible) section of the airfoil is a thing
sheet of poplar wood, designed to react the force of the lever
arm over a larger surface area so as not to
damage the foam section when the wing is extended.
4 Results
4.1 Specifications, Constraints, Standards
As the Martian reconnaissance drone only fulfils one extremely
niche function, creating a design that
demonstrated adherence with

e are still in flux due to the proposed mission
date, current mission specifications have been used to
demonstrate adherence with the proposed project,
thereby increasing the likelihood of having the senior design
project adopted for real world reconnaissance
applications.
When considering design parameters for the Martian drone, the
team first looked towards the rationale
behind NA -wing drone in lieu of multi rotor propeller copter
design.
From an aerodynamic standpoint the decision does not make
sense, as thrust is a reaction force created by
effect of pushing against the atmosphere. Since the Martian
atmosphere is considerably less dense than that
on earth, more thrust is required to generate lift on Mars. This
in turn directly necessitates an increase in
rotor revolution per minute (rpm) and thereby reduces available
battery substantially. Through wing design
analysis, see section 3.5) the team determined that fixed wing
flight is both more efficient and possible to
achieve on the Martian atmosphere, which lead the design team
to assume that the driving factor behind
to be a factor of form relative to available space. It stands to
reason that the
price of stowed equipment per unit volume is
specification became creating
a novel, foldable wing design that would not encroach upon the
planned footprint currently allotted by
NASA for their Martian reconnaissance drone prototype.
Having determined that the retractable wing design cannot
exceed currently proposed volumetric

limitations, the team then focused on optimizing energy saved
through fixed-wing flight. As currently
charged energy cell. Nas
13
drone achieves lift and scouts an accessible path that the
Martian land rover is capable of surmounting.
Having spent all available energy in just two minutes, the
Martian drone then lands and begins a lengthy
recharging process through use of solar cells. The
aforementioned steps are to be repeated indefinitely for
the duration of the mission. Noting that this order of events is
both tedious, which could lend to mechanical
failure due to excessive takeoff and landing, and extremely
laborious in time spent, the senior design team
optimized the process through mathematically proven velocity
and flight parameters using a fixed wing
approach.
While both the problem and potential solution made possible
through the retractable-wing Martian
drone are easy to define, competitive benchmarks against which
to compare methodology against is not
available. Mathematical models exist to demonstrate the
feasibility of flight in the Martian atmosphere, but
not practical applications have yet been used for this purpose.
Both ASME and AIAA standards denote best
practice for material and airfoil selection on earth but said
standards do not consider the complexities of the
Martian atmosphere, or the potential material damages unique to

Martian dust particulate or atmosphere
penetrating free radicals.
4.2 Concepts
As flight is predominantly a function of form and weight, the
senior design team first generated designs
that could meet the constrained volumetric and weight
limitations described as stakeholder needs in section
2.3. To meet this challenge, all designs considered were
required to meet the following specifications:
Designs must demonstrate a well-formed wing that
demonstrates mechanism to collapse.
complex to assure proposed methods could indeed demonstrate
the correct airfoil shape. Achieving flight
in the Martian atmosphere requires flight in Ultra-low Reynolds
number conditions. In stark contrast to
conditions found on earth, proposed Reynolds number
contingent upon fixed conditions resulted in a value
of roughly 8,300. While this number may seem insignificant, the
low value mandates use of a flat-plate
airfoil to achieve lift. Because flight in ultra-low Reynolds
number conditions requires a very specific shape
to achieve lift, all proposed designs needed to successfully
retract and expand with a high degree of accuracy
into the shape desired.
All designs must achieve velocity, and thereby lift, without
modifications to motor design.
To increase the likelihood of having NASA garner interest in
the groups project, volumetric footprint,

weight, and power requirements used in the senior design
project are meant to mirror that which is currently
planned for Martian reconnaissance. Wing size and shape that is
less reliant on ultra-
flight conditions were discarded due to the radical change
Further, more laxed flight conditions found through increased
velocity, and increased energy expenditure
as a function thereof, stands to skew the purpose of the
reconnaissance drone in general. Mandating use of
motors with similar energy expenditure thereby fixes the
available velocity for cruise flight, and once more
references the importance of well-formed shape in collapsible
wing design.
Materials used in the design must contend with the Martian
atmosphere.
14
Due to a lower atmospheric density found on Mars, the
likelihood of encountering free radicals or other
damaging forms of energy are relatively higher than that found
on earth. Further, the temperature on the
Martian surface is much colder than that on earth, posing
complication to both propulsion systems,
retractable wing mechanisms, and materials themselves.
To meet these design criteria, multiple retractable
wing mechanisms were envisioned and
discarded after demonstrating functional complications. Designs

that did not demonstrate mathematical or
theoretical complications are as follows:
Concept 1 is a tilt rotor drone with a foamed wing design. From
a conceptual standpoint, a polyurethane
reaction in the presence of water releases CO2 as a byproduct.
The amount of CO2 gas released is
stoichiometrically proportional to the amount of water in the
system during reaction and stands to
dramatically change the density of the polyurethane. This
reaction in mind, it has been envisioned that the
reactants responsible for this polymerization would be held
within the aircraft during transit to mars in
separated canisters. Upon deployment to the Martian surface, a
mechanism would release the reactants,
thereby creating low density polyurethane wings with CO2
byproducts. The result of the reaction would
expand through malleable rubber wing negatives that are affixed
to the side of the drone, thereby creating
a set of wings on the Martian surface. Concept 1 is appealing
due to the nature of the polyurethane reaction
in the presence of water. Increasing water stands to decrease the
density of the result and vice versa. By
choosing concept one, the Martian aircraft could be custom
tuned per strength specifications to be just
strong enough to maintain shape, while light weight to assuage
the difficulties with generating lift in ultra-
consider
Concept 2 does away with the need for causing a reaction on the
Martian surface by making use of a
telescopic wing. By creating an interlocking series of
increasingly smaller airfoils, it becomes possible to
completely contract both wings towards the center line of the
aircraft while in transit or take off. As opposed

