2014DBF_University_of_Oklahoma

University of Oklahoma 0

DBF Report Draft University of Oklahoma Team [Pick the date] Tba Matt


Table of Contents

1.0 Executive Summary ................................................................................................................................ 3

2.0 Management Summary ........................................................................................................................... 4

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

2.2 Project Schedule ................................................................................................................................. 5

3.0 Conceptual Design .................................................................................................................................. 5

3.1 Mission Requirements ......................................................................................................................... 5

3.2 Mission Scoring Analysis .................................................................................................................... 8

3.3 Translation to Design Requirements ................................................................................................... 8

3.4 Configuration Selection ....................................................................................................................... 9

3.5 Conceptual Conclusion ..................................................................................................................... 13

4.0 Preliminary Design ................................................................................................................................ 13

4.1 Design and Analysis Methodology .................................................................................................... 14

4.2 Design and Sizing Trades ................................................................................................................. 14

4.3 Aerodynamic and Structural Design Components ............................................................................ 16

4.4 Propulsion System ............................................................................................................................ 20

4.5 Mission Model ................................................................................................................................... 22

4.6 Aerodynamic Analysis ....................................................................................................................... 23

4.7 Preliminary Mission Performance Estimates .................................................................................... 26

5.0 Detail Design ......................................................................................................................................... 26

5.1 Dimensional Parameters ................................................................................................................... 26

5.2 Structural Characteristics .................................................................................................................. 27

5.3 Aircraft Systems Design, Component Selection, and Integration ..................................................... 32

5.4 Payload Systems Design .................................................................................................................. 34

5.5 Weight and Balance .......................................................................................................................... 35

5.6 Flight and Mission Performance ........................................................................................................ 36

5.7 Drawing Package .............................................................................................................................. 38

6.0 Manufacturing Plan and Processes ...................................................................................................... 44

6.1 Manufacturing Process Selection ..................................................................................................... 44

6.2 Manufacturing of Components .......................................................................................................... 46

6.3 Manufacturing Schedule ................................................................................................................... 48

7.0 Test Plan ............................................................................................................................................... 49

7.1 Propulsion Testing ............................................................................................................................ 50

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

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

7.4 Rough Field Taxi Test ....................................................................................................................... 54


8.0 Performance Results ............................................................................................................................. 55

8.1 Propulsion System Results ............................................................................................................... 55

8.2 Structural Performance ..................................................................................................................... 56

8.3 Taxi Performance .............................................................................................................................. 57

8.4 Flight Performance ............................................................................................................................ 57

References .................................................................................................................................................. 58

Nomenclature

AOA = Angle of Attack.

AR = Aspect Ratio

b = Reference Wingspan of Main Wing

CAD = Computer Aided Drafting

= Mean Aerodynamic Chord of the

Main Wing

CD = Coefficient of Drag

CG = Center of Gravity

CL = Coefficient of Lift

CLMAX = Maximum Coefficient of Lift

CNC = Computer

DBF = Design/Build/Fly

EW = Empty Weight

FOM = Figure of Merit

FTF = Fastest Time Flown

M1 = Mission One

M2 = Mission Two

M3 = Mission Three

MDF = Medium Density Fiberboard

MGTOW = Max Gross Take Off Weight

NBF = Number of Blocks Flown

NBFmax = Maximum Number of Blocks Flown

NLF = Number of Laps Flown

NLFmax = Maximum Number of Laps Flown

PFD = Primary Flight Display

RAC = Rated Aircraft Cost

RPM = Revolutions per Minute

Re = Reynolds Number

S = Reference Wing Area

TF = Time Flown

TFS = Total Flight Score

TM = Taxi Mission

TMS = Total Mission Score

TS = Taxi Score

WRS = Written Report Score

= 80% of Maximum Velocity

⁄ = Velocity at Maximum Lift-to-Drag

Ratio

= Maximum Velocity at which Aircraft

Can Fly due to Aerodynamic and

Propulsive Limits

XPS = Extruded Polystyrene Foam


1.0 Executive Summary

This report outlines the design, manufacturing, and testing processes for a radio-controlled

aircraft built by the University of Oklahoma Crimson Skies Design/Build/Fly Team for the 2013-2014

AIAA/Cessna Aircraft Company/Raytheon Missile Systems Design/Build/Fly competition. The goal of the

Crimson Skies Team is to design and fabricate an aircraft that will successfully complete all of the

missions, with the goals of maximizing the total mission and written report scores, and of optimizing the

rated aircraft cost. The objective of this year’s competition is to design a backcountry rough field bush

plane to complete three flight missions and a taxi mission.

