2014dbf_sandiegostateuniversity

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San Diego State University San Diego State University

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San Diego State University

Table of Contents

1.0 Executive Summary ................................................................................................................ 4

2.0 Management Summary ........................................................................................................... 5

2.1 Team Organization ........................................................................................................................ 5

2.2 Milestone Summary ...................................................................................................................... 7

3.0 Conceptual Design .................................................................................................................. 8

3.1 Mission Requirements ................................................................................................................... 8

3.2 Translating Mission Requirements Into Design Elements .......................................................... 11

3.3 Design Selection ......................................................................................................................... 12

3.4 Selected Conceptual Design ....................................................................................................... 17

4.0 Preliminary Design ................................................................................................................ 18

4.1 Design and Analysis Methodology .............................................................................................. 18

4.2 Mission Model ............................................................................................................................. 19

4.3 Propulsion ................................................................................................................................... 23

4.4 Aerodynamics.............................................................................................................................. 25

4.5 Stability and Control .................................................................................................................... 28

4.6 Aircraft Mission Performance ...................................................................................................... 30

5.0 Conceptual Design ................................................................................................................ 30

5.1 Dimensional Parameters ............................................................................................................. 31

5.2 Structural Characteristics ............................................................................................................ 32

5.3 System Design, Component Selection and Integration .............................................................. 33

5.4 Aircraft Component Weight and Balance .................................................................................... 36

5.5 Flight Performance Summary ..................................................................................................... 38

5.6 Mission Performance Summary .................................................................................................. 38

5.7 Drawing Package ........................................................................................................................ 40

6.0 Manufacturing Plan and Processes .................................................................................... 45

6.1 Time Table .................................................................................................................................. 45

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6.2 Materials Selection ...................................................................................................................... 45

6.3 Wing and Fuselage ..................................................................................................................... 46

6.4 Stabilizers and Control Surfaces ................................................................................................. 47

6.5 Landing Gear............................................................................................................................... 47

7.0 Testing Plan ........................................................................................................................... 48

7.1 Propulsion System Testing ......................................................................................................... 49

7.2 Structural Testing ........................................................................................................................ 50

7.3 Flight Testing ............................................................................................................................... 52

8.0 Performance Results ............................................................................................................ 52

8.1 Propulsion Testing Results ......................................................................................................... 52

8.2 Structural Testing Results ........................................................................................................... 56

8.3 Flight Testing Results .................................................................................................................. 56

9.0 References ............................................................................................................................. 60

Nomenclature and Abbreviations

Symbol Name Symbol Name

α Angle of Attack µK Coefficient of Kinetic Friction

CD Coefficient of Drag ρ Air Density

CL Coefficient of Lift RAC Rated Aircraft Cost

CM Pitching Moment Coefficient Re Reynolds Number

CFD Computational Fluid Dynamics S Wing Planform Area

CG Center of Gravity T Thrust

g Gravitational Constant T0 Static Thrust

J Advance Ratio T.O.Dist Takeoff Distance

l Length V Velocity

L/D Lift to Drag ratio VStall Stall Velocity

L/W Lift to Weight ratio VTO Velocity at Takeoff

µ Kinematic Viscosity W Weight

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1.0 EXECUTIVE SUMMARY

The purpose of this document is to present the analysis, manufacturing, and design methodology

performed by the San Diego State University team for the competition aircraft to be entered into the

2013-2014 AIAA Design Build Fly competition. The objective of the aircraft is to complete all of the

missions and requirements set forth by the contest rules while achieving the highest flight score and

lowest rated aircraft cost (RAC).

This year’s theme is a backcountry rough field bush plane. The aircraft will simulate a large cargo

payload and a medical emergency mission where two patients on gurneys and medical attendants

must be placed securely inside the fuselage. Mission one requires the aircraft to fly for four minutes

with no cargo. Teams are scored on how many laps they are able to complete. Mission two is a

payload mission in which the aircraft must carry as many blocks internally as possible. Mission three

is another payload mission where the aircraft must fly three laps as fast as possible. A separate taxi

mission is also included in which the aircraft must taxi an obstacle course with a corrugated panel

floor. All missions have a maximum runway length of 40 feet. The overall score is a function of

report score, mission scores, and unloaded aircraft weight.

Scoring analysis revealed that aircraft weight would be the most critical parameter so efforts were

made to ensure the design would be as light as possible. The optimal strategy would be to maximize

Mission one and three scores and to only carry three blocks for Mission two. The additional points

gained by carrying more blocks in Mission two would not compensate for the additional weight added

to the aircraft structure and the decrease in lap times for the other missions.

A stringent constraint of 15 amps and a large cargo bay volume is required for this year’s

competition. This led the team to perform propulsion testing to determine which combination of

propeller diameter, propeller pitch, motor, and battery configuration would be ideal.

Although several designs were considered, the team concluded a conventional tail dragger

aircraft using a single tractor motor would be best. For aerodynamic purposes the payload will not be

carried side by side but rather in a line. Due to the taxi mission, the aircraft will use large wheels to

navigate the course and the use of cowlings will be employed to decrease the amount of drag from

the wheels. The cowlings will also act as a skid platform for the taxi mission to increase the

maneuverability and to reduce the oscillation of the aircraft.

To accommodate the different types of cargo required by the competition and to allow for easy

loading of payloads, the aircraft utilizes a 20 inch by 8 inch removable bay door. The inside of the

bay door is lined with Velcro and Velcro straps are used to secure the cargo for the different missions.

The cargo is mounted onto the bay door and is inserted into the aircraft.

A variety of manufacturing techniques were used in the construction of the aircraft. The wing and

fuselage consists of a balsa buildup combined with a wet layup for the control surfaces. These

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manufacturing choices emphasized creating the lightest aircraft possible. A balsa buildup was the

lightest option available to the team and was the primary manufacturing technique.

In addition to the numeric analysis performed on the aircraft’s performance, physical testing was

performed to verify the numerical results and to help make design decisions. Multiple structural and

propulsion tests were run; along with a wide berth of actual flight testing.

Table 1.1 summarizes the performance prediction of the SDSU entry in this year’s competition.

We look forward to compete with the best institutions from around the world.

Table 1.1: Summary of Expected Score

Mission 1

Mission 2

Mission 3

Gross Weight

(lbs)

Predicted SDSU

Results 7 laps 3 blocks 102.10 seconds

3.08 Predicted Max

Results 7 laps 5 blocks 102.10 seconds

Flight Score Total Score (not including report)

10.4 3.38

2.0 MANAGEMENT SUMMARY

The SDSU DBF team is comprised of 11 students with varying backgrounds and experience levels.

For the first two weeks of the school year, the team met as a group to familiarize ourselves with the

rules and with each other. After this introduction period, the team divided into sub-teams based on

areas of interest.

2.1 TEAM ORGANIZATION

The team is divided into five major sub-teams: aerodynamics, propulsion, design, manufacturing,

and report. The aerodynamics team is responsible for deciding on the optimal wing configuration

based on analysis of lift, drag, and stability characteristics of the aircraft. The propulsion team is

responsible for analyzing propeller, motor and battery interdependencies, and to find the ideal

combination of the three. The data and analysis from these two teams provide the framework of

possibilities for the other teams to work in. The design team is responsible for designing the major

features of the aircraft and creating the computer models of the design. The computer models that

the design team created were not only instrumental for the report team, but also necessary for the

manufacturing team to be able to use computer guided cutting systems in various manufacturing

techniques. The manufacturing team fabricated the aircraft and is responsible for deciding and

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utilizing the best techniques and materials. The manufacturing team put emphasis on continuously

creating and modifying prototypes throughout the year. This allows the team to perform a large

number of flight tests. It is helpful to have a physical aircraft as reference for the other sub-teams.

The report team is primarily responsible for creating the final report. However the report team had the

additional responsibility of acting as a facilitator for the other sub-teams. Because the report team

has a wider view of the project as a whole, it is their responsibility to help gather information that the

other sub-teams require to execute their roles. The project manager and assistant manager ensure

all team members are kept on schedule, make final decisions, and to handle the non-engineering

aspects of the team. DBF regulation requires one third of the team must be underclassmen, which

was carefully observed.