to concept 1, which creates a wing on the surface, concept two
allows for the wingspan to be expanded and
contracted an indefinite amount of times, allowing for
contraction during rotorcraft take off to minimize
drag.
Concept 3 acts in similar fashion to concept two by presenting a
prefabricated wing that is folded for
stowage to Mars. By making use of a series of hinges on the
fixed wing, the expanded wingspan would be
able to fold in and against the body of the aircraft for a more
compact footprint. As the wing would be
required tilt and rotate to collapse against itself, a prebuilt
solution for increasing flight mobility by having
tilt-axis wings is thereby a byproduct of concept 3.
4.2.1 Decision Matrices
When comparing the potential design solutions for the Mars
reconnaissance drone, Concept 1 is
appealing due to the nature of the polyurethane reaction in the
presence of water. Increasing water stands
to decrease the density of the result and vice versa. By choosing
concept one, the Martian aircraft could be
custom tuned per strength specifications to be just strong
enough to maintain shape, while remaining light
weight to assuage the difficulties with generating lift in ultra-
Difficulties arise when consider the very nature of the reaction.
A study would have to be undertaken to
15

assess the reaction rate coefficients under the presence of the
Martian environment and atmosphere. Further,
the shelf life of polymer retarders or inhibitors may not meet
the required amount of time necessary to reach
mars. This places inordinate risk on the mission of the Mars
Drone, as improper deployment of the reaction
would result in catastrophic failure of the drone. If the wing
were created successfully, a study into what
material the outer sleeve should be made from would also have
to be assessed. Unlike on earth where most
thermoset polymers are near indestructible, the presence of free
radicals in space ultimately stands to make
the wing more brittle over time, thereby causing failure through
continued use.
Due to the inherent complications that arise when considering
the polyurethane wing design, it has
been decided that either a telescopic or foldable wing will best
fit the design parameters for the given
application. When considering a telescopic wing design
solution, the inherent shape of the telescoping
airfoils becomes an object of scrutiny. By function, telescoping
components must link together in smaller
iterative patterns to achieve the most compact size available
when retracted. Because the Martian Drone
requires flight in an ultra-low Reynold's number environment, it
becomes paramount to use a flat plate
airfoil capable of generating lift. The importance of shape in
design application implies a purely telescoping
wing may not be able to create the surfaces necessary to achieve
lift under known constraints.
Having assessed the difficulty associated with creating an
airfoil shape through a telescopic inner
mechanism, it has been determined that the most likely venue

for success would result from a foldable wing
that collapses under the drone body. This design inherently
presents complications due to necessitating both
a translational and rotational component of the fixed wing
design. While not inherently difficult, the control
schema of this concept design promises complications as rotors
responsible for generating thrust will be
situated on the polar lengths of the wings and would need to
account for translation in design. A physical
representation of the logic to use a foldable wing design can be
seen below.
Figure 2: Wing Design Decision Matrix
4.3 Detailed Designs
Having determined that a foldable wing offers the greatest
chance for project success, prototypes were
created to assess component balance, airfoil choice, propeller
choice, actuation mechanism, and assembly
mechanism. Prior to prototyping, the feasibility of generating
lift with the RAF 6 airfoil under Martian
conditions was analyzed. This was completed to assess whether
a telescopic wing was feasible, or if the
physical discontinuities in airfoil design would stand to hinder
lift generated.
16
To assess whether fixed wing flight was possible in the Martian

atmosphere with a foldable wing,
Given
that Reynold's number is a function of flight velocity, dynamic
viscosity, chord width and atmospheric
density, realistic assumptions were made relating motor RPM
with expected velocity output to determine
ation 1.
It was ultimately determined that to achieve lift, a velocity of
27 meters per second would have to be
extraordinarily low, but with a fixed footprint parameter
governing the width of the chord, it was not
possible to increase. Having determined that flight would take
place in ULR (Ultra-
conditions, the design team set out to find an airfoil capable of
generating lift at 8320. Assessing the lift
parameters of an airfoil is done exclusively through
experimental data and best trend fitting, and the online
application XFoil was used to find airfoil data that fit our
application. It was determined that one of few
airfoils with widely available data in ULR environments was the
RAF 6 airfoil, as seen below.
Figure 3 Raf 6 Airfoil
Experimentally derived values for coefficient of lift and
coefficient of drag were then plotted as a function
of the tilt angle of the wing, shown as alpha.

Figure 4 CL vs Alpha
17
Figure 5 CD vs Alpha
As is shown, the coefficient of lift and coefficient of drag at the
required flight parameter of 27 m/s shows
optimum lift at an angle of 6-8.5 degrees and minimized drag at
an angle of 4 degrees. For the purposes of
the design project, fixed wing flight is optimized by optimizing
Cl/Cd, thereby covering the greatest
distance for a fixed battery life.
By plotting optimized flight requirements, as shown below, it
was found that the greatest flight distances
could be achieved by using a dihedral wing angle of four to 5
degrees.
18

Figure 6 Cl/CD vs Alpha
Having determined the velocity requirements to achieve flight
with the RAF 6 airfoil, and the optimum
angle of attack in wing design to achieve optimum flight
conditions, coefficient of lift was calculated in
step sizes of .1 degree to assess lift from 4 to 6 degrees. Lift
was then calculated as a function of coefficient
of lift, angle of attack, wing length, wing chord, velocity, and
atmospheric density. As wing length was not
yet considered in equations, an iterative approach was taken to
assess lift generated relative to small changes
in angle of attack relative to wing size. Compromise was found
at an angle of five degrees with a total
wingspan of 2 meters on length. Having fixed necessary
parameters to achieve flight with a known weight
and known airfoil, design considerations turn to assessing the
feasibility of creating components that meet
mathematically derived equations of flight.
Having calculated wing parameters necessary to achieve lift,
power required to generate the necessary speed
of 27 m/s was calculated and stands to demonstrate the cyclic
systems of checks and balances intrinsic to
aircraft design. Given the power required is a function of both
thrust and velocity, a stepwise of 0.1 m/s
was used to assess power requirements through known values
found during wing design.
Equation 3: Power Required
Further, the stall velocity or velocity that must be overcome to
generate lift was calculated as a function of
wing dimension and ultimate value of generated lift.