For Mission 1 (M1), the aircraft must complete the maximum number of laps possible along the

flight course within four minutes without a payload. For Mission 2 (M2), the aircraft must complete three

laps while internally carrying as many 6x6x6 inch 1 lb wooden blocks as possible. For Mission 3 (M3), the

aircraft must complete three laps as fast as possible carrying two patients and two attendants,

represented by 9x4x2 inch and 6x2x4 inch wooden blocks respectively. All M3 blocks are ballasted to 0.5

lb. All flight missions require the aircraft to take-off within 40 ft with the propulsion system limited to 15

Amps and a maximum battery weight of 1.5 lb. The Taxi Mission (TM) creates a challenge since it is new

to the competition, and because the total score is directly dependent on the taxi score. The aircraft must

taxi a 40x8 ft corrugated course while maneuvering around obstacles. A score of 1 is given for the

completion of the TM, and a score of 0.2 for failure to do so.

Using Figures of Merit (FOM), a number of aircraft concepts are analyzed in detail, and are

narrowed down to the concept described within this document. Based on the results of the FOM analysis,

a tail dragger aircraft with two motors, and a low rectangular wing provides the best solution for the

required missions while keeping the rated aircraft cost (RAC) at a minimum. The twin-motor design is

used for differential thrust to maneuver around the obstacles on the taxi mission, to reduce the overall

motor weight, and to provide the thrust required to take off within 40 ft with the maximum payload. Three-

blade propellers allow for the required ground clearance, while also providing a more efficient propulsion

system due to the aerodynamic characteristics of a three-blade propeller. The low-wing design is

advantageous when compared to other wing configurations, since it provides maximum structural support

for the payloads in combination with a low fuselage spine, while minimizing structural weight. Here, the

main load bearing structures include a C-channel spar and tubular spine made of carbon fiber to

maximize the strength-to-weight ratio for this aircraft. The body of the aircraft is a composite design,

primarily comprised of extruded polystyrene (XPS) foam with added carbon fiber and balsa wood where

more structural support is required. The aircraft is designed to carry four 6x6x6 inch cargo blocks in a

streamlined fuselage. The U-tail empennage configuration allows the vertical stabilizers and rudders to

operate outside the wake generated by the fuselage.

The team’s goal is to design an aircraft to be competitive in all missions. Speed is the key

requirement for two of the three missions. Therefore, the wing structure is designed to sustain a 5 g turn

at maximum loading. The Maximum Gross Take-Off Weight (MGTOW) is designed around the M2 weight,


because the payload of four blocks results in the heaviest overall aircraft weight. The wing’s quarter-chord

is placed at the fuselage’s center of gravity (CG).

The final aircraft design is estimated to have an empty-weight cruise velocity ( ) of 60.8 ft/s

for M1, a fully-loaded maximum lift-over-drag velocity ( ⁄ ) of 44 ft/s, and a of 60 ft/s for M3.

For M1, the final aircraft is estimated to compete six laps within the four minute time frame at an empty

weight (EW) of 4.2 lb. With the 4 lb payload for M2, the aircraft’s estimated MGTOW is 8.2 lb. The aircraft,

loaded to 6.2 lb for M3, is estimated to complete the timed mission in 142 seconds.

2.0 Management Summary

2.1 Team Organization

The Crimson Skies team is a hierarchy of 5 seniors and 1 junior as the core leadership of the

team with assistance from student volunteers ranging from freshmen to senior undergraduates. Each of

the four sub-teams is led by an undergraduate student chosen for their experience and expertise with the

Design/Build/Fly Competition. These sub-teams are Aerodynamics, Systems and Payloads, Structures,

and Propulsion Systems. The leadership assigns student volunteers to sub-teams to assist the leads with

their tasks. Each sub-team focuses on the tasks for their discipline, and brings their recommendations to

the team for the overall design of the aircraft. The team is led by a Project Manager, whose job consists

of ensuring schedules are kept, communication is maintained between the sub-teams, and the project is

completed. A Chief Engineer heads the overall project design. The team leaders take direction from the

Faculty Advisor. Figure 1 below shows the team’s organizational breakdown.

Figure 1: Crimson Skies Team Organization

The responsibilities of the team leaders are as follows:

Project Manager - Oversees the entire project, plans meetings and flight tests, and keeps the

project on schedule within budget.

Chief Engineer - Oversaw all aspects of the aircraft design, the technical head of the project.

Aerodynamics Lead - Calculates the dimensions of the wing, empennage, and control surfaces,

selects the airfoil(s), and analyzes the lift, drag, and stability characteristics for the designed

aircraft.

