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Table of Contents 1. Executive Summary.................................................................................................................................. 3 2.

3.

Management Summary............................................................................................................................. 4 2.1

Team Organization ................................................................................................................................. 4

2.2

Milestone chart....................................................................................................................................... 5 Conceptual Design ................................................................................................................................... 5

3.1 Mission Requirements and Scoring................................................................................................................ 5 3.2 Design Requirements.................................................................................................................................... 7 3.3 Concept Selection Process ........................................................................................................................... 8 3.4 Selected Concept ....................................................................................................................................... 11 4.

Preliminary Design ................................................................................................................................. 12 4.1 Design and Analysis Methodology ............................................................................................................... 12 4.2 Design and Sizing Trades ........................................................................................................................... 13 4.3 Mission Model............................................................................................................................................. 14 4.4 Propulsion Characteristics ........................................................................................................................... 15 4.5 Aerodynamic Characteristics ....................................................................................................................... 17 4.6 Stability Characteristics ............................................................................................................................... 20 4.7 Mission Performance Estimates .................................................................................................................. 26

5.

Detailed Design ...................................................................................................................................... 26 5.1 Dimensional Parameters ............................................................................................................................. 26 5.2 Structural Characteristics ............................................................................................................................ 27 5.3 Aircraft Systems Design, Component Selection and Integration .................................................................... 29 5.4 Aircraft Component Weight and Balance...................................................................................................... 36 5.5 Flight Performance Parameters ................................................................................................................... 38 5.6 Rated Aircraft Cost...................................................................................................................................... 38 5.7 Mission Performance Summary ................................................................................................................... 39 5.8 Drawing Package........................................................................................................................................ 39

6.

Manufacturing Plan and Processes ........................................................................................................ 48 6.1 Manufacturing Process Selection................................................................................................................. 48 6.2 Subsystems Manufacture ............................................................................................................................ 49

7. Testing Plan .................................................................................................................................................... 52 7.1 Testing Schedule ........................................................................................................................................ 52 7.2 Flight Testing .............................................................................................................................................. 53 7.3 Propulsion Testing ...................................................................................................................................... 54 7.5 Miscellaneous Testing ................................................................................................................................. 56 8.

Performance Results .............................................................................................................................. 57 8.1 Subsystem Performance ............................................................................................................................. 57 8.2 Aircraft Flight Performance .......................................................................................................................... 59

9.

References.............................................................................................................................................. 60

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Acronyms PA - Production Aircraft MSA - Manufacturing Support Aircraft 1.

Executive Summary The following report details the design, analysis, manufacturing, and testing completed by Cornell

University's Design-Build-Fly project team to develop radio-controlled aircraft for the 2015-2016 Cessna/Raytheon/AIAA Design-Build-Fly competition. The goal of the team is to create aircraft that complete all competition missions while maximizing the total score. To accomplish this goal, the team created an aircraft solution that meets all competition requirements in the simplest manner possible. The competition profile is distributed manufacturing. In real-world aircraft design, components are not manufactured in a single location. Instead, they are created at specialized manufacturers and transported to a centralized assembly location. The three flight missions and the bonus mission model this process. The first mission is the arrival of an empty manufacturing support aircraft. The second mission is a delivery mission, during which all components of the production aircraft are transported within the manufacturing support aircraft as a series of sub-assemblies. The bonus mission is the assembly of the production aircraft within two minutes. The third flight mission is a production aircraft flight with the design payload, a 32 oz. Gatorade bottle. Analysis of the competition scoring demonstrated that aircraft battery weights would be the most important factor affecting our final score; this is due to the fact that battery size can be fine-tuned without significantly impacting the structure of the aircraft. The next most important factor for scoring was the number of production aircraft sub-assemblies, and then lastly the aircraft empty weights. An analysis of payload configurations, compared with expected mission score, suggested that the simplest and most efficient approach was to store the production aircraft as a single piece. Thus, the driving design requirement was minimizing the weights of the production and manufacturing support aircraft. We considered several configurations of production and manufacturing support aircraft. As the total mission score does not depend on mission performance (how quickly the tasks are completed) we chose a conventional design for both aircraft because it yielded the simplest payload configuration and best stability characteristics. The designs for both aircraft include a fuselage with the lowest possible profile, a conventional tail, high wing, single tractor propeller, and taildragger landing gear. After finalizing the conceptual design, the team broke up into four sub-teams: aerodynamics, stability and controls, structures and propulsion. Each group developed requirements for their subsystems, and then collaborated with each other to complete detailed design and analysis. The individual sub-systems and the aircraft as a whole were optimized to minimize mass, and key structural components were tested to validate their designs. The successful fabrication and testing of the first production aircraft prototype highlighted differences between expected and actual aircraft performance. In addition, it revealed limitations in our

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manufacturing process that were not anticipated. Four subsequent design iterations have optimized the design to maximize score. The completion and testing of the first manufacturing support aircraft prototype has begun the same process for that vehicle. By the competition date we hope to have completed two more iterations of the entire aircraft solution. Figure 1 shows the estimated score of our final aircraft design. We look forward to our fifth year of participation in the Design-Build-Fly competition! MF1

MF2

MSA Arrival Flight EW1 (lb) 2.02

2.0

MSA Delivery Flight

Wt_Battery_1 (lb) 0.32

MF3 4.0

N_Components 1

PA Flight EW2 (lb) 3.91

Bonus 2.0

Ground

Total Mission Score 18.0 2.0

Wt_Battery_2 (lb) 0.41

RAC 2.2346

Written Report Score Total Mission Score RAC SCORE (without WRS) 8.055 WRS 18.0 2.2346 Figure 1 – Scoring Estimates for the Final Design 2.

Management Summary The Cornell Design-Build-Fly team comprises 22 undergraduate students interested in aircraft

design. The team began the Fall 2015 semester by assembling airplane kits to develop an intuition for lightweight aircraft building. Combining experience from this and from previous Design-Build-Fly competitions, the team then approached the 2015-2016 challenge. Preliminary sub-teams consisting of a mix of new and returning members developed design requirements and preliminary designs for the aircraft. After adjusting the placement of new members based on their skills and interests, finalized subteams completed preliminary and detailed design of the production and manufacturing support aircraft. 2.1

Team Organization The aircraft design was completed by four sub-teams: aerodynamics, stability and controls,

structures, and propulsion. The teams were formed based on individual interest and experience. Subteam leaders managed individual tasks, collaborated with other sub-teams, and mentored new members.

Figure 2.1 – Team Organization

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2.2

Milestone chart The team used a milestone chart to manage the progress of the project. It ensured that deadlines

would be met and that the team would be ready for competition. The difference between the planned and actual completion of projects, such as the construction of the first iteration of the production aircraft, was used to refine estimates in later iterations. Design Week 9-Sep Aircraft Design Production Aircraft P1 Conceptual Design Preliminary Design Detailed Design Design Review ★ Manufacturing Support Aircraft Conceptual Design Preliminary Design Detailed Design Design Review Manufacturing Production Aircraft Iteration 1 Iteration 2 Iteration 4 Iteration 5 Final Iteration Manufacturing Support Aircraft Iteration 2 Final Iteration Flight Testing Production Flight Mission 1 Mission 2

12-Oct P2

9-Nov

7-Dec 4-Jan 1-Feb Winter Break P3 P4 P5



★ MS1



★ MS2

29-Feb ★ Event Planned Actual Production Aircraft P1 PA 1st Design Period P2 PA 2nd Design Period P3 PA 3rd Design Period P4 PA 4th Design Period P5 PA 5th Design Period

★★ Manufacturing Support Aircraft MS1 MSA 1st Design Period MS2 MSA 2nd Design Period

Figure 2.2 – Milestone Chart 3.

Conceptual Design During the conceptual design phase, the team closely examined the 2015-16 competition rules

and defined the specific requirements and constraints for each mission. We then performed an in-depth scoring analysis to better understand the effects of certain design parameters on our overall competition score. 3.1 Mission Requirements and Scoring The 2015-16 DBF competition requires the design of two aircraft: a production aircraft, which must carry a 32 oz. Gatorade bottle, and a manufacturing support aircraft, which must transport the production aircraft, disassembled. These aircraft must undergo a total of three independent flight missions and one bonus mission. A maximum score will be obtained by successfully completing all missions within the given time limits while complying with all mandated criteria.

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3.1.1 General Requirements Several requirements are constant for all flight missions. The aircraft must complete a rolling take-off within a one hundred foot long runway and land without sustaining significant damage. All payloads must be secured significantly without substantial variation of the aircraft center of gravity. There is no battery weight limit. The same course must be flown for each mission, as shown in Figure 3.1.1 below.

Figure 3.1.1 - Flight Course Layout

3.1.2 Scoring Summary The team's score will be determined by the following formulas: 𝑆𝐶𝑂𝑅𝐸 =

𝑊𝑟𝑖𝑡𝑡𝑒𝑛 𝑅𝑒𝑝𝑜𝑟𝑡 ∗ 𝑇𝑜𝑡𝑎𝑙 𝑀𝑖𝑠𝑠𝑖𝑜𝑛 𝑆𝑐𝑜𝑟𝑒 𝑅𝐴𝐶

𝑇𝑜𝑡𝑎𝑙 𝑀𝑖𝑠𝑠𝑖𝑜𝑛 𝑆𝑐𝑜𝑟𝑒 = 𝑀𝐹1 ∗ 𝑀𝐹2 ∗ 𝑃𝐹 + 𝐵𝑜𝑛𝑢𝑠 Judges determine the team's report score, which can vary from 0 to 100. The total mission score is the product of the manufacturing support and production aircraft flight missions plus the bonus mission. Rated Aircraft Cost (RAC) is calculated from the equation below, where EW1 and EW2 are the empty weights and BW1 and BW2 are the battery weights of the production aircraft and manufacturing support aircraft, respectively, and Ncomponents are the number of sub-assemblies the production aircraft is broken into for transport. 𝑅𝐴𝐶 = 𝐸𝑊1 ∗ 𝐵𝑊1 ∗ 𝑁𝑐𝑜𝑚𝑝𝑜𝑛𝑒𝑛𝑡𝑠 + 𝐸𝑊2 ∗ 𝐵𝑊2 3.1.3 Mission 1 - Manufacturing Support Aircraft Arrival Flight The objective of this mission is to fly three laps with the MSA within a total 5 minute time window. The time starts when the throttle is advanced for the first take-off attempt and it stops when the aircraft passes over the start/finish line in the air. The team must complete a successful landing to receive the full score. A score of 2.0 is given for a successful mission and a score of 0.1 is given otherwise. Cornell University