Figure 2.1 SDSU DBF 2013/14 Team layout. (Upperclassmen are in red boxes while

underclassmen are in gold)

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2.2 MILESTONE SUMMARY

Sept Oct Nov Dec Jan Feb Mar Apr

Design Projected Progress

Conceptual Projected Milestone

Actual Progress

Preliminary Actual Milestone

Current Date

PDR

CDR

Detailed Design

Manufacturing

Prototype 1

Prototype 2

Competition Model

Aircraft Built

Flight Testing

Prototype 1

First Flight

Prototype 2

Prototype 3

Report

Write Report

Edit Report

Report finished

Figure 2.2 Schedule and Milestone Summary

Planning centered on the goal to build three airplanes total: two prototypes and a competition

model. The first prototype’s purpose is to test the aerodynamic properties of our preliminary design.

The prototype matched the specifications of our intended design, but did not need to be constructed

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with the most efficient manufacturing process or meet the projected weight. The second prototype is

as close to the competition model as possible. This prototype’s purpose is to verify our final design,

provide an opportunity to practice the manufacturing techniques we will be using for the competition

aircraft, and to be able to test its limits without putting any unnecessary strain on our competition

model. The progress of these three planes, along with other important milestones, is recorded above.

3.0 CONCEPTUAL DESIGN

During the conceptual design phase, the team created design requirements based off the DBF

regulations and developed an overall plan for the aircraft.

3.1 MISSION REQUIREMENTS

The competition is divided into three flying missions and one taxiing mission. Scores for these

missions are combined with the report score and the Rated Aircraft Cost (RAC), which is the weight

of the empty aircraft, to create the total score for the competition.

3.1.1 General Requirements

Figure 3.1 Course Layout (AIAA 2013)

The course depicted in figure 3.1 consists of two 1000 feet stretches connected by two 180

degree turns, with the addition of a 360 degree turn in the second stretch. There is no minimum turn

radius for any of the three turns. The runway field length is the same for all three flying missions and

is limited to 40 feet. All flights must end with a successful landing and no major damage to receive a

score for that mission. The aircraft must not use a rotary wing or any kind of lighter-than-air design.

The aircraft must clear a height of 3.5 inches under each wing at a point that is half span from the

centerline. The battery pack must not exceed 1.5 pounds and the propulsion system is limited to a 15

amp draw by means of a fuse.

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3.1.2 Mission 1: Ferry Flight

Mission one requires teams to fly as many laps as possible within a four minute window with no

payload. Time starts when the throttle is advanced for the first time and only completed laps count

towards the score. The mission score is based on how many laps the aircraft completes compared to

the most laps completed by any team in the competition.

3.1.3 Mission 2: Maximum Load Mission

Mission two requires the aircraft to carry however many one pound, 6 in x 6 in x 6 in blocks the

team decides. A representation of one of the cubes can be seen in the figure 3.2.

Figure 3.2 Dimensions of Mission 2 cargo in inches (AIAA 2013)

Teams must be able to fly three laps carrying the cargo internally. There is no time limit for this

mission. The mission score is based on how many blocks our team carries relative to the maximum

number carried by any team.

3.1.4 Mission 3 Emergency Medical Mission

Mission three simulates a medical emergency and teams are scored on how fast they fly with a

specific payload. The load is meant to represent two “patient/attendant” systems. A representation of

the patient/attendant system can be seen in figure 3.3.

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Figure 3.3 Dimensions of Mission 3 Cargo (AIAA 2013)

Each block weighs 0.5 pounds for a total payload weight of two pounds. Because these are

meant to represent patients, there are restrictions as to how the cargo is allowed to be positioned.

From the official DBF rules:

The attendant shall be oriented vertically and the patient shall be horizontal and flat

(as shown in the figure).

The attendant must be immediately adjacent to the side of the patient

The patients must be separated by a minimum of 2” side to side or above/below

At least 2" space above the patient shall be an "air space" with no structure or

systems present.

The attendants must be separated by at least 2” from each other

Teams must fly three laps with the specified cargo and scoring will be based on the time it takes

to complete the mission relative to the fastest team at the competition.

3.1.5 Ground Taxi Mission: Rough Field Taxi

This year, a taxiing mission has been added to the competition. The aircraft must successfully

navigate a course immediately before performing either Mission one or two. A picture of the taxi

course can be seen in figure 3.4.

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Figure 3.4 Taxi Mission Corse Layout (AIAA 2013)

The aircraft must taxi over Palruf Roofing Panels, which have ridges 0.625 inches high at 3 inch

intervals. The course also features two obstacles which must be successfully navigated. These

obstacles are 3.5 inches high and extend half way across the course. It is permitted for the wings to

pass over the obstacles rather than navigating the aircraft completely around them. This mission has

the same cargo requirements as Mission three (two patient/attendant systems weighing two pounds).

The aircraft must travel 40 feet within five minutes without leaving the course, becoming airborne, or

sustaining any damage. Teams earn a score of 1.0 if they are successful and 0.2 if they are

unsuccessful.

3.1.6 Final Score

Teams will be given a final score based on mission scores, the report score, and the unloaded

weight of the aircraft. The report score is a number from 0-100 and is determined by the judges prior

to competition.

3.2 TRANSLATING MISSION REQUIREMENTS INTO DESIGN ELEMENTS

The following table serves to summarize the important requirements in section 3.1 and to detail a

thorough process as to how the aircraft’s features were selected and designed.

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Table 3.1 Mission Requirements Translated as Design Elements

Mission Requirement Design Element Potential Approach

Rated Aircraft

Cost

Lightweight aircraft Use lightweight materials in fabrication

Use the least amount of material without sacrificing

structural integrity

40x40 foot runway

High L/W ratio

High static thrust

Use airfoil with high CL max

Large wing area

Powerful propulsion system

Successful

landing to earn

score

Strong landing gear

Stable aircraft

Use tough material

Landing gear design distributes load

Effective control surfaces for static and dynamic stability

Complete laps as

fast as possible

Low drag during cruise

High maneuverability

Optimal propulsion

system

Streamlined fuselage design

Efficient propulsion system

Large bank angle in turns

Navigate taxi

mission course

Wings must be raised

3.5 inches off of the

ground

Wheels capable of

crossing corrugated

surfaces

Size landing gear to meet wing height requirements

Use large wheels/and or skis

3.3 DESIGN SELECTION

The SDSU DBF team considered all of the design requirements detailed previously during the

conceptual design phase. The following sections summarize the selection of the general aircraft

layout. Each section details what factors were considered and a figure of merit (FOM) table is

included to discuss the parameters affecting the design.

3.3.1 Overall Aircraft Configuration

Takeoff (30%): The runway length and current draw limitation means that our aircraft will need a

large wingspan.

Weight (30%): Weight is a crucial design parameter because of its effects on RAC, takeoff, and

flight performance. Designs requiring less structure are highly beneficial.

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Drag (20%): Configurations with a smaller profile and efficient aerodynamics will be able to fly

and takeoff faster, improving Mission one and three scores.

Stability (10%): Weather is an uncontrollable variable. If flight conditions are less than ideal on

competition day, an easy to control aircraft will not only be easier for the pilot, but will also be able

to fly more efficiently, decreasing the average lap time.

Manufacturing (10%): A simple manufacturing process not only allows the team to spend time

focusing on other aspects, but is less prone to failure and is easier to repair should the aircraft

sustain minor damage during flight or in transport to the competition.

Table 3.2 Figure of Merit Analysis of Overall Aircraft Configurations

Figure of Merit Weight Conventional Biplane Flying Wing Canard

Takeoff 0.30 4 5 3 2 Weight 0.30 5 3 2 5 Drag 0.20 5 4 2 5 Stability 0.10 4 5 4 3 Manufacturing 0.10 5 4 3 5 Total 1.00 4.6 4.1 2.6 3.9

3.3.2 Fuselage Selection

Weight (50%): The weight of the fuselage is the primary consideration and should be the first and

foremost thought during design of the aircraft.

Drag (40%): The fuselage design is also a significant component to the total drag of the aircraft,

following only behind the main wing according to CFD analysis.

Lift (10%): A fuselage design that offers some additional lift to be generated will allow for a

shorter takeoff and/or reduce the wingspan.