Equation 4: Stall
Velocity
19
Figure 7 Power Required Vs Velocity for known flight
conditions
As is seen, lift in a fixed wing approach begins at a speed of 24
m/s and requires reasonable power
requirements seen in comparably sized drones to the design
project. The power required to reach speeds
greater than 30 m/s are unrealistic given the weight constraints
of the system, but the earlier design
constraints of flight at 27 m/s is both realistic and
currently proposed pure rotorcraft design. Having proved
feasibility of project using small angle
approximation and determined necessary values to obtain as a
function of component selection, prototyping
was able to begin.
4.3.1 Prototype 1
4.3.1.1 Description
A plane's center of gravity, determined with precise
calculations, is a critical factor in guiding and
stabilizing the aircraft for a successful flight. Based on the
derived Static Margin, the center gravity was
determined to be located 5.6inch from the leading edge of the

prototype. To maintain the center gravity at
the exact location, the moment of force or torque that results
from an object's weight acting through an arc
must be centered on the zero point of the reference datum
distance. To this end, physical components have
been placed within the prototype following the schematic shown
in figure 8 to ensure proper cg to static
margin ratio.
Figure 8 : Physical Weight 2-D Schematic
20
4.3.1.2 Results
Due to the space and size of the components, the heavier objects
must be installed as close as
possible to the Cg (Center of gravity). Therefore, the battery of
the aircraft was installed right after the
tailing edge due to acting as the largest and heaviest component
in this design. The rotors and servo motors
are installed at the wing, which is located at the Cg. This can
also reduce the impact of heavy object affecting
the Cg shift. The lightest objects include the control board and
Arduino and are located at the leading edge
of the aircraft model.
The final product will need to account for minute changes to the
weight distribution as a factor of
the inconsistent density that the prototyping materials are

expected to have. The positioning of components
will have to be variable to allow the user to adjust the Cg
manually to adapt for the minute changes to the
interchangeable parts that comes with the manufacturing
process.
4.3.2 Prototype 2
4.3.2.1 Description
An airfoil prototype was created with the intention of
practically testing the theoretical aerodynamic
capabilities of the RAF-6. The prototype was designed to be a
scaled down version of the airfoil. The size
was geometrically scaled such that the low Reynolds number of
the Martian atmosphere at our intended
cruise speed within a wind tunnel. The prototype was 3D
printed and set at the designed angle of attack. To
measure the pressure differential about the top and bottom of
the airfoil, pressure taps were designed into
the print.
4.3.2.2 Results
tunnel has recently been very limited because of a misplaced
component.
The team has assisted the Drexel faculty in procuring the
necessary component however the process has
been slow, therefore testing has been limited to virtual fluid
dynamic modeling. This modeling showed that
by simulating the conditions of the Martian atmosphere, the lift
and drag produced are similar to what was
expected from the theoretical values, however practical testing
is still valuable tool for confirmation of
functionality because of how idealized the fluid dynamic model
is.

In an ideal case the performance of the fixed wing should be as
expected from specifications of the
RAF-6 airfoil. In order to confirm this however, the practical
testing will still have to take place to account
for surface roughness and manufacturing errors inherent to the
manufacturing limitations imposed on the
engineering team.
21
4.3.3 Prototype 3
4.3.3.1 Description
A prototype was also made to test the thrust achievable from
our rotor and motor assembly. This
prototype setup included a motor controller and potentiometer
that when connected to the motor allowed
for variable speed control. The assembly was affixed to a scale
reading the thrust as negative weight on the
system and an optical tachometer was used in tandem to read
rotor RPM values. This setup allows for a
relationship between signal sent to the motor, rotations per
minute and thrust output to be established.

4.3.3.2 Results
The rotor prototype testing was complicated by an unreliable
potentiometer connection and a failure of a
motor controller component. What information we were able to
establish from this testing was that the
motor is very sensitive to the input voltage and was
overpowered for our design at the time
This Prototype testing resulted in a better understanding of how
we need to control this mechanism such
that the resulting power is appropriate for the sizing of the
aircraft. As such the earthbound model was
increased in scale to make better use of the power that was
available.
4.3.4 Prototype 4
4.3.4.1 Description
In order to achieve the rotation of thrust necessary to produce
the desired flight profile, a
mechanism for rotating the motor and rotor assembly had to be
devised. The initial attempt at this was a
housing for the motor that would sit within the wing structure
on either side of the aircraft, both of which
would connect to a single high-torque servo at the center of the
fuselage by way of a shaft with mating ends.
One of these can be seen in Figure 11.
22

4.3.4.2 Results
The reason for designing the mechanism in this way was to save
on the weight that the mechanism
would add to the aircraft. Using a housing, we would be able to
rotate the rotor-motor assembly at its center
of mass, reducing the amount of torque required and therefore
the size and weight of the servo required. By
affixing the servo at the center of the fuselage, the same servo
could also be used to rotate the assembly for
each wing. This was successful for the design however it forced
the wing geometry to fit around the housing,
causing a discontinuity in the wing and leaving little room for
other components to be mounted at critical
points in the wing.
Embedding two servos into the wings themselves rather than the
fuselage, each with slightly higher
torque requirement than the original design allows, thereby
allowing the rotor to rotate about the continuous
wing construction without causing clearance issues. Doing so
also minimally impacts the weight of the
design because the larger quantity of ABS material and
fasteners that are made unnecessary when mounting
within the wing surface. Using a second servo also allows
individual thrust vector rotation which is
advantageous for control of z axis rotation when hovering. The
next iteration of the mounting assembly
which implements the lessons learned from the original design
can be seen in Figure 12.