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3.1.4 Mission 2 - Manufacturing Support Aircraft Delivery Flight This mission calls for the transportation of the disassembled PA. There is a 10 minute time window for this mission. The mission begins with the first sub-assembly installed. A sub-assembly consists of one or more parts of the production aircraft, not including the PA battery or payload. A single lap is flown with this sub-assembly and then the team must land the aircraft, taxi to the designated payload change area to swap the sub-assembly payload with the next, taxi back to the runway, fly the new payload, and then repeat the process until all sub-assemblies have been transported. All subassemblies must be carried internally. The mission time ends when the aircraft passes the start/finish line in the air at the end of the final flight. A mission score of 4.0 is given if the aircraft completes all subassembly group flights within the time window. A score of 1.0 is given if the aircraft completes less than all the sub-assembly flights within the time window but successfully transports at least one sub-assembly group. A score of 0.1 is given otherwise. 3.1.5 Mission 3 - Production Aircraft Flight This mission requires the PA to fly three laps in a five minute time window with a 32 oz. Gatorade bottle fully contained within the aircraft. The time starts when the throttle is advanced for the first take-off attempt and it stops when the aircraft passes over the start/finish line in the air. The team must complete a successful landing to receive the full score. A score of 2.0 is given for a successful mission, a score of 1.0 is given for completing less than the required number of laps or for exceeding the time period, and a score of 0.1 is given otherwise. If the team is given a score of 1.0 they are allowed a single attempt to improve their score. 3.1.6 Bonus Mission - Manufacturing of Production Aircraft This mission may be attempted after receiving a score of 4.0 on mission 2. This mission is ground based and starts with the PA unassembled. The team must assemble the PA and re-install the payload within two minutes. The assembled PA must pass the wing tip lift test and a control systems check. A score of 2.0 is given for a successful mission and a score of 0.0 is given otherwise. 3.2 Design Requirements The team conducted an analysis of this year's competition scoring criteria to highlight key design requirements. The analysis revealed which design requirements have substantial impact on the overall score. For each mission, a fixed score is awarded for completing the mission successfully within a given time window. Therefore, assuming that all missions can be completed, the important scoring variables are determined by the RAC. These variables are: PA empty weight, PA battery weight, number of subassemblies the PA is broken into for transport, MSA empty weight, and MSA battery weight.

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The team utilized MATLAB as a tool to more appropriately understand tradeoffs between the variables mentioned above. To do so, we sought to reduce the number of variables governing the RAC by imploring various estimates. First, we noted that the main driving force behind the battery power (and therefore, weight) required for flight is the aircraft empty weight; as such, PA and MSA battery weights were estimated as functions of PA empty weight and MSA empty weight, respectively. Estimating the PA empty weight further reduces the number of score-determining variables such that only two remain: MSA empty weight and PA sub-assembly number. The RAC was then estimated as a function of these two variables, as shown in Figure 3.2 below.

Figure 3.2 - RAC vs. MSA EW & Ncomponents (N) The team used the MATLAB analysis to determine the cost of adding additional sub-assemblies to the Production Aircraft. As can be derived from Figure 3.2, the RAC only benefits from adding to N if the MSA empty weight can be significantly reduced by doing so. As such, great consideration was given to the advantages and disadvantages of having larger or smaller values of N. This was the primary driving force behind our initial design processes. 3.3 Concept Selection Process The team began the concept selection process by conducting independent and collaborative brainstorming for the aircraft systems. After the scoring analysis, we determined that transporting the entire PA as a single sub-assembly (N = 1) was the best way to maximize our score. Therefore, we agreed that following an inside-out approach, by first building a successful PA then determining how to store it inside the MSA, was the most logical and effective strategy. Team members split up into subteams to focus on particular aircraft systems. Within their groups, members used decision matrices to help evaluate relevant concepts and select aircraft configurations. The categories in which concepts were evaluated were plane configuration, motor configuration, landing gear configuration, payload Cornell University

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configuration, and tail configuration. Because we decided to transport a single sub-assembly, we deduced that it would be most efficient if these configurations were identical for both aircraft. This way, we could fit the wing of the PA inside a hollow wing of the MSA and the same for the tail and other parts of the PA. Accordingly, for each subsystem, the sub-teams compared configurations by rating them based on weighted design parameters deemed critical to the system's success. We collectively weighted the parameters for each decision matrix, primarily based on results from our scoring analysis, and we normalized the weights to sum to one. This objective selection process enabled the team to select the configurations for both aircraft that would maximize our competition score. 3.3.1 Plane Configuration The categories considered for the selection of a plane configuration were lift-to-drag ratio, maneuverability, weight, stability, payload integration, and manufacturability. After the scoring analysis, the team decided that plane configuration should be aimed at achieving the least weight, the most maneuverability and stability, and the easiest payload integration. As shown in the decision matrix below, a conventional aircraft proved to be the most favorable combination of these requirements. The flying wing configuration was also strongly considered because of its low weight capability; however, it suffered because of its lack of stability and maneuverability and overall complexity.

Plane Configuration Lift/Drag Maneuverability Weight Stability Payload Integration Manufacturability Total

% Canard 10 2 15 5 30 4 15 3 20 2 10 3 100 3.10

Flying Wing 4 2 5 2 1 2 2.90

Lifting Body 4 4 2 2 4 2 2.90

Conventional 3 3 4 5 4 5 4.00

Figure 3.3.1 - Aircraft Configuration Decision Matrix

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3.3.2 Motor Configuration The team went through the process of selecting a motor configuration with a focus on obtaining a high thrust-to-weight ratio and good stability while minimizing size and weight. After the scoring analysis, the team found that having one tractor motor in front of the fuselage would best fit these requirements. The tractor motor configuration with one motor was found to give better thrust for its weight, as well as accommodate static stability, when compared to one pusher motor or two motors with the same total battery size. This finding was augmented by how well one tractor would integrate with other subsystems. Motor Configuration

%

Single Tractor

Single Pusher

Double Tractor

CounterRotating

Trust to Weight

50

5

3

4

3

Stability

30

3

1

4

2

Size/Weight

20

4

3

2

2

Total

100 4.2 2.4 3.6 Figure 3.3.2 – Motor Configuration Decision Matrix

2.5

3.3.3 Landing Gear Configuration The factors considered in choosing a landing gear configuration were the ground steerability, weight, and integration with the chosen payload configuration. Skids, though the lightest option, were eliminated because they could not provide enough ground control and would make it very difficult to store the PA in one piece in the MSA. Due a large propeller diameter chosen for both planes, a tricycle configuration would require a very tall, heavy nose gear, and due to the distance from the rudder servo, a dedicated servo for steering. A taildragger configuration was ultimately selected by virtue of being significantly lighter than a tricycle design while retaining the ability to be integrated with the steering and payload storage systems. Landing Gear

%

Skids

Taildragger

Tricycle

Weight

35

3

2

1

Steerability

35

1

2

3

Integration

30

1

3

2

Total

100 1.7 2.3 Figure 3.3.3 - Landing gear decision matrix

2

3.3.4 Payload Configuration The containment of the production aircraft inside the manufacturing support aircraft was the first payload configuration decision made during the preliminary design process. Given the large weight placed on the number of sub-assemblies the PA is broken into, a significant decrease in empty weight and battery weight would have to be made to justify each additional sub-assembly (see Figure 3.2 for a breakdown of this relationship). Based on structural analysis performed by several team members, it was decided that the PA would have to be split into several additional sub-assemblies before significant Cornell University

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structural changes could be made to the MSA. With this information, the decision was made to keep the PA in one piece. Once the conventional plane configuration was chosen for the production aircraft, the next immediate decision was how the plane would contain and support the Gatorade bottle. Our final decision for placement was to place the bottle in the front of the fuselage, with the cap end towards the tail. The orientation kept the center of mass of the plane as far forward as possible, which helped in pitch stability. It also meant that the Gatorade bottle sat directly below the wing, such that a single load bearing connection rib could hold the wing in place at the quarter-chord and directly support most of the weight of the bottle. This allows the plane to have minimal structure in other areas. 3.3.5 Tail Configuration The team considered four options for the tail configuration: conventional tail, H-tail, T-tail, and Vtail. The ideal tail would provide sufficient stability for the aircraft while minimizing weight. The conventional tail configuration was selected due to its well-known stability characteristics, ease of manufacturability, and low weight. Tail Configuration

%

T-Tail

V-Tail

Conventional

H-Tail

Weight

30

3

2

4

1

Stability

30

2

3

3

5

Maneuverability

15

3

5

2

4

Integration

15

3

2

4

2

Manufacturability Total

10

4

1

5

2

100 2.80 2.65 3.50 Figure 3.3.5 - Tail Configuration Decision Matrix

2.90

3.4 Selected Concept The final concepts for the PA and MSA are shown in figures 3.4a-b, incorporating all of the previously discussed component configurations. To maximize score it was decided that the PA should be one piece and fit inside the hollow MSA. The conventional fuselage and high wing provides an overall aircraft shape that allows for improved airflow, therefore reducing aerodynamic drag and maintaining stability. A single tractor motor configuration maximizes the aircraft’s overall thrust to power output. A taildragger landing gear configuration provides ease of integration with the rest of the systems, while being very light. A conventional tail configuration provides the stability required for both takeoff and steady flight.

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Figure 3.4a – PA concept

4.

Figure 3.4b – PA and MSA concept

Preliminary Design During the preliminary design phase the team used conceptual design ideas to produce

performance and size parameters for systems across both aircraft. 4.1 Design and Analysis Methodology Our team relies on an iterative design process in order to develop our aircraft. The diagram in Figure 4.1 illustrates the general methodology. The team defines the requirements from the competition guidelines, brainstorms and researches concepts to meet the requirements, develops system configurations, and creates rough design sketches. We begin our preliminary design by conducting design trades, determining initial sizing and performance values, producing CAD models, and selecting materials. Next, aspects of the designs are analyzed and revised using analytical, virtual, and experimental techniques, and a detailed design is established. Finally, prototypes are manufactured and more practical tests are performed to fully evaluate the design based on the competition requirements. The process allows for frequent iteration and rapid prototyping to optimize design features and ultimately maximize our competition score.

Requirements

Brainstorm and Research

Conceptual Design

Preliminary Design

Design Analysis

Detailed Design

Manufacturing

Testing

Iterate

Sizing and Performance Optimization

Figure 4.1 – Iterative Design Process Diagram Cornell University

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4.2 Design and Sizing Trades In past years, the team has conducted several trade studies to analyze how various aspects of the aircraft design affect our final competition score. Due to the nature of this year’s competition, and particularly due to the lack of a speed element in scoring, our team completed just one design and sizing trade study aimed at determining the optimal number of sub-assemblies for the production aircraft. As described in the scoring analysis of Section 3.2, the score-determining variables outlined in the RAC can be reduced to the MSA empty weight and PA sub-assembly number. Weighing these two considerations, many concepts were considered for configurations with anywhere from 1 to 4 sub-assemblies; two ideas are shown in figures 4.2a-b.

Figure 4.2a – Conceptual Design, N = 3

Figure 4.2b – Conceptual Design, N = 4 Early design work demonstrated that approaches with N>1 would increase the RAC without significantly decreasing complexity or weight. Based on this analysis, the team decided to use N=1. The configuration design selection, Section 3.3.1, yielded a conventional aircraft design for PA. With the determined N=1 configuration, the team then planned a design and sizing trade to determine the best way to store a conventional aircraft in one piece inside another conventional aircraft. However, it soon became apparent that the only feasible solution was a complete component-inside-component approach. Therefore, the competition solution was designed for wing-inside-wing, tail-inside-tail, and fuselage-inside-fuselage.