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Table 3.3 Figure of Merit Analysis of Fuselage Design

Figure of

Merit

Weight Square

Cross Section

Rounded

Cross section

Lifting Body

Weight 0.50 5 4.5 3 Drag 0.40 3 5 2 Lift 0.1 0 0 5

Total 1.00 3.7 4.25 2.8

3.3.3 Propeller Configuration and Location Selection

Weight (40%): Designs with multiple motors will weigh more due to the accompanying structure

required as well as the weight of the additional motor itself.

Power (30%): More power allows for a higher top speed as well as a shorter takeoff.

Efficiency (20%): A two motor system is allowed to run on two separate fuses, so the 15 amps

do not have to be shared between the motors. However, the efficiency of the motors is reduced

by a factor of two.

Takeoff (10%): The position of the motor in regard to the location of the wing will induce a

moment around the CG that will help the aircraft rotate and take-off.

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Table 3.4 Figure of Merit Analysis of Propeller Configuration and Location

Figure of Merit

Weight Tractor Pusher Tractor + Pusher Dual Tractor

Weight 0.40 5 5 3 3

Power 0.30 3 2 2.5 2.5

Efficiency 0.20 5 5 2 2

Take-off 0.10 5 4 3 3

Total 1.00 4.4 4.0 2.65 2.65

3.3.4 Empennage Design

Weight (40%): Weight is the same critical factor here as it was in the previous section. However,

weight in the tail has a much larger impact on the aircraft’s CG than weight in other subsystems.

A large weight increase in the empennage will result in a greater static margin, deterring the

ability of the aircraft to rotate about its CG.

Stability (30%): Obtaining high tail authority is critical to controlling a large aircraft in

unpredictable weather.

Drag (20%): A high drag design will impact the speed and lower the score for Missions one and

three.

Manufacturing (10%): Ease of manufacturing is important not only in the fabrication of the

aircraft but in case any damage is sustained that must be repaired quickly.

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Table 3.5 Figure of Merit Analysis of Empennage Design

Figure of Merit Weight Conventional Cruciform V-Tail

Weight 0.40 5 3 4 Stability 0.30 4 5 4 Drag 0.15 5 4 4 Manufacturing 0.10 5 4 3.5 Total 1.00 4.45 3.7 3.75

3.3.5 Landing Gear Wheel Design Selection

Taxiing Mission (30%): This year, the most important role of the landing gear is to perform the

taxiing mission. The landing gear must be able to navigate over the roofing’s grooves.

Takeoff (25%): The landing gear should not hinder the aircraft as it reaches takeoff speed.

Drag (20%): Drag on the wheels is smaller in magnitude than drag caused by other components,

but causes a larger moment on the aircraft. Thus, the pilot must trim the aircraft, which will add

more drag to the system.

Landing (15%): The landing gear must be strong enough to survive landing with no damage.

Weight (10%): Designs requiring less structure will reduce weight and are thus beneficial to

overall score.

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Table 3.6 Figure of Merit Analysis of Landing Gear Wheel Design

Figure of Merit Weight Wheel Skids Wheel Cowlings

Taxiing

Mission 0.30 2 4 5

Takeoff 0.25 5 3 5 Drag 0.20 3 5 4 Landing 0.15 5 3 5 Weight 0.10 4 5 2

Total 1.00 3.6 3.9 4.5

Wheel cowlings score higher on the taxi mission compared to simple wheels because they can be

utilized in an unconventional way. As the aircraft passes over a trough in the roofing, the cowlings

provide points of contact and prevent the aircraft from descending into the corrugation. The behavior

of the cowlings compared to a simple wheel can be seen in the figure below.

Figure 3.5 Simple Representation of Landing Gear over a Corrugated Surface.

3.4 SELECTED CONCEPTUAL DESIGN

Figure 3.6 is a conceptual design that incorporates the decisions made in section 3.

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Figure 3.6 Conceptual Designed Based on Preliminary Decisions

4.0 PRELIMINARY DESIGN

The team focused on sizing the selected conceptual design during the preliminary design phase.

The iterative process allowed the team to optimize the aircraft in order to satisfy the design

constraints and mission requirements.

4.1 DESIGN AND ANALYSIS METHODOLOGY

An iterative process was utilized to make design decisions and maximize score at the final

competition. The first task is to estimate the relevant dimensions of the aircraft and develop a mission

model. The mission model is used to determine the performance capabilities of a theoretical aircraft,

such as velocity, takeoff distance, drag, and load factor. The mission model provided guidance by

allowing continuous iterations on the design by adjusting preliminary assumptions. This process

continued until the inputs and outputs converged on an ideal configuration. Despite the best efforts of

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the team, the model still contains uncertainties. Actual flight testing is required to verify the results of

the analytic iteration and complete the design process.

4.2 MISSION MODEL

The mission model used was developed via Excel and MATLAB to determine performance

estimates and mission scores in accordance to aircraft geometry and system capabilities. The

MATLAB software allowed the team to initially estimate how many blocks to carry for Mission two by

generating a 3-dimensional plot based on sets of values for aircraft weight and Mission three times.

M3

Time

Empty

Weight

Overall

Score

Figure 4.1 MATLAB Scoring Analysis

Next, an Excel sheet was created to further analyze the performance of the aircraft in each of the

four flight phases: takeoff, climb, cruise, and turn.

Table 4.1 Mission Model Inputs/Outputs

Inputs Outputs

Wing Area

Aspect Ratio

Airfoil CL, CD

Static Thrust

Component Weights

Number of blocks for Mission two

Takeoff Distance

Takeoff Velocity

180° Turn Time

360° Turn Time

Mission one Laps

Mission three Time

Total Flight Score

Overall Weight

As the aircraft was tested and analyzed, changes could be made by inserting different input

values into the model to show the effect on performance.

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4.2.1 Scoring Analysis

Using the mission model developed in Excel, the team was able to see the variation of overall

score based on the number of blocks carried for Mission two. The team determined the minimum

number of blocks a team would carry is two; thus, the payload weight would be the same for Missions

two and three. From there, the team investigated what effects carrying additional blocks would have

on scoring. Wing area must be increased in order to takeoff within 40 feet. The fuselage must be

increased in length to accommodate the additional payload. The overall drag of the aircraft will be

increased due to skin friction as well as trim drag. Trim drag is increased because the elevator and/or

rudder must be deflected to counter act the moment caused from the main wing. As the wing area is

increased, more lift is produced. In the case of Mission one and three, where payloads would be less

than Mission two, the net force on the aircraft is greater. This increase in drag ultimately decreases

the aircraft maximum speed and increases lap time.

Table 4.2 Effects of Carrying More Than Two Blocks for Mission Two

Parameter Result Effect

Wing Area Increase 1.25 feet 0.1 lbs of weight gained

Fuselage length Increase 6 inches 0.05 lbs of weight gained

Overall Drag (Skin and Trim) Increase 6.25%

Mission 1 speed Decrease 3.5 mph Lap time increased by 3 seconds

Mission 3 speed Decrease 3.5 mph Lap time increased by 3 seconds

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As a result, the team determined the

following:

Figure 4.2 Effects of Mission 2 Payload on Mission Performance and Aircraft Weight

The overall effect on score is shown in figure 4.3.

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Figure 4.3 Normalized Scoring Curves for Various Maximum Mission two Payloads

The team concluded three blocks would be the optimal amount of payload to carry for Mission

two based off of the number of blocks other teams carry.

4.2.2 Mission Model Uncertainties

Despite efforts to make the mission model as accurate as possible, the model is limited to

estimated inputs, such as propulsion performance, and result in approximated outputs due to a

variety of factors such as:

Weather Data: The mission model was unable to account for actual wind and weather conditions

during the competition. Uncertainty in weather data has the potential to affect the performance of

the aircraft the most.

Other Team’s Scores: Although the team attempted to approximate the design of other aircrafts

in the competition, actual results are impossible to predict.

Turn Rate: Turning analysis was performed for an aircraft in steady level flight (no change in

altitude). However, the pilot will most likely use a wing over maneuver to reduce the turn radius

and quickly reverse direction. As a result, actual turn rates and thus flight times will be less than

anticipated in the analysis.