4.3.5 Prototype 5
4.3.5.1 Description
The final prototype produced was a complete wing assembly.
This included foam airfoils, rotor
connections and folding wing component. Several airfoils, each
one quarter of the total wingspan of the
designed product were produced. This was done because in
subsequent testing it is expected that some of
23
these will be lost to failure. Then support structures were added
to these airfoils and electrical components
were embedded into the foam. this prototype can be seen in
Figure 13.
4.3.5.2 Results
This prototype was successful in proving the fit and function of
the tiltrotor components, it also
allowed the team to effectively produce airfoil shapes from the
foam insulation material at low cost. It also
made apparent the complications of the folding wing which was
designed to be actuated on a hinged joint
by a low torque servo and stiff wire assembly that was
unsuccessful due to an unexpected limited range of

motion in the joint.
The manufacturability of the prototype is very reasonable and
producing replacement parts for
testing is very achievable. It was also determined that the
tiltrotor mechanism devised after the first
l. The sizing
of the hinge used as well as the torque specification of the servo
used to actuate the folding wing will need
to be increased for the next build to allow for seamless rotation.
A specialized hinge is being created to
allow the servo to directly actuate the hinge rather than using a
rigid connection.
4.4 Additional Analysis
At the end of the winter quarter, the senior design team had
manufactured all of the requisite
components to begin experimental wind tunnel testing to
validate analytically derived values for lift and
drag. Unfortunately, due to global affairs the wind tunnel
testing was canceled, and the data needed to
prove the original calculations correct was not acquired.
Without the ability to collect experimental data
from physical means, A 3D model of the proposed Martian
aircraft was created and ran through
simulation software designed to replace the wind tunnel. With
the software a full model was able to be
tested instead of in parts like the wind tunnel models would
have been. This gave more accurate numbers
for the full model including the approximate lift generated and
the drag coefficient in both the airplane
and helicopter flight modes, though it must be stated that

assessing values of lift from a computational
model considers perfectly smooth surfaces, as well as little to
no gap between foldable wing sections. In
practice, ensuring either of these parameters has proven more
than difficult than previously considered,
and once more mandates a reliance on construction assumptions
that do not fully capture the physically
created prototype.
24
Having finalized the CAD model to properly articulate from a
vertical take-off to fixed-wing flight
orientation, the design was uploaded into AutoCAD flow
design, and known values for Martian
atmosphere, and calculated velocity as a function of motor and
rotor design were assigned as constants.
Given the limitations of the simulation software, it was not
feasible to fully recreate a vertical take-off
to fixed wing transition, so the analysis was broken into two
distinct sections. In the first section of
analysis, the model was uploaded in a vertical take-off
orientation, and was used to assess the ground
effects of vertical take-off. The resultant data was used to
determine the induced vertices of the aircraft
when close to the surface, and ultimately validated the aircrafts
ability to remain level during liftoff when
subjected to its own reactionary forces, as seen in figure 14.
Figure 14 Vertical Take-off ground effects

Having proved the aircrafts ability to take off in a vertical
orientation, the senior design team moved
to assess the expected lift and drag of the entire vehicle when in
a fixed-wing orientation. For the
purposes of this analysis, the fixed-wing orientation of the
design was uploaded into AutoCAD flow
design and has been analyzed to assume a successful transition
stage. Using the pressure gradient as
shown in figures 15 and 16, the drag coefficient given by the
simulation program; the original design
specifications could be checked to make sure the aircraft design
will fly under the original design
conditions. From analysis, it was determined that the velocity
that would have to be met to achieve flight
purely as a function of the airfoil was slightly greater than
expected, but still within an acceptable range
of deviation. Such as dsicrepency was however expected, as all
previous calculations were conducted on
airfoil alone, and did not take into account contributions of the
finalized aircraft.
25
Figure 15 Fixe Wing Flow Conditions
Figure 16 Fixed Wing Flow (2).

From analysis, it was ultimately determined that vertical take-
off and fixed wing flight was possible
in the Martian atmosphere with the current design. Further
analysis would include finding a balance
between increased motor and rotor size to achieve the slightly
higher velocity necessary to generate lift,
while balancing induced ground effects as a function of larger
rotors that are still controllable. Further,
the transition stage between vertical take-off and fixed wing
flight has gone unaddressed due to an
26
inability to find software that can simulate such advanced flow
simulations. It can only be surmised that
as the rotors increase in size, the transition will become harder
to achieve smoothly, but such an iterative
design process can only truly be completed through physical
experimentation.
5 Discussion
In assessing the totality of accrued design and test data during
senior design, it becomes apparent
that some avenues of analysis proved more fruitful than others,
and that the assumptions made during
preliminary calculations are subject to further scrutiny. At
conception, the Martian VTOL design called for
lift generation purely as a function of airfoil design, while
maintaining as small of a volumetric footprint as