Structures Lead - Designs the structural framework of the aircraft and determines the materials,

performs weight and balance calculations, performs structural testing and analysis, and heads the


aircraft construction.

Systems/Payload Lead - Designs and manufactures the payload constraints, selects the servos,

and designs the landing gear for the aircraft.

Propulsions Lead - Designs the motor mounts, conducts all propulsion system testing, and

selects all propulsion system components: batteries, motors, speed controllers, and propellers.

2.2 Project Schedule

A project schedule lays out the entire year from the first meeting until the competition. The

schedule keeps track of deadlines and ensures that all leads are aware of the overall project status. This

includes the design, fabrication, and testing status for the project throughout the year. The schedule

shown in Figure 2 is a Gantt chart of the planned and actual progress of the project from start to finish.

Figure 2: Project Milestone Chart

3.0 Conceptual Design

3.1 Mission Requirements

The 2013-2014 AIAA Design/Build/Fly Competition calls for the design of a backcountry rough

field bush plane to complete one ground mission and three flight missions: a rough field ground taxi

mission, a speed ferry mission, a maximum load mission, and an emergency medical mission. The

payload for the maximum load mission, M2, is as many 6x6x6 inch 1 lb wooden blocks as the aircraft can

carry. The payload for the emergency medical mission, M3, consists of two patients on gurneys and an

attendant positioned beside each patient. This payload is simulated by 9x4x2 inch 0.5 lb wooden blocks

oriented flat and lengthwise and 6x2x4 inch 0.5 lb wooden blocks standing upright beside them. All

payloads must be carried internally and must be properly secured [1].

Aircaft Design

Conceptual Design

Preliminary Design

Detail Design

Optimization

Competition

Final Design Freeze

Aircraft Fabrication

Prototype 1

Prototype 2

Competition Aircraft

Testing Processes

Structure

Propulsions

Prototype 1 in Flight

Prototype 2 in Flight


4/11

2/17

Mar. Apr.Sept. Oct. Nov. Dec. Jan. Feb.

Estimated Actual


3.1.1 Ground Mission One: Rough Field Taxi

The aircraft is required to taxi across a 40x8 ft course consisting of corrugated fiberglass roofing

panels oriented normal to the direction of the course, as seen in Figure 3. Two obstacles consisting of 4 ft

standard 2x4 obstacles span from the edge of the course to the centerline at 1/3 and 2/3 of the length of

the course. The aircraft must taxi through the entire course and clear the obstacles without leaving the

course or becoming airborne. Successful completion of the taxi mission results in a score of 1 and failure

results in a score of 0.2 [1].

Figure 3: Taxi Mission Course Layout.

Figure 4: Competition Flight Course Layout.

3.1.2 Flight Mission One – Ferry Flight

The first flight mission is an EW ferry flight in which the aircraft flies as many laps as possible in a

four minute time limit following the flight course shown in Figure 4. Timing starts when the pilot advances

the aircraft’s throttle. The score for M1 depends on the number of laps flown by the aircraft, NLF, and the

maximum number of laps flown by any team at competition, NLFMAX, as seen in Equation 1 [1].

(1)

3.1.3 Flight Mission Two – Maximum Load Flight

The second flight mission is a three lap, maximum cargo flight in which the aircraft is internally

loaded with as many 6x6x6 inch wooden blocks as possible. The aircraft must take off within the

prescribed field length, complete three laps, and land safely to successfully complete the mission. The


scoring for M2 is directly dependent on the number of blocks flown in the aircraft, NBF, and the maximum

number of blocks flown by any team at competition, NBFMAX, as shown by Equation 2 [1].

(2)

3.1.4 Flight Mission Three – Emergency Medical Flight

The third flight mission is a timed three-lap flight, carrying two attendants and two patients in the

fastest time possible. Wooden blocks simulate the two patients and two attendants at 9x4x2 inch and

6x2x4 inch. Requirements for the layout and orientation of the attendant and patient are as follows:

The attendant shall be oriented vertically adjacent to the side of the patient

The patients shall be oriented horizontally and flat

There must be at least two inches of space above and between each patient

The scoring of the third flight mission, M3, depends on the time flown for three completed laps, TF, and

the fastest time flown by any team at competition, FTF, as shown by Equation 3 [1].

(3)

3.1.5 General Aircraft Requirements

In addition to the mission requirements, there are several constraints to which the aircraft must

adhere to:

The battery pack weight limit is 1.5 lb.

The ground clearance of the aircraft is measured by passing a standard 2x4 on edge under each

wing no further than the half span from the centerline during inspection with payload from M3.