Maximum Mission

Two Cargo Flown

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The use of CFD, wind tunnel, and flight testing helped reduce the level of uncertainty and obtain

accurate results.

4.3 PROPULSION

The propulsion team was responsible for motor, battery, and propeller testing as well as selecting

the most efficient system to be used in the aircraft. Since the competition limitations for battery

weight and fuse amperage are 1.5 lbs and 15 amps respectively, the team had to ensure that the

system could still provide enough thrust to complete each mission and takeoff within the allowed

distance. The following table illustrates the requirements for the propulsion system.

Table 4.3 Propulsion System Requirements

Requirement Approach

Low Weight

Select lightweight motor without

sacrificing performance

Determine necessary amount of

batteries

High Thrust High static thrust motor and

propeller

Efficient Select efficient system that will not

draw too much current.

4.3.1 Battery

The goal of the battery analysis was to determine which type of batteries would perform best in

the competition as well as to determine how many pounds of batteries to carry. The team determined

the propulsion system would have to operate as close to the 15 amp limit as possible to optimize

performance.

Table 4.4 Battery Analysis

Battery Type Capacity

(mAh)

Cell

Weight

(oz)

# of Cells

Pack Volts

Under

Load

Energy

Density

(W-hrs/oz)

4 Minute

Watts

4 Minute

Average

Amps

Elite 1500A 1500 0.81 20 26.10 1.67 587.25 22.50

Elite 1700AA 1700 1.00 16 20.70 1.53 527.85 25.50

Elite 2000AA 2000 0.68 24 18.00 1.55 540 30.00

Elite 2200 2200 1.46 10 13.5 1.29 445.5 33.00

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Based on the analysis, the team determined that the Elite 1500A batteries would be the most

effective and reliable selection for the competition. Prior DBF experience has proved the batteries to

be reliable and capable of exceeding the amperage limit for brief periods of time without causing

problems.

4.3.2 Motor Analysis

The goal of the motor analysis was to determine which motor would allow the fastest lap time

without requiring too much power from the batteries. The initial motor selection process focused on

choosing an efficient motor that was lightweight that operates at the desired continuous watts.

Table 4.5 Motor Analysis

Motor Continuous

Watts Kv (rpm/volt) Weight (lbs)

Neu 1105-6D 200 3050 0.143

Neu 1110-2.5Y 500 1814 0.27

Neu 1110-2Y 500 2250 0.27

Neu 1112-2Y 600 1750 0.294

The team selected the 1110-2.5Y due to its optimal power draw as well as its high thrust based

on the results from Motocalc and testing described in section 8.

4.3.3 Propeller Analysis

Once the motor and battery pack system was selected, the team focused on choosing a propeller

that would provide the necessary thrust for each mission. The team tested a wide range of

propellers, as discussed in section 7 and 8 of the report. The 13”x10” propeller was selected due to

its performance capabilities for missions one and three while the 14”x8.5” was selected for mission

two due to the higher amount of static thrust.

Table 4.6 Propeller Analysis

Propeller Static Thrust

(lbs)

Efficiency (%)

12”x12” 2.75 50

13”x10” 3.00 80

14”x8.5” 3.75 75

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4.3.4 Selected Propulsion System

The figure below summarizes the selected propulsion system.

Table 4.7 Overall Propulsion System Component Selections

Item Name Weight (lb)

Motor Neu 1110-2.5Y 0.27

Batteries Elite 1500A 1.00

Propeller APC 13” x 10” (M1 and M3)

APC 14” x 8.5” (M2 backup) 0.05

4.4 AERODYNAMICS

After selecting a propulsion system, the team focused on sizing the main wing and empennage,

selecting airfoils, and predicting the lift and drag characteristics of the aircraft. The sub-team

identified the following requirements:

Table 4.8 Aerodynamic Requirements for the Main Wing


Low Drag Low drag airfoil at takeoff and cruise

Minimal parasitic drag from fuselage and empennage attachment

Low Weight Find an airfoil with high CLmax to acquire the least amount of wing

area

4.4.1 Airfoil Selection

Airfoils were analyzed using both XFLR-5 and JAVAFOIL to ensure accuracy with the results.

The team considered only pre-existing airfoils versus creating an original shape would be too time-

consuming and unnecessary due to the effectiveness of what is already available. Airfoil coordinates

were obtained from the University of Illinois database in addition to airfoiltools.com. Reynolds

numbers used for analysis were determined using expected speeds at takeoff and cruise as well as

initial estimations on chord length.

Airfoils were selected based on the following criteria:

CLmax (50%): A larger lift coefficient will allow the wing to produce greater lift, reducing necessary

wing area.

CDmin (30%): A smaller drag coefficient will allow the aircraft to fly faster in Mission one and three.

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Cm (10%): A smaller pitching moment coefficient will reduce the control surface area needed to

stabilize the aircraft. The level of experience of the pilot diminishes the weight of this

requirement.

Airfoil thickness (10%): A thinner airfoil will reduce drag and save weight compared to a thicker

airfoil with similar aerodynamic characteristics.

In the table below, a score of 1-5 is awarded in each category with five being the best.

Table 4.9 Airfoil Figure of Merit Table

Category Weight NACA 1210 sd7034 E210 GOE 396

CL max 0.50 2 4 5 3

CDmin 0.30 5 4 4 4

Cm 0.10 3 2 2 3

Thickness 0.10 5 4 3 3

Total 100 3.3 3.8 4.35 3.36

Figure 4.4 E210 Airfoil Polars

27


4.4.2 Aerodynamic Sizing

Focus turned to sizing the main wing once the airfoil was selected. The team determined the

runway length is the most critical parameter for wing sizing. Figure 4.5 was generated using the data

from the propulsion system and the equation below:

[

] (

(

)

(

)

[

]

)

Table 4.10 Takeoff Equation Input Parameters

Parameter Origin of Value

Using results from propulsion system testing

Using results from propulsion system testing

Using velocity from acceleration sled

Roskam (Vol 6)

Figure 4.5 Takeoff Distance vs. Wing Area

The figure above depicts the necessary wing area for an airfoil with a given maximum lift

coefficient to takeoff within the prescribed runway length. It can be seen that airfoils with larger

30

31

32

33

34

35

36

37

38

39

40

3 3.5 4 4.5 5

TO D

ista

nce

(ft

)

Area (ft^2)

1.4

1.3

1.2

1.1

Effective

CL

28


maximum lift coefficients reduce both takeoff distance and wing area leading to a lower R.A.C. and

better performance during takeoff.

4.4.3 Drag Build Up

Component drag build up was completed using methods in Roskam. The main wing contributed

the most to the overall drag, as expected.

Figure 4.6 Aircraft Component Drag Build-Up

4.5 STABILITY AND CONTROL

The aerodynamics group also managed the stability and control analysis and was responsible for

sizing the control surfaces on the aircraft.

4.5.1 Horizontal Stabilizer

The horizontal stabilizer was sized to provide the pilot with a high level of maneuverability. Given

the pilot’s level of expertise and experience, the team felt comfortable using a static margin of 5%.

In order to keep the aircraft as light as possible, the team decided to employ the use of a span-wise

elevator rather than two separate control surfaces on either side on the horizontal stabilizer as that

would require two additional servos. Based on estimates from Raymer, the elevator area would be

25% of the horizontal stabilizer area.

4.5.2 Vertical Stabilizer

Based on estimates from Raymer, the rudder area should be 25% of the vertical stabilizer area.

It will also run along the span of the vertical stabilizer.

4.5.3 Main Wing

The control surfaces on the main wing were sized to adequately control the aircraft in cruise,

turns, takeoff, and landing. Although the use of dihedral would improve the aircraft roll stability,

manufacturing proved too difficult to implement and was not used. Instead, a high mounted wing was

Component CD0 % of total

Main Wing 0.017292 60.54

Fuselage 0.00925 32.38

Horizontal

Stabilizer 0.0011189 4.16

Vertical

Stabilizer 0.000648 2.26

Landing Gear 0.00142 0.49

Prop Nose 0.000054 0.17

Total 0.0286 100

29


used for increased roll stability. A wing mounted atop the fuselage also requires less structure to

attach and thus weighs less than a mid or low mounted wing.