possible. Initial calculations undertaken to assess project
feasibility made use of a small angle assumption
to assess lift and drag in leu of experimental testing in a wind
tunnel. While a good starting place, an absence
of physical wind tunnel data required the team to continue to
make use of a small angle assumption in
assessing lift throughout the length of the exercise, only
allowing for further refinement by non-dimensional
computer analysis undertaken after the winter quarter. While the
statements used to assess lift as a function
of small angle assumption are well founded, it does not go
without notice that the entirety of the assembly
design, both physical dimensions and mass, have been assessed
under calculated and not experimental
values of lift. The ramifications of this assumption therefore
assume perfect design of airfoil with little to
no asperities, perfectly level wing fixation to fuselage body, and
smooth laminar flow left unhindered by
gaps in the foldable wing sections. Without experimental means
to assess values of lift as a function of
unique design, these criteria will have to be considered met to
validate calculated and computer simulated
values demonstrating proof of concept. The inclusion of
physically tested data may very well have further
refined the proposed aircraft and airfoil shape, or possibly
mandated movement of propeller actuation
mechanism as a function of hindered flow.
While assumptions have been made in leu of experimental data,
it must be reiterated that the criteria
used to determine what assumptions are valid is sound, thereby
lending credence to the idea that the
proposed design would be able to achieve lift in the Martian
atmosphere given proposed design criteria.
Regardless of wind tunnel testing, the project by definition
required computer simulation and assumptions

to calculate lift values in order to assess lift conditions under
the Martian environment, which simply cannot
be replicated to perfection given financial limitations. Through
assumptions and assessing wing and rotor
components individually, it has been determined that the
proposed RAF-6 airfoil could be used to achieve
lift under ultra-low Reynolds number conditions as a function of
volumetric and component limitations
proposed for the Martian VTOL aircraft. Further, the proposed
propeller size, dimension, and location
proved effective in both analysis and simulation to achieve
vertical take-off. The sole component not
accounted for during senior design is the downwash effect of
rotor impeding airflow on the leading edge of
the aircraft during transition from vertical flight to fixed-wing
mode. Preparations for assessing downwash
effect have been completed large in part due to the assistance of
Mr. David Harding, and were expected to
start directly after senior design was cut short due to global
circumstances. A device was created to measure
downwash effect for the current design and can be seen below.
27
Figure 17 Propeller Downwash Test Assembly
If any lesson has truly been imparted during the duration of
senior design, it would have to be not
to underestimate what may seem like a simple physical task

before the actual attempt. The very nature of
the design project required seamless completion of task A
before attempting task B, simply because any
dimensional changes of one section greatly affected the
proposed feasibility of another. Countless foam
boards and sections of wooden dowel have been scrapped during
the design process due to wanting to make
sure step A was perfect before attempting the following design
build. In hindsight, the design tolerances
and stipulations imparted on initial aspects of the design were
too high to be successfully replicated given
the teams toolset and expertise, which in turn caused many
wasted attempts in an effort to meet self-imposed
stipulations on shape and angle. If the project was started again,
the level of perfection in physical design
that was attainable would first be assessed for each individual
design component, thereby allowing the
engineering team to assess which steps required absolute
perfection, and which steps could have a less strict
tolerances.
6 Context and Impact
6.1 Economic Analysis
The size and shape of the aircraft have multiple impacts on the
total cost of the mission. The choice to
develop a prototype model to demonstrate the critical functions
on earth will help to develop a final design
that could potentially be sent to Mars on the next mission. The
size of the aircraft has large impacts on the
overall cost of the mission due to the costs of sending a large
payload into space and transport it to other
planets. If the cost per unit is kept low, then possibly a fleet of
these aircraft could be sent on a mission and

cover an even larger area of the surface of Mars. The design
was centered around the overall weight and all
the individual components designed around this weight concern
6.2 Environmental Impact Analysis
The main environmental impacts of the proposed vehicle will
not be on Earth, but on Mars. The main
impact will be at the end of the useful life of the aircraft it will
end up abandoned on the surface of the
planet. It would be difficult to retrieve the aircraft at the end of
its useful life since there is no plan for a
return trip to earth. This is a similar method that is currently
used by other extraterrestrial surveying devices.
They are abandoned at the end of their useful life and just waste
away. Our model life is no different, but
28
if given more time we might be able to come up with a new
method to retrieve the aircraft at the end of its
useful life.
6.3 Social Impact Analysis
Some of the potential impacts would be the access to new data
about the surface of mars in a timely
manner. More data would be able to be collected at a closer
range than an orbiting satellite and more rapidly
than the current unmanned vehicles currently on Mars. This new
data can help with new research mission
proposals, and aid the current long-term missions taking place.
Finally, it could also help with planning for
future manned missions to Mars. Overall, this has shaped the

design of the aircraft because a long flight
time with high detail data calculation would allow better use of
time and money on this mission.
6.4 Ethical Analysis
little more freedom with the design. The major ethical concern
is sending this aircraft to Mars. There is
always a concern when sending devices from Earth to other
celestial bodies due to the potential for
contamination of these foreign areas. Further, sending a
scouting vehicle to the Martian surface may stand
to anger religious leaders who preach geocentric doctrines. That
said, these are risks the engineering design
team deems acceptable.
7 Project Management Update
7.1 Team organization
The team organization has not changed since conception, so
each team member has the same
roll. Each member serves a critical role in this senior design
group. Nate has been the main point of contact
for our advisor and most of the outside contacts, this has led
him to be the Team lead for the group. Daniel
has put us in contact with the materials department professors to
help us determine the materials needed for
the design. This has led him into the materials expert in our
group. Rex has helped with the aerodynamic
calculations and helped to determine the flight characteristics of
the aircraft, so he is the aerodynamic expert
of the group. Patrick has helped to crate and visualize the
design and find components that would benefit
the design, this has given him the position of designer.
Throughout the design process, online cloud folders
have been utilized to share information quickly and be