Missions must be flown in order, and mission assembly time cannot exceed five minutes.

All payloads must be secured sufficiently to assure safe flight without possible variation of aircraft

CG during flight [1].

3.1.6 Overall Team Score

The overall score for each team depends on the written report score (WRS), the RAC, and the

taxi score (TS). The RAC this year is directly equal to the empty weight (EW) of the aircraft. The total

mission score (TMS) is based on the total flight score (TFS) and the TS. The TFS is the sum of the

mission scores: M1 score, M2 score, and M3 score. Equations 4-6 summarize the overall scoring

contributions [1].

(4)

(5)

(6)


3.2 Mission Scoring Analysis

The derivative of each variable in Equation 6 expands and analyzes which is most important. This

is accomplished by normalizing the equation and varying each variable individually from 5% to 100% with

the exception of EW. EW is assumed to be a minimum of 1 lb and a maximum of 4.8 lbs in order to

encompass all possible weights of the aircraft. Figure 5 shows the results of the analysis.

Figure 5: Mission Scoring Analysis.

The slope of each line in the figure corresponds to the sensitivity of the total score to a change in

each variable. It is evident that the empty weight of the aircraft is the most important factor.

3.3 Translation to Design Requirements

After the conclusion of the mission scoring analysis, this analysis and the competition constraints

are translated into aircraft design requirements. These were a starting point for the team to produce a

successful aircraft for competition.

● The aircraft’s design weight is to be minimized without hindering the ability to carry the

designated payloads for M2 and M3. This requirement derives from the fact that the RAC is equal

to the highest empty weight of the aircraft after the completion of each mission and that the empty

weight is the most important factor for scoring.

● The aircraft’s propulsion system must produce the required thrust while being constrained by the

15 Amp fuse and the 1.5 lb. battery weight limit.

● The aircraft’s landing gear design must be able to handle taxiing along the rough surface in the

0

2

4

6

8

10

12

14

0102030405060708090100To

tal F

ligh

t Sc

ore

% of Best Score

TS

WRS

M1

M2

M3

EW


taxi mission while also providing the ability to maneuver around standard 2x4 obstacles.

● The aircraft’s aerodynamics must be optimized for flying empty in M1 and at maximum loading in

M2 without increasing the drag significantly.

3.4 Configuration Selection

The transition from design requirements to aircraft configuration is accomplished through a series

of metrics that objectively evaluate potential configurations. A FOM analysis is conducted to hone in on

superior design elements for a conceptual aircraft configuration. Within the FOM, each parameter is given

a weight and each configuration an impact value to determine which configuration scores the highest. The

following scoring parameters are used for the FOM analysis. Some parameters are used only for a

specific FOM.

● Speed – Due to the scoring of M1 and M3, a competitive score necessitates that the aircraft be

able to complete missions in a short amount of time, making speed an important factor.

● Volume – Given the rather large geometry of the payload for M2 and M3, having adequate

payload space is imperative.

● Weight – As the RAC is dependent only upon the weight of the aircraft, weight is considered the

most important metric in evaluating any component.

● Thrust – Given the constraint on current draw from the batteries and the short take-off

requirement, thrust is heavily considered in propulsive considerations.

● Efficiency – Propulsion systems add more weight to the aircraft and ultimately raise the RAC;

therefore, an efficient motor at a low weight is an important parameter to consider.

● Gear Placement – As the landing gear needs a specific set of structures in a particular location

to be properly attached to the airplane, the ease of incorporating those structures is considered.

● Drag – The amount of drag created by a configuration heavily influences the performance of the

aircraft, so minimization of drag is considered for some analyses.

● Manufacturability – Ease of manufacturing is a significant factor in configuration analysis

because it is related to the rate of construction. Faster construction allows for multiple prototypes

and greater optimization through iteration.

● Stability – Different configurations have different inherent levels of stability, so the relative

stability of various configurations is considered.

● Ruggedness – Due to the rough taxi mission, it is imperative that the aircraft be able to

structurally handle bumps and vibrations while taxiing on the course.

3.4.1 Overall Aircraft Configuration

The team began the process of contemplating design configurations through several

brainstorming sessions, where ideas and concepts were introduced and discussed. Using the above

design requirements, each configuration was examined to explore its advantages and disadvantages. In


each category, different concept configurations were weighted according to their importance to the design

of the aircraft. Configurations were given values between 1 and 5, with 5 being the optimal choice. After

analysis, the BWB and flying wing configurations were deemed too volumetrically inefficient when

considering payload storage, while the biplane and twin-boom configurations were deemed too heavy to

be competitive. In the end, a conventional aircraft configuration was chosen.