4.5.4 Stability Derivatives

Once the empennage, ailerons, elevator, and rudder were sized, AVL software was used to

determine the stability characteristics for the aircraft. AVL uses geometric inputs to calculate the

stability and control derivatives for the aircraft as shown in the figure below. The results are

consistent with that of a statically stable aircraft. Flight tests were used to ensure the pilot felt

comfortable controlling the aircraft.

The team used XFLR-5 to model the main wing and empennage to be exported into the AVL

software.

Figure 4.7 XLFR-5 Model of Main Wing and Empennage

Table 4.11 Stability Derivatives

Angle of Attack Pitch Rate Elevator Deflection

CLα 4.753 CLq 5.75 CLδα 0.006

CMα -0.88 CMq -4.40 CMδα -0.013

Side Slip Roll Rate Yaw Rate

Aileron Deflection CYβ -0.15 CYp -0.61 CYr 0.15

Clδα 0.0075 Clβ 0.006 Clp -0.45 Clr 0.062

Cnδδ -0.0003 Cnβ 0.051 Cnp 0.003 Cnr -0.05

30


4.6 AIRCRAFT MISSION PERFORMANCE

Using the mission model, the team input the preliminary design geometry, predicted propulsion

performance, and stability/maneuverability parameters to obtain overall mission performance

predictions. Section 5 includes a refined analysis of mission performance.

Table 4.12 Preliminary Flight Performance Parameters

Parameter Mission 1 Mission 2 Mission 3

CL Cruise 0.51 0.51 0.51

CL Takeoff 1.4 1.4 1.4

L/D Cruise 21.24 21.24 21.24

Cruise Speed (mph) 57.00 50.00 55.00

Takeoff Speed (mph) 16.21 24.80 22.24

Empty Weight (lbs) 3.08 3.08 3.08

Loaded Weight (lbs) 3.08 6.08 5.08

5.0 DETAIL DESIGN

The purpose of the detail design phase was to convert the proposed design into a physical

structure. Once a physical structure was manufactured, changes could be made to bridge the gaps

between the theoretical model and actual design. As always, reducing weight was the highest

priority.

31


5.1 DIMENSIONAL PARAMETERS

The figure below summarizes the dimensions of the aircraft and key subsystems.

Table 5.1 Aircraft Dimensional Parameters

Main Wing Horizontal Stabilizer Vertical Stabilizer

Airfoil E210 Airfoil NACA

0006 Airfoil

NACA

0006

Span (ft) 6.00 Span (ft) 2.00 Span (ft) 1.125

Root

Chord (in) 10.00

Root

Chord (in) 9.00

Root

Chord

(in)

9.00

Overall Layout Tip Chord

(in) 7.00

Tip Chord

(in) 8.00

Tip

Chord

(in)

8.25

Length

(in) 49.23 Area (ft

2) 3.54 Area (ft

2) 0.15 Area (ft

2) 0.797

Width (in) 72.00 Aspect

Ratio 7.05

Angle of

Incidence 0.00

Angle of

Incidence 0.00

Height of

main

wing (in)

17.27 Angle of

Incidence 0.00

Distance

from Main

(in)

42.50

Distance

from

Main (in)

42.5

Aileron Elevator Rudder Fuselage

Span (in) 16.00 Span (ft) 2.00 Span (ft) 1.125 Length

(in) 36.35

Chord (in) 2.50 Chord (in) 2.00 Chord (in) 2.00 Width (in) 8.00

Max

Deflection +/-15.00

Max

Deflection +/-15.00

Max

Deflection +/-15.00

Height

(in) 8.00

32


5.2 STRUCTURAL CHARACTERISTICS

The aircraft had the following structural design requirements:

Table 5.2 Aircraft Structural Requirements


Minimal Weight Use least amount of structure possible

Secure payload Have sufficiently large internal cavity and

fastening system

High Maneuverability Handle high wing loading

Hard landing Strong landing gear

5.2.1 Load Paths

The spar in the main wing is the primary structural member in carrying the aerodynamic loads.

Stringers also run perpendicular to the main wing through the fuselage and connect to the tail boom

via the aft bulkhead to create the main load paths on the aircraft.

Figure 5.1 Aircraft Load Paths

Lift

Payload Weight

Empty Aircraft CG

33


5.2.2 Operation Flight Envelope

Telemetry data from section 8.3 determined that the max g’s the airplane must withstand in

Mission three is 5.21. This translates to a maximum load limit of 8.66 g’s for the empty configuration

of Mission one. It was assumed that the maximum negative loading for Missions one and three is

four g’s and two g’s respectively. These values were used to generate the V-n Diagram seen in

figure 5.2

Figure 5.2 V-n Diagram for Missions 1 and 3

5.3 SYSTEM DESIGN, COMPONENT SELECTION AND INTEGRATION

The following subsections summarize the payload, aerodynamic, landing gear, loading,

propulsion, and control systems for the aircraft.

5.3.1 Fuselage

The design of the fuselage was driven by the size of the blocks for Mission two and the need to

keep the fuselage as light as possible. The main structural elements of the fuselage are two carbon

stringers to carry the load. The structure of the aircraft is characterized by 17 balsa ribs, separated

by three inches each. There are an additional five bulkheads in the fuselage. These bulkheads are

used as mounting surfaces for the motor mount block, wing, and tail boom.

-6

-4

-2

0

2

4

6

8

10

0 5 10 15 20 25 30 35 40 45 50 55 60 65

Load

Fac

tor

(g)

Velocity (ft/s)

Mission 3

Mission 1

Load Limit

Positive stall limit

Negative stall limit V Max

34


Figure 5.3 Fuselage Layout with Side View (Left) and Top View (Right)

5.3.2 Door

The bay door was designed to allow quick loading of the payload for Missions two and three. The

door is located on the bottom of the aircraft and is made from one pound density expanded

polystyrene foam. Packaging tape (3M brand) is used to secure the door to the fuselage during flight.

Epoxy was applied to the perimeter of the door. This layer of epoxy creates a much stronger bond

with the tape than the untreated foam does. Testing and analysis was performed to ensure the

tape/door could handle the loads as described in Section 7.2.3.

Figure 5.4 Bay Door

5.3.3 Mission Cargo Securing

The inner part of the door is lined with Velcro. Before each mission, strips of Velcro are wrapped

around the cargo to form loops. Two loops are used on each piece of cargo to restrict motion in all

six degrees. The blocks are securely placed onto the bay door and is inserted and taped into

position. A sample block wrapped in Velcro loops can be seen in figure 5.5.

Birch Wood Bulkheads

35


Figure 5.5 Velcro Straps Securing Mission 2 Payload

This method of securing the payload is also applied to the blocks in Mission three in a similar

manner.

5.3.4 Main Wing

The main wing is composed of balsa ribs and two spars: one carbon and one birch. The 0.5”

carbon spar runs span-wise throughout the wing for strength and rigidity. The birch ribs primary

purpose serves as a second mounting point for the ribs. The balsa ribs are 0.02” thick and spaced 3”

apart from one another. The one-piece wing is wrapped in Monokote to create the skin and provide

additional stiffness. The wing is attached using pylons mounted atop the fuselage.

Figure 5.6 Main Wing Structural Layout

36


5.3.5 Empennage

The horizontal and vertical stabilizers are made using two pound per cubic foot density foam and

are attached to each other using epoxy. A tail boom is inserted into the horizontal stabilizer at mid

span and connects into the fuselage.

Figure 5.7 Empennage Layout and Tail boom Attachment

5.3.7 Power and Control Systems Integration

The aircraft uses a total of four servos: one on each side of the wing for the ailerons, a third one

to control the elevator, and the fourth one to control the rudder.

Table 5.3 Overall Control Systems Components

Component Description Reason

HS-65MG Mighty Feather Servos Lightweight, sufficient torque and speed

Optima 9 Receiver Long range, lightweight, 8 channel

Kan 700 Receiver Battery Light, adequate power for one mission

Phoenix Ice2 HV 40 Speed Controller Has a factor of safety of 2 for current and

rated at 42 volts.

5.4 AIRCRAFT COMPONENT WEIGHT AND BALANCE

The following figure summarizes the aircraft component weights and locations. The Z-coordinate

is based on the aerodynamic top of the wing while the X and Y directions are longitudinal and lateral,

respectively.

