Final Report

Tiltrotor Gold Team

November 30, 2022

Executive Summary

Tiltrotor aircraft are a relatively new development, utilizing the lift capabilities of a

rotorcraft like a helicopter while maintaining the cruise flight of a fixed-wing aircraft.

The versatility this vehicle provides makes it very useful in challenging environments.

surveillance and mapping over rough and uneven terrain would be a difficult area for a fixed

wing aircraft to lift off, while a rotorcraft vehicle might not have the range needed to travel the

entire distance. A tiltrotor aircraft would provide Vertical Takeoff and Landing (VTOL)

capabilities with the required range of a system over this distance.

Delivery is another major use for tiltrotor aircraft, most notably in more remote locations.

In crisis situations where a smooth, preferred landing surface might not be possible,

VTOL-capable vehicles will be required. A tiltrotor aircraft could carry a small payload of

medical and survival items to its required destination and land on most surfaces.

This project utilized the tiltrotor capabilities and implemented them into an Unmanned

Air Vehicle (UAV) design. A few concepts were created to better understand the requirements a

tiltrotor would have to perform to lift vertically, and fly in cruise mode.

Contents

Executive Summary        2

Contents        3

List of Figures        4

List of Tables        5

Glossary        7

1. Introduction        8

2. Project Objectives & Scope        8

3. Assessment of Relevant Existing Technologies and Standards        8

4. Professional and Societal Considerations        8

5. System Requirements and Design Constraints        8

6. System Concept Development        8

7. Design Analysis        9

8. Final Design and Engineering Specifications        9

9. System Evaluation        9

10. Significant Accomplishments and Open Issues        9

11. Conclusions and Recommendations        10

References        11

Appendix A: Customer Requirements        12

Appendix B: System Evaluation Plan        12

Appendix C: User Manual        12

Appendix D: Cost Analysis and Manufacturability Analysis        12

Appendix E: Expense Report        12

Appendix F: List of Manuals and Other Documents        12

Appendix G: Design Competencies        12

List of Figures

List of Tables

Terms and Abbreviations

CAD - Computer-Aided Design

CBO - Community Based Organization

CFD - Computational Fluid Dynamics

ESC - Electronic Speed Controller

FAA - Federal Aviation Administration

FEA - Finite Element Analysis

FPV - First Person View

LiPo - Lithium-Polymer

PWM - Pulse Width Modulation

TRUST - The Recreational UAS Safety Test

UAS - Unmanned Aircraft Systems

UAV - Unmanned Air Vehicle

VTOL - Vertical Takeoff and Landing

MPa - Megapascals

BEC - Battery Eliminator Circuit

1. Introduction

Background

Figure 1: Bell V-22 Osprey [1]

For most products in today’s market, there exists an abundance of variety, with each option  bringing its own advantages and disadvantages to the table. Although this style of the economy is  ideal to fit a world of diverse needs, consumers often find themselves split between options that  force them to trade one preference for another. Major corporations have overcome this paradox by  introducing a product that upholds the “best of both worlds” ideology. This ideology is the guiding  value of how this team tackles the state-of-the-art challenge of designing an aircraft that  incorporates the strengths of both fixed and rotary-wing aircraft into a single vessel.  

Fixed-wing crafts lead the market in the payload to weight capacity, efficiency, and top end speed. Still, they have the downside of requiring great lengths for take-off and landing and are  often limited in maneuverability. In contrast, rotary-wing crafts have high maneuverability and  can land and take off in just about any terrain but come up short when considering traveling far  distances or carrying significant payloads. The culmination of their advantages results in what is  called a Tiltrotor Aircraft, where the body and wings of a fixed-wing craft are maintained, but with  the addition of rotary propulsion systems that can rotate about the wing axis and provide thrust in  both the vertical and horizontal direction. A practical example of this in industry, which the team  will closely model, is the Bell Boeing V-22 Osprey used in the U.S Marine Corps Aviation for amphibious assault transport of troops, equipment, and supplies from assault ships and land bases  [1].

The following report intends to provide insight into how the Tiltrotor Gold Team will  accomplish this hybrid objective, starting with recognizing the desire for this technology in the  field and what kind of interpretations are already in use. The sections that follow will then go into  the specifics of theoretical design concepts and how the process will be verified, accompanied by  relevant analysis data, and describe the team's finalized path for the execution of the project. This  report will end by reflecting on the significant findings and accomplishments thus far and the future  plans for this project.  

End User Need

Remotely controlled aircraft provide users with an immense opportunity to survey large or  hard-to-reach areas. However, fixed-wing aircraft and rotorcraft both come with their respective  disadvantages. Fixed-wing aircraft provide excellent flight range efficiency at the cost of takeoff  and landing versatility and maneuverability. Rotorcraft, in terms of benefits, are essentially an  inverse of fixed-wing aircraft, with low flight range efficiency but great takeoff and landing  versatility as well as high maneuverability. A tiltrotor aircraft is a rotorcraft and fixed-wing hybrid,  providing a rotorcraft’s maneuverability and takeoff and landing versatility, as well as containing  the flight range of a fixed-wing aircraft. A tiltrotor affixed with a camera to capture images and  videos from long ranges can be immensely versatile and valuable in many different applications.  

Food delivery services have recently begun experimenting with the concept of using  remotely controlled systems as their delivery method. A tiltrotor aircraft would prove to be the  perfect delivery method for such a service, as it provides ample flight range compared to a  quadcopter counterpart. Additionally, they can safely provide the necessary maneuverability to  drop off the payload. They can potentially make deliveries more quickly that a car delivery service  could, as UAVs are not hindered by traffic congestion and speed limits.

Surveillance groups such as building inspectors and land surveyors can also benefit from  such a system. A tiltrotor can provide the camera stability that a rotorcraft does while having the potential to view many locations through the duration of one flight, requiring fewer return trips  due to energy depletion. A company can invest in many tiltrotor aircraft knowing that they are  incredibly versatile and able to adapt to many different situations, requiring the stocking of only  one model of aircraft rather than investing in multiple other models. This also reduces the  complexity of repairs, as parts storage for only one model type would be required.

Tiltrotor aircraft also have many applications in the military as well. The Golden Team’s  tiltrotor aircraft takes inspiration from the MV-22 Osprey, a tiltrotor aircraft used in the military  for delivery of supplies and infantry, as well as performing quicker extractions than other fixed wing aircraft can provide. The military can use a tiltrotor for remote surveillance operations, as it  can reach further destinations for a fraction of the energy cost than a rotorcraft could and remain  stationary in the air for remote imaging or video capture.  

2. Project Objectives & Scope

Objectives:

• Final simulation testing
• Finalize design
• Part acquisition through lab inventory
• Implement coding for controls
• Live thrust testing
• Design tweaks based off testing results
• Housing development for fuselage and motors
• Flight testing with motors
• Flight stability testing
• Final drone assembly
• Final drone testing after assembly

3. Assessment of Relevant Existing Technologies and Standards

Wing

The wing of an aircraft acts as the main lift generator and allows the aircraft to stay in the  air. The shape of a wing is designed to make air move faster over the top surface which decreases  air pressure on the top of the wing while the bottom surface of the wing will have a higher pressure.  When the air pressure on the bottom of the wing is higher than the air pressure on the top of the  wing, the wing will generate lift. The wing shape plays a vital role in the design of the aircraft.  The shape of the wing depends on the mission of the aircraft. A rectangular wing is the most basic  among personal planes due to the ease of manufacturing. An elliptical wing, seen on most WWII  fighter planes, benefits from having the best aerodynamic properties compared to other wing  shapes. The wing shape used by most modern-day aircraft is the swept wing, seen on popular  jetliners. The swept wing has the tip of the wing angled behind the wing root. This angle reduces drag at higher speeds, allowing for more fuel efficiency and speed while reducing the adverse  effects at higher speeds. This wing shape is optimized for high-speed aircraft. Wing configuration  plays a crucial part in designing a plane based on the mission objective. A low wing configuration, mainly used by most jets and aircraft, offers more maneuverability in flight. A high wing  configuration is popular among larger aircraft such as cargo planes due to the need to have the  fuselage closer to the ground for easier loading. High wing configuration is also popular among  smaller personnel planes such as the Cessna 172. The wings prioritize lift at lower speeds for a  tiltrotor aircraft as the plane will have less speed than a conventional jet. For most tiltrotors, aircraft  will employ a fixed-wing system where the propellers will be the moving system. When in  helicopter mode, the propellers serve to bring the aircraft to optimal heights. When in flight mode,  the propellers move forward and provide additional prop wash that acts onto the wing, generating  more lift. That additional lift generation will make up for a wing's ability to produce lift at lower  speeds. Therefore, the shape and design of the wings must consider how the propeller system will  be integrated into the aircraft. One such example includes the Bell V-22 Osprey employed by the  U.S military.

Figure 2: Bell A821201 airfoil of Bell V-22 Osprey [2]

The Bell V-22 Osprey uses a Bell A821201 airfoil with a wingspan of 50.92 ft (15.52m)  including nacelles, and 84.50 ft (25.77 m) with rotors turning. The design of the wing allows for  the propeller system to be positioned at the wingtips. The Bell V-22 Osprey can fly with a range  of 505 nm (935 km) and a max payload of 20,000 lb (9,090 kg). The Bell V-22 has a max speed  of 115 mph (185 km/h) at sea level in helicopter mode. The max speed increases to 390 mph (630  km/h) in airplane mode at altitude. The United States Marine Corps employs the V-22 Osprey. to  transport supplies and soldiers based on the mission. The V-22 Osprey has replaced the CH-46  Sea Knight since 2007, being able to hold a crew of four and up to 24 troops.

Figure 3: Flaps, ailerons, and winglets on a modern wing [3]

Wing Systems 

Ailerons are used by the aircraft to change roll and alter the aircraft's flight path. Ailerons  work in a tilting fashion, changing the airfoil's shape at the rear of the wing. Changing the angle at  the rear of the airfoil will change the amount of lift generated for each wing. A greater downward  angle will increase the lift generated, which will cause the wing to move upwards, while the  opposite occurs in the other wing where the pitch is upward, lowering the lift and moving the wing  down. With both wings acting opposite of one another, the plane's movement moves about the  center of the aircraft. The ailerons will be controlled using servos and motors. Aircraft also use  flaps to change the stall speed and lift potential. Figure 4 showcases the application of ailerons on  a plane and shows how each configuration changes the aircraft's roll. Flaps are located mainly on  the wing's trailing edge and act according to pilots' motives. Planes will use them primarily for  take-off and landing. Compared to ailerons, flaps will mainly be located closer to the root of the  wings and act in unison to generate more lift or slow down the aircraft. Ailerons act opposite of  each other to affect the plane's roll, while the flaps serve to change the aircraft's speed. Flaps work  by altering the shape of the wing, mainly about the trailing edge in a way that will affect the amount  of lift generated. Flaperons combine the functionality of both the ailerons and flaps. Flaperons can  control the bank angle much like ailerons and can be lowered to serve as flaps. Flaperons can be  found mainly on large commercial jets.

The pilot can separately control all wing systems. Another component employed by many  modern-day aircraft is winglets, which serve to reduce induced drag, reducing the strength of the  tip vortex and leading to a reduction in fuel use. Winglets are part of the wing, located mainly at  the tip with an upward shape. Flight tests at the NASA Dryden Flight Research Center have found  that using winglets led to a 6.5% reduction in the fuel use of a Boeing 707.  

Figure 4: Application of ailerons on a wing [4]

Empennage

An aircraft’s tail is an integral part of the vehicle. The empennage provides enough lift to  stabilize the back of the plane, negating any weight that the fuselage tail adds to the main body.  The primary purpose of the empennage is to provide the aircraft with stability in both the pitch and  yaw direction. A horizontal stabilizer is used on the empennage to provide support in the pitch  axis, that is, the up-and-down direction of an aircraft’s nose. A vertical stabilizer is used to  maintain a direction in the yaw axis, which is the left-to-right movement of the nose. Both are  required to hold a particular direction for the plane, as one will stabilize the aircraft’s elevation  (horizontal stabilizer) while the other keeps the plane from straying its course and turning (vertical  stabilizer). Figure 5 shows a conventional empennage design, with a singular vertical stabilizer  that is placed perpendicular to the horizontal stabilizer. Control surfaces in the form of a rudder  system and elevator system are used in tandem with the horizontal and vertical stabilizers to  provide additional support to the aircraft’s stability and direction.

Figure 5: Conventional Empennage of a Boeing 707 [6]

The elevator system can be found on the horizontal stabilizer. Much like the ailerons on a  wing, the elevator is a system that can be controlled independently of the solid stabilizer using  servos and motors. Where the horizontal stabilizer works to maintain a constant elevation, the  elevator is used to change a plane’s pitch and regulate the aircraft’s lateral direction. By changing  its angle of attack the elevator creates a trailing edge flap on the stabilizer’s airfoil, changing the  amount of lift the tail generates. This would then change the angle of attack the aircraft’s nose  would face and alter the pitch.