Table 1: Plane Configuration Design Selection Matrix

Plane Configuration Weight Conventional BWB Flying Wing Bi-Plane Twin Boom

Speed 0.23 3 3 5 2 2

Volume/Drag 0.12 3 2 1 4 3

Weight 0.31 3 4 3 2 2

Manufacturability 0.12 4 2 3 2 4

Stability 0.12 4 2 1 3 4

Ruggedness 0.12 4 2 2 2 4

Total 1 3.42 2.89 2.92 2.4 2.88

3.4.2 Fuselage Shape

The basic shape of the fuselage was the driving factor behind the design of the body of the

aircraft. Weight was considered the most important factor of the design. Minimizing the weight of the

design was tied to minimizing the amount of excess structure in the design. Since the payloads for M2

were blocks and the payloads for M3 would fit in the same area as three in-line blocks from M2, the best

fuselage shape was rectangular. This created a structurally efficient fuselage while minimizing the cross-

sectional area of the fuselage. Since this design was also the easiest to manufacture, a rectangular

fuselage was selected for this plane.

Table 2: Fuselage Shape Design Selection Matrix

Fuselage Shape Weight Blended Cylinder Rectangle

Speed 0.17 5 3 1

Volume 0.13 3 2 4

Weight 0.44 3 3 4

Manufacturability 0.1 2 3 4

Ruggedness 0.07 2 4 3

Stability 0.09 1 1 1

Total 1 2.9 2.67 3.06

3.4.3 Propulsion Configuration

The propulsion configuration analysis placed major emphasis on physical weight and available

thrust. Since the motors will be purchased from a manufacturer, their weights are predetermined. Due to

the short field length for take-off and the rather heavy payload, it was imperative that the propulsion

system provide the required thrust without using more than 15 Amps of current draw while weighing as


little as possible. Previous Crimson Skies aircraft favored a single motor with a gearbox and propeller, but

it was postulated that two smaller outrunner motors could provide the same thrust, while weighting less

than a single larger inrunner motor. The current limit on the propulsion systems also came into

consideration. Having two motors meant that each motor could draw 15 Amps from its own battery,

allowing for more overall thrust to be produced, rather than one large motor being limited by a 15 Amp

fuse. Table 3 depicts the weighted analysis of different propulsion configurations. It was decided that the

design would use a twin tractor configuration.

Table 3: Propulsion Configuration Design Selection Matrix

Propulsion Weight Tractor Pusher Twin Tractor Twin Pusher

Speed 0.23 4 3 5 5

Drag 0.04 4 4 2 2

Weight 0.6 2 2 3 3

Manufacturability 0.13 4 2 4 2

Total 1 2.8 2.31 3.55 3.29

3.4.4 Wing Configuration

Wing configuration is a critical aspect in the overall aerodynamic characteristics of the aircraft. It

was established that the wing would need to withstand high loading with the least amount of structure. A

rectangular wing would be simple and easy to manufacture. A trapezoidal wing would have more

elliptical, and therefore more favorable, lift distribution, but would be more challenging to manufacture.

Other wing shapes were also compared to the rectangular wing, such as delta and elliptical, and scored

as shown in Table 4. After analysis, the team decided to utilize a rectangular wing configuration.

Table 4: Wing Shape Design Selection Matrix

Wing Shape Weight Rectangular Delta Trapezoidal Elliptical

Speed 0.24 2 3 4 4

Volume/Drag 0.14 2 4 4 5

Weight 0.48 4 2 2 3


Ruggedness 0.03 5 3 3 3

Total 1 3.38 2.66 2.79 3.3

3.4.5 Wing Placement

Vertical wing placement greatly affects the structure and stability of an aircraft. A mid-wing

design, while stable, would require a larger fuselage, since any wing carry through would most likely

interfere with the payload space. This design was deemed unfavorable. A high wing configuration might

be favorable because the aircraft would be stabilized by the pendulum effect of the fuselage and payload

hanging under the wing, but would require additional structure to support the payload under the wing. A

low wing configuration, on the other hand, would be much lighter because the payload weight could be


transmitted directly down into the wing and spine. After much debate, the team decided on a low wing

configuration.