37


Table 5.4 Aircraft Component Weight and CG Location

Item Quantity Weight (lbs) Total Weight (lbs)

X –

position

(in)

Y –

position

(in)

Z-

position

(in)

Wing Servos 2 0.02 0.04 -20.81 +/- 14.80 0.00

Elevator Servo 1 0.02 0.02 -48.31 0.00 -1.20

Rudder Servo 1 0.02 0.02 -48.31 0.00 -6.67

Speed Controller 1 0.15 0.15 -2.00 0.00 0.00

Receiver 1 0.05 0.05 -3.00 0.00 0.00

Receiver Battery 1 0.17 0.17 -2.00 0.00 0.00

Wiring 0 0 0 -20.81 0.00 0.00

Landing Gear 1 0.2 0.2 -15.68 0.00 -6.00

Batteries 1 1 1 -25.198 0.00 -5.00

Motor 1 0.27 0.27 -2.00 0.00 0.00

Prop 1 0.06 0.06 0.00 0.00 0.00

Wing 1 0.25 0.47 -20.27 0.00 0.00

Horiz. Stabilizer 1 0.064 0.063 -50.75 0.00 -6.20

Fin 1 0.064 0.063 -50.75 0.00 -2.67

Fuselage 1 0.81 0.5 -21.97 0.00 0.00

Blocks (M2) (total) 3 1 3 -20.54 0.00 0.00

Blocks (M3) (total) 4 0.5 2 -20.54 0.00 0.00

CG Location - - - -20.54 0.00 -2.00

Table 5.5 Aircraft Total Weights

38


Grouping Weight (lbs)

Non Structural 1.78

Structural 1.30

Mission 1 3.08

Mission 2 6.08

Mission 3 5.08

5.5 FLIGHT PERFORMANCE SUMMARY

The table below summarizes the aircraft performance for each mission.

Table 5.6 Aircraft Flight Performance Summary

Parameter Mission 1 Mission 2 Mission 3

CL max 1.5 1.5 1.5

CL takeoff 1.4 1.4 1.4

CL cruise 0.51 0.51 0.51

L/D T.O 56 56 56

L/D Cruise 34 34 34

Stall Speed (mph) 17.50 24.83 22.52

Takeoff Speed (mph) 21.29 24.30 23.29

AOA @ T.O (°) 9 9 9

T.O. Distance 30.00 38.00 35.00

Cruise Speed (mph) 56.60 50.00 53.30

Turn Time (°/s) 109.10 75.30 85.31

Bank Angle (°) 80 60 70

Wing Loading (lbs/ft2) 1.03 1.03 1.03

5.6 MISSION PERFORMANCE SUMMARY

The course has been separated into different sections as shown below. Each section has its own

optimal flight characteristics. For laps that do not involve take-off or landing, that portion of the track

is considered to be cruise mode.

39


Figure 5.8 Course Layout with Mission Phases

Below is the performance summary for one lap of each mission:

Table 5.7 Mission One Performance Summary

Phase Segments Distance (ft) Time (s) Avg. Velocity (ft/s)

Take-off 1 40 3 35.21

Cruise A 1 460 7.80 59.20

Cruise B 12 1000 12.05 83.00

180° Turn 14 50.91 2.17 67.10

360° Turn 7 101.82 4.34 67.10

Totals 16129.12 216.94 62.28

Total Laps 7

Table 5.8 Mission Two Performance Summary


Take-off 1 40 3 35.21

Cruise A 1 460 7.80 59.20

Cruise B 5 1000 13.25 75.47

180° Turn 6 55.49 2.39 67.10

360° Turn 3 168.02 4.78 67.10

Totals 6342.88 105.73 60.86

Total Blocks Carried 3

Cruise

Take-off

Climb

Landing Descent

360° Turn

180° Turn

40


Table 5.9 Mission 3 Performance Summary


Take-off 1 40 3 35.21

Cruise A 1 460 7.80 59.20

Cruise B 5 1000 12.05 83.00

180° Turn 6 56.45 2.26 86.19

360° Turn 3 110.26 4.34 86.19

Totals 6158.86 97.63 69.96

Lap Time 32.54

5.7 DRAWING PACKAGE

The following drawing package includes three-views of the aircraft, structural arrangement,

component layout, and payload arrangement for various missions. Drawings and parts were created

using SolidWorks.

41


42


43


44


45


6.0 MANUFACTURING PLAN AND PROCESSES

The manufacturing process began once the overall design of the aircraft was finalized.

6.1 TIME TABLE

The following figure depicts the predicted, actual schedule and timeline in the manufacturing process.


Manufacturing Projected Progress

Prototype 1 Projected Milestone

Actual Progress

Tooling Actual Milestone

Current Date

Prototype 2

Prototype 3

Figure 6.1 Manufacturing Time Table

6.2 MATERIALS SELECTION

The first step in the manufacturing process was to carefully consider the materials available. A

summary of the most relevant materials is below.

Foam: A very light material that the team has experience working with. It can be carved from a

large block using a hotwire (either by CNC hotwire cutter or by hand) or can be shaped via a

sanding block. In addition, there are two types of foam: a lighter 1 lbs/ft3 and a stronger 2 lbs/ft

3.

This makes foam a versatile option.

Balsa: Slightly heavier than the 1 lbs/ft3 foam and not as easy to work with. However, balsa does

lend itself a “build up” technique. This will allow for most of the wing or fuselage to be hollow, and

is the lightest realistic design. A balsa buildup is covered with MonoKote to provide a shell that

preserves the aerodynamic characteristics of the design while adding some stiffness to the

structure.

Carbon fiber: Significantly stronger than the previous two materials and has a wide array of

applications. However, it is too heavy to be a realistic primary material. It’s best used as

reinforcement in critical areas made of lighter and weaker materials.

Aluminum: Due to its weight, aluminum is a costly material to use. It should only be used in

specific applications where the other materials simply are not strong enough. Historically SDSU

has used it as a mounting surface and in landing gear.

46


A summary of the selection process for deciding the primary material of the fuselage and wing

can be seen in the following FOM table. The reasons for each of the variables of consideration are

consistent with their counterparts in section 3.

Table 6.1 Figure of Merit Summary for Materials Selection

Figure of Merit Weight Foam Balsa Carbon Fiber

Weight 0.70 4 5 3 Strength 0.20 3 2 5 Manufacturability 0.10 5 3 3 Total 1.00 3.9 4.2 3.4

6.3 MAIN WING AND FUSELAGE

The wing was made using a balsa buildup. The first step was to take the model seen in section 5

and convert it into a collection of 2D drawings. These drawings are the schematics for each of the

ribs in the wing. Each piece is cut from a large sheet of balsa by a high precision laser cutter and

then aligned on a long jig. The jig standardizes the spacing between the ribs, and holds them in

place so that the spars can be fastened to each rib simultaneously. A picture of the jig can be seen in

the figure below.

Figure 6.2 Ribs Aligned Using Placement Jig

The carbon spar was glued in place while in the jig and the lower half of the leading edge was

wrapped by 1/16” balsa sheeting. The wing was then turned over and re-attached to the jig so that

the shear members, birch spar, and upper half of the leading edge were wrapped again in 1/16” balsa

sheeting. The wing skeleton was then completed with shaped pieces of foam for the trailing and

leading edges. These pieces were cut using a computer guided hotwire to ensure that the foam

pieces matched the missing pieces of the airfoil exactly.

After the structure of the wing was complete, mounting platforms for the two aileron servos were

installed. The finished surface was then completed by wrapping the wings in MonoKote.

47


The fuselage is also a balsa build up and similar manufacturing techniques used for the main wing

were employed in fabricating the fuselage.

6.4 STABILIZERS AND CONTROL SURFACES

The empennage and control surfaces were manufactured using a wet lay-up process. Due to the

fact that the empennage and control surfaces are significantly thinner than the wing or fuselage, it

becomes difficult to construct them with a buildup in a precise and reliable manner. Instead, these

were fabricated using foam. A CNC hotwire machine cut the desired airfoil shape out of a large foam

block. Carbon fiber was applied to the control surfaces to create a crisp trailing edge. Epoxy is

prepared and applied onto a piece of carbon fiber cut to match the size of the trailing edge. The

carbon is placed on the foam and more epoxy is applied to ensure saturation. The carbon is then

dabbed with absorbent material to rid the piece of excess epoxy and then Mylar is placed on top.