accessible to all members. This allows all members
to communicate even when we are not formally meeting
together. We also keep in contact through
messaging and setting up weekly internal meetings along with
weekly meetings with the team advisor.
7.2 Schedule and milestones
The original schedule was to create and test a full-scale model
that was able to fly and demonstrate
all critical parameters on earth. The schedule was changed
dramatically, and the only prototype was the
partial wing design created at the end of last term. Also, the
wind tunnel testing was cancelled so this term
was dedicated to running simulations on the 3D models to
substitute the wind tunnel and get the required
data. An updated Gantt chart is given in a table in Appendix A.
7.3 Project Budget
The original plan for the budget was to create a prototype
design to demonstrate the critical functions
of the aircraft on earth. The second part of the budget was to
purchase software that would allow simulation
of the Martian atmosphere and simulate the aircraft dynamics in
another atmosphere. Due to current events,
the required software was not acquired and used for testing. A
substitute fluid dynamics simulation was run
instead. The materials for the earth prototype were used to
create a partial prototype which showed some
of the key aspects of the wing design. The entire budget has
been provided in the Appendix.

29
8. Summary and Conclusions
8.1 Project Reflection
Upon reflection of the accomplishments that the engineering
team has accomplished over the course of
their senior year, it has been determined that the team
successfully proved the concept of fixed wing lift
under Martian atmospheric conditions, while maintaining a
volumetric footprint that remains within the
ts of the design project were completed in terms of a
physical deliverable, the concept has been proven both
analytically and through computer simulation,
leaving all but the final physical prototype complete. As the
scope of the project was to prove concept and
considering the physical earth-based deliverable a secondary
project to more accurately demonstrate
vertical take-off to fixed wing transition for an audience, the
final project is viewed as successful.
8.2 Senior Design During Covid-19
At the tail end of the winter quarter, and for the entirety of the
spring quarter, the United States deemed
it necessary to enact social distancing measures to prevent the
spread of Covid-19. As a function of decree,
the winter quarter that had been set aside by the engineering
design team to finalize a physical prototype
was forced to change, thereby relegating final assessment to be
completed through computer simulation.
While the senior design group set out to complete that which
they had intended, it would be a farse to state
that current lockdown effects had not directly hindered the

project. When assessing change in direction, the
most severe change manifested as the group tried to share
responsibilities to complete the project. Prior to
social isolation, the majority if not all group members were
present for every physical build or design
change. This in turn made sure that all group members were
heard and acted to catch many small
inconsistencies in thought process that would escape a single
group member. While working together to
finish the project over electronic means, it was far more
difficult to share partially completed solutions,
receive feedback, or spot logical inconsistencies. Difficulties
were assuaged by an increase in video-based
meetings and weekly group phone calls.
8.3 Insights Gained During Project
Throughout the senior design project, it was made increasingly
clear that
examinations used to present information in a test-taking
scenario did nothing but minimize the effort that
went into retrieving useful information. Throughout the senior
design project, the group had to assess their
own mechanical and material properties, assess whether or not a
mathematic approximation acted as an
acceptable means to quantify information, an
that did not already have a clearly defined solution. The critical
thinking gained as a function of completing
the process that is senior design will most likely act as an
invaluable tool in transferring academic
knowledge to practical knowledge in a working environment.
Further, the act of creating so many
prototypes made evident the complexity and nuance associated
with physical design creation. Too often in
engineering college has a physical design been relegated to a
perfect approximation of computational

design, thereby diminishing the art behind physically creating
something. Having gone through the process
of attempting to create prototypes that match design, the senior
design group has been left with an
appreciation for more technical work than previously
demonstrated. Understanding the limitations of tools
and tolerances will more likely than not aid the engineers in
their occupations, as it greatly reduces the
chance of presenting designs or tolerances that are outside of
the realm of feasibility.
8.4 Project Conclusions
Based on the prototype testing and the computer simulations for
the proposed aircraft design it was
found that flight should be achievable on Mars with the current
design. The lift generation by the 3d design
in the wind tunnel simulation was like the expected calculated
results from the initial design phase. The
30
prototype testing for the critical aspects of the wing design
proved to be a success and the folding wing
design can fold and unfold according to specifications. The
controllability of the proposed design was not
tested so further experiments would need to be conducted to see
if the control surfaces on the proposed
design will be enough to be able to fully control the aircraft in
both flight modes.
9 Future Work

The group had hoped to be able to complete a fully functioning
model at scale for use in an earth
atmosphere to prove the functionality of the proposed
mechanisms that would be integral to the flight of
the Martian model. Prototype 5 as described in earlier sections
incorporates the mechanisms that were yet
to be flight tested. Flight testing these mechanisms would allow
for the qualification of reliability of these
mechanisms under in flight conditions and how controllable
they would be remotely. Practical wind tunnel
testing is another required next step to corroborate the fluid
simulation data that has already been recorded
ent. After this information
has been collected, any adjustments to the geometry that are
necessary should be made and testing should
be reiterated. More complete flight simulations would be
appropriate to produce a robust control system
prior to a Martian environment prototype. After parameter
requirements have been satisfied, a full-scale
Martian prototype can be constructed for testing. While this
prototype would most likely not fly on earth
ity, however, it would still allow for progress in the form
of full-scale wind tunnel testing and the construction of
components. Once the design has been finalized,
the project could then shift to making the aircraft autonomous
as is the state of the Mars 2020 scout
helicopter.
10 References
[1] T. Greicius, "NASA's Mars Helicopter Attached to Mars
2020 Rover", NASA, 2019. [Online].