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4.3 Mission Model A comprehensive model for the missions was developed using MATLAB programming software to predict mission performance given inputs for the geometry and propulsion system of the design. The software simplifies the mission course into four flight segments: takeoff, climb, cruise and turning. The model outputs parameters such as takeoff and cruise velocity, cruise angle of attack, and coefficient of lift for the four segments. These parameters were then used to estimate the mission score.

Figure 4.3a – Course segment divisions The MATLAB model uses the numerical methods for analyzing governing differential equations of motion along with constraints unique to each different segment of flight, summarized in Table 4.3a. The geometric inputs also varied depending on the mission being modeled given the two different aircraft being used. The model accounted for the accelerated flight between the climb and cruise segments as well. For the turning segments, the load factor varied so that the loading was the same across all missions and aircraft. Flight Segment

Description

Takeoff

Acceleration on the ground until lift force equals weight

Climb

Constraints AoA=constant

Steady climb up to safe cruise altitude

AoA=constant L = W, T = D Cruise Level flight at constant speed and angle of attack AoA =constant Changing direction, either for 180 or 360 degrees, high L = nW, T = D Turning wing loading altitude = constant AoA=angle of attack, L=Lift, W=Weight, T=Thrust, D=Drag, n=Load factor Figure 4.3b - Mission segment constraints The model has multiple shortcomings and uncertainties that limit its accuracy and reliability. Since some inputs to the model are only estimates (such as turning radius) and not representative of the actual aircraft performance, the outputs obtained from the program are only estimates as well. The model does not account for wind conditions during flight; doing so adds significant variability to performance estimates. Battery voltage is also treated as constant and equal to the average cell value throughout the flight. Finally, the model neglects ground rolling friction and interference drag. In order to account for these uncertainties, we use the model as a tool in conjunction with extensive testing, which ensures that we ultimately attain the best results.

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4.4 Propulsion Characteristics The propulsion sub-team is responsible for selecting all of the electronics used to power and control the aircraft. Operating under the design requirements listed in Figure 4.4, the sub-team's objective is to determine the system of batteries, motors, and propellers that would yield the highest overall score at competition. Due to the RAC's direct inclusion of battery weight as a crucial factor in our score, the principal objective for our propulsion system design was to use as few battery cells as possible while still having enough power to fly. Additionally, the competition rules state that each aircraft must take off within 100 feet and that we may not exceed 5 minutes of flight for missions 1 and 3, and 10 minutes for mission 2; since our design for the Production Aircraft will yield N = 1, we maintain that mission 2 can adhere to the 5 minute flight time limit. Overall Design Requirement

Sub-team Design Requirement Minimize number of battery cells used

Minimize battery weight

Optimize system components for mission requirements Maximize efficiency of motor and propeller system

Maintain stable flight

Select a system with an appropriate thrust/weight ratio

Take off within 100 ft. Provide enough power to take off within 100 ft. Fly for a maximum of 5 minutes Size batteries to have just enough capacity for flight Figure 4.4 - Propulsion design requirements

4.4.1 Power Requirements The first step in our propulsion system design was to determine the power needed by each aircraft to fly; these power requirements were analyzed using a custom-developed MATLAB program. The program takes inputs for aircraft parameters and competition-defined constraints and outputs the power to weight ratio needed to fly under various flight conditions. Using weight estimates for each aircraft, the total power needed for flight can be derived from the maximum power to weight ratio of all considered flight conditions. Figures 4.4.1a-b highlight the power to weight ratios required for flight for several different conditions. Figure 4.4.1c summarizes the total power required for each aircraft.

Figure 4.4.1a - PA Power Curve Cornell University

Figure 4.4.1b - MSA Power Curve Page 15

Aircraft Production

Power Required (Watts) 143

Manufacturing Support Figure 4.4.1c - Power Requirements

154

4.4.2 Battery Selection Only two types of battery chemistries are allowed in the competition: nickel metal hydride (NiMH) and nickel cadmium (NiCd). The NiMH battery chemistry was chosen over NiCd due to its higher energy density and better resistance to memory effects. Figure 4.4.2 highlights a simple comparison of several NiMH cells used to determine the specific cells that best suit our needs. The study incorporated factors such as capacity, energy density, and maximum discharge current. The results suggest that the Tenergy 1600 cells are have the best energy density available; we know from experience, however, that this cell is not as reliable as the Elite 1500, the next best in the study. As such, Elite 1500 cells were chosen for both the Production Aircraft and the Manufacturing Support Aircraft. Brand

Capacity (mAh)

Weight (oz.)

Energy Density (mAh/oz.)

Max. Discharge Current (A)

Kan

700

0.49

1429

7

Elite

1500

0.81

1852

15

Tenergy

1600

0.81

1975

16

Elite

2100

1.15 1826 Figure 4.4.2 - Battery Comparison

21

4.4.3 Motor & Propeller Selection Extensive propeller and motor databases were generated for analysis and system optimization. The motor database was limited to include only mid-to-low Kv (RPM/volt) motors and motors with gearboxes (effectively mid-to-low Kv). Since the load on the batteries is determined by the size and rotational speed of the propeller, only motors with lower Kv values are able to spin larger, more efficient propellers without pulling high currents. An additional limitation put on the motors in the database was that they must have a maximum power rating of 500 surge watts; anything more would be unnecessarily powerful and heavy. The database included the following brands: AXI, Hacker, Neu, and Tiger. The propeller database included strictly APC brand propellers. Through the years, our team has come to find that this brand is the most reliable. Propellers with diameters ranging from 12" to 22" were included. The database was further refined by removing propellers with a pitch/diameter ratio less than 0.5; a minimum 0.5 ratio was established such that sufficient airspeed could be maintained for a given amount of thrust. With extensive motor and propeller databases prepared, the team then utilized a custom MATLAB program to optimize the entire propulsion system. The power requirements, as defined in Section 4.4.1, were input to the program which returned a ranked table of motor and battery combinations

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for a chosen propeller. Many propeller sizes were selected to undergo optimization. The potential systems were ranked by their weight, efficiency, and number of batteries cells required such that light, efficient systems were preferred. MATLAB calculations were later verified using the online program eCalc to ensure accuracy. Figure 4.4.3a below highlights the process used in the optimization code.

Figure 4.4.3a - Propulsion System Optimization Process Due to the significant impact of battery weight on the RAC, large propeller/small battery systems were chosen for each aircraft. The optimization also revealed that Neu brand motors with gearboxes proved to be the most efficient. Due to the different weights of the Manufacturing Support Aircraft in missions 1 and 2, different propellers were chosen for each mission. Mission 1 requires less thrust and more flight time, so a smaller propeller was selected. Mission 2 requires a more thrust to compensate for the added payload, so a larger propeller was chosen for this mission. The battery configuration remains the same for missions 1 and 2. The chosen systems can be seen below.

Aircraft

Motor

Gearbox

M1 Propeller

M2 Propeller

M3 Propeller

PA

Neu 1105/2.5Y

P29 (6.7:1)

-

-

APC 18x12E

MSA

Neu 1105/6D

P29 (6.7:1)

APC 18x12E

APC 22x12E

-

Batteries 6 Elite 1500 cells 7 Elite 1500 cells

Figure 4.4.3b - PA & MSA Propulsion Systems

4.5 Aerodynamic Characteristics The aerodynamics sub-team was responsible for determining the aerodynamic characteristics of the plane that optimized the design of the wings given the mission parameters. Tasks included airfoil filtering and selection, aspect ratio selection, and analytical determination of lift and drag characteristics of the aircraft. Overall Design Requirement High wing area

Subteam Design Selection Size wing for takeoff Choose a low drag airfoil Low drag at cruise Decrease plane parasitic drag Figure 4.5 – Aerodynamic design requirements

4.5.1 Airfoil Selection The selection of the airfoil for an aircraft's wings is a crucial component to ensuring the performance of the aircraft is good. The airfoil databases located at airfoiltools.com and airfoildb.com Cornell University

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were utilized to begin the selection of the airfoil for the aircraft’s wings. The 1,756 airfoils available were initially filtered using the following conditions: 

Max CL/CD > 40: to eliminate low efficiency airfoils.



Max Thickness 7% - 15%: to eliminate thick airfoils since they tend to be inefficient and thin airfoils which would be difficult to manufacture. The percentages were determined from sample airfoils and team experience.



Max Camber < 9%: to eliminate extremely high camber airfoils, ensuring manufacturability. The remaining 742 airfoils were then filtered again. Low lift airfoils and airfoils that had an onset

of stall between 0 and 8 degrees or strange CL curves were eliminated. Additionally, airfoils with sharp trailing edges or otherwise difficult to manufacture shapes were eliminated. After this iteration of selection, 40 airfoils remained. The 2D CFD software XFLR5 was then utilized in order to quantitatively test each of the remaining airfoils at takeoff and cruise conditions, with estimated Reynolds numbers of 100,000 and 200,000, respectively. The desired airfoil would perform well in the following categories: 

Maximum CL > 1.0 at 75% of the stall angle under takeoff conditions (40%). An increase in the CL reduces the necessary wing area, reducing the total weight.



Maximum CL/CD during cruise conditions at low angles of attack (30%). Increased cruise efficiency reduces the necessary thrust, reducing the weight due to the propulsion system.



Minimum CD during cruise conditions at low angles of attack (20%). Reducing drag leads to a reduction in the necessary thrust, reducing the weight due to the propulsion system.



Stall angle during takeoff conditions (10%). A high stall angle during takeoff reduces the chances of failure during takeoff. The data gathered from XFLR5 was used to rate each airfoil in the four categories listed above.