Table 5: Wing Placement Design Selection Matrix

Wing Placement Weight High Wing Mid Wing Low Wing

Speed 0.17 3 2 3

Drag 0.06 4 1 3

Weight 0.23 4 1 5


Stability 0.14 3 2 2

Ruggedness 0.09 4 2 3

Gear 0.11 2 3 3

Payload 0.14 2 1 3

Total 1.00 3.17 1.63 3.37

3.4.6 Empennage Configuration

The empennage configurations considered for the aircraft were conventional tail, T-tail, V-tail,

forward canard, and U-tail. The team decided that the best tail would be the one to provide good stability

and maneuverability with the lowest weight. When looking at the design, it became apparent that the

large frontal area of the fuselage required the control surfaces of the tail to stay out of the wake of the

fuselage. This meant that most configurations would be prohibitively heavy and only those that efficiently

removed the control surfaces from the wake, the U- and T-tails, would be feasible. Of these, the U-tail

was deemed advantageous.

Table 6: Empennage Configuration Design Selection Matrix

Empennage Weight Conventional T-Tail V-Tail Canard U-tail

Speed 0.25 3 3 3 3 3

Volume/Drag 0.05 3 3 3 3 3

Weight 0.55 2 3 2 2 4

Manufacturability 0.15 4 3 2 2 3

Total 1 2.60 3.00 2.30 2.30 3.55

3.4.7 Landing Gear Configuration

The landing gear configuration was considered a very important design aspect, primarily due to

the TM. The landing gear must successfully withstand vibration, carry loads, and maneuver while taxiing

over the corrugated surface. A conventional layout would provide better propeller clearance, have a

greater angle of attack for take-off, weigh less, and most likely produce less drag. While the conventional

layout is inherently unstable on the ground, it could provide some flexibility with maneuvers on the rough

taxi course, and the twin propellers could be used for differential steering. The conventional layout was

ultimately selected as the best landing gear configuration.


Table 7: Landing Gear Design Selection Matrix

Landing Gear Weight Tricycle Conventional Reverse Tricycle

Dual Axle

Drag 0.13 4 3 4 5

Weight 0.44 4 4 2 2


Ruggedness 0.07 4 5 4 2

Total 1 2.86 3.00 1.98 1.97

3.5 Conceptual Conclusion

The conceptual design phase began with the analysis of the competition rules and design

specifications to understand what was required of the aircraft to compete successfully. A scoring analysis

was completed to determine the most important aspects of the aircraft design for a maximum TFS. From

that analysis, a set of design requirements was created. The design requirements were analyzed and

developed into a specific list of design configurations. A FOM analysis was conducted to objectively

select the best configuration. The resulting conceptual aircraft design included the following

characteristics:

Low-wing, rectangular fuselage

Twin tractor motor configuration

U-tail empennage

Tail dragger landing gear configuration

Figure 6 shows an early conceptual model. From the conceptual model, further engineering analysis

began in the form of preliminary design process.

Figure 6: Concept Plane

4.0 Preliminary Design

In the preliminary design phase, conclusions from conceptual design were translated into aircraft

dimensions and performance parameters. Through design iterations, the aircraft design was refined to

maximize mission performance. This section outlines the design efforts and selected configurations in

aerodynamics, structures, propulsion, and aircraft systems. The preliminary configuration was then

evaluated in a flight mission simulation for further design optimization.


4.1 Design and Analysis Methodology

The design phase of this project in its entirety can be illustrated by the design wheel shown in

Figure 7 [2]. All of the steps below begin at a low level of sophistication of the design concept, and then

iterate to move forward to a more complex concept design until the end-goal is reached.

Figure 7: Project Design Wheel

Before the competition rules had even been released, studies were already underway for

reducing weight. New materials, manufacturing processes, and methods for collecting data were all

researched. Then, the research efforts and experience from previous years were used to break down the

mission requirements, leading to the initial design concept. Once this concept was determined,

calculations of performance, aerodynamics, and propulsions were completed as part of the design

analysis. This transitioned into sizing, where a rough weight estimate of the aircraft was obtained. Trade

studies for various parameters were employed at this point in the cycle. The iterations from this point on

lasted over weeks of redesigning the aircraft to best fit the mission requirements for the most successful

concept possible.

4.2 Design and Sizing Trades

The team looked back at previous successful competition aircraft designs and highlighted crucial

design parameters that greatly influenced the overall score. Successful aircraft designs provided a

starting point for the present design process.

4.2.1 Variations in Maximum Gross Take Off Weight and Thrust

In order to select a competitive target weight, two historical trends were examined. The first was

the trend of payload fractions, the weight of payload divided by the aircraft’s MGTOW. The second trend

was thrust-to-MGTOW ratio. Figure 8 shows the variation of payload fraction vs. payload, as well as the

linear relationship between thrust and MGTOW.