The rudder system works similarly to the elevator system, but in the yaw axis on the  Vertical stabilizer. Controlled by motors and servos, the rudder system helps correct excess drag  that an aircraft might experience during a turn. Ailerons on the wings are used to roll the plane and  create an imbalance of lift on either side of the fuselage, which in turn alters the aircraft’s flight  path. During this roll, drag is also affected, pulling the aircraft’s nose off course. The rudder is  used to correct this drag, changing the angle of attack of the stabilizers trailing edge. This will  generate lift to maintain a steady direction while the aircraft is turning in its roll. Figure 6 helps  visualize what happens during this maneuver. As shown, the increased lift on the outside wing  generates more lift and drag than the inside wing due to volume and pressure differences. This induced drag that is generated pulls the outside wing-back, which will begin to pull the aircraft’s  nose up. The rudder is used to generate more lift and drag on the inside wing now, canceling the  excess drag that was generated on the outside.

Figure 6: Rudder correction during an aircraft's turn [7]

Most, if not all, aircraft use both a rudder system and elevator system and consist of at least  both a single vertical and horizontal stabilizer on its empennage. Multiple different configurations  of the empennage are used depending on the required stability and assistance a plane might require.  Figure 3.5 shows one of the more common setups of an empennage, the conventional tail. This  consists of a vertical stabilizer and a horizontal stabilizer which is split into two parts who meet at  the fuselage tail end. The amount of support and stability this tail configuration supplies tends to  be enough for most aircraft. A couple of designs that take inspiration from the conventional tail  are the cruciform tail and the T-tail. Both tails, much like the conventional tail, have a singular  vertical stabilizer as well as a horizontal stabilizer split in two parts. The largest variance is the  location of the horizontal stabilizer. A cruciform tail has the horizontal stabilizer moved up slightly  on the vertical stabilizer. This design is used mainly in high-speed jets and aircraft and is used to  keep the horizontal stabilizer away from the wing wake region as well as exhaust from the engines.  The T-tail has its horizontal stabilizer up at the top of the Vertical stabilizer, creating a “T” shape.  The “T” shape of the tail helps the horizontal stabilizers efficiency by placing it outside of any propeller wash and the wing wake region. This gives the engineer leeway on the size and weight  of the horizontal stabilizer. A downside to this configuration is the additional strength and stability  the vertical stabilizer would require. Since the Vertical stabilizer is the only object holding the  horizontal stabilizer, a moment is placed on it. This will create stress and require a stronger body  for it, which would potentially increase the weight of the empennage system.

Figure 7: List of possible tail configurations [8]

 The three tail configurations listed previously are popular designs in today’s aircraft,  though many different designs are possible. Figure 7 shows a list of different tail designs that stray  from the conventional tail. Some designs, like the V-tail and inverted V-tail, replace both vertical  and horizontal stabilizers with a hybridized stabilizer, accounting for both pitch and yaw axes.  These designs help reduce the drag an empennage might face and could lighten the weight. While  these properties are helpful, the V-tail has difficulty stabilizing the aircraft in a turn due to the  inverse angles each control surface would face.

 The twin-tail design was looked at as well, as it shares many similarities to the conventional  tail with the addition of an extra vertical stabilizer. This would increase the stability an aircraft  could achieve and wouldn’t require as large of a surface as the conventional tail. This benefit is  countered by one of the main disadvantages the twin-tail design faces, which is added weight. To  fit two vertical stabilizers, a wider fuselage tail would be required, increasing the material needed  on the back end of the aircraft. This design is mainly used for fighter jets such as the F-18, which  don’t have a tapered fuselage tail and instead have a rather compact body.

Fuselage

Tiltrotor vehicles are hybridized aircraft that combine the vertical liftoff capabilities of a  helicopter with the range of fixed-wing turboprop aircraft. This could greatly benefit commercial  air flight in densely populated communities; Vertical takeoff could save space, with tiltrotor  aircraft lifting from pads instead of runways and such. A major component of the tiltrotor aircraft  is the fuselage and main body of the vehicle. The fuselage needs to be streamlined during cruising  flight (With the rotors positioned in flight mode) while maintaining weighting limitations of  vertical takeoff.

A major component of an aircraft’s fuselage is its payload capabilities. Long-range  traveling coupled with vertical takeoff means that a tiltrotor can carry cargo further than helicopters  with the ease of loading and stowage that such rotorcraft bring. The Bell Boeing V-22 Osprey has  a load capacity of 24 individuals, or a payload weight capacity of around 20,000 pounds inside the  cargo hold. An additional 15,000 pounds could be fitted to a winch below the aircraft. In total, the  max weight of the V-22 while performing a vertical takeoff is approximately 52,600 pounds. This  is a minimal difference to its rotorcraft counterpart the Boeing CH-47 Chinook, which has a  payload capacity of 24,000 pounds and the capability to hold upwards of 50 individuals in its  cabin. The V-22’s true benefit is its range and speed, according to Boeing, the Chinook’s maximum  speed is estimated to be about 170 knots (approximately 302 km/h) and has an effective mission  range of 200 nautical miles (approx. 370km). The Bell Boeing V-22 has an effective mission range  of 428 nautical miles with a full payload and holds a maximum speed of 270 knots (approx. 500 km/h). This shows that with a similar payload, the V-22 can travel double the distance to that of  the Chinook CH-47 while maintaining a faster speed.

 For a fixed-wing counterpart, the Lockheed C-130J Super Hercules was researched;  specifically, the C-130J-30. At a payload weight of 46,700 pounds, it boasts a significant increase  in payload capacity to that of the V-22. The maximum range of a C-130J-30 is around 2,160  nautical miles (approx. 4,000 km) with a maximum cruising speed of 365 knots (approx. 675  km/h). These readings prove to be much better than that of the Osprey, though there are many  benefits to counteract these differences. The C-130J-30 requires around 900 meters of flat runway  to safely land and come to a full stop. Utilizing Vertical Take-Off and Landing (VTOL), the V-22  needs no runway, requiring only enough space for itself. Another major benefit is size and stowage  capability. The Super Hercules has a length of 113 feet, double the sizeof the V-22’s 57 feet.  Compounded with its 133-foot wingspan, the Super Hercules requires a large portion of land  dedicated to its storage.

A tiltrotor UAV holds many similar benefits to larger aircraft such as the V-22. With no runway  required, a tiltrotor drone could deploy practically anywhere. This could provide reconnaissance  over terrain at a more efficient rate than that of a quadcopter drone. For civilian/commercial use,  a tiltrotor could be very useful for shipping companies such as Amazon. Drones could be deployed  from any warehouse and provide a faster delivery over a wider distance.

 The fuselage of the tiltrotor UAV is dependent on wing design and materials selection. The  placement of the propellers will play a significant role in fuselage design. Too wide of a fuselage  and the body could interfere with lifting; too high of a body design and it could prove detrimental  to the aerodynamics when in cruise mode. Strakes placed along the nose of the fuselage as well as  the body could help provide aerodynamic stability. The fuselage will hold most vital components  of the drone and will serve as the cargo hold for a payload. Each component must be weighed to  understand the most efficient placement and setup. This means weight must be evenly distributed  to maintain a steady lift.

 Taking the Max Gross Takeoff Weight of the V-22 at 52,600 pounds and comparing it to  the maximum payload capacity of the fuselage at 20,000 pounds, the approximate payload mass  fraction of the V-22 could be calculated to be 38%. To maintain this, a drone that is based off the  V-22 would be required to hold a minimum of 30%-40% of the vehicle’s full weight.

Figure 8: Variation Of Fuselage Structures [9]

There are various structure designs used for a fuselage that each have their own highlights  and drawbacks. The four most common styles used in industry are: Geodesic, Truss variations,  Monocoque, and Semi-Monocoque [10] which can be seen in Figure 8 above, respectively. The  decision of what style is most useful for the tiltrotor application must be determined considering  structural strength and integrity, ease of manufacturing, internal space efficiency and material  weight cost.  

Figure 9: Geodesic Airframe Example [11]

Figure 10: Truss Style Airframe Example [12]

To begin with, Geodesic structure, which was pioneered in the 1930s by Barnes Wallis  [10], is considered one of the strongest in the market due to its massive interconnection of force members. This style (Figure 9) is most useful in the tasks of carrying the heaviest payloads and  flying through environments with extreme external forces from high winds or debris without major  deformation to the main body. The trade-off to this style of structure is that the gain equally  matches the gain in strength and rigidity in weight and loss of internal real estate. The next  structural configuration to be considered is the truss-style (Figure 10). This method is mostly used  for aircraft that prioritize lightweight characteristics and don't expect to be under great  pressurization. The benefits this brings are a lightweight aircraft, easy to construct using slender  members, and great structural strength and rigidity. These benefits are countered by the drawbacks  of the inability to incorporate stiffeners in the assembly, the lack of additional supports for  streamlined shapes, and it does not handle great pressurized loads well [13].  

Figure 11: Monocoque Airframe Example [14]

Monocoque structure (Figure 11) design is the most used type behind Truss. This structure  is used for its uniform strength and rigidity and has great form for fitting wings. Its advantages are  contrasted primarily by its low strength to weight ratio, followed by heavier design and frequent  maintenance. Finally, the Semi-Monocoque is like the previous design, but with the addition of  linear reinforcements around the circumference of the fuselage called langerons [16] which add an  increase in bending resistance (Figure 12). It also has a better strength to weight ratio and is used  where strong and rigid streamlined construction is necessary. Out of all the structures mentioned,  this design is the most expensive to make and the most difficult to manufacture.

Figure 12: Semi-Monocoque Airframe Example [15]

The primary trait of an airframe structure for the tiltrotor fuselage is one that has the  strength and rigidity to support the weight of the propeller motors, which are applying a bending  stress from the length of the wings. The structure of choice will also need to be lightweight, easy  to manufacture, and leave sufficient space in the cavity for avionics and payload. From the designs  briefly overviewed above, it is most likely that the UAV Tiltrotor for this project will be  constructed using a Truss style airframe.  

Materials and Structures

In engineering any system, there is always the understanding of how the combination of  material and how it is assembled affects the intended performance. This section briefly reports on  the various materials to be used for the drone to be built, considering assembly methods, internal  structure, and external forces.

Choosing of material for any component should be decided based on the expected forces  the UAV body will experience and the efficiency of weight to power desired. Realistically, the  cost of material also needs to be considered. For example, one of the material types investigated  for this application is fiber cloths. Generally, fiber cloths are lightweight, offer excellent  temperature resistance, and have an incredible weight-to-strength ratio [17]. One of the best  options for this is Carbon Fiber cloth, cured with epoxy. As shown in Figure 13, a composite  numerical and experimental comparison research from the University Tenaga Nasional [18] shows  how Carbon Fiber cloth in certain applications can handle pressures of up to 1200 MegaPascals  with little to negligible deformation. Although this is an incredible feat of material, stronger than  steel and lighter than aluminum, the cost for carbon fiber as well as working it, may not be justified for a drone design when fiberglass cloth is significantly cheaper, more widely available, and can  offer similar advantages over other plastics or Styrofoam alternatives. Another cost of fiber cloth  is that it impedes on radio frequency (RF) signals, which may lead to one of the advantages of  using plastics instead. Plastic is by far the most versatile material available, with an extensive  variety of forming methods.  

Figure 13: Carbon/Epoxy Stress-Strain Curve [18]

The most popular methods of plastic forming for drone application are heat forming or  molding from thermoplastic sheets and 3D printing using PLA or PETG filament. Plastic molding  is better for mass production, given that it is faster and lays easily. It also offers a not as rough  finish compared to 3D printing. On the other hand, additive manufacturing can produce greater  precision from Computer-Aided Design (CAD) files, which can be friendlier when it comes to  designing assembly points and mechanisms for serviceability. In addition, it can always be sanded  or finished to offer a smoother surface to reduce fluid drag for aerodynamic advantage. For the  sake of single instance production and better alignment for design intent, 3D printing seems like  the preferred method of use when choosing plastic as the body material. The most common  filament used for this is PLA, also known as polylactic acid or polylactide. Figure 14 from the  Michigan Technological University [19] shows the experimental stress-strain curve for PLA  printed at various temperatures: 215 C, 190 C, and natural PLA at 190 C. Although this isn’t in  any way competitive with the strengths and weight reduction of Carbon Fiber or Fiberglass, we can establish that PLA can more than handle any range of forces expected for this UAV drone  application.

Figure 14: PLA Stress-Strain Curve with varying Temperature [19]

From the data established, with the reduction in price point and advantage of precision, 3D  printing seems like the best route to take for most components, when compared to fiber cloths or  plastic molding. It is predicted to withstand the compression and tensile strength we need while  giving us the advantage of better budget distribution and practical part creation. With that said, the  team is not restricted to only one type of material or manufacturing for every system or part.  Therefore, this tilt-rotor drone should use a combination of materials depending on its own  individual advantages. Ideally, 3D printed PLA - Carbon Fiber composite will be used for critical  components such as the airframe, propeller housing, and internal structure of the fuselage, while  carbon fiber or fiber-glass cloth for the wings, rudders, and other external compressive load bearing surfaces.

Avionics/UI

FPV Implementation

A first-person view camera is quintessential to a remote UAV design, as it allows the pilot  to appropriately navigate the space around the drone and view its surroundings. For this design, a  fixed camera with a wide field of view mounted on the tip of the aircraft’s nose will be used. This  will provide the pilot with a view directly in front of the aircraft, allowing for flight further from  the pilot’s visual line of sight, expanding the effective range of the aircraft. To successfully  implement an FPV, a receiver and antennae are also required.  