The ratings were then input into the decision matrix below (Figure 4.5.1a) to compute each airfoil's score. Figure 4.5.1b displays the CL vs. α graph and Figure 4.5.1c displays the CL vs. CD graph for the top three airfoils. Figure 4.5.1a summarizes the top three performing airfoils based on the decision matrix. Based on the results of the scores from the decision matrix, the MH 114 airfoil was selected. Airfoil

%

DAE-31

MH 114

SD7062

CL

40

4.5

5

4

CL/CD

30

4

4.5

4

CD

20

4.5

4

4.5

Stall Angle

10

4

4

4

Total

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100 4.3 4.55 Figure 4.5.1a - Airfoil selection decision matrix

4.1

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Figure 4.5.1b - CL vs. α

Figure 4.5.2c - CL vs. CD

4.5.2 Airfoil Sizing and Performance Parameters Sizing and performance estimates were determined using lift and drag calculations tabulated in the mission model. Propulsion parameters and projected weight estimates were used as the bases for the aerodynamic calculations. The goal was to design the wing to be as light as possible while safely withstanding the forces during flight. Therefore, a safety factor of 1.5 was introduced to account for approximations and uncertainties in the calculations. Figure 4.5.2a outlines the equations used by the mission model. These formulas are taken from sections of Raymer’s text. Figures 4.5.2b and 4.5.2c illustrate the parasitic drag buildup of the PA and MSA. As expected, the majority of the drag for both aircraft is caused by the wings, followed by the landing gear and tail. Inputs Term

Description

𝜌

Density of air (slug/ft3)

U 𝑎𝑜𝑤 𝐴𝑅𝑤 𝛼𝑧𝑙 m.t. 𝑑𝑓

Term

Description

𝑐̅𝑤

Mean aerodynamic chord of wing (ft) Viscosity of air (slug/fts)

2D airfoil lift curve slope (1/deg) Aspect ratio of wing

𝜇 u s

Correction for non-elliptic loading Fuselage correction factor

Angle of attack for zero lift of wing (deg)

𝑆𝑤

Planform area of wing (ft2)

Maximum airfoil thickness (%)

𝐶𝐷𝑝𝐿𝐺

Landing gear parasite drag coefficient

Diameter of fuselage (ft)

𝐿𝑓

Length of Fuselage (ft)

2

Airspeed (ft/s )

Formulas 2

𝑞 = 0.5 ∗ 𝜌 ∗ 𝑈 𝑎𝑜 𝐶𝐿𝛼,𝑤,𝑡 = 𝑎𝑜 ) 1 + 57.3 ( 𝜋𝐴𝑅𝑤,𝑡 𝐶𝐿,𝑤,𝑡 = 𝐶𝐿𝛼,𝑤,𝑡 ∗ (𝛼 − 𝛼𝑧𝑙,𝑤,𝑡)

Dynamic Pressure (lb./ft2)

𝐿 = 𝐶𝐿𝑤 𝑞𝑆𝑤 + 𝐶𝐿𝑡 𝑞𝑆𝑡 𝑚. 𝑡. 𝐾𝑤,𝑡 = 1 + 2.61 ∗ ( ) 100 𝐿𝑓 𝐾𝑓 = 1.4 − 0.05 ∗ ( ) 𝐷𝑓

Lift from wing and tail (lb.)

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3D airfoil lift curve slope Lift coefficient

Wing, tail parasite pressure drag correction factor Fuselage parasite pressure drag correction Page 19

𝑅𝑒 = 𝜌𝑈𝑐̅/𝜇

Reynolds Number

𝐶𝑓 = 1.328/√𝑅𝑒 𝐶𝐷𝑝,𝑤.𝑡 = 2𝐾𝐶𝑓

Skin friction coefficient (laminar assumption)

𝐶𝐷𝑝𝑓 = 𝐾𝑓 𝐶𝑓

𝐶𝐷,𝑤,ℎ = [𝑘𝑤,𝑡 + (𝜋𝐴𝑅𝑤,𝑡 𝑢𝑠) ] 𝐶𝑙2 + 𝐶𝐷𝑝

Fuselage parasitic drag Wing and tail induced pressure drag correction factor Wing and tail drag coefficient

𝐶𝐷𝑝𝐿𝐺 = 0.25 ∗ (𝑓𝑟𝑜𝑛𝑡𝑎𝑙 𝑎𝑟𝑒𝑎)/𝑠

Landing gear drag coefficient

𝑘𝑤,𝑡 = 0.38𝐶𝐷𝑝 −1

Wing and tail parasitic drag

𝐷 = (𝐶𝐷𝑤 + 𝐶𝐷𝑝𝐿𝐺 )𝑞𝑆𝑤 + 𝐶𝐷𝑡 𝑞𝑆𝑡 + 𝐶𝐷𝑓 𝑞𝑆𝑓 Total drag of aircraft Figure 4.5.2a – Lift and drag mission model calculations

Parasitic Drag Distribution of PA Wing 35%

Landing Gear 30% Tail 24%

Fuselage 11%

Component Wing Fuselage Tail Landing Gear Total

Parasitic Drag Coefficient 0.0096 0.0031 0.0065

% of Total 34.9% 11.3% 23.7%

0.0083 0.0275

30.2% 100%

Parasitic Drag Coefficient 0.011 0.0032 0.0074

% of Total 36.3% 10.6% 24.4%

0.0087 0.0303

28.7% 100%

Figure 4.5.2b – Parasitic drag buildup of PA Parasitic Drag Distribution of MSA

Landing Gear 29%

Wing 36%

Tail 24%

Fuselage 11%

Component Wing Fuselage Tail Landing Gear Total

Figure 4.5.2c – Parasitic Drag buildup of MSA 4.5.3 Aspect Ratio Selection The aspect ratio of the wing greatly affects both aerodynamic efficiency and the weight contribution of the wing. A MATLAB optimization code was developed to address these tradeoffs within the context of our competition requirements. Emphasis was placed on obtaining a high lift coefficient, high CL/CD ratio, and minimizing wing weight. Factors such as takeoff ability and the maximum speed of the aircraft were less emphasized. The resulting range of allowable aspect ratios was between 5.5 and 8. 4.6 Stability Characteristics The stability and controls sub-team was responsible for sizing the tail and control surfaces so that both aircraft were adequately stable and could successfully maneuver the flight course. Because the scoring analysis demonstrated that the aircraft weights would be the determining factor in competition Cornell University

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score, the stability sub-team aimed to create a stable PA and MFA while minimizing any additional weight to the aircraft. The sub-team's design requirements are shown in Figure 4.6. Since the competition score does not depend on the speed at which the missions are completed, the driving design requirement was stability rather than maneuverability. The biggest challenge for the sub-team was sizing the two tails, given that the PA tail needed to be enclosed by the MSA tail. Initial design was completed using a MATLAB code, which analyzed the static and dynamic stability characteristics of the aircraft. The program takes in all aircraft dimensions, moments of inertia, and CG locations, and outputs plots that can be used to refine the aircraft sizing. Iteration was crucial to reaching the final design. Overall Design Requirements

Subteam design requirement

Adequate Static and dynamic stability

Size horizontal and vertical tail for adequate stability

Ability to maneuver the flight course Size control surfaces for adequate maneuverability Figure 4.6 - Stability and controls design requirements

4.6.1 Horizontal Tail Sizing Parameters The initial sizing of PA tail was completed using rules of thumb defined by Raymer. The elevator was chosen to be 25% of the horizontal tail chord and the rudder was chosen to be 40% of the vertical tail chord. The rules for tail size are based on volume ratios, as in Figure 4.6.1a, and Raymer defines standard volume ratios for a variety of aircraft (Figure 4.6.1b). The production aircraft is designed to transport a heavy payload concentrated in the fuselage, so it was modeled as a military cargo aircraft. Term 𝑆𝑉 , 𝑆𝐻 𝑙𝐻 , 𝑙𝑉 𝑏𝑊

Inputs (new) Description Vertical tail and horizontal tail reference areas Distance between aerodynamic centers of wing and horizontal, vertical tail Wing span Formulas 𝑆𝑉 ∗ 𝑙𝑉 Vertical tail volume ratio 𝑉𝑉 = 𝑆𝑊 ∗ 𝑏𝑊 𝑉𝐻 =

𝑆𝐻 ∗ 𝑙𝐻 Horizontal tail volume ratio 𝑆𝑊 ∗ 𝑐̅𝑊 Figure 4.6.1a - Tail volume ratios

Figure 4.6.1b - Standard volume ratios (Raymer) Cornell University

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The MSA horizontal tail was designed to enclose the PA tail. Therefore, the initial sizing was completed so that it was just large enough to fit the PA tail. The final sizing of both horizontal tails was completed using the MATLAB code. Starting with the initial sizing from Raymer’s prescribes, we iterated on horizontal tail geometry, elevator size, tail location, tail incidence angle, and CG location. We sought a configuration that had an adequate static margin and was able to be trimmed with minimal elevator deflection. Past performance demonstrated that a static margin within 5-25% would yield a longitudinally stable aircraft. The calculation is shown in Figure 4.6.1c. Inputs (new) Term

Description

𝐴𝑅𝐻

Aspect ratio of horizontal tail

𝑎0 𝐻

Lift curve slope of tail

Λ𝑊𝑐4 , Λ𝐻𝑐4

Sweep angle at quarter chord for wing and tail

𝜂, 𝑉𝐻

Horizontal tail effectiveness and volume ratio

𝜕𝜀 𝜕𝛼

The effect of downwash due to the wing

𝑋𝐶𝐺

Location of aircraft CG relative to wing leading edge, positive aft Formulas 𝐶𝐿 𝛼 = 𝑤

𝐶𝐿 𝛼 = 𝐻

𝜋 ∗ 𝐴𝑅𝑊 2

𝜋 ∗ 𝐴𝑅𝑊 ) 1 + √(1 + ( 𝑎0 𝑊 ∗ cos(Λ𝑊𝑐4 )

Wing lift curve slope

𝜋 ∗ 𝐴𝑅𝐻 2

𝜋 ∗ 𝐴𝑅𝐻 ) 1 + √(1 + ( 𝑎0 𝐻 ∗ cos(Λ𝐻𝑐4 )

𝐶𝐿𝛼 = 𝐶𝐿 𝛼 + 𝐶𝐿 𝛼 ∗ 𝜂 ∗ 𝑊

𝐻

𝐶𝑚 𝛼 = 𝑓

𝑋𝑁𝑃 = 0.25 +

Horizontal tail lift curve slope

𝑆𝐻 𝜕𝜀 ∗ (1 − ) 𝑆𝑤 𝜕𝛼

Aircraft lift curve slope

2 ∗ 𝑉𝑓 𝑆𝑊 ∗ 𝑐̅𝑊

𝜕𝜀 𝜂 ∗ 𝑉𝐻 ∗ 𝐶𝐿 𝛼 ∗ (1 − 𝜕𝛼 ) − 𝐶𝑚 𝛼 𝑡

Fuselage pitch stiffness Location of Neutral Point behind wing leading edge

𝑓

𝐶𝐿 𝛼 𝑆. 𝑀. =

𝑋𝑁𝑃 𝑋𝐶𝐺 − 𝑐̅𝑊 𝑐̅𝑊

Static Margin (% of root mean chord)

Figure 4.6.1c – Static Margin Calculation The static margin gives an approximate measure of longitudinal stability, but the aircraft also needs to have adequate longitudinal control. Figure 4.6.1d outlines calculations for elevator control authority, as well as the coefficient of moment due to the wing, tail, and fuselage. For adequate

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maneuverability, the elevator must be effective enough to overcome the total aircraft coefficient of moment. Inputs (new) Term

Description

𝛼0

Angle of attack at which to calculate the aircraft coefficient of moment and required elevator deflection. Angle of attack for zero lift on the aircraft

𝜀0

Angle of downwash coming off the wing

𝑖𝐻

Incidence angle of the horizontal tail

𝛼

𝐶𝐿 0

Wing lift coefficient at zero angle of attack

𝑊

𝐶𝑚 𝑎𝑐

Coefficient of moment about the aerodynamic center of the wing

𝑤

𝑋𝑎𝑐

Location of wing aerodynamic center relative the leading edge (~quarter chord) Formulas 𝑋𝑐𝑔 𝑋𝑎𝑐 𝐶𝑚 0𝑊 = 𝐶𝑚 𝑎𝑐 + 𝐶𝐿 0 ∗ ( − ) 𝑊 𝑊 𝑐̅𝑤 𝑐̅𝑤 𝐶𝑚 0 = −𝜂 ∗ 𝑉𝐻 ∗ 𝐶𝐿 𝛼 (𝑖𝐻 − 𝜀0 + (1 − 𝐻