Figure 8: Historical Aircraft Weight and Thrust Trade Studies

The average payload fraction for the past several years has been approximately 0.5. From the

maximum payload this year’s design would potentially carry, i.e., 4 lb, an estimated MGTOW was

determined to be around 8 lb. This would translate into an aircraft with an empty weight of 4 lb.

Examining the plot of thrust vs. MGTOW, it is clear that for a higher MGTOW, higher thrust will be

required. Thus, the thrust required to be competitive this year will be between 3.5 and 4 lb.

These conclusions formed the basis for preliminary design calculations for the aerodynamic,

structural, and propulsion teams.

4.2.2 Variation of Score with Number of Blocks Flown in M2

The mission requirements for M2 require that two or more 6x6x6 inch blocks be carried within the

aircraft for three laps. Given the dimensions of the payload for M3, it was decided that the payload

volume and dimensions of the patient and attendant would fit inside three 6x6x6 inch blocks, as seen in

Figure 9.

Figure 9: M2 and M3 Payload Dimension Consideration

Thus, it was determined that the minimum number of blocks an aircraft would carry was three, so

most planes would carry three. Analysis of the initial structure of the aircraft resulted in the conclusion that

adding an additional block of payload would result in a higher score in M2 without such an increase in

weight as to obviate that score increase. There was a small concern that the aircraft would be able to lift

four blocks but when flown empty would produce too much lift, resulting in unnecessary drag during the

timed flight missions, but it was determined that this could be mitigated though aerodynamic optimization.

0

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4.3 Aerodynamic and Structural Design Components

4.3.1 Airfoil Selection

A airfoil selection is one of the most crucial design choices. Initial analysis showed a need for a

high lift coefficient (CL) as well as a medium thickness airfoil for more favorable stall AOA characteristics.

These criteria are then used to select airfoils using the applet JavaFoil [3], a program that utilizes

numerical analysis to determine the aerodynamics of different two-dimensional (2-D) airfoils. Then, trade

studies of manufacturability and aerodynamic characteristics are used to further refine the selection of

airfoils. The two airfoils selected during the design phase, the NACA 67015 and 57015, are compared in

Table 8. Both have a large leading edge radius, a maximum thickness of 15% of the chord as well as a

maximum camber location at 35% of the chord. However, the NACA 67015 airfoil has a design CL of 1.0,

while the NACA 57015 airfoil only has a design CL of 0.8.

Table 8: 2-D Airfoil Comparison at Re = 300,000

Airfoil CL at AOA =0° Stall AOA CL/CD at AOA = 0°

NACA 67015 0.539 1.854 16° 39.068

NACA 57015 0.402 1.751 16° 31.86

As shown in Table 8, the NACA 67015 airfoil outperforms the NACA 57015 airfoil in three of the

four criteria:

High max CL – The NACA 67015 airfoil has a higher CL at every AOA due to its increased

camber compared to the NACA 57015 airfoil.

High stall AOA – Both airfoils have the same stall AOA.

High CL/CD – The NACA 67015 airfoil has a higher CL/CD at 0° AOA than the NACA 57015 airfoil

due to lower parasite drag which is shown in Figure 10.

High CL at AOA = 0° – The NACA 67015 airfoil shows a higher CL than the NACA 57015 airfoil

due to increased camber.

A higher CL at an AOA of 0° is desirable in order to shorten the wingspan and lower the overall

weight of the aircraft, to satisfy the short take-off requirement, and to reduce the power required from the

propulsion system for flight. The NACA 67015 airfoil also has a higher CL at each AOA than the NACA

57015 airfoil. This will allow the aircraft to accomplish the same three objectives described above.

Another reason for the superiority of the NACA 67015 airfoil for this aircraft is the higher CL/CD at an AOA

of 0°. This results in more lift and less drag at cruise speeds, which decreases strain on the propulsion

system and allows for more weight to be carried per unit span of the wing. The NACA 57015 and NACA

67015 airfoil results are graphically compared in Figure 10. These graphs were calculated using a

Reynolds Number of 300,000, with an approximated velocity of 70 ft per second, and a chord of nine


inches. The NACA 67015 airfoil was chosen due to its high CL at all AOA and low drag coefficient (CD) at

most CL values despite the increased Cm1/4.

Figure 10: 2-D Airfoil Comparison

The airfoil for the tail of the aircraft is also carefully considered. The idea of using a flat plate was

contrasted with a symmetric airfoil. The flat plate would decrease weight but decrease performance and

have inadequate volume for any supporting structure. Because a U-tail was chosen in the preliminary

design, internal volume was very important in the consideration of the tail airfoil. The NACA 0015 airfoil

was chosen for the vertical and horizontal tail surfaces since the internal volume of the horizontal tail was

sufficient for structurally supporting the vertical tails and able to store the tail servos internally.