Controls

A tiltrotor aircraft has two modes of functionality, fixed-wing mode, and rotorcraft mode,  as shown in Figure 15. Both come with their own unique methods of control. In fixed-wing mode,  three primary flight control surfaces are required to control direction of flight: ailerons, elevators,  and rudders. The drone will implement one aileron on each wing for roll control, one for pitch  control, and for added design simplicity, a differential in propeller frequency for yaw control, as  opposed to rudders. With typical rotorcraft, forward and lateral motion are achieved by the tilting  of a swashplate, creating unequal lifts in different angular sectors of a propeller. However, due to  the lack of a swashplate mechanism in this drone, in rotorcraft mode, a forward and backward  motion will be achieved by rotating the propellers toward the tip or tail, as well as slightly  increasing propeller frequency to account for the decreased lift from an angled propeller. To  achieve lateral motion, a difference in propeller frequency will be used to create an imbalance of  lift on either side of the drone, causing it to drift laterally. Yaw in rotorcraft mode can be achieved  by rotating the propellers in opposing directions. To achieve a rightward yaw, the right propeller  should tilt toward the rear, while the left propeller should tilt toward the front of the aircraft. All  these functions will be achieved with a two-joystick controller as tabulated in Table 1.

Figure 15: Fixed-wing mode (left), rotorcraft mode (right) [20]

Table 1: Aircraft response mapped to controller input

Neural Network and Wiring

As shown in Table 2, the Arduino Uno, Micro, and Nano feature similar performance and  negligible size and weight differences with respect to the tiltrotor’s fuselage design. However, the  Uno is superior to the Micro and Nano due to the Uno’s built-in solderless pin connections. This design facilitates testing of the microcontroller and different wiring configurations due to not  requiring connections to be soldered on. For these reasons, the Arduino Uno will be used as the  drone’s microcontroller. Its small dimensions of 69 mm by 53 mm, combined with its lightweight  of 25 grams, makes it an excellent contender compared to other microcontroller options available.  The Uno can operate on an input voltage anywhere from 6 to 20 volts, allowing for a wide range  of batteries to be used. The Uno also features 14 digital I/O pins, 6 of which allow for PWM output,  allowing for a maximum of 6 servo motors to be used. The drone’s wiring will initially include a  breadboard for ease of testing; however, the drone’s final design will incorporate wires that are  soldered together. As opposed to connections with a breadboard, soldered connections are much  more resilient to impact and much less likely to fail due to unexpected disconnections mid-flight.  There is also a lower electrical resistance across a soldered connection than across a breadboard,  as resistance between the connections in a breadboard are highly dependent on the manufacturing  method of the breadboard, and testing of a breadboard’s quality requires its partial destruction.

Table 2: Comparison between Arduino Uno [21], Micro [22], and Nano [23]

Gyroscope

 The drone’s programming will include an auto-stabilization system to keep flights more  efficient and safer. To accomplish this, a gyroscope is required for the drone to recognize changes  in its angular position relative to the ground in flight and adjust accordingly. For this, the MPU6050  IMU will be used. This sensor features a 3-axis accelerometer and a 3-axis gyroscope, all in one  chip. The sensor requires an input voltage of 3V, while the Arduino Uno requires a minimum input  voltage of 6V, so a voltage divider circuit must be implemented to prevent damage to the sensor.  The gyroscope’s range can be +/– 250, 500, 1000, or 2000 degree/sec, and the accelerometer’s  range can be +/– 2, 4, 8, 16 G [24], where an increased range causes a decrease in precision. For this design, the smallest range is most ideal to use as the aircraft is not intended to perform high G aerobatics.  

Propulsion 

Engine/Engine Configuration

 In terms of engine selection, several factors must be determined to make an ideal selection.  These key factors are engine configuration/count, propellor size, payload, fuel/power efficiency,  and cost. Without a specified payload, it must be determined which combination of factors results  in the greatest power-to-weight ratio for the aircraft while being affordable and retaining a long  flight time.

 When reviewing possible engine configurations, there was a surprising array of designs  outside of the traditional two puller engines of the V-22 Osprey or XV-15. A more straightforward  way of changing tiltrotor design involves adding additional motors, such as the three-engine AVX  TriFan 600 and the four-engine Bell QTR. However, while this appears to be very simple, adding  additional motors results in several complicating factors.

In the case of a three-engine configuration, the third engine would be mounted along the plane's  longitudinal axis and towards the rear to keep things symmetric. However, the position of this  engine results in it usually being fixed, only used for hovering or standard flight. Because of this,  the engine would be dead weight during the phase in which it is not in use. Even if it can tilt, there  still lies some issues with an odd number of engines. To start, torque from a helicopter's main  engine tries to spin the aircraft itself, but this is counteracted by a tail rotor or adding another  engine/propeller spinning in the opposite direction. However, the third engine would have nothing  to counteract it, resulting in a slight spin while hovering. Additionally, its location on the aircraft may cause issues when trying to move while hovering by creating undesirable shifts in its center  of thrust. While a four-engine configuration would not run into this issue, it faces increased power  consumption. Several more unique ideas were proposed by Young L. A. in his 2018 conference  paper: "What is a tiltrotor? A Fundamental Reexamination of the tiltrotor aircraft design space".  [25]

Figure 16: Various possible tilt-rotor configurations from left to right: (top) Bell Boeing V-22 Osprey, Bell Boeing QTR,  (bottom) PIVOT, Tiltrotor Oblique Wing (TOW) [25]

 To start, a pusher-style configuration would eliminate the wing load from prop wash  produced by a puller configuration. However, ground clearance would be an issue, in addition to  problems with payload loading and engine exhaust. Another unique design proposed by Young  was that of an asymmetric pusher-puller aircraft. As seen in Figure 16, this configuration called  for a single puller engine and a single pusher engine at either end of the aircraft's lateral axis. While  this does create a unique imbalance due to unequal thrust coefficients, it opens the options for  things such as an oblique flying wing aircraft. While these designs are merely theoretical, they  could provide alternatives to the traditional configuration should the need arise. [25]

Figure 17: Different options for RC Powerplants (left to right): (top) Electric, Nitro, (bottom) Gas, Turbine

 For the most part, modern tiltrotors use turboprop or turboshaft engines; however, some  past examples have used piston or even jet engines. However, for UAVs, the engine selection  options are a lot different. Currently, there are four engine types available: electric, nitro, gas, and  turbine. Electric engines are by far the simplest, having one of the highest power to weight ratios  and just needing electricity to run. However, they are held back by heavy lithium batteries and  limited flight time. Additionally, electric motors come in two types: innruner and outrunner. The  two types differ on whether the rotor lies inside or outside of the sator, affecting the engines  performance. The next type, nitro, are powerful and durable engines that have minimum downtime  between flights and can save money in the short term due to the cost of high-capacity batteries.  The problem with them is that they are messy, need fuel tanks to store fuel, and are a lot more  complicated compared to electric engines. Gas engines are like nitro engines; however, they are  cheaper, make less of a mess, and use gasoline which is easy to get. However, they have the lowest  power to weight ratio, are large, and are more expensive than the previous options. Lastly, turbines  are the most realistic option, which tend to be very reliable thanks to their engine computer and  produce minimal vibrations (approximately the same amount as electric). Unfortunately, they are  the most expensive, fragile, and complex option available. [26]

Rotation Mechanism

Like other VTOL aircraft, tiltrotor aircraft can generate active and passive lift using their  engines and wing, respectively. To transition between the two means of lift, the aircraft's engines  are rotated from a vertical position (where they used the thrust generated to produce lift) to a  horizontal configuration where they solely produce thrust to move forward. On traditional tiltrotors  such as the V-22, the entire propeller engine turns about the wing. While this is a straightforward  solution, it runs into some issues regarding shifting the aircraft's center of mass (CoM) and added  complexity. To solve this, the Bell Boeing V-280 and the University of Maryland's Excalibur  resolve this issue by keeping the engine stationary. In the case of the Excalibur, only the propeller  moves, minimizing the movement of the CoM. To facilitate this movement, a Tilt-Independent  Transfer Joint was created to maintain engine power to the propeller regardless of tilt angle. [27]

Figure 18: Excalibur Engine in its standard configuration and hover configuration [27]

Hover Configuration Movement

 On aircraft, movement is usually controlled by the flight surfaces on the wings and  stabilizers. These flight surfaces are reliant on an aircraft's airspeed to be effective, meaning that  they are useless while the tiltrotor is hovering. As such, it is essential to find another means of  control for this mode. Pitch, or rotation about the lateral axis, is traditionally controlled by an  aircraft's elevators located on its horizontal stabilizers. However, in tiltrotors, this is achieved by  tilting the engines forward or backward in the same direction. Roll, defined as rotation about the  longitudinal axis, is typically done by the ailerons that move in opposing directions. Tiltrotors  replicate this by producing an uneven thrust coefficient between the two engines, though unlike  regular aircraft, rolling is typically limited to a set roll angle in addition to some difficulties with maintaining a constant altitude. Yaw, rotation about the vertical axis, tends to be controlled by an  aircraft's rudder located on the plane's vertical axis. Typically, yaw is to adjust for a crosswind  during landing or for maintaining coordinated flight; however, the tiltrotors can complete a 180  degree turn almost on the spot just using yaw, just like helicopters. Helicopters typically achieve  this by reducing their tail rotors rpm; tiltrotors either tilt a single-engine or by tilting both in  opposing directions instead. This does produce an issue where a roll moment may be produced  due to uneven thrust coefficients, or altitude may be lost due to a reduction in total vertical thrust.  These challenges must be dealt with before the team's tiltrotor production. [28]

Figure 19: Movement while hovering: (top to bottom) pitch, vertical thrust, roll, and yaw [28]

Propeller Design

 While creating a propeller may appear to be a straightforward process, it is more complex  than it seems. This is due to the airplane and helicopter configurations having different  performance objectives that they aim for and conditions that they operate under. While hovering,  the propeller needs to produce the maximum amount of lift while dealing with a low inflow rate  (due to slow airspeed). By contrast, the propeller needs to produce horizontal thrust while taking  a high inflow rate while it is in an airplane configuration. [29] Tiltrotors typically solve this by  adjusting the pitch of the blades to meet the specified requirements. This adjustable blade pitch is controlled by a variable-pitch propeller system. Also referred  to as a controllable-pitch or constant-speed propeller, it allows for pilots to adjust the aircraft's  speed without the need for adjusting the aircraft’s RPM. An advantage of this system is allowing  for more efficient flight operations which helps to reduce energy consumption. An additional  advantage is that it allows for optimum performance even at altitude, allowing for an aircraft to  handle higher altitudes. [30]

Table 3: Performance of a fixed-pitch propeller vs a variable-pitch propeller [30]

 Several other factors regarding propeller design include propeller length, material, blade  count, and Mach Criteria. Propeller length is typically determined by the engine power/max rpm;  however, on airplanes, it is also limited by ground clearance. However, because tiltrotors can keep  the propeller vertical while on the ground, it isn't limited by this, allowing for larger blades sizes.  Engine rpm is the only issue as increasing it would increase the speed of the blade tips. As the  propeller tip approaches supersonic speeds, its performance decreases due to compressibility  effects; a relationship referred to as the Mach number Criteria. To counteract this loss, the tips of  the propeller blade can be swept. [29] Due to the forces the propeller experiences, material  selection is essential to keep it intact. While standard propeller blades are made of wood or steel,  it is made of plastic for a small UAV. Given the limitations to a propeller blade's length, adding  additional blades may be a possible solution if performance requirements are not met. However,  the addition of more blades limits their width and adds additional weight for the engine to spin.

Table 4: Different conditions a propeller with a radius of 3.7m would experience at various flight conditions [29]

Power Distribution

Since batteries are the main source of power for drones, you must look at the pros and cons  of various batteries. Two potential power units are Lithium-polymer (Li-Po) and Lithium-Ion (Li ion) batteries. Li-Ion batteries are used for more long-term power sources, due to their longer  charge retention and ability to be recharged hundreds of times. They have a lower self-discharge  rate, though, which causes a lower thrust force; these batteries still have a high-power efficiency  while maintaining a cheaper price. On the other hand, the disadvantage of this battery is a low  lifespan per charge, which is equivalent to having a phone on 10 percent after it is charged every  time. The biggest drawback for Li-Ion batteries is their weight, since they carry the same voltage  as lighter batteries but weigh more than over two times Li-Po causing a higher payload on the  drone [30].

On the other hand, Lithium Polymer batteries are more beneficial for a RC drone due to  their high-power capacity and reduced weight. Li-Po batteries are also very thin and malleable, as  seen in figure 3.19, which allows the battery to take up less space and be placed where it is less  likely to get punctured on the drone. Another bonus of using Li-po batteries is the short duration  it takes to charge the batteries while still maintaining a longer lasting power life. Although these  batteries have many advantages over Li-Ion batteries, there are still minor downsides to using  Lithium Polymer batteries, which include having a shorter recharge life that lasts around 400  cycles, and the ability to easily catch on fire if punctured. With that being said, since these batteries are rated as the highest quality when it comes to using them for RC drones, they are also on the  more expensive side.