This assembly is then placed inside a vacuum bag, sealed, and pressurized to remove lingering air

bubbles and any resin that was not absorbed. The part is then left to cure, under pressure, for a

period of 24 hours. A picture of a stabilizer after being removed from the vacuum bag can be seen in

figure 6.3. Once the epoxy is completely cured, the Mylar is removed and the control surface is

measured and cut out. Hinge tape is then used to reattach and position the ailerons, elevator, and

rudder to the lifting surfaces in which they were cut from.

Figure 6.3 Finished Horizontal Stabilizer

6.5 LANDING GEAR

Balsa and foam are too weak to make effective landing gear, making carbon fiber the better

choice. A protective plastic bag is placed over a mold to prevent the carbon from adhering to it. A

total of 12 layers of carbon were used in fabricating the landing gear using the ply layup

[0,90,+45,-45]3. These layers are designed to withstand the loads that act on the landing gear. The

completed landing gear can be seen below.

48


Figure 6.4 Finished Landing Gear with Mold

7.0 TESTING PLAN

After all preliminary design, detail design, and initial manufacturing had been carried out, each system

on the aircraft was tested to validate performance predictions. The experimental data provided allowed

further changes to the design, ultimately leading to a well-refined final product.


Propulsion Testing Projected Progress

Propeller Testing Projected Test

Actual Progress

Motor Testing Actual Test

Current Date

Battery Testing

Acceleration Sled

Structural Testing

Wing Spar Testing

Payload Testing

Landing Gear Testing

Flight Testing

Prototype 1

Prototype 2

Prototype 3

Figure 7.1 Testing Schedule. Floating stars indicate single day tests.

49


Table 7.1 Testing Checklist

Takeoff Thrust: Determine which propeller is able to generate the most static thrust to meet

the 40 feet requirement.

Dynamic thrust: Determine which propeller to use for each mission by measuring power

consumption, thrust at various speeds, and stalling speed.

Flight time: Verify that batteries can last for required time (four minutes plus landing time),

and to measure performance over time. Motor should not exceed 15 Amp limit.

Wing Structure: Verify that wings can withstand 6g turns with Mission three weights.

Payload: Verify that loading door is strong enough to hold 2 lbs cargo during 6g turns

Landing Gear: Verify that the landing gear will not break on impact.

Stability: Verify that the aircraft is able to maintain a constant angle of attack and ensure

that control surfaces are influential enough for easy aircraft control.

Performance: Measure performance characteristics of the aircraft including max g’s,

speed, and turn radius.

7.1 PROPULSION SYSTEM TESTING

The propulsion team conducted a series of tests to size, validate, and optimize the motor, battery,

and propeller.

7.1.1 Propeller Testing

The propulsion team considered a wide range of potential propellers. The tests were performed

in the wind tunnel lab at San Diego State University and measured thrust and power draw across a

velocity sweep, including static thrust and the propeller’s stall speed. This would in turn allow the

team to determine the maximum velocity for the aircraft.

Figure 7.2 Propeller Testing Setup

50


7.1.2 Battery Testing

Battery testing was performed to determine how many batteries would be ideal for the

competition. Since the contest rules have a 15 amp limit on the fuse, the team investigated whether

or not the full 1.5 pounds of batteries would be necessary. Testing was performed using 1.5, 1.25,

and 1 pounds of Elite 1500A batteries. This test involved connecting each battery pack to a

motor/propeller and increasing the motor thrust until it drew 15 amps. The team then measured how

long the batteries lasted at that amperage.

7.1.3 Motor Testing

The motor testing was performed by increasing each motor to approximately 6000rpm, which is

required for takeoff whilst measuring the amount of current each motor drew.

7.1.4 Acceleration Sled Testing

Acceleration sled testing was performed in order to confirm that the selected propulsion system

could reach the necessary ground speed for takeoff. Using the chosen propulsion system

configuration, the team ran multiple tests using the sled in figure 7.3 on a mock runway. The weight

of the sled was ballasted to the heaviest takeoff weight as that would be the critical case.

Figure 7.3 Acceleration Sled Setup

7.2 STRUCTURAL TESTING

The purpose of structural testing was to ensure that various structures on the aircraft could

successfully withstand expected loads. Testing consisted of carbon spars, the bay door, and landing

gear loads.

7.2.1 Carbon Spar Testing

Because the load bearing in the wing is principally performed by a single carbon spar, testing was

done to find the breaking strength of three different spars: ¼ inch, ½ inch, and ¾ inch diameters. The

spar was clamped to two tables. In the gap between the two tables, a box was tied to the carbon

spar. This box was loaded with lead weights until failure.

51


7.2.2 Bay Door Testing

The bay door was tested to ensure that the payload would not fall out during flight or landing.

Since tape is being used to secure the door to the fuselage, the team tested three different types of

tape to determine if and which kind should be used. A sheet of foam was cut into three pieces with

one being the size of the bay door. The configuration was then rejoined using the tape (a different

kind for each test) and loaded until failure. The results were converted in terms of PSI to allow for

scaling to a full sized door.

Figure 7.4 Bay Door Tape Testing

7.2.3 Landing Gear Testing

To verify that the landing gear would not fail on impact on landing, a sample pair of landing gear

was created. This landing gear was attached to the acceleration sled and weight was loaded on the

nose until the landing gear was pushed past the elastic region and sustained permanent deflection.

Figure 7.5 Landing Gear Testing

52


7.3 FLIGHT TESTING

Flight testing is an important step in the design process, which enables the designers to combine

all factors that went into the aircraft rather than analyze them in isolation. Not only did the flight

testing allow the analysis methods to be verified and updated, but it also allowed the pilot to become

familiar with the aircraft and aware of its performance capabilities. A preflight checklist was used to

ensure the safety of both the aircraft and the team.

Table 7.2 Flight Test Checklist

Preliminary Checklist

Unpack components, visually check for damage

Rotate landing gear wheels to ensure uninhibited motion.

Attach 4 servo arms. Make sure that the default position for control surfaces is parallel to wing.

Secure wing to fuselage, make sure bolts are tight without damaging wing.

Attach all servos to receiver. Left and right ailerons plug into slots 2 and 3 respectively. Elevator

plugs into 4 and rudder plugs into 5.

Turn on controller and receiver, test if servos are plugged into right spot.

Attach motor to receiver port 1. Plug in battery.

Test motor functionality.

Unplug battery.

Mission Preflight Checklist

Set battery to designated location for CG purposes (varies between missions)

Plug in battery.

Attach payload to bay door (Missions 2 and 3 only)

Attack bay door to bottom of aircraft.

Final Visual inspection for any damaged parts.

8.0 PERFORMANCE RESULTS

The performance results section is used to compare preliminary analysis data to actual

performance data. The actual recorded data is then used to revise the final design of the aircraft.

8.1 PROPULSION TESTING RESULTS

The following propulsion system testing results were generated using the 1110-2.5Y motor and 1

lb of 1600 Elite batteries unless otherwise specified.

53


8.1.1 Propeller Testing Results

The figure below summarizes the thrust versus velocity for the range of propellers tested. From

the data, the team determined that the 13”x10” propellers would be the best for Mission one and three

due to its high thrust in the critical 60-70 mph range while maintaining high efficiency. The 14”x8.5”

propeller would best serve Mission two, as seen in figure 8.1; it has the highest static thrust and does

not need peak performance in flight.

Figure 8.1 Thrust vs. Velocity Results for Various Propellers

Data gathered confirmed initial assumptions about the propellers made in previous sections.

While the gap in efficiency between the 12”X12” propeller and 13”X13” propeller was not as large as

initially anticipated, the 13”X10” propeller is still more efficient and better to use in the competition.

8.1.2 Battery Testing Results

Using the selected propulsion system, the team tested if using 1.5 pounds of batteries was

necessary for this competition.

Velocity (ft/s)

Th

rust

(lb

f)

54


Table 8.1 Battery Test Results

Pounds of Batteries Time Lasted (s)

1.5 377

1.25 342

1.00 315

Based on this data, the team determined that only one pound of batteries would be necessary for

the competition. This test was necessary because the weight of the battery influences the current

draw. The ramifications of selecting a smaller battery will be discussed in the next section.

