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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.

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

4.3.1.1 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

4.3.2.3 Lessons Learned ...................................................................................................................... 20

4.3.3 Prototype 3 ..................................................................................................................................... 21

4.3.3.1 Description .............................................................................................................................. 21

4.3.3.2 Results ..................................................................................................................................... 21

4.3.3.3 Lessons Learned ...................................................................................................................... 21

4.3.4 Prototype 4 ..................................................................................................................................... 21

4.3.4.1 Description .............................................................................................................................. 21

4.3.4.2 Results ..................................................................................................................................... 22

4.3.4.3 Lessons Learned ...................................................................................................................... 22

4.3.5 Prototype 5 ..................................................................................................................................... 22

4.3.5.1 Description .............................................................................................................................. 22

4.3.5.2 Results ..................................................................................................................................... 23

4.3.5.3 Lessons Learned ...................................................................................................................... 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

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

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

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

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

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.

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

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.

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

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.

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

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.

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

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.

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

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

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. 4.3.1.3 Lessons Learned

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.

Figure 9 : CAD Model of Printed Wind Tunnel Testing Airfoil

4.3.2.3 Lessons Learned 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.

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.

Figure 10 : Rotor Thrust Test Prototype

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 4.3.3.3 Lessons Learned 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.

Figure 11 : Motor Housing Original Prototype

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. 4.3.4.3 Lessons Learned

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.

Figure 12 : Rotor-Motor in Wing Mounting Assembly

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

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.

Figure 13 : Wing Prototype Assembly

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. 4.3.5.3 Lessons Learned

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.

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.

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

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.

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

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.

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

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.

11 Appendices 11.A Detailed Project Management

Figure 18 Fall Quarter Gannt Chart

Figure 19 : Updated Gantt chart depicting winter and spring term progress

Figure 20 : Budget for the overall project