Figure 20: Li-po vs. Li-Ion battery design [31]

Accessibility / Design

Figure 21: Bottom-Mounted Battery [33] Figure 22: Top-Mounted Battery[32]

While determining what power source needed to operate a RC UAV plays a huge factor,  as well as deciding where the battery should be positioned. Because Li-po batteries are easily  flammable if punctured, a casing should be made for it to protect it from falls on failed test runs.  There are advantages to mounting the battery either on top or on the bottom of the RC. For instance,  a top mount as seen in Figure 22 allows the drone to have a higher center of gravity causing the  drone to become more stable, responsive, and allows for cool “touch and go” landing tricks.  However, it is more susceptible to getting hit by the propellers if not secured properly.

Alternatively, having a battery mounted to the bottom results in the center of gravity being  below the center of thrust which could lead to a larger rotational inertia [33]. In contrast, a bottom  mount allows for a lower profile, meaning the UAV will have less drag resulting in a smoother  and faster flight. In all, each mount has its advantages and disadvantages, but for this project, we  are only tasked with making a lap around a track carrying a payload with no time specifications,  making a bottom mount more appropriate for these tasks.  

Maximizing Battery Performance

The component that could have the most drastic effect on your drone’s performance is  being able to properly care for and maintain the battery without overcharging or depleting it.  According to 3D Insider, to avoid any damage to the batteries, always: store them where they are  not in any direct sunlight or extreme temperatures, never let the battery get fully drained, never fly  the drone when the battery is less than twenty percent charged and most importantly, do not try to  puncture or alter the battery because it could result in a fire[34].  

In all, maximizing a battery’s performance is not solely based on how the battery is handled  but also how the drone is designed and constructed. To get the most out of a battery, you must  remove any unnecessary weight on the drone which could include any mounted cameras or any  additional designs that don't have any impact on the drone [35]. This also includes installing lighter  and more efficient propellers that can still maintain a smooth flight, otherwise it could result in  quickly draining the battery. Lastly, before you put your drone to the ultimate test to see if it flies,  you must have a fully charged battery. However, be careful to not overcharge it or it could result  in overheating, or it could generate electrolytic decomposition which ruins the battery.  

Regulations and Standards

FAA Small Recreational UAS Rules

§   Regulatory Standards & Codes

• §14 CFR Part 107

• Applies to non-recreational UAS under 55 lbs
• Maybe applicable (Does not address operations as part of a university project)

• Small Recreational UAS Rules (as directly stated from FAA.gov)

• Fly only for recreational purposes (enjoyment).
• Follow the safety guidelines of an FAA-recognized Community Based Organization (CBO).
• Keep your drone within the visual line of sight or use a visual observer who is co-located (physically next to) and in direct communication with you.
• Give way to and do not interfere with manned aircraft.
• Fly at or below 400' in controlled airspace (Class B, C, D, and E) only with prior authorization.
• Fly at or below 400 feet in Class G (uncontrolled) airspace.
• Take The Recreational UAS Safety Test (TRUST) and carry proof of test passage.
• Have a current registration, mark (PDF) your drones on the outside with the registration number, and carry proof of registration with you.
• Do not operate your drone in a dangerous manner. For example:

• Do not interfere with emergency response or law enforcement activities.
• Do not fly under the influence of drugs or alcohol.

4. Professional and Societal Considerations

        When considering the project's impact on the world, there are several notable issues to denote. Namely, economic, environmental, sustainability, manufacturability, ethical, health and safety, social, and political issues. Considering the project’s scope and design, the tiltrotor will likely have little impact on these issues as a whole. However the technology, as a whole, may aid in the resolution of several issues. In particular, the technology should help improve several economic and environmental issues.

Aviation has significantly altered the transportation industry, allowing supplies to be transported across vast distance around the globe. The Boeing 747-400ER cargo plane, for example, could carry a maximum payload weight of 248,000lbs over 4,450nmi (5,120.969 miles) [Boeing]. Despite its seemingly impressive range and payload weight, the payload it carries only makes up 27.25% of its total maximum takeoff weight. By contrast helicopters like the Sikorsky S-64E Skycrane can carry up ~20,000lbs of payload, making up 47.62% of its maximum takeoff weight. However, this comes at the cost of its maximum flight range being 230nmi (264.679 miles) [S64]. Referred to as the aircraft’s payload mass fraction, maximizing the amount of payload an aircraft can carry relative to its takeoff weight without sacrificing flight range is crucial to lowering emissions and transportation costs.

Tiltrotors are an ideal solution to resolving this issue by being able to exploit the benefits of both fixed-wing aircraft and rotorcraft. The Bell Boeing V-22 Osprey is prime example of this capable of carrying ~20,000lb and a flight range of 2,100nmi (2,416.637mi) [V22]. The payload mass fraction for the V-22 is 33.06%. In addition to having a high carrying capacity while maintaining a long flight range, tiltrotors can land and takeoff from hard to reach spots. This is because of its VTOL capabilities.

5. System Requirements and Design Constraints

Table 5: Engineering Requirements

6. System Concept Development

Changes made in design between M3 and M6

Wing

For the wing rib design, it was important to cut as much weight as possible while maintaining structural integrity. To do this, holes had to be cut out of the rib. The figure below showcases an initial design of each wing rib, made with stringer notches to place stringers in a way that the stringers would be tangent to the curves of the airfoils.

Figure xx: Initial Rib Design for the Wing

             The rib design was then implemented onto FEMAP for an FEA to view the stress distribution acting on each rib. The figure below reveals where the most stress was acting on the rib.                

Figure xx: FEA on Initial Rib Design

               The FEA shows that the majority of stress was occurring towards the center spar and weight reduction holes. This prompted that the holes were too close to the edges and spar holes, which in turn, left a structural problem on the rib. A redesign of the rib was necessary to ensure structural integrity. Movement and resizing of the holes were made to maintain the minimum weight while leaving enough room between the edges and holes to minimize stress.        

Figure xx: Redesign of Rib

               The figure above represents the new design of the ribs after performing an FEA and viewing the stress distribution throughout the rib. The holes were moved and reshaped to allow for more strength in the areas with the most stress. An FEA was performed on the new rib design and as shown in the figure below, the areas that held the most stress saw a decline in stress compared to the previous design.                

Figure xx: FEA on New Rib Design

                 After confirming the results on the FEA, the next step was to finalize the design of the wing assembly to be used for the drone. The wing structure consists of seven ribs on each side to minimize the weight while maintaining the structural integrity during flight. The figure below shows the final assembly of the wing, fitted with ailerons and servos, which act as the control surfaces.

Figure xx: Assembly of Wing Structure in SOLIDWORKS

Airframe - Fuselage

The airframe underwent a major redesign due to size complications with the manufacturing method. Like originally planned, the frame used a carbon fiber-PLA composite and additive manufacturing to print, however, the 3D printing bed had a maximum area of 30x30cm. This

would require the team to break up the frame four to five times and use a high volume of material, increasing risk and weight. The team decided to move forward with an airframe design consisting of a CF-PLA base connector for the wing, tail connector for the empennage, and a slender boom that brings them together (Figure 5XX). The base connector was modeled to match the same curvature of the airfoil of the wing, then using the stress analysis shown in Figure 6 XX,  areas of low or no stress were removed from the geometry, reducing material use and weight without affecting the rigidity and strength of the frame. This final geometry was then 3D printed, cleaned and readied for assembly (Figure 1XXX).  The tail connector was designed and manufactured in a similar fashion as the base, by matching the airfoil of the empennage, and also including the vertical stabilizer in the model to make it one piece, also being printed in CF-PLA (Figure 2XXX).

        For the most part, the fuselage shell design remained the same throughout the process, only being tweaked for manufacturing purposes. A 2D cross section of the side profile was outlined and turned into a dxf file to be laser cut out of 1/8in plywood. The two halves were then mounted together with a flat rectangle of plywood at the bottom, and two spare balsa dowels in the middle for rigidity. The remaining out geometry was made up of thin balsa sheet, only being used as a foundation for the monokote to be applied at a latest step in the process.

Empennage

        Not many redesigns were utilized for the Empennage. Taking a look at the final CAD assembly shown in figure x1x, the horizontal and vertical stabilizer needed a base that could connect to the tail boom as well as the rest of the assembly. To do this, a 3D printed frame was used that could hold onto the tail boom as well as provide structure for the stabilizers to form around. This can be seen in figure x2x, where the metal boom tail fits in the front of the tail connector. After a little more tweaking and modeling, the design moved towards printing the vertical stabilizer with the tail connector to ease the difficulties that could be faced when assembling the tail. With a 3D printed vertical stabilizer, the center body of the connector could be used to run the horizontal stabilizers sparring through without any interference.

Avionics

Some of the avionic components included in the design were changed to accommodate for part malfunctions. The initial design was meant to implement the Arduino Mega, however, the Arduino Mega was malfunctioning during testing, requiring us to use the Arduino Uno. Due to hardware limitations the Arduino Uno has, the gyroscope was not able to be implemented in the final circuit.

Propulsion

        The propulsion system concept was altered drastically between its original conception and final production. Designed to be powered by twin electric motors, variable-pitch propellors were chosen to improve efficiency and flight performance due to the two different flight modes. However, issues regarding wiring and complexitiy resulted it the design pulled back down to a fixed-pitch propeller. After further simulations and refinement, a 14x7 3-blade propeller was selected, to be mounted on one of there motors: an E-flite 32 or 60, or a Hacker A40-10S V4 14 Pole.

Figure #: Finalized Powerplant Ver. 1. Hacker A40 Electric Motor (left) and a Master Airscrew 14x7 Propeller (Right).

        Final engine selection was to be determined during thrust tests prior to the manufacturing process. Unfortunately, delays in the parts acquisition process resulted in the timeline being delayed significantly, resulting in the testing and manufacturing periods to be combined. During testing, torque and powerplant issues we discovered, resulted the design being further changed. Despite achieving 10 lbs of thrust, the motors would overheat rapidly due to the strain on the engine caused by the propeller size. As such, a higher torque Cobra C-4120 was selected with the blade count being reduced to a APC 15x4E Propeller.

Figure #: Cobra C-4120/16 (Left), and the APC 15x4E Propeller (Right)

        The engine housing experienced similar changes throughout the design process. Originally designed as an aerodynamic shell with an inner framework for mounting components, it was changed to a near solid rod. These changes were made to account for the limitations of additive manufacturing as well as structural integrity issues. However, despite the simplified rod design being less aerodynamic, the total weight of the propeller housing decreased slightly and structural integrity increased exponentially.

Figure #: Initial 3-Part Shell Housing (Left), Finalized “Rod” Housing with Engine (Right)

7. Design Analysis

 Wing

        For the wing design, each decision was made for accomplishing the goal of creating a wing that is capable of achieving the desired lift while using lightweight materials and maintaining structural integrity during flight. To do this, the team used multiple analysis and modeling programs to satisfy design requirements. These programs looked to achieve data for CFD, FEA, and 3D modeling. These programs included XFLR5, ANSYS Fluent, STAR-CCM+, FEMAP, and Solidworks. Hand calculations were done to gather results for lift and drag using different velocities at different angles of attack. With the results acquired through computation and hand calculations, the team received data which proved the wing design was capable of achieving flight at certain velocities and with a maximum load while maintaining structural integrity.

        The first step to designing the wing involved the selection of an airfoil that would provide the drone with enough lift, taking into account the weight of the drone and expected flight speed. The program used for this analysis was XFLR5 to simulate the capabilities of multiple different airfoils and then narrow the selection until the proper geometry was best fitted for the aircraft’s objectives. The first airfoil tested was the NACA 2412 which held an adequate camber with a thin design. After analyzing the lift data, it was found to have an impressive lift to drag ratio, however, the lift coefficient values were under the desired capabilities for the drone. The NACA 6412 airfoil was analyzed as the next potential choice. While the lift values were impressive, the airfoil also presented a lower lift to drag ratio which could hinder the performance of the drone at desired speeds. The next choice of airfoils involved the NACA 4412. The data acquired for the airfoil revealed impressive lift values while also maintaining less drag compared to the NACA 6412. This aspect was important because a tilt rotor drone typically does not fly faster than conventional aircrafts, thus, being able to generate more lift at lower speeds would benefit the tilt rotor drone more extensively. XFLR5 is capable of simulating lift and drag computations with customizable settings. The overall geometry was developed within the program and data for lift was acquired for different angles of attack to view which angle would provide the best lift coefficient at the desired speeds while minimizing drag. It was decided that an angle of attack of 5 degrees was best suited for the drone’s objective of flight at 20 m/s. The dimensions were configured at a 1.4m wingspan with a 0.3m chord length to meet design requirements. The first design included an elliptical wing shape to help reduce the drag coefficient. After thorough discussion with the team, the decision was made to give the wing a rectangular shape to make manufacturing more simple. Changing the shape to the rectangular shape was shown to improve the lift to drag coefficient ratio compared to if the elliptical shape had remained in the design.

Figure XX: Lift and Drag Ratio Vs Angle of Attack for Various Airfoils

  Figure XX: Lift Coefficient Vs Angle of Attack for Various Airfoils

Based on the data in figure xx and figure xx, it was decided that the NACA 4412 airfoil was best suited for the design as it offered the required lift values while maintaining a relatively low drag coefficient for the angle of attack chosen.