𝐻

𝜕𝜀 )) ∗ 𝛼0 𝜕𝛼

𝐶𝑚 0 = 𝐶𝑚 𝛼 ∗ 𝛼0 𝑓

𝐶𝑚 0 = 𝐶𝑚 0 + 𝐶𝑚 0 + 𝐶𝑚 0 𝐶𝑚 𝛼 = (

Coefficient of moment due to the tail at zero angle of attack Coefficient of moment due to the fuselage at zero angle of attack

𝑓

𝑊

Coefficient of moment due to the wing at zero angle of attack

𝐻

𝑓

𝑋𝑐𝑔 𝑋𝑎𝑐 𝜕𝜀 ) ∗ 𝐶𝐿 𝛼 − 𝜂 ∗ 𝑉𝐻 ∗ 𝐶𝐿 𝛼 ∗ (1 − ) + 𝐶𝑚 𝛼 − 𝑊 𝐻 𝑓 𝑐̅𝑤 𝑐̅𝑤 𝜕𝛼

Coefficient of moment due to the aircraft at zero angle of attack Coefficient of moment with respect to angle of attack of the aircraft

𝐶𝑚 𝑐𝑔 = 𝐶𝑚 0 + 𝐶𝑚 𝛼 ∗ 𝛼

Coefficient of moment about the aircraft center of gravity

𝑐𝑜𝑛𝑡𝑟𝑜𝑙 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎 𝜏 = 𝑓( ) 𝑙𝑖𝑓𝑡𝑖𝑛𝑔 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 𝑎𝑟𝑒𝑎

Flap effectiveness parameter

𝐶𝑚 𝛿 = −𝜏 ∗ 𝑉𝐻 ∗ 𝜂 ∗ 𝐶𝐿 𝛼 𝑒

𝐻

Coefficient of moment due to elevator deflection

𝐶𝑚 𝑐𝑔

Required elevator deflection to counter coefficient of moment of 𝐶𝑚 𝛿 𝑒 aircraft Figure 4.6.1d – Coefficient of Moment and Elevator Calculations 𝛿𝑒 = −

The results of the iterations are shown in Figures 4.6.1e-g. For all aircraft configurations, the coefficient of moment vs. angle of attack plot (left) has a negative slope for the aircraft, indicating longitudinal stability. The elevator deflection plots (right) demonstrate that the airplane can be trimmed to zero angle of attack with minimal elevator deflection. In addition, the manufacturing support aircraft will only require a 2o control adjustment to transition from the unloaded to loaded missions.

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Figure 4.6.1e – MATLAB Plots for PA

Figure 4.6.1f – MATLAB Plots for MSA

Aircraft PA MSA MSA Loaded

Static Margin 24.56% 15.64% 17.27%

Elevator trim for 0o AoA 1o -2o -4o

𝑋𝐶𝐺 4.75 in 5.25 in 5.00 in

𝑙𝐻 27.0 in 27.5 in 27.5 in

𝑆𝐻 127 in2 272 in2 272 in2

Figure 4.6.1g – Iteration Summary

4.6.2 Dynamic Stability In order to evaluate the lateral performance of the aircraft, a dynamic stability analysis was completed using the MATLAB program. The analysis assured adequate lateral and longitudinal handling

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characteristics and was used to complete a final sizing of the vertical tail. Figures 4.6.2a-b show the results of stability derivative calculations for the PA and MSA. Figure 4.6.2c describes the relevant variables.

Sideslip

Roll Rate

Yaw Rate

-0.680

𝐶𝑦 𝑝

-0.052

𝐶𝑌 𝑟

0.304

-0.091

𝐶𝑙 𝑝

-0.649

0.161

0.214 Angle of Attack 𝐶𝐿 𝛼 4.17 𝐶𝑀 𝛼 -1.15

𝐶𝑛 𝑝

𝐶𝑙 𝑟

-0.018 Pitch Rate -5.50 -17.7

𝐶𝑛 𝑟

-0.341

𝐶𝑌 𝑟

Yaw Rate 0.090

𝐶𝑦 𝛽 𝐶𝑙 𝛽 𝐶𝑛 𝛽

𝐶𝑍 𝑞 𝐶𝑀 𝑞

Figure 4.6.2a – PA Stability Derivatives

Sideslip 𝐶𝑦 𝛽

-0.380

𝐶𝑦 𝑝

Roll Rate -0.030

𝐶𝑙 𝛽

-0.034

𝐶𝑙 𝑝

-0.671

0.157

0.078 Angle of Attack 𝐶𝐿 𝛼 3.43 𝐶𝑀 𝛼 -0.537

𝐶𝑛 𝑝

𝐶𝑙 𝑟

-0.021 Pitch Rate -1.89 -3.30

𝐶𝑛 𝑟

-0.242

𝐶𝑛 𝛽

𝐶𝑍 𝑞 𝐶𝑀 𝑞

Figure 4.6.2b – MSA Stability Derivatives

Figure 4.6.2c – Aircraft Coordinate System In general, the calculated stability derivatives represent a stable design. The same is true of the loaded MSA aircraft (values for shown). Note that both aircraft have a negative value for the roll stability derivative, Clβ, indicating roll stability. This met the primary design requirement for the stability and controls sub-team. Figure 4.6.2d demonstrates the influence of each component. The production aircraft is more stable (has a larger negative derivative) because it has 1.5o of dihedral and a larger vertical tail relative to the wing size. Dihedral was not chosen for the MSA because it has much larger wings, which

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lead to a larger moment of inertia and a larger roll-damping derivative. Pilot input will be sufficient to stabilize the MSA in roll. Aircraft

Wing

Vertical Tail

Fuselage

Total

PA

-0.026

-0.043

-0.022

-0.091

MSA

2.77x10-4

-0.026

-0.008

-0.034

Figure 4.6.2d – Breakdown of Roll Stability Derivative 4.7 Mission Performance Estimates Figure 4.7 provides estimates of our mission performance. The PA will be carried as a single part, so only one lap is needed during mission 2. During the bonus mission we only need to install the payload and batteries, so it will be completed easily. Given the generous time windows for missions 1 and 2, and our own pilot’s experience flying the course, it is predicted those missions will be completed as well. Mission

1

2

3

Bonus

Score 2 4 2 2 Figure 4.7 – Mission performance estimates

5.

Detailed Design

5.1 Dimensional Parameters Figure 5.1a lists the pertinent dimensional parameters for the final design of the PA. It includes overall dimensions as well as the dimensions of key subsystems. Horizontal Tail Fuselage Span 20" Length 36.20" Overall Dimensions MAC 6.35" Width 4.47" Length 41.37" Area 124 Height 5.03" Width 65.47" Airfoil Flat Plate Main Landing Gear Height 12.18" Incidence 0º Length 1.00" Wing Vertical Tail Width 10.20" Span 65.47" Span 10" Height 5.60" MAC 7.6" MAC 7.12" Wheel Diameter 1.50" 2 2 Area 3.208 ft Area 71.2 in Ground AoA Aspect Ratio 8 Airfoil Flat Plate 20º Airfoil MH114 Wing apex to horizontal tail apex distance 26.69" Incidence 0º Wing apex to vertical tail apex distance 27.11" Ailerons(2) Elevator Rudder Span 24" Span 18.5" Span 10" Percent Chord 20% Percent Chord 29.29% Percent Chord 35.11% Maximum defl. ± 30º Maximum defl. ± 30º Maximum defl. ± 30º Figure 5.1a – PA Dimensional Parameters Production Aircraft

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Figure 5.1b lists the pertinent dimensional parameters for the final design of the MSA. It includes overall dimensions as well as the dimensions of key sub systems. Manufacturing Support Aircraft Overall Dimensions Length 51.67" Width 66.60" Height 17.62" Wing Span 66.60" MAC 15.5" Area 7.169 ft2 Aspect Ratio 5 Airfoil MH114 Incidence 0º Ailerons(2) Span 22" Percent Chord 20% Maximum defl. ± 30º

Horizontal Tail

Fuselage Span 21.5" Length 36.20" MAC 12.75" Width 4.47" Area Height 5.03" Airfoil NACA 0012 Main Landing Gear Incidence 0º Length 1.00" Vertical Tail Width 10.20" Span 10.33" Height 5.60" MAC 13.55" Wheel Diameter 1.50" Area 137.8 in2 Ground AoA Airfoil NACA 0012 20º Wing apex to horizontal tail apex distance 28" Wing apex to vertical tail apex distance 28" Elevator Rudder Span 21.5" Span 9.36" Percent Chord 21.18% Percent Chord 25.83% Maximum defl. ± 30º Maximum defl. ± 30º Figure 5.1b – MSA Dimensional Parameters

5.2 Structural Characteristics 5.2.1 Load Paths The PA has two tapered 3/18” thick basswood spars that compose the wing’s main spar and run the span of the wing. The structure of the fuselage is composed of a series of ribs held in place by 1/8 and 1/16” thick balsa sheeting. The front landing gear is manufactured from carbon fiber and epoxy resin. These structure support the bulk of the loads during flight and landing. The wing spar and wing-fuselage connection support the loads imposed by the lift as well as some of the payload load. The fuselage ribs support the loads created by the weight of the payload and batteries as well as the thrust created by the motor. The combination of sheeting and ribs creates a rigid structure that helps to transfer the load from the tail lift and the associated moment across the entire fuselage. During landing the front landing gear takes the bulk of the associated loads. The landing gear is located near the theoretical center of mass for this reason, underneath the wing connection rib to help dissipate the load up through the fuselage. The design of the PA heavily influenced the load paths of the MSA. Because the MSA is meant to enclose the PA in its entirety, the load paths for the MSA are nearly identical with the exception of the landing gear having to be further from the center of mass in order to accommodate for the PA.

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Figure 5.2.1a – Load Paths of PA

Figure 5.2.1b – Load Paths of MSA

5.2.2 Structural Analysis We analyzed the wings of both the Production and Manufacturing Support aircraft under a wingtip loading scenario using ANSYS structural. Since the majority of the load in the PA wing is directed through the wing spar, we chose to only model this member as opposed to modeling the entire wing. We believed that this would accurately represent the true loading condition on the most critical component of the wing. Since the MSA wing has a less conventional geometry, and since the load is not directed through the span at quarter cord, we chose to model the entire wing geometry. For both of the ANSYS models created, we utilized symmetry about the center of the aircraft to reduce computation time.