Available: https://www.nasa.gov/feature/jpl/nasas-mars-
helicopter-attached-to-mars-2020-rover.
[Accessed: 09- Nov- 2019].
[2] "Mars Exploration, Mars Rovers Information, Facts, News,
Photos -- National
Geographic", Nationalgeographic.com, 2019. [Online].
Available:
https://www.nationalgeographic.com/science/space/space-
exploration/mars-exploration-article/.
[Accessed: 09- Nov- 2019].
[3] Science Science
Buddies, 09-Aug-2017. [Online]. [Accessed: 05-Nov-2019].
[4] "RAF 6 AIRFOIL (raf6-il)", Airfoiltools.com, 2019.
[Online]. Available:
http://airfoiltools.com/airfoil/details?airfoil=raf6-il. [Accessed:
09- Nov- 2019].
[5] J. Anderson Jr, Aircraft Performance and Design. Boston,
Mass: McGraw-Hill Higher education,
2012.
31

11 Appendices
11.A Detailed Project Management
Figure 18 Fall Quarter Gannt Chart
Figure 19 : Updated Gantt chart depicting winter and spring
term progress
32
Figure 20 : Budget for the overall project
Mars Extended Range Scout (MERS) Close-Out Document
1
Purpose
The purpose of this document is to mark the completion of the
Mars Extended Range Scout (MERS) Project by identifying the
location of all assets, the disposition of materials, reconciling
the budget and identifying key analysis that have to be
completed.
1.1

Background
The quest to observe and document the red planet has been an
ongoing mission for the better part of the past five decades.
Beginning with the myth of lost civilizations on its surface that
fed popular culture and science fiction, Mars has been focus of
public interest. This interest has persisted as researchers made
discoveries such as indications of water having once existed in
a liquid state on the planet at some point in its history. The
existence of liquid water on the Martian surface is among one of
the many aspects of Mars that makes it a treasure trove of
scientific insight. A planet which once was able to sustain
liquid water may have also been able to sustain life in some
form. Understanding what has led to the change that Mars has
undertaken from its days of liquid water to its current state of
frigid barrenness has the potential to answer questions about the
formative years of our solar system as well as the formative
years of life on Earth. While several potentially habitable Earth
analogs have been identified and numerous are hypothesized to
exist, the fact that Mars is in Earth’s own solar system has made
it the most accessible specimen that we believe may have had
earth like environment. This makes geological samples and
high-quality data recovered from the planet very valuable to for
the progress of scientific efforts. Martian exploration is also a
necessary steppingstone in human space travel, similarly to the
way in which the in which the lunar landing was a milestone for
human exploratory capability in the late 1960’s.Mars is a
probable destination for a manned mission in the future and
fully understanding the planets resources and hazards is critical
to the success of such an endeavor. Today’s efforts to explore
Mars’ environment, terrain, geology and history are going to
shape exploratory missions of the future, presumably to even
more distant parts of our galaxy.
These are all the reasons for which billions of dollars have been
spent in the effort to reach Mars and retrieve meaningful data
on its environment and conditions. Devices including satellites,

landers and rovers have been deployed in the past to image,
collect atmospheric and topographical data, and samples. The
Spirit and Curiosity rovers, delivered in 2004, being the most
recent of these deployments were sent to collect high resolution
images from the surface of the planet and conduct field studies
on surface samples utilizing onboard geological laboratories.
The next generation of rover is planned to land on Mars in
February of 2021. NASA’s Mars 2020 mission rover is the first
to be equipped with a drill to probe beneath the surface of Mars
in search of signs of life supporting conditions. It is also the
first rover to be accompanied by an aerial scout. This scout is a
solar powered, coaxial helicopter and is slated to be the first
aircraft to fly on another planet.
The mobility of land rovers, like the 2020 rover, is fundamental
to exploratory efforts. Although they are designed with
considerable regard to this need, there are still complications
created by the highly varied and ultimately unknown geography
of the Martian surface. Land surveyance vehicles are subjected
to movement and directional constraints due to topography,
thereby necessitating a means to assess viable routes prior to
rover dispatch to optimize time spent moving between sites of
interest. Unmanned rotocopters such as the aforementioned
helicopter, have been devised for scouting operations to allow
for aerial assessment of potential routs for the rover. Aerial
vehicles provide similar resolution images to those collected by
rovers but can more easily access remote areas and can traverse
distances much more quickly than a land rover. The intention of
including this helicopter to the Mars 2020 mission is to prove
that flight in a Martian atmosphere is possible and a viable
option for future missions all while the aircraft fills a valuable
scouting role in the 2020 mission.
The 2020 Scout’s design, while revolutionary for the
groundbreaking accomplishment it is poised to achieve, is not
without its own limitations. Achieving flight on Mars presents

complications due to the planet’s carbon dioxide atmosphere
providing reduced atmospheric density and viscosity relative to
that on Earth. Reduced atmospheric conditions require
significantly greater power to produce sufficient thrust for the
generation of lift than comparable terrestrial rotorcrafts. Low
density and viscosity, CO2 rich atmosphere contributes to ultra-
low Reynolds numbers, compared to Reynolds numbers
experienced by conventional aircrafts flying in standard Earth
altitudes. Early attempts to combat atmospheric complications
resulted in fixed wing drone prototypes but were ultimately
discarded in favor of the more maneuverable, transportable co-
axial helicopter design NASA’s Jet propulsion Laboratory
produced. This increased power requirement results in a
diminished battery life of mere minutes per reconnaissance
mission. While this is sufficient for the current scope of the
aircraft’s mission, in order to be effective, future surveying
aircraft will need to be efficient utilizing power to produce lift
in order to extend their range. Increasing the range of unmanned
aerial vehicles (UAV’s) on Mars while maintaining
maneuverability and takeoff/landing/hover capabilities of
rotorcraft has the potential to advance extent and capabilities of
exploratory missions beyond what is currently possible with
landbound rovers, accelerating timelines and saving money
simultaneously.
2
Project Completion Work
2.1
Work Completed
Prototype 1
An airfoil prototype was created with the intention of
practically testing the theoretical aerodynamic capabilities of
the RAF-6. The prototype was designed to be a scaled down
version of the airfoil. The size was geometrically scaled such
that the low Reynolds number of the Martian atmosphere at our
intended cruise speed within a wind tunnel. The prototype was
3D printed and set at the designed angle of attack. To measure