Figure XX: NACA 4412 Airfoil Geometry at 5 deg Angle of Attack

The next phase for the wing design was creating a 3D model of the wing. To accomplish this, the geometry of the NACA 4412 airfoil was transferred into Solidworks as a sketch. Solidworks allows users to create computer-aided designs, or CAD, which may portray a model of a structure. The CAD file can then be used in various printing methods to create a specified structure. Solidworks also allows users to configure the material used for the model and evaluate the mass properties of the structure to aid in design. The first iteration of the wing design was a simple extrusion of the airfoil to create the elliptical wing shape as previously designed. It was here that the team decided a rectangular shaped wing would be more plausible for manufacturing, thus, the design of the wing became a straight rectangular extrusion of the airfoil.

FigureXX: Extruded NACA 4412 Airfoil in Solidworks

The figure above shows the simple extruded airfoil in Solidworks with a 1.4m wingspan and 0.3m chord length. This model would be used for future operations pertaining to computer analysis. With this model, the team was able to gauge at how the wing would look like connected to the fuselage to help better understand how the wing would fit onto the drone. Further iterations involved changing the material to balsa wood, a lightweight and soft material often used in crafts and model airplanes, as the main material for the wing. This material will keep the wing light but be able to maintain structural integrity. Final iterations of the wing design involved using wing ribs and spars to develop the internal structure of the wing and keep weight levels to a minimum. Each rib will be ½ in. thick. Stringers were introduced to maintain the structural shape of the wing between the gaps of the wing ribs.

Figure XX: Model of Wing Rib

Various holes were implemented to reduce the overall weight of the wing and allow internal components to fit through the wing. Notches were implemented around the edges of the rib to allow slots for stringers, which will assist in maintaining the wing structure between the gaps of the ribs. The design intent was to coat the wing ribs with Monokote to give that solid structure of the wing while maintaining the lighter weight. The Monokote also serves to house all internal components within the wing structure.

FigureXX: Closer Look of Wing Design

Figure XX shows a closer look at the right wing structure. Each wing side uses 12 ribs in total throughout the structure. 4 of the ribs have been cut towards the trailing edge to implement the aileron, which will act as the main control surface for changing flight directions.

Figure XX: 3D Model of Wing Structure in Solidworks

The figure above showcases the internal structure of the wing with ribs, spars, and stringers all a part of the design. The two extended spars serve to connect the wing to the fuselage connector piece and hold all wing ribs in place. The third spar towards the trailing edge serves as an axle for the ailerons. In the final assembly of the wing, monokote will be applied over the wing to give it a solid shape and retain the wing shape while also minimizing the weight.

To print the wing ribs, a laser cutter will be used to accurately cut a balsa sheet until all necessary ribs are accounted for. To print the ribs, a DXF file was used, which included an outline of the wing ribs as well as the empennage ribs.

To ensure the wing meets design expectations, extensive analysis of the wing structure was performed. To verify the fluid flow of air around the wing, a computational fluid dynamics analysis, or CFD, was performed. CFD is a type of fluid mechanics which uses numerical analysis to simulate the fluid dynamics acting on the wing during flight. This is important to gauge how free stream fluid flows around the wing and justifies whether the wing is generating enough lift to carry the aircraft. One program used for this analysis was ANSYS Fluent, a highly credible fluid simulation software.  

Figure XX: CFD Analysis of a Wing Section in ANSYS Fluent

The figure above showcases a CFD analysis of a cut section of the wing with 20m/s wind speed acting on the wing. The results provided further evidence that the wing was capable of generating the required lift for the drone.

Another important wing analysis includes the finite element analysis, or FEA to understand how the loads and forces are affecting the wing structure during flight. A stress analysis is important to gauge the risk factors for the design. The program used for this analysis was FEMAP, a software created by Siemens Simcenter.

Figure XX: Wing Rib FEA in FEMAP

        The figure above shows the FEA of a single wing rib computed in FEMAP. FEMAP allows users to view the stress of applied loads on an object. This analysis is necessary to understand how the structural integrity will hold up with the applied loads to ensure the wing does not break apart during flight. During flight, wings will experience different forces acting on the structure, such as wind force hitting the leading edge, and the weight of the plane acting in the downward direction. The wings are designed to carry a weight of 50 N, which is the expected weight of the aircraft at speeds of about 20 m/s.

Empennage

        The empennage is an important part of the modeling phase, as this piece helps stabilize the system. With a front-heavy drone, the empennage would require enough downforce to correct any dip the system might have in its pitch axis. Further consideration for pitch variation was considered in the elevator system which would be located on the empennage configuration. Two propellers would be added to either end of the wing, which the empennage design needed to take into account; the prop wash this system would experience needed to hold minimal effect on the empennage configuration. A boom tail was used to connect the empennage to the fuselage. Figure X.X.1 shows the design of the empennage which was analyzed, utilizing a single vertical stabilizer that will be printed onto the connector. The horizontal stabilizers will form the main body of the empennage.

Figure X.X.1: SOLIDWORKS design of empennage piece.

        Taking the wing’s dimensions and design into account, a few different equations were used to size and shape the empennage.

Figure X.X.2 shows the two formulas used to find the correct dimensions. Following an average horizontal tail coefficient of 0.5, the horizontal and vertical tail areas were gathered. From there, the areas were used with the chord length of the empennage airfoil to find relative wingspans for each stabilizer.

        After the initial sizing and modeling for the empennage, a CFD analysis of the empennage with respect to the full drone was analyzed. Streamlines were analyzed to show how the empennage may react, as shown in figure X.X.3, while a 2D velocity contour was taken to show the effects on the vertical stabilizer as shown in figure X.X.4.

Figure X.X.3: Streamlines of Empennage

Figure X.X.4: Velocity Contour of Drone.

These images help show how the empennage could affect the final design; too much drag could hamper the drone enough to cause trouble with the amount of lift generated, or potential downwashwash catching on the empennage. Looking back at figure X.X.3, the downwash seems to have little effect on the empennage as it dips below the upturned stabilizer. Prop wash would be another consideration to the design of the empennage, though the main wing’s span and distance of the propellers have little effect on the sole vertical stabilizer.

        Finite Element Analysis was the final step in the analysis of the Empennage. With one 3/16” diameter dowel as a spar, one 5/16” diameter dowel as the main spar, and a 3/16” diameter dowel for the elevator shaft, structural analysis will need to be utilized to test the stress capabilities of the empennage. Balsa hardwood was chosen as the main choice of material for both the main wing and the empennage for its relative lightweight capabilities. FEMAP was utilized to help show where stress was most prevalent. This information could be used to better place ribs and spars to help diminish any possible errors or faults.

The empennage is a critical component in the static stability of an aircraft, and must be tested to ensure its structural integrity during flight conditions. To test this, a similar process to that of the wing was performed. A full scale model was constructed for this test. The test consisted of securing the empennage upside down, so that the vertical stabilizer is under the assembly, and a 0.5 kg weight was placed upon either side of the horizontal stabilizer, to simulate a load force of approximately 10 N experienced by the horizontal stabilizer. This is significantly greater than the design lift of the empennage, providing a factor of safety of greater than one.

Propulsion

        Propulsion system modeling was primarily focused on three components; propeller, motor, and engine housing. While the engines were not outright, they served as the basis for the operating conditions operated at. Using a suite of CFD software, a propeller configuration consisting of three rotor blades of 0.350 m length was tested based on the thin airfoil theory. The primary objective was the verification of the propeller's effectiveness while operating at 6000 rpm, with an intake speed of ~15 m/s and a blade pitch angle of 20 degrees. The three software used for testing were QBlade, Ansys Fluent, and Star-CCM+. Software complications and testing goals varied throughout the design process, prompting the need for a range of programs.

Figure 36: QBlade Results for a Straight Blade and V-22 Style Design

QBlade was the first software used for propeller design. Based on XFOIL and XFLR5, QBlade is an open-source software that is purpose-built for propellers and turbines. This software allowed for streamlined initial propeller design development. Over ten different designs were tested using a variety of airfoils throughout them. It was found that a 0.350 m at ~5000 rpm and a blade angle of 20 degrees would produce up to 400 N of thrust individually, well over the required ~45 N total. The rotor blades were initially intended to be custom-made; however, it was later to buy aftermarket blades per the request of the Senior Design Advisor. Using basic symmetric airfoil blades to obtain conservative results, confirmation tests were initially done through Ansys Fluent. While tests done of the full propeller were erroneous and unusable, single-blade approximations found that the total thrust for a propeller blade at 5000 rpm was ~226.71 N.

Figure 37: STAR-CCM+ Results (top to bottom): Velocity Gradients, Total Thrust Monitor

High-quality CFD was done through STAR-CCM+. While it is the most accurate software of the ones used, meshing and simulation calculation times were obscenely high, incentivizing more minor adjustments to be tested in other software prior to testing on STAR-CCM+. For a propeller at 6000 rpm, with an intake speed of ~15 m/s and a blade pitch angle of 20 degrees, the following was found: Thrust of ~264 N, required Torque of ~22.5 N-m, and Propwash at ~45 m/s. It is worth noting that the Torque may be higher than the actual value due to the lack of any selected material.

Table 5: Tabulated Initial Thrust Measurements

After further development of the avionics and powerplant, it was decided to reduce the propeller size and switch to a fixed-propeller. As a result, additional CFD work was done to revise the propulsion system given the new design constraints. Given this, two propellers were determine to be ideal candidates for the aircraft. A heavy, 14x7 3-Blade propeller could operate at low rpms while producing significant amounts of thrust, however it would place notable strain on the engines. Alternatively, a 15x4E 2-Blade propeller would put little strain on the engines thanks to its lightweight design, but would have to operate at significantly higher rpms to achieve the desired effect. Q-Blade tests found the force produced to be 88.4 N at 6000 rpm and 88.3 N at 9000 rpm respectively. CFDs were repeated again on Fluent and Star-CCM+ with their results being approximately half of those found on Q-Blade.

Figure #: 15x4E Star-CCM+ CFDs

        These initial simulations were done based on the Hacker A40-10S V4 motor; however, lack of availability has resulted in plans to redo the simulations using updated engine specifications. However, using the above table as a reference, an early simulation of the updated propeller still shows desired thrust performance. As mentioned early, these results were designed to give a conservative estimate of the produced thrust values. Propellers typically use a cambered airfoil to produce large amounts of thrust; however, lack of information on the specific airfoil used resulted in a symmetrical airfoil being used as a worst-case scenario.

In spite of these extensive CFDs, it was deemed necessary to verify the results of the simulations. Because the tiltrotor is required to be VTOL capable, having the necessary thrust-to-weight ratio is crucial. Additionally, it is necessary to determine the optimal power settings for the engine for several reasons. Firstly, while the engine will be operating at max power settings during takeoff, it can not operate at these settings for the majority of the flight due to engine strain. The higher the power settings and the bigger the propeller, the more work the engine has to do resulting in overheating. While this can be somewhat alleviated with proper air airflow via intakes, returning to a lower throttle setting is ideal for cruise flight. Additionally, operating at maximum throttle results in increased power consumption. By lowering the power settings for long durations, the aircraft's endurance will increase allowing its flight time to be longer.

Table 22: Tabulated Thrust Measurements

        To secure the motor during testing, a 3D-Printed Thrust Stand was produced to allow for vertical testing. Placed on a scale, the weight of the stand and engine would be zeroed out to allow for accurate thrust measurements. Four different motors were initially selected for testing: the E-flite 32, E-flite 60, Hacker A40-10S V4 14 Pole, and the Cobra C-4120/16. During these thrust tests, several factors determined our final engine selection:

• The E-flite 60 was found to be too heavy given the weight constraint: it weighed in at 0.838 lbs while the other engines weighed around ~ 0.64 lbs
• The E-flite 32 and Hacker A40 could reach 10lbs with thrust but would fail soon afterward due strain from both the propeller and the battery.
• The Cobra-C4120/16 could achieve the desired thrust values using either blade, allowing for interchangeability.

Table 22: Tabulated Final Thrust Measurements

Modeling of the engine housing was done through Solidworks with FEA analysis down through MSC Nastran. While it may appear to be an auxiliary component, it plays a crucial part in ensuring the aircraft is tiltrotor capable. In addition to serving as the engine mount, the housing also houses a crucial servo that serves to change the engine’s orientation from vertical to horizontal.

Figure ##: Engine Pod Assembly with motor and 14x7 propeller.

        In order to allow for ease of assembly, the housing is split into three components, two outer shells to help mitigate drag and a central, load-bearing section which the engine and servo mount to. While the design is asymmetric to allow for the addition of gears, it can be flipped to allow for use on both sides of the aircraft.

Figure ##: Deconstructed Engine Housing - Outer (Top Left), Mid (Top Right), Inner (Bottom).

        FEA was performed on the middle section to ensure it could take the needed amount of load. Initially, MSC NASTRAN was to perform these tests with the results seen below. A 10lb load was applied on the servo mounting holes to account for the weight of the aircraft and an additional 10lb load was applied to account for the front engine mounting plate to account for the pull of the motor while it was producing thrust.

Figure ##: FEA on the Mid Section of the Engine Housing.