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Figure 5.2.2a – PA Spar Modeling

Figure 5.2.2b – MSA Wing Modeling The PA simulation results align well with the expected behavior of the spar. The MSA also aligns well with the expected behavior of the structure. Additionally, we note that nearly all of the stress present in the MSA wing is found in the carbon fiber spar – the strongest and most resilient part of the structure; this alleviates our concerns about the unconventional design of the wing. We later validate the structure of each wing through testing. Summaries of the ANSYS results are in Figure 5.2.2c below. PA Spar

MSA Spar

MSA Wing

Material

Basswood

Carbon Fiber

Various

Maximum Deflection (mm)

29.9

28.6

40.7

Maximum Stress (MPa)

33.8

60.3

-

Factor of Safety in Failure 1.77 9.95 Figure 5.2.2c – Structural Analysis Summary

-

5.3 Aircraft Systems Design, Component Selection and Integration 5.3.1 Fuselage Design The PA fuselage had several design constraints: enough tensile strength to withstand the forces present at the wing connection and support the Gatorade, enough torsional rigidity to withstand torques created by the empennage, ready access to the Gatorade and propulsion systems, and minimal weight. In the Gatorade containment section, “U” shaped ribs are used to support the Gatorade, transferring the load to the wings. Each rib has a different shape, fitting the grooves of the Gatorade bottle. The tight fit of the ribs fully constrains the bottle with normal and frictional forces, while the open Cornell University

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top provides accessibility. The wing connection rib has more structure, as it transfers the load of the fuselage and Gatorade weight to the wing. The rest of the Gatorade containment structure is minimal, as it does not bear a significant load. Torsional and bending rigidity are obtained with pocketed balsa sheeting and MonoKote®. The front section of the fuselage is reserved for the propulsion system, and a magnetic hatch was designed for easy access to this region. A simple wing connection as described in Section 5.3.3 allows for easy payload access.

Figure 5.3.1a – PA Front Fuselage Similar to the PA, the MSA utilizes ribs that support the payload and allow for easy access. However, the length of the MSA’s payload requires that the entire fuselage be used. Because of this, payload containment ribs are used throughout the main fuselage and the tail boom. As in the PA, pocketed balsa sheeting and MonoKote® provide torsional and bending rigidity. As in the PA, the nose space is used for the propulsion system; a Velcro® hatch is used for convenient access.

Figure 5.3.1b – MSA Fuselage Interior

5.3.2 Wing Design They PA wings were designed to be lightweight, aerodynamically efficient, and stiff. In order to achieve a near-elliptic lift distribution and minimize induced drag, a taper ratio of 0.45 and an aspect ratio of 8 was selected. (Section 4.5 and Figure 5.3.2a). By designing the tapered wings with a flat top, we gained a 1.5° dihedral angle, which contributed to roll stability. The wing structure uses a balsa build-up consisting of 1/8” balsa ribs, leading and trailing edge stringers, and a main quarter-chord spar. To achieve stiffness, a wooden D-box was constructed with 1/16” balsa sheeting between the leading edge stringer and central spar. The two sections of the main spar are attached in the middle via a semi-circular Cornell University

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connection piece (figure 5.3.2b). This reduces stress concentrations at the interface and provides more surface area for the spar sections to attach. A 1/32” thick carbon fiber strip is attached to the bottom of the main spar, across the connection, to hold the bulk of the tensile load due to lift during flight.

Figure 5.3.2a – PA wings

Figure 5.3.2b – PA wings spar connection

The MSA plane wings (Figure 5.3.2c) also use a balsa build-up, and are designed to enclose the PA wings by sliding over the left and right sides. Each of the 1/8” ribs has a cutout slightly larger than the cross section of the PA plane’s wing at that point along the span. A rectangular cross section allows for a full-span, leading edge, ½” diameter, carbon fiber tube to provide bending strength for the entire wing. Torsional strength is achieved by sheeting the entire wing with 1/16” balsa and MonoKote®.

Figure 5.3.2c – MSA wings (without sheeting)

5.3.3 Wing Attachment Design The design considerations for the production aircraft’s wing attachment were tensile strength, bending stiffness, connection rigidity, and quick removal. A dowel connection between the fuselage and the spar at the wing’s quarter cord was implemented for quick and easy payload access. A Velcro® strap was used to fix the trailing edge. Two section views of the wing connection are shown below, before and after it is in place. The MonoKote® covering the Velcro® strip are not pictured.

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Figure 5.3.3a – PA Wing Attachment (Off)

Figure 5.3.3b – PA Wing Attachment (On)

The MSA wing connection is similar to that of the PA. Dowels were again chosen to affix the wing to the fuselage, though they slide on laterally instead of from behind as in the PA. Velcro® is again used to prevent the dowels from sliding out. Two views of the MSA wing connection are shown below, picturing the structure before and after the wing is put in place. The MonoKote® covering and the Velcro® attachment strip are not shown.

Figure 5.3.3c – MSA Wing Attachment (Off)

Figure 5.3.3d – MSA Wing Attachment (On)

5.3.4 Empennage Design The PA empennage is designed to provide pitch and yaw stability while maintaining a lightweight structure. Sized by Raymer’s volume ratios, the empennage is constructed of plywood and implements a strutted structure for reduced weight. The vertical and horizontal components are tapered, reducing weight and providing an aerodynamic profile. The vertical stabilizer is notched and glued into the horizontal stabilizer, and both are then affixed to the fuselage with screws. Rudder and elevator servos are mounted to the horizontal stabilizer.

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Figure 5.3.4a – PA Empennage The MSA empennage is designed so that it can slide over the PA empennage. It is comprised of three separate pieces: two horizontal tail halves, and a vertical tail. All of these components are made of hollow symmetric airfoil ribs that create an aerodynamic profile, and have enough volume to contain the PA tail. Dowels are used to connect both halves of the horizontal stabilizer and to the fuselage, and screws are used to affix the vertical stabilizer to the horizontal stabilizer. Small boxes are built into the empennage to mount the elevator and rudder servos.

Figure 5.3.4b – MSA Empennage

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5.3.5 Landing Gear Design The landing gear were designed to be as light as possible while retaining steering capabilities and facilitating the storage of the PA within the MSA. The tail gear of the PA attaches directly to the bottom of the rudder and extends less than 1” below the end of the PA fuselage, allowing it to fit easily into the fuselage of the MSA. The MSA landing gear is much longer, so a plastic support is attached to the bottom of the MSA fuselage, extending outward for the MSA tail gear shaft to pass through it, and be constrained as it rotates.

Figure 5.3.5a – Both tail gear in Mission 2 configuration Each landing gear is sized to provide ½” of clearance to the propeller when the central body of the fuselage is parallel to the ground (an extreme case, as the landing gear on both planes are mounted well in front of the centers of gravity). The front landing gear of the MSA is taller to ensure the containers around the front landing gear of the PA never contacted the ground. Each front landing gear is mounted directly onto the balsa sheeting of the fuselage below support ribs to ensure that the load is transferred to a rigid part of the aircraft.

Figure 5.3.5b – Both front gear in Mission 2 configuration

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5.3.6 Electronics Selection All electronics were selected to be as light as possible while still meeting requirements for safe operation. The aircraft uses standard servo motors to control the ailerons, elevator, and rudder during flight. Three servo motors were considered: Servo

Futaba S3111

Futaba S3114

Hyperion DS11-AMB

Weight [oz.]

0.23

0.27

0.37

Torque [oz.-in.]

8.3

20.8

29.2

Width [in.]

0.433

0.433

0.453

Speed [s/60 deg.]

0.12 0.1 Table 5.3.6a - Servo Selection

0.18

The Futaba S3114 servo was selected because it offers the most torque and speed compared with its weight. The required torque was determined through hinge moment calculations (Section 4.6), which demonstrated that the additional torque available from the DS11-AMB was not necessary. The competition requires that a separate battery power the receiver and servos. Using the servos that were specified above and the known maximum flight time of five minutes for both aircraft, we calculated an appropriate required battery capacity as shown in Figure 5.3.6b. Servo Current (mA) 300

Total Current (mA)

Flight Time (hr)

1200 0.083 Figure 5.3.6b - Receiver Battery Sizing

Capacity (mAh) 100

In addition to the conservative estimate of full time maximum current, we implemented a factor of safety of 1.5 to arrive at a capacity of 150 mAh. We chose the KAN 160 battery cell to fill this role in a five cell pack. We chose the E-flite 40-Amp brushless ESC as the speed controller for both the PA and the MSA because it was the lightest option available that met the requirements of our propulsion systems. We selected a Spektrum AR6210 6-channel receiver for both aircraft based on its weight and compatibility with all other electronic components. As per competition rules, an arming device is required to "safe" the motor when handling the plane; a deans-style connector was chosen to fulfil this requirement.

Component

Name

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Servo Motor Futaba S3114

Electronic Speed Receiver Controller (ESC) E-flite 40-Amp Spektrum KAN-160 Brushless ESC AR6210 Figure 5.3.6c - Selected Electronic Components Receiver Battery

Arming Device Deans connector

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5.4 Aircraft Component Weight and Balance Figure 5.4a lists the breakdown of the production aircrafts empty weight and center of mass by component. All X and Z coordinates are relative to the plane’s aerodynamic center. Figure 5.4b displays the location of the calculated center of mass, as well as the measured center of mass of the built aircraft, aerodynamic center, and the direction of the axes. Figures 5.4c and 5.4d display the same information for the manufacturing support aircraft.

Component

Weight (lbf)

Tail Boom Tail Front LG Fuselage/Nose Motor Propeller Battery Wing Mission 3 Payload Empty Weight Empty Weight w/ Batteries Mission 3 Total Weight

0.11 0.138 0.154 0.138 0.324 0.25 0.316 0.588 2.25 1.702 2.018 4.268

CG Location (in) X Z 0 -15.1 -0.04 -29.7 0 0.33 0 -2.5 0 4.12 0 6.79 -0.42 1.51 0 -0.65 0 -2.68 -0.003 -2.000 -0.069 -1.450 -0.032 -2.099

Figure 5.4a - PA Component Weights and Center of Mass

Figure 5.4b - PA Center of Mass Diagram

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Component Tail Boom/Back Fuselage Tail Front LG Front Fuselage/Nose Motor Propeller Battery Wing Mission 2 Payload Empty Weight Empty Weight w/ Batteries Mission 2 Total Weight

Weight (lbf) 0.334 0.478 0.164 0.394 0.324 0.3 0.408 1.512 2.25 2.402 3.914 6.164

CG Location (in) X Z 0.06 -8.9 0.09 -31.7 0 3.17 -0.12 2.87 0.03 12 0.03 13.41 -0.01 8.01 0 -1.91 0.000 -2.893 0.014 -4.767 0.008 -2.091 0.005 -2.384

Figure 5.4c - MSA Component Weights and Center of Mass

Figure 5.4d - MSA Center of Mass Diagram

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5.5 Flight Performance Parameters Figure 5.5 lists relevant flight performances calculated by the mission model and from MATLAB codes discussed in Section 4 for each of the competition’s three missions. The values shown reflect the final weight and geometry estimates developed during the detailed design phase. Parameter Mission 1 Mission 2 CLmax 1.42 1.42 CLtakeoff 1.01 1.08 CLcruise 0.288 0.413 CD0 0.03 0.03 (L/D)cruise 4.89 5.62 Stall speed (ft/s) 22 24.4 Takeoff speed (ft/s) 27.5 26.2 Takeoff distance (ft) 42 51 Takeoff angle (o) 7.2 7.9 Cruise speed (ft/s) 36.7 32.3 Cruise angle (o) -4.82 -3.94 Turn rate (o/s) 138 120 2 Wing loading (lbf/ft ) 0.55 0.8 Total flight time (s) 177 77.8 Figure 5.5 – Flight Performance Parameter Estimates

Mission 3 1.42 1.13 0.587 0.028 9.78 29.7 35.2 72 8.4 39.6 -3 225 1.38 154

5.6 Rated Aircraft Cost The rated aircraft cost (RAC) is determined by the empty weight and battery weight of the two aircraft, and the number of components that the production aircraft is split into. The RAC is calculated by the following formula:

𝑅𝐴𝐶 = 𝐸𝑊1 ∗ 𝐵𝑊1 ∗ 𝑁𝑐𝑜𝑚𝑝𝑜𝑛𝑒𝑛𝑡𝑠 + 𝐸𝑊2 ∗ 𝐵𝑊2 With the calculated weights from Figures 5.4a and 5.4c, the RAC was calculated to be 1.97. EW1 (lb.)

