Mission and Design

Systems Engineering Approach

Mission Requirements Analysis

The SUAS competition has three main sections [1]: The technical paper, the flight readiness review, and the mission demonstration. The competition is scored out of 100% in which the first two sections are worth 20% of the final score, and the final section is worth 60% of the final score. All scores are presented as their percentage of the final score.

The purpose of this portion of the document is to determine the sections of the mission demonstration on which to focus in order to maximize the competition score. The intent is to create hardware requirements for the vehicle design process. Table 1 breaks down the mission demonstration sections and highlights the aspects of the vehicle required to effectively complete those sections. Operational excellence is omitted as it is highly subjective.

Design Rationale

Prior to the start of the design process for the AUVSI SUAS competition, the UCF Robotics Club intended to create a multirotor vehicle. The aerial team deemed that a multirotor platform would provide the most utility to future club endeavors compared to other vehicle types such as fixed-wing, quad plane, helicopter, etc. This decision was made due to the nature of multirotor vehicles, notably, the ability to hover, decoupled yaw, pitch, and roll control, and relatively ease of upgrade mechanically and electrically. As previously shown in Table 1, the ideal vehicle for the AUVSI SUAS competition will be one that has high maneuverability, long range, and the ability to hover. The tradeoff is that the vehicle will not be relatively fast. The only competition element that suffers from reduced speed is the Timeline, but this portion is worth the least of the Mission Demonstration score. The team’s decision to create a multirotor vehicle and the values of the competition were well aligned, and as such a multirotor vehicle was designed for the purposes of the competition. 

From the decision of creating a multirotor vehicle, preliminary research was conducted as outlined in section 1.1.1 From the preliminary research, the project budget, and preliminary mass budget were created. The budgeting process is largely ignored in this document as it was not a major contributor to the design and construction of the vehicle. From the mass budget and preliminary research, appropriate batteries, motors, and propellers were selected. With the majority of the vehicle systems determined, the vehicle frame was designed and constructed. In tandem with the physical design process, the vehicle software was begun. The first stages of software development focused on flight-critical items such as the flight controller implementation and basic autonomy. Further development was then done for more advanced software systems such as path planning, computer vision, etc.

Systems Design

Aircraft

Vehicle Configuration

The team began the vehicle configuration generation by creating scripts to calculate the thrust given the propeller diameter, propeller pitch, motor RPM, and vehicle speed. The intention was to create a system optimizer with these parameters and other necessary vehicle parameters as inputs in order to output vehicle performance estimates. The thrust script was based on theoretical equations [12, 14, 16, 18, 20, 21, 23]  as well as empirical data from the University of Illinois at Urbana-Champaign (UIUC) Propeller Data Site [11].  During the development process of the optimizer, an online resource xcopterCalc [17], was discovered. This resource inputs multirotor parameters and outputs expected performance data. xcopterCalc boasts 9000 motors and 7 million visitors [17].  The team independently verified results using data from the UIUC Propeller Data Site as well as the beginnings of the system optimizer. Values tested in the calculator matched the UIUC data as well as the system optimizer within a reasonable degree. The decision was made that because xcopterCalc was deemed reliable, in the interest of time xcopterCalc would function as the team’s system optimizer. 

It was at this time that the decision was made to utilize NCR battery chemistry. NCR cells are roughly 15% lighter per unit energy, at the cost of decreased power per unit capacity, and increased volume per unit capacity. Typical multirotor vehicles and especially racing drones require relatively high power [24] and suffer significantly from increased drag due to increased volume. The intended mission for the vehicle has a particularly enhanced range of requirements. Due to this, any battery option selected will have a required capacity on the order of 80,000 mAh. Additionally, due to the enhanced range requirements, minimizing weight is of significant concern. Since the NCR cell chemistry is lighter per unit energy, an estimated 1.2 kg can be removed from the vehicle assuming an 80,000 mAh battery at 6S. The lower power per unit capacity can be sidestepped by the shear capacity required. Additionally, the increased volume of the cells can be ignored as the planned mission for the vehicle does not require speeds such that drag is of significant concern, and the size of other components makes the battery volume not the driving factor.