the pressure differential about the top and bottom of the airfoil,
pressure taps were designed into the print.
Prototype 2
A prototype was also made to test the thrust achievable from
our rotor and motor assembly. This prototype setup included a
motor controller and potentiometer that when connected to the
motor allowed for variable speed control. The assembly was
affixed to a scale reading the thrust as negative weight on the
system and an optical tachometer was used in tandem to read
rotor RPM values. This setup allows for a relationship between
signal sent to the motor, rotations per minute and thrust output
to be established.
Prototype 3
In order to achieve the rotation of thrust necessary to produce
the desired flight profile, a mechanism for rotating the motor
and rotor assembly had to be devised. The initial attempt at this,
was a housing for the motor that would sit within the wing
structure on either side of the aircraft, both of which would
connect to a single high-torque servo at the center of the
fuselage by way of a shaft with mating ends. One of these can
be seen in Figure 5.
Embedding two servos into the wings themselves rather than the
fuselage, each with slightly higher torque requirement than the
original design allows, allows the rotor to rotate about the
continuous wing construction without causing clearance issues.

Doing so also minimally impact the weight of the design
because the larger quantity of ABS material and fasteners that
are made unnecessary when mounting within the wing surface
displace some of the added weight of the additional servo.
Using a second servo also allows individual thrust vector
rotation which is advantageous for control of z axis rotation
when hovering. The next iteration of the mounting assembly
which implements the lessons learned from the original design
can be seen in Figure 6.
Prototype 4
The final prototype produced was a complete wing assembly.
This included foam airfoils, rotor connections and the folding
wing component. Several airfoils, each one quarter of the total
wingspan of the designed product were produced. This was done
because in subsequent testing it is expected that some of these
will be lost to failure. Then support structures were added to
these airfoils and electrical components were embedded into the
foam. this prototype can be seen in Figure 7.
Financial Closure
The group received a grant worth $500 from the MEM
department through the Boeing Grant. The team has spent $340
through Drexel orders, leaving $160 in unused funds. A team
member also requires reimbursement for the purchase of a $40
sheet of plexiglass for the department’s wind tunnel.
Asset Transfer
Assets purchased include prototypes, raw materials that has not
been used, including several sheets of foam board insulation,
some wood struts and several electrical components. These are
all currently being stored at a member’s home and are intended

to be delivered to the Spacelab on the 4th floor of the main
building before the end of the term.
Information Management
All assets that have been used over the course of senior design
have been compiled in a shared drive, accessible by any
individual through the following link:
https://drexel0-
my.sharepoint.com/:f:/g/personal/pjs322_drexel_edu/En2nGYQ
Zy6BLqVPdms-
QWasBiq_8RyYv8S5YA0Cm5JO7JA?e=1tSN4H
The master file contains deliverables required by both Drexel
engineering, and those required to meet individual requests of
the faculty advisor. Further, excel spreadsheets and word
documents that have been used internally during analysis have
been annotated to better facilitate information transfer to an
audience less familiar to the project. The following table
denotes the contents of the shared folder, along with a brief
description of intended use.
Deliverable (File Name)
File Type (Electronic)
Description
Progress Report
Folder
Winter Term Deliverables
Poster Presentation
Folder
Winter Term Deliverables
Elevator Pitch Video
Folder
Video Link to Elevator pitch
Propeller Stuff
Folder
Analysis leading to propeller and motor sizing
Weekly Reports
Folder

Links to weekly updates throughout Senior Design
MEM_43_State_of_project
word document
state of project, Spring quarter
Fall_Proposal_Presentation
PowerPoint
Presentation, fall quarter
Component Weight Distribution
.dwg
Schematic of component weight distribution
Budget
excel
Gantt charts and budget information
formulas for Airplane design
excel
workbook denoting physical design limitations of aircraft
mem-43 abstract
word document
abstract requirement for winter quarter
Airfoil & Power Requirements
excel
workbook assessing power requirements as a function of airfoil
/ flight feasibility
Technical Progress Report
word document
document denoting work done over winter quarter
senior Design Gantt Chart
excel
Gantt Chart
on mars
.png
picture of cad model
Top view
.png
iso view

.png
MEM-43_Proposal_fallterm
PowerPoint
fall proposal of project
Aircraft parameters
word document
denotes geometric boundaries of model
Boeing Funding Request
word document
Request for Funding
2.2
Proposed final analysis
Before the end of the term, the team intends to produce a
finalized computer aided design model which is to be used to
run a computational fluid dynamic simulation. This simulation
will be used to establish the aerodynamic properties of the
assembled model in each stage of flight. The environment of the
simulation will emulate the properties of the Martian
atmosphere as reported by NASA, most importantly, the air
density, air viscosity and subsequently the Reynolds number of
the flow must be accurate to understand how the designed
geometry functions aerodynamically. The first stage of flight to
be tested is an airplane configuration where the rotors are
oriented forward, the wings are fully extended, and the
simulated airflow flows from the nose tip to the tail. The second
configuration that must be tested is vertical takeoff and landing.
In this configuration the wings are in their retracted position,
the rotors are rotated upward, and the airflow orientation is
from the rotor tips downward. The final configuration for
testing is a dynamic one which simulates the transition from a
vertical takeoff and landing configuration to an airplane
configuration. To do this, the rotors with move from a vertical
to forward orientation in sync with the expansion of the wings
and a coordinated shift in the airflow from top-down to nose-
tail. This combination of these events give insight into a crucial

portion of the flight. This analysis will complete the project to
the satisfaction of the team.
MEM-43 State of the Project
12

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