        The design was ultimately redesigned during the manufacturing stage for several reasons. To start, reduced manufacturing time due to parts delays limited the means of manufacturing available. As such, the housing had to be redesigned to be compatible with standard 3D printers. Alongside this, changes made to allow for compatibility reduced the housing’s structural integrity. As such, a lighter shell design was abandoned for a stronger but heavier rod design. Because of this, many of the avionics that would have been stored within were now exposed to the outside. Even so, the new housing's solid design allowed it to support the load caused by the engine.

Figure #: Finalized Housing with Engine and ESC

Fuselage

The fuselage system for this design is a combination of an airframe, an outer shell, and landing gear. The airframe serves as a structural skeleton in which all major subsystems are mounted to, such as the wing, fuselage shell, and tail connecting the empennage. The fuselage outer shell has three main functions: to house the internal avionics, power supply, and control surface mechanisms, providing an aerodynamic body to protect said internals and to reduce the drag around the chassis. Connected to the outer shell is the landing gear which adds a light suspension system for the UAV to rest on so that the fuselage belly is not exposed to a risky impact when landing. The satisfaction of these three systems will result in a solid design for the UAV fuselage to handle various stresses during flight and landing while maintaining a drag efficient geometry.

 With these functions established, the team developed each system and their components beginning with design requirements of withstanding a payload of 10 lbs, providing a lift of about 40 N and an overall structural deflection of less than 5mm. Each part was then modeled using Solidworks (Figures 1, 2 , 3, 4 and 5) and mated piece by piece to verify fitment and function as the assembly gradually developed.

Figure 1 XX - Airframe Base Connector, Solidworks

Figure 2 XX - Airframe Boom, Solidworks

Figure 3 XX - Airframe Tail Connector, Vertical Stabilizer Attached, Solidworks

Figure 4 XX - Fuselage Shell, Solidworks

Figure 5 XX - Fuselage Assembly with Empenage, Solidworks

Once the part designs were sufficiently optimized to be practical, the team moved on to analysis, simulating the expected environment and loads each subsystem would endure. The airframe, fuselage shell and landing were processed through a finite element analysis using Siemens FEMAP program with the computational method of Simcenter Nastran (Figures 6, 7 and 8)  applying respective shear and torsion forces determined through basic static distribution equations. With the airframe being a mostly internal component, and the landing gear having a significantly small geometry compared to the overall drone assembly, they were exempted from being put through CFD analysis. The fuselage shell, however, is crucial to the drones airflow and needed to be verified in order to reassure that it does not impede on the flow of air to the wings while also contributing minimal drag (Figure 9).

Figure 6 XX - Airframe Base Connector Stress Analysis, Siemens FEMAP

Figure 7 XX - Airframe Boom Stress Analysis, Siemens FEMAP

Figure 8 XX - Airframe Tail Connector Stress Analysis, Siemens FEMAP

Figure 9 XX - Fuselage Fluid Flow Analysis, Siemens Star CCM+

When designing the fuselage, the main parameters to keep in mind are the stress concentrations from shear forces and torsion, as well as the skin drag coefficient for the fuselage outer shell. Given the criticality of these parameters, the modeling and simulation analysis needs to be carefully executed. The model is ensured to represent an accurate depiction since every part generation has been developed around pre existing market material, and for the custom parts, the manufacturing method has been verified with respective uncertainty values. The other half to maintaining accuracy in our modeling is to represent the proper material parameters in our simulation, such as ultimate tensile strength, shear strength and Youngs Modulus at environmental temperature, as can be scene for Carbon Fiber infused PLA in Figures 10 and 11, in combination with appropriate external conditions like the previously calculated shear and torsional forces happening on the body, as well as the designed wind speed. With all these bases covered, the preliminary data from modeling and simulation prove to be reliable and appropriately used for the design phase.

Figure 10 XX - Material Properties of Carbon Fiber - PLA, (Maqsood,  Rimašauskas, 2021) [XX]

Figure 11 XX - Material Properties of Carbon Fiber - PLA, (Maqsood,  Rimašauskas, 2021) [XX]

The fuselage housed most internal electrical components, and it was absolutely crucial that it remained intact, even in the case of a crash. It is for this reason that a test of the fuselage’s crash resistance is necessary. A partial prototype was built for this test. The fuse was constructed at full scale, and a weight of 5 kg was placed upon the top of the fuse. This simulated if the fuse could hold the weight of the entire aircraft while it is grounded then, this fuse with the weight attached will be drop tested at a height of 10 feet. The experiment was repeated until the structure could pass the test and was finished after 3 successful attempts. A small scale fuse was not produced for this experiment as the fiberglass layering cannot be scaled down with the fuselage, because the results of the test would be unreliable for a larger model.

Avionics

The aircraft, being a VTOL, had two flight modes: fixed wing, and rotorcraft. In fixed wing mode, the roll was controlled by the ailerons, the pitch by the elevator, and the yaw by a differential in motor rpm, creating a torque on the aircraft resulting in a level turn. In rotorcraft mode, roll was controlled by a differential in motor RPM, pitch by tilting both motors forward, and yaw by tilting both motors in opposite directions by equal amounts. The ailerons and elevators were both controlled with the use of servo motors. The servo motors controlling the ailerons were placed within the wing ribbing, directly mated to the aileron axis of rotation, and the servo motor controlling the elevator was within the fuselage, connected to the elevator through a lever arm.

The microcontroller used for the circuit is an Arduino Mega. This microcontroller was chosen due to its abundance of interrupt pins, allowing for a larger amount of interrupts in the program, making it more responsive. The Arduino Mega will take inputs from a six channel FS-IA6B receiver. To operate the aircraft, only 5 inputs, and thus, are needed: roll, pitch, yaw, RPM control, and flight modality switch, meaning that at least 5 radio controlled channels are required for operation. It is for these reasons that a six channel receiver was chosen. The controller selected to send signals to said receiver is the FLYSKY FS-i6X 10 channel 2.4GHz controller. This controller was optimal as it provides an intuitive controller layout for the pilot and is compatible with the chosen receiver. The gyroscope used in this circuit will be an MPU6050, a 6 degrees of freedom accelerometer. It will provide three linear acceleration and three rotational acceleration measurements. For the FPV implementation, a 120 degree camera and 5GHz transmitter will be used. This will provide the pilot with a clear view of what is ahead of the aircraft allowing for safer control at further distances. The transmitter will send video to a 4.3” monitor on the ground that the pilot will have access to.

The avionic system will be powered by a 4S LiPo battery. The 4S LiPo battery will act as the power source for both of the electronic speed controllers, and one electronic speed controller will utilize a battery eliminator circuit to power the remainder of the accessories. As pictured in Figure XX, the 4S LiPo will be connected to two 60 A electronic speed controllers in parallel, which in turn are each connected to a motor in parallel. One of the ESC’s will deliver power to the rest of the avionics wired in parallel. The circuit was designed in this manner to provide each component with its required input voltage. The decision to power the system with one battery and a BEC was to save weight on the final design.

Figure XX - Circuit Pictorial Diagram

 Each motor will draw a maximum of 60 A according to their specifications. This, in combination with the distance between the ESC and the battery stored in the fuselage factored into the decision to use a 10 gauge copper wire between the two. This is because the resistance within a wire will increase the longer the current has to travel for, which has the potential to cause major inefficiencies or even component failure without the proper wire gauge. The size for the wiring was chosen according to the chart in Figure XX.

Figure XX - Wire Gauge Chart

All components will be mounted securely on a wooden plank within the fuselage to allow for increased modularity and ease of access. The batteries, being the heaviest components, will be placed ahead of the aerodynamic center to achieve static stability requirements.

It is crucial to prototype the avionic components of the aircraft to ensure proper functionality of each part, as well as to measure their limits. The prototyping and testing of these components will involve many different methodologies.

        To choose an optimal battery capacity, testing of the batteries will consist of measuring current draw from a battery during normal operating conditions. To test the battery, a propeller motor and ESC must be connected to the battery and the motor must be spun at the average operating RPM. Then, current drawn from the battery would be recorded via a multimeter. With this value, the necessary battery capacity for a design requirement operating time can be calculated with the formula:

After a battery of the required theoretical capacity is acquired. Both motors will be mounted and powered on to their average operating RPM until the battery voltage drops to 3.2 V per cell, the lowest safe voltage for a LiPo battery, and would therefore be considered drained.  The time it takes for this to occur will be measured with a timer to ensure design requirements are fulfilled. A partial prototype of the circuit only involving the propeller motor and ESC will be sufficient as the current draw from the avionic component is negligible in comparison to the motors.

Control

In order to control pitch, the elevator must be controlled by a servo motor. To do this, a servo will be mounted within the fuselage, and control a lever arm extended to the elevator connected by an adapter, as shown in Figures XX and XX. The action of the servo rotating will cause the mounting points to be laterally pulled toward or pushed away from the fuselage, causing a pitch down or pitch up response, respectively.  

Figure XX - Elevator Control System, Pitch Up

Figure XX - Elevator Control System, Pitch Down

Figure XX - Elevator Control System, Assembly

An FEA on these parts was done to ensure they would theoretically be able to withstand the pulling force being applied by the servo. To do this, a theoretical drag calculation was done on a rectangle of equal area to the elevator at the maximum operating speed oriented normal to the air flow, maximizing drag, giving a theoretical maximum drag of 5.16 N acting on the entire elevator. This was found by assuming a coefficient of drag of 1.28, that of a rectangle oriented normal to a flow, and applying the drag equation:

This drag would be equivalent to the force applied by the servo motor to keep the elevator stationary. Then, an FEA on the elevator was done with a 15N force applied where the lever arm connects. This force was chosen to guarantee the elevator and associated linkage components remain structurally sound even under a stress higher than the theoretical maximum. The results of the FEA as shown in Figures XX and XX shows that the maximum stress experienced by the elevator will be 1,280 Pa, and by the servo lever arm will be 1196.6. The parts will be 3D printed with PLA at a 100% infill, which has a tensile strength of 37 MPa, which is significantly greater than the stress experienced by the parts. Therefore, the parts will theoretically remain structurally intact even under the highest possible stress. To guarantee this, and account for imperfections in the material caused by manufacturing processes, further physical testing must be done.

Figure XX - Elevator Stress Analysis, Forward Facing Force, Siemens FEMAP

Figure XX - Elevator/Aileron Servo Lever Stress Analysis, Siemens FEMAP

        In order to control the ailerons, they must be controlled by servo motors. To do this, a similar lever system applied to the ailerons will be used as shown in Figure XX. The servo will be mounted within the ribbing of the wing, and be attached to an extrusion on the aileron by a lever arm, as shown in Figure XX. To test the theoretical structural integrity of this system, an FEA analysis on the aileron will be conducted. To do this, a similar theoretical drag calculation  to that of the elevator was done, resulting in a maximum of 10.78 N of drag per aileron at a flight speed 25 meters per second, As shown in Figures XX and XX, FEA done on this component with an applied force of 15 N shows that the 936.1 Pa of stress experienced by the aileron and the 11452 Pa experienced by the aileron-servo linkage due to the servo pulling against a drag force is significantly less than the tensile strength of 100% infill PLA. The model the FEA was performed on does not include the actual aileron control surface, as there were issues meshing the geometry. However, it is expected that the lever arm will experience the majority of the stress, and the stress acting directly on the aileron is negligible. The results of this stress analysis will have to be confirmed with a physical experiment.

Figure XX - Aileron Control System, Assembly

Figure XX - Aileron Stress Analysis, Forward Facing Force, Siemens FEMAP

Figure XX - Aileron-Servo Linkage Stress Analysis, Siemens FEMAP

        In order to control the propeller angle, a gearing mechanism will be employed as shown in Figure XX. The propeller housing was designed to be strong enough to withstand large torques applied on the surfaces making contact with the high torque servo (red) powering the gearing mechanism. The assembly includes a blue motor and ESC for mockup purposes. Figure XX features the first iteration of gears designed, however, further trial and error was conducted to ensure proper meshing of the gear teeth. The iteration process involved modeling the gearset on Fusion 360, a CAD software, using a spur gear tool to automatically generate gears based on specific parameters. Initially, both gears were designed with a pitch diameter of 38mm and 32 teeth, however, further iterations involved keeping the spar gear’s module constant while increasing the tooth count to slowly increase the pitch diameter of the spar gear until a perfect mesh was achieved in the real assembly. In the end, a gear ratio of 32:35 proved to be sufficient to create a perfect tooth mesh.

Figure XX - Propeller Pitch Control System        

To guarantee the integrity of the elevator control system, an FEA of the adapter connected to the elevator was conducted to guarantee its structural integrity when a drag force is applied to the elevator. To test its structural integrity, the elevator was secured such that a weight hanging from the mounting hole for the lever arm would apply a force toward the leading edge of the elevator. Then, a 1.5 kg weight was secured to the lever arm hole. This simulated a stress caused by a drag force experienced by the elevator causing a reaction force in the lever arm acting on the mounting hole. The maximum theoretical true drag force experienced by the elevator is significantly below the test weight that was applied, and the part experienced no significant damage, therefore giving the part a factor of safety greater than 1.

To guarantee the integrity of the aileron control system, a similar experiment to that of the elevator was conducted. A calculation similar to that of the elevator was done on the aileron control surface and a maximum theoretical drag force of 10.78 N was found. To test the true structural integrity of this part, the aileron was secured such that a weight hanging from the mounting hole applied a force toward the leading edge of the aileron. Then, a 1.5 kg weight was attached to the mounting hole. The part experienced no significant deformation or damage, therefore granting the part a factor of safety of greater than 1.