A preliminary mass estimate for the vehicle was created in order to properly scale the vehicle for the required mission. Research was conducted for all components necessary for the vehicle and mass estimates for each component were tabulated. The base weight is the same as that previously defined. The mass of the UGV was assumed to be the entirety of the allotted 1.36 kg (3 lb) by the competition rules. Table 2 shows the mass estimated for the valid vehicle types (quad, hex, octo) both with and without the UGV.

Table 2: Estimated System Masses

From preliminary mass and range estimates, it was found that the vehicle would need a propeller diameter between 15 and 21 inches. It was at this time that the team decided to purchase consumer off the shelf (COTS) propellers in lieu of manufacturing custom components. Experimental data for COTS propellers is largely unavailable on both general hobbyist and propeller manufacture sites. Possible vendors of these large hobbyist propellers included Quanum, XOAR [13], and Falcon, etc. Of these vendors considered, only XOAR provided significant experimental data [23]. Performance estimates from the team's calculations, the UIUC Propeller Data Site, xcopterCalc, and XOAR all matched within a reasonable tolerance. 

It was decided that the number of rotors would need to remain under 8 in order to ensure compatibility of the system with available and well characterized flight controllers. Additionally, there were concerns with power distribution, frame size, and vehicle weight, should the number of rotors exceed 8. The number of rotors was further constrained to only the even numbers in order to again ensure compatibility with the available flight controller software. Therefore, the number of rotors on the vehicle is to be 4, 6, or 8.

From this concept space, consisting of 13 motors, 10 propellers, and 3 vehicle types, there were a total of 390 potential vehicle configurations available. For each of the 3 vehicle types, a matrix was created to show a rough battery percentage required to complete the 4 mile waypoint traversal.  The goal of this phase of the concept selection was to narrow the search space from 390 to some more manageable number. The quadcopter concept matrix is omitted as none of the quadcopter configurations were able to lift off within the ratings of the selected motors. Additionally, only 10 of the propellers are shown as none of the other 10 were able to lift off within the ratings of the selected motors. These matrices are shown in Tables 3, and 4.

Table 3: Hexacopter Setup Motor and Propeller Configurations

Table 4: Octocopter Set-Up Motor and Propeller Configurations

Within each matrix, there was an envelope of configurations that were able to complete the waypoint traversal. Most configurations suffered from excessive power required to hover, low thrust resulting in no maneuverability, excessive current required at cruising speed, or excessive power required at cruising speed. Of the initial 390 configurations, only 55 were deemed feasible. 

An additional requirement was placed on the vehicle at this time to ensure than in the event of a single rotor failure the vehicle is able to continue controlled flight. With this requirement considered, only 16 configurations remained viable. All of these configurations were octocopters. The remaining configurations were ranked based on the estimated post waypoint traversal range. Six of the configurations had approximately 3 miles of post traversal range and the remaining 10 had under 2.5. This was a clear separation that narrowed the configurations available to 6. 

Of the 6 remaining configurations, there were 2 propellers options: 18.5 x 6.7 and 19.5 x 7. Given that all the performance estimates are based on the preliminary mass estimate, a margin of safety should be considered in the event that the mass estimate was an underestimate. Therefore the 19.5 x 7 propeller was selected. With the 19.5 x 7 propeller selected, only 4 configurations remained with only the motor differentiating them. 

The motors remaining were the 5010, 5012, 5015, and 6012. The 6012 was removed as it is significantly oversized for the propeller selected. The remaining 3 motors had performance estimates from xcopterCalc that were too similar to make a final decision upon. Therefore, the empirical data from XOAR was used as more accurate source. Ultimately, the 5010 motor was selected because it had the lowest power requirements.