Wt_Battery_1 (lb.)

1.70

0.32

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N_Components

EW2 (lb.)

Wt_Battery_2 (lb.)

1 3.51 Figure 5.6 – RAC Calculation

0.41

RAC 1.97

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5.7 Mission Performance Summary In Figure 5.7, we document the aircraft’s performance for each of the three missions and for each of the phases of flight modeled. We generated these results using the mission model and the MATLAB codes discussed in Section 4. Mission One Summary Flight phase Qty. Final speed (ft/s) Distance (ft) Time (s) Takeoff 1 27.5 42 3.3 Climb 1 31 320 11.8 Cruise 3 36.7 1500 41.5 180 turn 6 28.7 36 1.3 360 turn 3 28.5 72 2.6 Accelerate 9 36.7 82 2.4 Mission One Totals 6032 177 Mission Two Summary Flight phase Qty. Final speed (ft/s) Distance (ft) Time (s) Takeoff 1 24.2 51 4.7 Climb 1 28.8 320 12.1 Cruise 1 32.2 1500 46.9 180 turn 2 26.9 39 1.5 360 turn 1 26.5 78 3 Accelerate 3 32.2 87 2.7 Mission Two Totals 2288 77.8 Mission Three Summary Flight phase Qty. Final speed (ft/s) Distance (ft) Time (s) Takeoff 1 29.7 72 5.1 Climb 1 33.4 320 10.7 Cruise 3 39.6 1500 38.3 180 turn 6 31.1 25 0.8 360 turn 3 29.5 50 1.7 Accelerate 9 39.6 56 1.5 Mission Three Totals 5696 154 Figure 5.7 – Mission performance summary

Capacity (mAh) 13.4 74.6 147.3 38.4 71.2 14.5 1104 Capacity (mAh) 17.2 79.5 164.3 43.9 79.3 20.4 489.3 Capacity (mAh) 20.6 63.1 155 31.7 63.8 15.4 1069

5.8 Drawing Package In this section, we present a full drawing package. The package includes a three view drawing, structural arrangements, systems layout, and the payload configuration for both the production aircraft and the manufacturing aircraft.

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Manufacturing Plan and Processes Planned Actual

Date 23-Jan 24-Jan 25-Jan 26-Jan 27-Jan 28-Jan 29-Jan 30-Jan 31-Jan 1-Feb 2-Feb 3-Feb 4-Feb 5-Feb 6-Feb 7-Feb 8-Feb 9-Feb 10-Feb 11-Feb 12-Feb 13-Feb 14-Feb 15-Feb 16-Feb 17-Feb

6.

PA Iteration 5 Laser Cut Parts Tail Boom Assembly Main Body of Fuselage Assembly Empennage Assembly Wing Assembly Monokote Mount Motor Attach Landing Gear MSA Iteration 2 Laser Cut Parts Main Fuselage Assembly Nose Assembly Empennage Assembly Wing Assembly Monokote Integration of Sub-Assemblies Mount Motor Figure 6 - Manufacturing Milestones Chart: Planned and Actual

6.1 Manufacturing Process Selection Manufacturing and construction methods play a crucial role in the weight and strength of an aircraft design. Our team considered three major fabrication methods for our two designs, with the following characteristics: 

Wood Build-Up – With exceptional strength-to-weight ratio, this method provides great scoring characteristics and a modest amount of manufacturing complexity.



Foam – This method provides a fast manufacture process but also usually results in heavier components than a comparable wood build-up.



Carbon Fiber – This method yields extremely strong components that are often heavier than either wood or foam. It provides flexibility to create nearly any shape that is desired, or can be purchased in a number of preformed shapes.

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We compared these methods based on aircraft weight, strength, and manufacturability in Figure 6.1a. We decided to manufacture the majority of both aircraft using a balsa, plywood, and basswood build-up, primarily due to the scoring impact of aircraft weight. In areas of significant structural loading, both aircraft utilize a small number of carbon fiber components for added strength and stiffness.

Category Weight Strength Manufacturability Total

% 50 30 20 100

Wood Build-Up 5 4 3 4.3

Foam 3 3 3 3

Carbon Fiber 3 5 3 3.6

Figure 6.1a – Manufacturing Process Selection 6.2 Subsystems Manufacture In the following sections, we document the manufacturing process for the wing, fuselage, empennage, and landing gear of both aircraft. 6.2.1 Fuselage Manufacture The fuselage of the PA is manufactured as two modules. The nose and Gatorade containment sections form the main fuselage module while the tail boom and empennage connection comprise the tail boom module. This allows fuselage components to be manufactured in parallel, decreasing lead time and construction periods. The two modules are joined using 6-32 nylon screws and t-nuts. The MSA fuselage is manufactured in three segments to allow for parallel construction; nose, main fuselage, and tail boom. These are joined using dowels and Velcro.

Figure 6.2.1a – The front and back modules of the PA fuselage The bulk of fuselage parts are cut using an on-site laser cutter, including the ribs, side sheeting, and empennage interface. The primary consideration for material choice is stiffness. The default choice is balsa, as it has the highest strength-to-weight ratio of woods. Fuselage ribs and fuselage side sheeting are cut from 1/8” thick balsa. Balsa sheeting is used for the tail boom, 1/16” in the PA and 1/8” in the MSA. As wood is highly anisotropic, balsa struts (cut by hand) are used to strengthen and support

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components that need additional “against-the-grain” strength. Structural components under significant loads, such as the wing connection rib and motor mount rib, are cut from 1/8” thick plywood. Sheeting and internal structural components are designed to notch at joints, minimizing construction time and ensuring proper alignment. Sheets that require bending are soaked in water for 24 hours, constrained to produce the required bend, and left to dry. The fuselage and tail boom are assembled by aligning all joints and bonding the interfaces using instant CA glue. Balsa and basswood struts run from the motor mount to the first fuselage rib to reinforce the nosecone. Once the skeleton of the aircraft is completed, the segments named above are separately coated. MonoKote® is adhered to the surface of the aircraft using a hot iron. Due to the size of the fuselage, the MonoKote® is applied in sections. Once the plane has been covered, a heat gun is used to smooth the wrinkles and tighten the skin. The last parts added are the miscellaneous connectors.

Figure 6.2.1b – MSA Fuselage Manufacturing Process

6.2.2 Wing Manufacture The PA wing is composed of laser cut balsa ribs, one main basswood spar running through the span at the quarter chord, a balsa leading edge spar, and a balsa dowel running through the first three ribs toward the trailing edge. The ribs are laser cut with notches at quarter chord, which slide into the small notches at the top of the main spar to ensure alignment. Wetted balsa sheets are bent between the leading edge spar and main spar to maintain the airfoil shape and provide torsional rigidity through a balsa D-box. Similarly, the trailing edge is covered with 1” wide balsa sheets. CA glue is used to bond all wood joints together. The carbon fiber strip below the spar connection is attached with epoxy. The entire wing is covered using MonoKote® to ensure laminar flow over the wing. The MSA wing is composed of laser cut balsa ribs, a carbon fiber rod at quarter-chord, a wooden dowel near the trailing edge, and balsa spars along the leading and trailing edges. Due to the infeasibility of notching the carbon fiber spar, a wing construction rig was built from laser cut balsa to ensure alignment. The ribs were aligned using the notches in this construction rig, then the leading and trailing

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edge spars were attached using CA glue. At this point, the PA wings were slid into the MSA wings (figure 6.2.2) to ensure proper fit. Next, the rear dowel was attached to the outer 6 ribs on each side using CA glue. Finally, the 1/16” balsa sheeting is wetted, pressed to the form of the ribs, and attached using CA.

Figure 6.2.2 – Construction of MSA Wing

6.2.3 Empennage Manufacture The PA and MSA empennage parts are laser cut from plywood for the outer frames and spars, where the stress is the highest, and balsa for the inner struts and airfoil ribs. Balsa sheeting is soaked and then bent overnight using clamps and weights to construct the continuous L-shaped beams in the MSA. Wood parts are joined with CA glue in softwood-to-hardwood connections and wood glue in hardwood-to-hardwood connections to provide high-strength, low-weight bonds. All wood edges are sanded down to a smooth finish and the entire tail is covered in MonoKote® to provide tensile and compressive strength and block internal airflow. CA hinges are slotted into the vertical and horizontal stabilizers to attach the control surfaces. Control horns, steel rods, and plastic clevises are used to link the servos to the control surfaces.

Figure 6.2.3 - Empennage Manufacturing Process 6.2.4 Landing Gear Manufacture The front landing gear are built from a carbon fiber and felt-core composite. A male mold is cut from insulation foam with a hot-wire cutter and sanded down. The mold is covered with two layers of overlapping packing tape and coated with mold-release wax. Layers of carbon fiber cloth and epoxy are Cornell University

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applied, with felt core in the middle of the layers. Several pairs of landing gear of varying composition are produced at once to enable weight optimization.