Once all partial prototypes were manufactured and tested, a full scale prototype was constructed to test the functionality of the control surfaces and tilt rotor mechanism. The theory behind the construction is simple, but a full prototype was necessary to ensure the fitment between parts and functionality of all subsystems was harmonious.

8. Final Design and Engineering Specifications

        After several revisions in design based on testing and manufacturing constraints, the tiltrotor’s design was finally complete. Manufacturing was done over the span of two weeks with wheeled landing skits being built from leftover parts to reduce budget costs. In total, the tiltrotor cost $684.53 to build which falls within the stakeholders budget (See Appendix B for a detailed breakdown). While most of the costs were offset by the Senior Design Parts Request process, a sizable portion was funded independently by the team itself.

Figure #: Finalized CAD (Left) with the propless full assembled design (Right)

        After complete assembly, the UAV weighed in at 13.5 lbs and had a wingspan of 4.59 ft (1.4 m). Despite exceeding the anticipated weight, the design was still capable of achieving lift thanks to its twin Cobra engines which produced a combined thrust of ~19.08lbf. Because of this, the final was VTOL capable, meeting one of the most crucial design requirements. While many of the changes made to the design did come at the cost of additional weight, they served to significantly improve the structural stability of the aircraft.

Table #: Final Wants & Requirements Status

        In spite of several setbacks with acquiring parts, in the end  the majority of required parts were obtained. FPV equipment was not ordered due to concerns regarding weight and complexity. Meanwhile the gyroscope was ordered but never received. In total 40 components made up the tiltrotor UAV with 29 being a prebuilt or aftermarket. Even manufactured parts were reliant on parts orders in order to obtain the needed materials and tools for manufacturing. As such the design was heavily affected by obstructions in the parts acquisition process. Manufacturing was done at the Innovation Lab and at the homes of certain Team Members which help minimize the cost associated with this process. Laser Cutting and 3D-Printing made up the bulk of the manufacturing done, allowing it do be done in a variety of facilities.

Table 6: Project Supplies

9. System Evaluation

Thrust Tests

Goal:

        Thrust tests were conducted to finalize the engine configuration for the tiltrotor. Testing a variety of engines with different propeller setups, thrust values could be recorded to find the ideal solution. In addition to finding the highest thrust configuration, these tests were done to study the strain and power drain of the engines. This was done to ensure the engines would reliably operate during the duration of the flight.

Procedure:

        Motor would be mounted on the thrust stand which would then be secured to the used scale. Afterwards, the propeller would be mounted onto the engine prior to wire connections. Connected to the battery through a potentiometer, it would then be turned on at the minimum power settings. Gradually increasing the power, it would eventually reach its maximum power settings which would be used for testing. While increasing the power of the motor, the scale will be observed to determine the amount of thrust the motors generate.

Equipment:

• Engines:

• E-Flite 32
• E-Flite 60
• Hacker A40-10S V4 14 Pole
• Hacker A40-10S V4 8 Pole
• Cobra C-4120/16

• Propellers:

• Master Airscrew 14x7
• APC 15x4E
• Master Airscrew 16x10

• Powerplant:

• 5S 5000 mAh LiPo

• Measurement:

• 300lb Analog Scale
• 400lb Digital Scale
• Custom Printed Thrust Stand
• Caliper

Hover Tests

Goal:

        To test hover stability prior to final flight tests given the risk and difficulty of maintaining VTOL mode as it would be used for takeoff and landing. It would also allow for controller and program refinements in addition to verifying the weight and balance of the aircraft.

Procedure:

        The tiltrotor was placed in an open space to ensure the drone did not crash into any objects. After standing a safe distance away, the motors were turned on while in hover mode and slowly increased in thrust until having enough power to generate lift. If the drone did not produce enough thrust or started to drift, fine tuning was done to the program before testing again.  

Equipment:

• Completed Tiltrotor
• Remote Controller
• Fire Extinguisher

Stability Testing

Goal:

In hover mode, the aircraft must be able to be supported solely by the wing spar without a significant change in pitch relative to the ground. To ensure this, the center of gravity must be directly below the center spar of the wing, where the propeller housing is mounted.  

Procedure:

To test the stability of the drone, the full assembly was placed on the main spar. This was used to test the stability based on where the center of gravity would be located during hover mode. The initial reaction would be observed, whether the drone leaned forward (Front-heavy) or backward (Tail-heavy). After noting the lean, weight would be added to compensate for the change in moment. This process was repeated until static stability was achieved.

Equipment:

• Completed Tiltrotor
• Extra Weights

Failure Mode Risk Assessments

Risk A: Drained/Damaged Battery

• Risk Description: The battery used could be damaged from excessive testing and could also be drained mid flight.
• Possible Causes: Overcharging past 4.2 volts per cell, draining the battery below 3 volts per cell, an uncontrolled crash, lack of caution with the battery, operating the aircraft for longer than the maximum flight time, or a faulty balance charger
• Possible Consequences:  Damage to the battery could end up puffing the battery or igniting and possibly cause an explosion. Overcharging could lead to the battery draining at a quicker pace or can lead to puffing. A puffed LiPo battery is at a significantly higher risk of catching fire.
• Mitigation Plans: Batteries require proper care so in order to mitigate the risks, they must be handled with caution, charging and discharging them periodically, keeping the cell voltage between 3.2 and 4.2, , and keeping them stored at recommended temperatures to avoid overheating.

Risk B: Propulsion

• Risk Description: The propulsion system of the drone has a major role on whether the flight can be successful. The biggest risk regarding this system is that it is not able to supply enough thrust. Also, this system must be able to provide enough thrust for a long enough flight time to be able to perform the functions. If either of these risks occur, it could lead to failure in lift off or possibly result in a crash. A smaller risk that could lead to propulsion failure but can be easily mitigated would be proper installation of the propellers, motors, ESC, and servos.
• Possible Causes: There are multiple causes of propulsion failure, the main ones resulting from motor or propeller failure. Since the propulsion system uses an electric motor, the motor or the ESC could result in overheating and draining the battery resulting in failure to supply enough thrust to the props.
• Possible Consequences: As stated above, if proper installation on the propellers is not successful it could lead to having a defective prop that may be subjected to not reaching the target RPM’s and not generating enough thrust needed for takeoff. Another consequence that could affect the propulsion is from the overheating of the motor or ESC’s being used, that would lead to a drained/dead battery resulting in failure to supply any thrust for propulsion
• Mitigation Plans: To minimize the possibility of improper installation on the propulsion system components like the motors, ESC, and batteries,  they will be secured in the fuselage and strapped down to ensure minimal movement to allow connections to stay secure during the flight. To mitigate the risks of the motor or ESC from overheating, the components will be tested at max thrust to ensure they are capable of handling more than enough voltage and current supplied.  

Risk C: Structural Integrity

• Risk Description: Structural integrity is the key to having a successful flight, the biggest challenge to maintain the durability of the drone is having a lightweight structure but still strong enough to withstand high stresses and strains acting upon it and avoiding material deformation.
• Possible Causes: Structural failure can occur from a variety of different things but the primary reasons for it may occur from weak structures that are not strong enough to withstand the loads it is faced with. Structural deterioration is another key factor that leads to structural failure which could be deteriorated from wear or fatigue from too much stress acting on one spot. Another cause of structural failure could be from manufacturing errors or defective material that cannot support the high stress it was designed for.
• Possible Consequences: The consequences of structural failure are substantial because it increases the risk of flight failure for the whole drone. If failure occurs in the wing the result could be critical because it is the main component of our drone and will lead to a crash. If structural failure may occur in the spar or part of the ribs, it could result in an unstable flight but may be manageable to have a safe landing.
• Mitigation Plans: To decrease the possibility of structural failure, the group must calculate the integrity of structures before deciding on a material. This includes running finite element analysis simulation on various materials with high loads acting upon it to determine if the material can withstand high stress. After running stress simulations and deciding on a preferred material the structural integrity of the material can be tested by running a bending test which is used to evaluate the ductility and soundness of it.

Risk D: User Pilot Error

• Risk Description: One of the biggest risks that could cause the most damage to the drone or bystanders around, is the possibility of the aircraft crashing into obstacles within the airspace. A lot of factors out of the pilot’s control could occur when flying the drone.
• Possible Causes: There are a variety of possible causes that could lead to pilot error when manning a drone, many of them out of their control. Some source errors that could occur could be from wiring errors, controller incorrectly configured, or even moisture on the circuit board that could affect the whole system. Inexperienced pilots might also cause issues with the aircraft by stalling the system due to excessive pitch or turning. Some other causes while flying could be from a gust of wind from bad weather, a pilot oversteering the aircraft, or running into an obstacle they didn't know was there. Many of these cases are not in the pilot’s control but still must be accounted for.
• Possible Consequences: With many of these causes stated above, a lot of them result in the same way, which is the possibility of damaging the drone. In the case of bad weather, heavy winds may affect the flight of the drone and possibly affect the flight course.
• Mitigation Plans: To mitigate the risk of crashing the drone, the airspace should be observed preflight to consider any factors that could be a danger to the drone. Also, to avoid flying in bad weather conditions, the weather will be monitored during the week of testing to ensure proper scheduling. In the case that during the flight weather conditions disrupt the drone’s flight stability the pilot will have fail-safe scenarios to ensure safety to bystanders.

Risk E: Avionics Error

• Risk Description: Without testing, a line of code could be written wrong which might result in a piloting error or change in engine capabilities. Another risk could come from interruption between the receiver and the remote or even just a program glitch. Hardware risks of avionics error could also include disconnection of wires, various parts overheating, and servo gears or different parts breaking.
• Possible Causes: These risks could cause the user inputs on the remote to not communicate with the plane while trying to fly it. Another possible cause could come from outside signal interference that disrupts the connection due to the fact the system is wireless.
• Possible Consequences: If the remote and receivers lose communication the drone could possibly shut down due to no connection and possibly result in a crash based on the amount of time the connection is lost.
• Mitigation Plans: To mitigate the risks of losing connection to the receiver, careful examination on all avionic components will be tested and examined numerous times to ensure a secure connection.

Risk F: Lift Capabilities

• Risk Description: One of the main requirements for the drone’s flight is its lift capabilities. Without enough lift and thrust, the system will never leave the ground. Two different lifts needed to be calculated for this: The lift of the system during cruise flight and the thrust of the engines during liftoff/takeoff. 45 Newtons was the expected lift requirement.
• Possible Causes: Possible causes for underperforming lift are ripped monokote/compromised wing surface, and inefficient thrust generation by the motors. A compromised wing could drastically diminish the lift that the piece can generate, as well as create an unstable surface along the wingspan. Inefficient thrust would generate less thrust force than needed to lift the aircraft, which would stop the test flight at the beginning.
• Possible Consequences: Possible consequences vary for this risk. As stated above, an inefficient force of thrust wouldn’t be able to lift the entire system off the ground during takeoff. While the dangers of this are low, the system would prove moot if it cannot achieve the desired lift. With a tear or break in the wing, a less severe consequence would be a noticeable dip during flight. The pilot would be required to perform a controlled crash to safely get the drone onto the ground for repairs and analysis. A dangerous break or tear could cause unstable flight which would be difficult to control. This could lead to crashes on the ground as well as into obstacles. This could damage the interior of the system as well as prove dangerous to civilians/onlookers.
• Mitigation Plans: One major plan to mitigate this failure mode is to use a strong, abrasive-resistant material on the wing to better protect against any weathering and wearing. Monokote was used for its lightweight and durable properties, as well as its relative ease of application to the wing; in the case of an unforeseen tear, repairs wouldn’t be too difficult. Motor tests and analyses were performed to benchmark the ability of the system’s engines. This helped show that the motors used could achieve the desired lift.

Risk G: Flight Stability

• Risk Description: Flight stability is extremely important when flying the drone, more specifically during liftoff as well as stationary in hover mode. Hover drift can hold many consequences when it comes to proper landing and takeoff, while an unstable flight while in cruise could cause unwanted movements.
• Possible Causes: An unstable flight can be caused from a high center of gravity or a low moment of inertia which may come from calculation errors or the placement of your components on the drone. A weak signal may also affect the stability of the flight due to certain obstacles that may affect the drone’s signal connection.
• Possible Consequences: If the aircraft has a poor connection, the pilot may not have full control which could result in the drone not responding efficiently to the controller and could fly erratically. If the center of gravity is off the drone’s flight performance could be diminished due to uneven distributed weight causing tilt throughout the drone
• Mitigation Plans: To minimize the risks of an unstable flight a gyroscope will be implemented in the drone to measure the rate of rotation and help keep it balanced and maintain a reference direction. Also, the center of gravity will be calculated before and after assembling while making sure weight is evenly distributed throughout the drone to prevent unwanted tilting.