Therefore, the final vehicle configuration is an octocopter with XOAR 5010 motors and XOAR 19.5” x 7 propellers. From this systems level design of the vehicle, all further vehicle properties can be determined and designed. Further CFD analysis was conducted on the propellers to confirm these thrust and power calculations, but the details of this analysis are omitted for brevity.

Materials

The materials in aerospace engineering must fit a very specific set of requirements in order to be used for the construction of aerospace parts. One of the main concerns is that the material be capable of handling the loads associated with flight while also being lightweight. This ratio is known as specific strength ratio and gives an idea of the materials ability to hold loading per density of the material. Another important factor is a high stiffness modulus, or the stiffness of the material per density. Lastly, manufacturability and resistance to corrosion direct materials choice. This leaves certain classes of composites and metals for the manufacturing of aerospace parts, such as: carbon fiber reinforced polymers (CFRP), aluminum alloys, and titanium alloys.

For this application, rigidity and weight are the two most crucial factors and will make up the selection criteria. As the preliminary research showed, carbon fiber is often chosen because it has the highest strength to weight ratio of these materials. The Table 5 [39] compares the specific tensile strength and specific tensile modulus of carbon fiber, aluminum, and titanium. 

Table 5: Specific Tensile Strength/Modulus of Discussed Materials 

The final decision on material for the chassis plates is to use carbon fiber with Nomex cores due to the high strength to weight ratio and rigidity of the carbon fiber. Additionally, the decision was made to utilize circular carbon fiber tubes for the vehicle arms for similar reasons.

Finite element analysis (FEA) was conducted on both the chassis plates and arms in order to ensure the structural integrity of the vehicle. The details of this analysis are omitted for brevity. Additionally, material samples of the carbon composite sandwich were testing in bending tests, and tensile tests to confirm manufacturer data and analytical calculations.

The manufacturing process of the carbon fiber was completed in two main phases, the plate manufacture and machining process. The plate manufacture was done with pre-preg and the nomex honeycomb core in an autoclave. The plates were oversized for the chassis plates for later machining. The machining process was completed by a third-party manufacturer on a CNC in a water bed for safety reasons. 

Frame Design

The design of the aircraft frame was driven by the results of the aerodynamics analysis and electronics selection. The size of the chassis plates were determined as the minimum size possible while still able to package all electronics with appropriate space.  Some key requirements of other structural components were derived from the team’s intent to make the aircraft easily transportable, which requires ease of disassembly due to its large size. These include the arm clamps remaining attached to the plates and in alignment when the chassis is disassembled to access the electronics, as well as the need to ensure the entire system can be taken apart with minimal tooling in less than 10 minutes. It was decided that having two plates would allow for these goals to be accomplished. The lower plate allows for mounting of hanging components, such as the UGV and camera gimbal. The majority of the electronics and the arm clamps are housed in the middle layer, between the two plates, and the top deck allows for space to mount the batteries and GPS receivers. The two plates, joined together by the arms clamps and threaded standoffs, also create a very rigid chassis that can resist the bending moments the arms are subjected to by the rotors and landing gear. The frame layout can be seen in figure 3. The fully equipped aircraft is also shown in figure 4.

The length of the arms was determined with CFD such that the increase in thrust due to  propeller efficiency due to the increase in arm length is equal to the increase in weight due to the increased mass of the arm. A more detailed view of one of the arms and rotors can be seen in figure . The clamp designs for the rotor mounting plate and the landing leg mounting plate are the same. The clamp design for mounting the arm to the chassis differs. All clamps are custom machined from 6061 aluminum. Multiple clamps are used to mount each component and are spaced out so as to limit the effects of stress risers when subject to high loading, such as a rough landing or crash. Test flights have shown the ability of the current design to withstand substantial impacts without major damage. The modular design of the system also allows for easy replacement or repair of parts in the case that serious damage does occur. For example, when a landing leg fails, the weakest point is the small carbon fiber tube, which can simply be pulled out and replaced with a spare in a matter of minutes.