Figure 6.2.4 – Front landing gear carbon fiber layup 7. Testing Plan A rigorous test regime was designed to validate the estimated performance characteristics generated during the design process and compare them with actual performance parameters. Figure 7.1a summarizes the goals of testing and the data each test was designed to provide. 7.1 Testing Schedule Structural Testing  Wing Loading: The wing must be able to sustain the loading applied during a 2g fully loaded turn without failure or substantial deflection. This will be simulated by wingtip and spar loading tests.  Landing Gear Strength: The landing gear must be able to sustain and absorb a landing shock of the fully loaded aircraft without failure. Propulsion Testing  Static Thrust Test: Static thrust testing will be performed to validate theoretical calculations of current draw and thrust values.  Receiver Battery: The receiver battery must be able to power the receiver and all servos for a minimum of five minutes. Payload Testing  PA Containment: The PA must be capable of containment inside the MSA, including small protrusions such as servos and control horns.  MSA Assembly: The components of the MSA must fit together precisely in order to properly constrain the PA. This capability will be validated by testing the tolerances of crucial components, e.g. the holes constraining the carbon fiber spar. Flight Testing  Stability and Control: Both aircraft must have acceptable performance in roll, pitch, and yaw, as observed during flight testing.  Taxi: Both aircraft must have the capability to maneuver in a straight line to facilitate takeoff.  Mission Performance: Both aircraft must fly the required number of laps and land safely within five minutes. Figure 7.1a – Subsystem and complete aircraft testing checklist

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Date

6-Sep 13-Sep 20-Sep 27-Sep 4-Oct 11-Oct 18-Oct 25-Oct 1-Nov 8-Nov 15-Nov 22-Nov 29-Nov 6-Dec 13-Dec 20-Dec 27-Dec 3-Jan 10-Jan 17-Jan 24-Jan 31-Jan 7-Feb 14-Feb 21-Feb 28-Feb 6-Mar 13-Mar 20-Mar 27-Mar

Planned Actual

Production Aircraft Wing Stucture Thrust Rig Battery Landing Gear Manufacturing Support Aircraft Wing Spar Thrust Rig Battery Containment Landing Gear Flight Testing Production Flight 1 Production Flight 2 Production Flight 3 Mission 1 Mission 2

Winter Break

Figure 7.1b – Testing Schedule 7.2 Flight Testing An iterative approach was taken to optimize the aircraft—the performance of the full aircraft solution was measured and improved. Each prototype test simulated the competition requirements. Procedures were designed to ensure the functionality of each subsystem prior to testing the full aircraft. This minimized risk to the aircraft and facilitated evaluation of individual systems.

1

Test Item Connect batteries, receiver, speed controller, control surfaces

 

2

Deflect ailerons, elevators, and rudder

  

3

Full throttle test while the aircraft is held in place

 

4

5

Taxi test



Objectives Test functionality of the batteries, receiver battery, and arming switch. Check that the angle of deflection of the control surfaces is adequate Ensure that the servos are still properly zeroed. Check that the rear wheel will rotate well. Test functionality of the propeller and motor Check for any excessive strain or vibrations due to throttling up. Ensure that rear-wheel steering is adequate Observe landing gear deflection while rolling Test stability and control

Takeoff and fly laps that approximate  mission requirements Table 7.2 – Flight Test Procedure

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One of the primary goals of a flight test is to evaluate the stability of the aircraft in flight and how easy it is to control. To this end we take into account pilot comments about stability and control, particularly in regards to ease of maneuverability while completing the mission tasks.

Figure 7.2 – Production Aircraft Iterations 1 and 4 in flight 7.3 Propulsion Testing The propulsion sub-team completed a series of tests on propulsion components in order to confirm theoretical predictions and optimize system selections. We applied the principles of moment balance and designed a custom thrust rig to transfer motor and propeller thrust values to a digital scale. Electrical characteristics such as voltage, current, and power were stored using an Eagle Tree data logger. A watt meter was used to view these characteristics live; this was a failsafe in the event that a test brought currents too high for the safety of the batteries. The pitch speed of the propeller was determined using an anemometer placed directly downstream of the propeller. Tests were performed for all designed systems and variations thereon by ramping the motor from zero to full throttle and then back to zero.

Figure 7.3 – Propulsion Test Rig Cornell University

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7.4 Structural Testing After conducting finite-element analysis on the aircraft structures, thorough structural testing was performed on the subsystems that experience the greatest loads: the wings and the landing gear. 7.4.1 Wing Structure The bending strength of the wing is based on the main spar. Basswood was selected based on specific strength and density. To determine the best connection design, several configurations were subjected to a three point beam bending test. The results are shown in Figure 7.4.1a. Note that short sections were tested, so the values do not correspond to flight loads. It was determined that the semicircle configuration reinforced with a carbon fiber strip was optimal, as it bore the most load before breaking. As expected, the failure occurred along the joint, which was bonded with epoxy. Connection Type Notched Semicircle Notched

Maximum Spar Connection Load (lb.) Without Carbon Fiber Strip 6 12

With Carbon Fiber Strip 26 32 Semicircle

Figure 7.4.1a – Spar Connection Testing For the MSA, an entirely wooden wing structure was considered undesirable due to the hollow rib shape, given the requirement that the wing must slide over the PA wings. A carbon fiber rod meets both of these requirements, and tests were conducted to determine the best spar configuration. A 1/2” rod and a ¼” rod were tested on the load test rig, along with a combination of several differently sized sections of rods in a "telescoping" configuration that was thickest at the root of the wing and thinnest at the tip.

Figure 7.4.1b – Carbon Fiber Rod Load Test Cornell University

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The test demonstrated that all of the configurations would be strong enough to withstand the approximately nine pounds needed for a loaded flight. The ½” rod was chosen because it took the loads with minimal deflection, so the MSA wing would avoid contact with the PA wing during transport. Spar Type (diameter) 0.25”

Deflection (in) 1.48

Telescoping (0.25” - 0.5”) 0.48” Figure 7.4.1c – Carbon Fiber Spar Deflection at 10 lb. Loading

0.827 0.512

7.4.2 Landing Gear The carbon fiber landing gear were tested to optimize weight. A set of landing gear with nearfinal-design dimensions were manufactured with thicknesses varying from four to ten layers of carbon fiber. These were tested to determine the minimum number of layers needed to ensure mission success.

Figure 7.4.2a – Landing Gear Test Rig The lightest gear which prevented deflection significant deflection (0.75” of deflection in an 11.5” width) when loaded with 10 pounds (approximately twice the weight of either aircraft) was selected. Number of Carbon Fiber Layers 4 6 8 10

Load (lbs.) 0 1.72 5.098 Deflection (inches) 0 0.36 0.91 0 0.18 0.27 0 0.09 0.18 0 0.09 0.14 Figure 7.4.2b – Carbon Fiber Landing Gear Data

6.73 1.36 0.36 0.23 0.18

10.1 2.27 0.55 0.27 0.23

7.5 Miscellaneous Testing 7.5.1 Electronics The system of sliding the MSA wings over the PA wings prompted a redesign of how the servos interacted with the PA ailerons. Specifically, neither the control horns on the ailerons nor the servos on the wings could protrude more than 1/8”. Two solutions were envisioned: a direct drive system and a Cornell University

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system of lines in tension to pull and push the ailerons. A proof of concept was built and tested for each idea by modifying old sets of wings. Ultimately the system of tension lines was selected because it produced a higher angle of deflection and demonstrated greater reliability. 8.

Performance Results

8.1 Subsystem Performance 8.1.1 Propulsion Figure 8.1.1a shows a sample of the data collected from propulsion system testing using the Eagle Tree data logger. The output in this figure shows battery pack voltage, current, and power in purple, yellow, and red, respectively.

Figure 8.1.1a – Sample Propulsion Data The team extensively tested the propulsion systems detailed in Section 4.4 in addition to systems that deviated slightly from those described. The reason for the extra testing was to see which systems provided the best true performance. We found, however, that the systems of Section 4.4 proved to be the best after all; this strongly validated our theoretical models and preliminary analysis. Figure 8.1.1b details the results of our tests; the shaded systems were those selected for competition. Motor Neu 1105/6D/P29 Neu 1105/6D/P29 Neu 1105/6D/P29 Neu 1105/6D/P29 Neu 1105/6D/P29 Neu 1105/6D/P29 Neu 1105/2.5Y/P29 Neu 1105/2.5Y/P29 Neu 1105/2.5Y/P29

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Power Pitch Speed (W) (ft/s) APC 17x10E 8 16.2 129 43 APC 18x12E 7 17.1 117 43 APC 20x13E 6 18.2 105 38 APC 18x12E 9 23 188 50 APC 20x13E 8 25.1 177 44 APC 22x12E 7 24.6 152 35 APC 16x10E 8 22.4 164 51 APC 18x12E 7 21.7 140 49 APC 18x12E 6 22.3 122 44 Figure 8.1.1b – Propulsion System Testing Results Propeller

# Cells

Current (A)

Thrust (oz.) 46 46 44 62 61 55 54 50 48

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Receiver battery testing indicated that the receiver was powered continuously for six minutes. These results were verified during flight tests in that the receiver never lost power during flight. 8.1.2 Structures To validate our wing design and construction methods, a pair of wings was tested to failure with a three-point beam bending test. The wings held the expected loads with minimal deflection and performed well to a factor of safety of at least 2. Although a 19 lb. load is excessive, the resistance to deflection—the design goal—validates what would otherwise be considered an overbuilt connection design. Load (lb.) 5 10 19

PA Deflection (cm) MSA Deflection (cm) 1.5 1.3 2.7 2.5 5.1 Figure 8.1.2a - Wing Loading Test

Expected load during flight Factor of safety of 2 Failure

For both the PA and MSA wings, observed deflection was approximately half the predicted values. This was expected because neither the sheeting nor the MonoKote® were modeled in the ANSYS simulations.

Figure 8.1.2b – Wingtip Loading for PA (top) and MSA (bottom)

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8.2 Aircraft Flight Performance Figure 8.2a summarizes the issues we encountered during each test and the steps we took to improve the design as a result. Figures 8.2b and 8.2c display the PA during a flight test. Flight 1 2 3 4

Issue Solution High weight, low speed Pocket sheeting, increase wing efficiency, reduce wing size. Lack of control during taxi Increase tailboom length Roll instability Offset CG, reduce manufacturing defects in wings Pitch and Yaw instability Increase tail size, wing dihedral, increase tailboom length Figure 8.2a – Flight Test Issues and Solutions

Figure 8.2b – Production Aircraft Flight Test

Figure 8.2c – Production Aircraft Flight Test Cornell University

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9. [1]

References "2015/2016 Rules and Vehicle Design." AIAA DBF. Web. 31 Oct. 2015.

http://www.aiaadbf.org/2016_files/2016_rules_final_20151031.html [2]

Raymer, Daniel. Aircraft Design: A Conceptual Approach. 3rd. Washington, D.C.: American Institute

of Aeronautics and Astronautics, Inc., 1999.

[3]

Anderson, John. Introduction to Flight. 6th. New York: McGraw-Hill, 2005.

[4]

Caughey, David A. Introduction to Aircraft Stability and Control Course Notes for MAE 5070. Ithaca,

NY: Cornell University, 2011. MAE 5070 Dynamics of Flight Vehicles. 12 Apr. 2011. Web. 27 Sept. 2013. [5]

Etkin, B., & Reid, L. D. Dynamics of Flight Stability and Control Third Ed. Hoboken, New Jersey:

John Wiley and Sons, 1996

[6]

Gilruth, R. R., and M. D. White. "Analysis and Prediction of Longitudinal Stability of Airplanes."

NASA Technical Reports 711th ser. 1941.

[7]

Nelson, Robert C. Flight Stability and Automatic Control. 2nd ed. Boston, MA: WCB/McGraw Hill,

1998. Print.

[8]

Phillips, Warren F. Mechanics of Flight. Second Edition. Daryaganj, New Delhi: John Wiley & Sons,

2010. Print. [9]

Muller, Marküs. “propCalc – Calculator for Propeller, Powered by Neu Motors”. Neu Motors. Web.

Oct. 2012. http://www.ecalc.ch/motorcalc_e.asp?neumotors

Cornell University

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