8.1.3 Motor Testing Results

As seen in the graphs below, the 1110-2.5Y was drawing a maximum of 14 amps while the 1105-

6D drew 30 amps at similar rpms. This data makes it clear that the 1110-2.5Y is the best motor to use

for this competition. Adding more cells decreases the amperage draw by the propeller; so the 1105

could be made viable by adding more battery cells. However, results from section 8.1.2 show that

additional cells are not explicitly required and the additional weight of the batteries needed would be

greater than the weight saved by using the smaller 1105.

Figure 8.2 1105-6D Test Results

55


Figure 8.3 1105-2.5Y Test Results

8.1.4 Acceleration Sled Testing Results

The results from the acceleration sled tests are shown below. Using this data the team was able

to further validate propeller choice as well as aerodynamic sizing. The maximum takeoff velocity for a

prescribed runway length of 40 feet is 22.30 mph for the given wing area of the aircraft. The

propellers tested below all successfully meet that requirement.

Table 8.2 Acceleration Sled Test Results

Propeller Run Time (s) Velocity (ft/s) Velocity (mph)

12”x12”

1 5.7 27.234 18.569

2 5.1 27.419 18.695

3 5.7 27.234 18.569

Avg. 5.5 27.296 18.611

13”x10”

1 5.1 33.343 22.734

2 4.8 33.060 22.541

3 4.8 32.174 21.937

Avg. 4.9 32.859 22.404

14”x8.5”

1 4.1 35.121 23.946

2 4.2 35.778 24.394

3 4.1 35.484 24.194

Avg. 4.1 35.461 24.178

56


It is important to note that the sled for this test was ballasted for a Mission two load. The 13”X10”

propeller was not able to reach takeoff speed in one of the runs; however, it will be required to carry

one pound less for Mission three. This reduction in weight allows the 13”X10” propeller to take off

successfully in all but extreme cases.

8.2 STRUCTURAL TESTING RESULTS

The test results presented here summarize the structural capabilities of certain aircraft

subsystems. All components tested had a safety factor of at least 1.5 to be conservative. The

maximum g’s needed was determined multiplying the telemetry data in section 8.3.1 by the safety

factor.

The carbon spar testing is not listed in table 8.3 because testing was not finished. The team

began by testing the smallest carbon spar. This ¼ inch diameter spar was able to withstand an

excess of 50 lbs without permanently deforming. The load is nearly double of what was needed on

the aircraft. Because this spar was strong enough, no further testing was performed and our designs

were updated to reflect the use of a ¼ inch spar. This saves 0.36 lbs in aircraft weight relative to the

largest spar.

Table 8.3 Structural Testing Results

Subsystem Maximum

Load carried

Maximum

Pressure

Maximum

g’s carried

Minimum g’s

needed

Bay Door- Masking Tape 2.75 lbs. .086 PSI 4.42 9

Bay Door- Packaging

Tape 8.25 lbs .258 PSI 13.41 9

Bay Door- Duct Tape >12 lbs >.375 PSI >19.5 9

Landing Gear 27.5 4.5 4

Both packaging tape and duct tape will be strong enough to hold the door in place and operate

outside of the envelope described in section 5.2.2. Packaging tape was chosen due to ease of

removal after flight.

8.3 FLIGHT TESTING RESULTS

The results in this section describe the performance of the aircraft both observed in flight and

through the use of a 3DR Pixhawk telemetry unit combined with a ground station system. This allows

for a live telemetry feed to log aircraft performance data in flight. The aircraft was also equipped with

a GPS system to track its flight path. Several issues were discovered during the flight testing and

required changes to be made as shown in the figure below.

57


Table 8.4 Flight Test Issue Observations

Flight day Issue Resolve

1 Stability Issues Increase moment arm of tail

2 Stability Issues Increase horizontal and vertical stabilizer area

2 Flow Separation Change fuselage shape

2 Takeoff Distance Change airfoil

3 Aileron Stall Increase control surfaces

During flight testing of the prototype, the team installed black tufts on the aircraft to observe flow separation during flight.

Figure 8.4 Prototype 1 In-Flight with Black Tufts

Table 8.5 Predicted vs. Actual Flight Times

Sequence Actual Predicted

Mission 1 Laps Completed 7 7

Mission 3 Total Time (s) 102.43 96.84

Total Blocks Carried 3 3

Table 8.5 summarizes our prototypes performance compared to our prediction in section 5. The

aircraft performed similarly to expected, however, was slightly slower. This is due to the prototype

having a longer wingspan and less aerodynamic fuselage than the competition model.

58


8.3.1 Telemetry Data

The team decided to use a telemetry system to better understand the performance of the aircraft.

Table 8.6 Telemetry Data

Parameter Result

Max Ground Speed 62.81

Turn G’s 5.21

Impact G’s 2.07

Using this data, the team was then able to determine the expected loads of the aircraft. These

expected loads then drove the structural testing and design for the final aircraft. This data also

allowed the team to accurately compare predicted results versus actual flight test data.

Figure 8.5 Onboard GPS tracking system.

59


Figure 8.6 Telemetry Data. Star indicates maximum g’s experienced in figure 8.5

0

10

20

30

40

50

60

0

1

2

3

4

5

6

0 2 4 6 8 10 12 14 16 18

gro

un

d S

pee

d -

mp

h

Acc

eler

atio

n -

g's

Time - sec

Telemetry Feed - Acceleration vs Velocity

Accel

Velocity

60


9.0 REFERENCES

1AIAA. (2013 04-November). 2012/13 Rules and Vehicle Design. Retrieved 2013 04-Nov. from AIAA

DBF: [http://www.aiaadbf.org/2014_files/2014_rules_31Oct.htm]

2Abbott and A.E. Von Doenhoff. Theory of Wing Sections, New York: Dover 1959

3Anderson, John D. Introduction to Flight. New York: McGraw-Hill, 1989

4Anderson, John D. Fundamentals of Aerodynamics. New York: McGraw-Hill, 1991

5Bandu N. Pamadi, Performance, Stability, Dynamics and Control of Airplanes. AIAA Education

Series, 2004.

6Drela, Mark. XFOIL 6.96 user Guide. Boston MIT, 1986

7Etkin, Bernard. Dynamics of Flight. New York: John Wiley & Sons, 1996

8Kroo, Ilan Aircraft Design: Synthesis and Analysis, Version 0.9.

http://adg.stanford.edu/aa241/AircraftDesign.html

9Nicolai, Leland. Fundamentals of Aircraft Design. San Jose: Mets, 1984

10

Raymer, D. (2008). Aircraft Design: A Conceptual Approach. Reston, Virginia: American Institute of

Aeronautics and Astronautics.

11Roskam, Jan. Airplane Design: Part VI. Lawrence: DARcorporation, 2000

12Motocalc 8. s.l. :Capable Computing inc. http://www.motocalc.com/index.html

13Muller, Markus. eCalc. http://wwww.ecalc.ch/motocalc_e.htm?ecalc

14UGS Corp., ComponentOne, DriveWorks Ltd., Geometric Ltd., Microsoft Corporation, Spatial

Corp., Luxology, Inc., The University of Tennessee, Siemens industry Software Limited, and

Siemens Product Lifecycle Management Software inc. SolidWorks 2012. Vers. Student.

Waltham: Dassault Systemes, 1993. Computer Software.

15Star-CCM+. By CD-adapco. http://www.cd-adapco.com/products/star_ccm_plus/

16UGS Corp., ComponentOne, DriveWorks Ltd., Geometric Ltd., Microsoft Corporation, Spatial

Corp., Luxology, Inc., The University of Tennessee, Siemens industry Software Limited, and

Siemens Product Lifecycle Management Software inc. Abaqus 2010. Vers.

Waltham: Dassault Systemes, 1993. Computer Software.

http://www.motocalc.com/index.html

http://wwww.ecalc.ch/motocalc_e.htm?ecalc

2014dbf_sandiegostateuniversity

Documents

design selection

system design

design elements

preliminary design

mission performance

selected conceptual

aircraft mission performance

executive summary