Risk H: Flight Transition

• Risk Description: A main component of a Tilt-Rotor UAV is the ability to transition from its vertical lift mode to its cruising flight mode. This is done by rotating the propellers and motors from a 90° position all the way back to the origin (0°), or vice versa. Servos will be mounted on the engine housing and will provide the required torque and mobility needed for this transition.
• Possible Causes: Possible causes for improper or incorrect transitions could be faulty servos for rotating the engines as well as a lower traveling speed than required while transitioning. If calculations were incorrect or the required torque isn't achieved, there is a possibility of the engines being at an improper angle for cruising flight. Connection issues and battery issues will also dampen the servo and could freeze the engines in lift mode.
• Possible Consequences: The consequences of an error in flight transitioning could be pretty crucial to the whole drone’s integrity. Improper torque could affect the drone’s ability to maintain either mode. This effect would lead to improper lift with the possibility of a crash, or the inability to fly the drone.
• Mitigation Plans: Mitigation of this failure would be proper testing and analysis of the servos used to transition the engines. Understanding how much torque each servo can withstand as well as the range of motion they can achieve can help determine the ability of the servo to achieve proper rotation.

Risk I: Connection Error

• Risk Description: A connection error is a dangerous failure during the flight of the system. A severed connection will break all communication between the pilot and the drone, so proper equipment and measurements are required. This type of error can be both machine oriented as well as a human error. The likelihood of this failure mode is average, as most causes require multiple steps to occur. A disconnected system could prove extremely severe, as the pilot could lose all control of the system. This could cause a crash at worst. Sudden connection errors are rather detectable as the pilot will most likely be working with the system at the time of failure.
• Possible Causes: Multiple causes of this failure are possible. Physical interference could take the form of clouds, trees and buildings. These all have the potential to block the signal of the controller. Other causes could be secondary signals disrupting the controller, low battery causing an unstable connection, and poor quality of equipment.
• Possible Consequences: Depending on when the connection is severed, there are multiple different consequences. A connection error mid-flight would disconnect any control the pilot might have. This could be dangerous with less open airspace, as there is potential for the drone to crash into a tree, pole, or any obstruction. With an open air space, the system will have time to glide while waiting for the connection to reestablish.
• Mitigation Plans: Proper equipment and a weather forecast are the best plans for mitigation. Understanding the best time to fly the drone could diminish any weather interference, while high-end equipment provides better connections. Understanding the drone’s flying space is important to make sure no object has the potential to disrupt a signal. If the controller disconnects the designed program will include disconnect detection, causing the motors to turn off to prevent a runaway plane or possibly programming to auto land by transitioning to hover mode and landing if disconnected.

Risk J: Faulty Manufacturing

• Risk Description: With certain parts, precise manufacturing could be crucial. Minor blemishes to parts could prove to create unnecessary risk during flight. A warped rib, or fractured spar increases the potential stress on the specific parts, which could in turn heavily decrease the weight-carrying capability of the wing. Manufacturing also has human error capabilities, as improper application of parts could waste the material and incur unnecessary costs towards the total budget through parts reacquisition and secondary orders. The likelihood of this potential failure is low since most dimensions are measured digitally and will be machined for precision purposes. The severity of this fault is dependent on the faulty part, as some deformities hold little impact on control and performance. Minor blemishes on the fuselage might only affect appearance while a splintered rib can jeopardize the entire wing. This failure mode is easily detectable without the need for software or machines.
• Possible Causes: As stated above, there are multiple causes that could create a manufacturing error. Human error will increase the likelihood of improper design, as well as unmeasured pieces. Incorrect measurements are another potential cause, either through human measuring or incorrect use in modeling software. The most probable cause of faulty manufacturing would be natural deformities and defects with the ordered parts.
• Possible Consequences: The consequences of this fault range from inconsequential to detrimental.
• Mitigation Plans: Machined parts will be measured through software to mitigate any potential difference in parts. A laser cutter will be used for any wooden design, so proper measurements will need to be taken in the modeling software. Units must be double-checked to account for any changes as well as a single unit for any one form of measurement. Metric measurements were first used before switching to imperial units. This was for parts ordering as well as conforming to a single measurement standard.

10. Significant Accomplishments and Open Issues

        The design and manufacturing process was fraught with countless issues which drastically altered the project's timeline. From difficulties with parts acquisition to the need for constant design revisions and updates. Despite these issues, the tiltrotor came together and hover flight was achieved. The following sections will detail the issues encountered and what solutions, if any, were implemented.

Part Acquisition Problems:

        The inability to acquire requested parts had a catastrophic effect on the design and manufacturing process. The initial manufacturing timeline was significantly delayed, pushing back critical testing needed to finalize the design. This issue was the result of several external factors which could only be corrected by the team personally funding the project. Of note, there were three critical issues plaguing the acquisition process: orders going unfulfilled for extended periods, UCF closing down several times, and wrong parts being sent. Anticipating issues with parts requisition in addition to several critical planned tests, an initial parts request form was submitted as early as September 14th. However, these initial parts were finally received on October 13th, pushing back the UAV’s development by a whole month. While the cause of this delay is uncertain, it is believed to have been caused by turmoil with the handling of the Senior Design Inventory. As a result, the final design could not be frozen due to the required tests being subsequently delayed.

This issue was further compounded by the occurrence of two natural disasters in the following period of time: Hurricane Ian and Nicole. The two storms forced UCF to shutdown the campus for several days at a time, resulting in the Senior Design Inventory to be closed as well. During this period no orders could be processed or received, delaying vital parts requests at a critical point in development. These problems were worsened by improper parts delivered, leading to the wrong parts being received. In particular, motor testing was done with the assumption that the motor was the type we had ordered, in addition to being the same as the one we had previously acquired. However, upon further investigation, it was discovered that we had received a significantly weaker motor than was believed.

Redesigns of Parts and Assemblies:

        One issue faced with developing the drone included weight distribution of various components. After some components of the drone changed materials, weight became an issue for the wing, fuselage, and tail. The design of the wing had to lose a significant number of ribs as the material and thickness of the wood used for the rib changed from ⅛ inch balsa wood to ¼ inch plywood. Redesigns of the ribs changed significantly to account for stress acting on each rib. It was important that the 1.4 m wingspan remain as an intended design choice. More open holes had to be added to the fuselage after realizing that the material used for 3-D printing would be heavier than intended. Some weight alterations were made for the tail and tail boom to reduce the weight. Confirming with the team and careful attention to detail will ensure the designs meet expectations. Design tweaks are natural when testing parts for weight and structural integrity.

11. Conclusions and Recommendations

         Although the drone was successfully built, and all parts assembled, testing of the ailerons saw multiple failures in the functionality of the servos. The servos, being mounted within the wing, became difficult to reach for fixing. Earlier tests showed the successful functionality of the control surfaces and guaranteed the ailerons would work after applying the monokote over the wings. Failure in the servos happened right before an initial flight test would occur. Without spare servos to replace the dead ones within the wing, fixed wing flight testing was in question. Until a resolution is found, hover testing seemed to be the next best option for testing the functionality of the motors and viewing the stabilization of the aircraft. What we found was that the overall weight of the drone far exceeded the intended payload, thus, calling for weight reduction in the tail, to best suit the position of center of gravity, which we wanted to be under the main spar that held the rotors. To decrease the weight of the tail, some of the tail boom was cut off.

Thrust testing of the motors on the drone saw that the right motor was producing more lift than the left motor, causing a spinning effect when attempting to fly. This led to a stabilization issue that could have been solved if gyroscopic functionality was implemented into the drone. This and raw inputs into the controller proved the most effective way to better control the drone once lift was achieved, though there were still some difficulties with drift and roll.  

With an efficient amount of lift achieved in hover mode, transition mode became the largest hurdle. This mode, as well as cruise mode, were considered too dangerous to test with the drift that was seen. To counter this, extensive testing would need to be performed during hover mode to better counter the instability that was seen.

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Appendix A: Customer Requirements

Table 5: Customer Wants/Functional Requirements

Appendix B: Expense Report

Appendix C: List of Manuals and Other Documents

http://www.innov8tivedesigns.com/images/specs/Cobra_4120-16_Specs.htm

https://content.arduino.cc/assets/UNO-TH_Rev3e_sch.pdf?_gl=1*1la8fea*_ga*MTMyNzQwNDc5MS4xNjY5OTUzNjMy*_ga_NEXN8H46L5*MTY2OTk1MzYzMi4xLjAuMTY2OTk1MzYzMi4wLjAuMA..

https://store-usa.arduino.cc/products/arduino-uno-rev3 

Appendix D: Design Competencies

Include the AE and/or ME Design Competencies and Competence Criticality Matrices

Table 25: Aeronautical/Astronautical ABET Table

Table 26: Mechanical ABET Table

Table A3: Mechanical/Aeronautical/Astronautical ABET Page Number Table

Individual Tasks:

• Jasper Gong: Designed wing structure and assembly. Tested multiple airfoils in XFLR5, developed data on angles of attack best suited for the drone, modeled wing geometry.Performed wing simulations and analysis such as CFD in AnsysFluent and FEA in FEMAP. Did CFD on wing to analyze flow characteristics around airfoil. Did FEA on wing ribs to identify stress distributions throughout rib structure. Assisted in wing assembly with Chris and Nicholas, provided wing structure model and measurements. Did laser cutting of wing ribs with Chris and Nicholas. Conducted thrust testing with Nicholas, Chris, and Kevin.  

• Christopher Robbins: Worked on power supply and helped select an appropriate battery, ran FEA simulations in FEMAP on various systems of the drone to have a better understanding of where the stress concentrations were as well as ran CFD simulations in Star-CCM+  on the whole assembly to see how airflow acts on the drone. Worked with Kevin Soto to write the code for the arduino uno. Regarding manufacturing, built and assembled the wing with Nicholas which included laser cutting the ribs, gluing together the stringers and spars to the ribs, and applying the monokote. Conducted thrust test with Nicholas Kevin and Jasper.

• Alexander Rodriguez: Worked on the propulsion systems focusing on the engine and propeller setup. Studied various propeller design and investigated possible usage of variable-pitch propeller. Conducted extensive CFD work, designing and testing various propeller designs and configurations utilizing QBlade, Ansys Fluent, and Star-CCM+. Did Propeller selection based off CFD results. Designed the initial thrust test setup (Vertical Thrust Tests) selected engines to be tested. Additionally, did light FEA work on the engine housing using MSC NASTRAN and Femap. Designed the initial engine housing and layout prior to manufacturing limitations. Designed the initial avionics shell design to house the avionics while being able to mount to the fuselage which the lasercut design is based off.

• Nicholas Solaski: Worked on modeling of empennage as well as initial airfoil research. Utilized CFD for empennage analysis as well as wing and full assembly analysis (Understanding lift and streamlines across designs). Studied flow over airfoils in XFLR5 as well as ANSYS. Utilized hand calcs to compare observed data between NACA coefficient of lift, ANSYS lift and XFLR5 lift. Helped teach Christopher Robbins and Abraham Tazi how to use Star-CCM+ for their own analyses. Compared own findings with their simulations to better understand data. Used STAR-CCM+ to simulate fluid flow over the entire assembly, noting how the airstreams interacted with the model as well as how much lift was achieved. Viewed  Manufactured the wing with Christopher Robbins and the tail with Kevin Soto. Helped Christopher Robbins with laser cutting. Worked on monokoting the full drone assembly with Christopher. Assisted Kevin in tuning most controls and mechanisms including a static stability test and conducting flight tests.
• Kevin Soto: Designed the aileron/elevator/propeller tilt control systems, the finalized propeller housing, unused initial iterations of the landing gear, and the circuit diagrams. Wrote and tested the arduino code with Christopher, Nicholas, and Abraham. Worked with Chris in performing the necessary electrical hand calculations along with the necessary component specification research to select a battery. Selected other avionic components required for the design. Performed FEA using FEMAP on the wing connector, tail boom, empennage connector, aileron/elevator control system. Performed CFD on the wing in ANSYS and XFLR5, as well as on the wing plus empennage assembly on XFLR5. Performed hand calculations relating to center of gravity, aerodynamics, and electrical properties. Designed all but the initial propeller thrust test stands, as well as manufacturing their setups. Conducted the propeller thrust testing with Nicholas, Chris, Abraham, Jasper. Manufactured the propeller housing, aileron/elevator/propeller tilt control systems. Helped Nicholas manufacture the empennage. Worked with Nicholas, Jasper, and Christopher to manufacture the wing. Worked with Abe to construct the circuitry wiring and wing connector. Collaborated with Abraham, Nicholas, Christopher, and Jasper in the final manufacturing and assembly process, performing static stability tests, tuning the program parameters, and conducting flight tests. Purchased the motors, ESCs, high torque servos, and battery used in our final design out of pocket.

• Abraham Tazi: Lead the design and development of all structures and material selection. Fully modeled the Airframe and Fuselage systems using Solidworks to meet realistic market supply and design requirements. Put the various components making up these subsystems through FEA to better understand the stress distribution and use that information to optimize the design. Put the Fuselage shell through CFD analysis to better understand the air flow around the body and verify that the shell geometry did not impede the performance of the system. Held a workshop for other team members to learn how to do FEA using Siemens FEMAP. Assisted team with producing a Bill of Materials based on modeled parts. Oversaught actual manufacturing of the UAV, directly manufactured and processed the making of the Airframe, Empennage base, Fuselage shell, landing gear and assisted in the avionics wiring and implementation. Calculated by hand the desired location of center of gravity in fixed wing mode with the empennage distance as the manipulating variable.