4.3.2 Wing Design

A wing area (S) of 3.75 ft2 was initially derived from historical data of payload size vs. wing area.

Then, an initial wingspan was estimated to be 4 ft, resulting in an aspect ratio of 4.27. The resulting chord

length ( ) was therefore 11.25 inches. In order to reduce drag due to lift, the wingspan was eventually

lengthened from 4 ft to 5 ft while maintaining constant wing area. Through similar design choices such as

reduced weight, increased efficiency and manufacturability constraints, the wing was dimensioned to the

values shown in Table 9.

-1

0

1

2

-10 0 10 20 30

CL

AOA

CL vs. AOA

-0.15

-0.1

-0.05

0

-10 0 10 20 30

Cm

1/4

AOA

Cm1/4 vs. AOA

-1

0

1

2

0 0.05 0.1 0.15

CL

CD

CL vs. CD


Table 9: Wing Design Values

Wingspan 5.0 feet Tip Chord Length 9.0 inches

Root Chord Length 9.0 inches 9.0 inches

Wing Area 3.75 square feet Aspect Ratio 6.7

An initial study and breakdown of each mission was performed in order to find the worst case

inflight loads. Each mission was analyzed to find the estimated mission weights, maximum mission load

factor, and the associated moments and forces. The load factors were calculated from estimated

maximum bank angle the aircraft would perform for each mission. The wing was modeled as a distributed

load along the half span, and the associated bending moment from this loading was calculated using

beam theory. Table 10 shows a calculation of a half-span C-channel carbon fiber beam under a uniform

load. This provides a good estimation as to the forces seen on the wing in flight under a maximum load.

Table 10: C-Channel Carbon Fiber Beam Calculation Based on Beam Theory

Uniform Load on C-channel Carbon Fiber Beam

Load (lb) 20

Length of C-channel (in) 31.9

Moment of Inertia (in4) 0.00792

Modulus of Elasticity (msi) 33

Load Distance from Center (in) 15.95

Maximum Moment (lb·in) 240

Maximum Deflection (in) δ = F L3 / 48 E I 1.46

In order to enhance the maximum strength for minimum weight, the spar height was maximized to

increase the cross sectional moment of inertia. However, this height was limited by the medium thickness

airfoil section. A C-channel cross section took advantage of the limited height by maximizing the amount

of material away from the neutral axis. This increased the moment of inertia of the spar for enhanced

stability. To minimize weight, carbon fiber was used because of its excellent strength to weight ratio.

4.3.3 Fuselage Design

The fuselage was not expected to experience intense loads, so the main goal of their design was

weight savings. XPS foam built-up structures were chosen because they provided the lightest structure,

easiest manufacturability, and lowest cost of material compared to balsa or composite materials. The use

of foam would greatly reduce the number of parts pieced together with epoxy, thereby reducing the

overall weight of the aircraft. XPS foam can be easily sanded to improve aerodynamic performance and

reduce non-structural weight.


4.3.4 Spine Design

The main objective of the spine was to provide support to the tail. Different carbon fiber tubes with

various diameters were compared to determine the idea spine size. From this comparison and simple

calculations, a carbon fiber tube with a diameter of 0.75 inches was found to provide enough strength and

stiffness to withstand bending and torsional loads from the tail. A tubular spine provided an added benefit

in ease of manufacturing by laying up a tubular sock over a cylindrical mold.

4.3.5 Tail Design

The tail was sized using historical trends on tail volume coefficients. The target tail volume

coefficient was 0.66 for the horizontal tail and 0.044 for the vertical tail. These coefficients were obtained

by proportionally scaling historical trends outlined in Raymer [2]. Values obtained from this process are

shown in Table 11.

Table 11: Tail Volume Coefficient Definition

Length from wing quarter chord to vertical

tail quarter chord.

Length from wing quarter chord to

horizontal tail quarter chord.

Vertical tail wing area. inches

Horizontal tail wing area.

Vertical tail volume coefficient.

Horizontal tail volume coefficient.

The position of the tail was found with weight control in mind. The longest distance to provide the

smallest tail size without a noticeable difference in control authority was chosen using the above

equations. The final shape of the tail was decided upon using historical perspective as well as

qualitatively viewing the velocity flow field around it through the Autodesk Flow Design program. The

velocity flow field around the fuselage (Figure 11) also confirmed a conventional tail configuration would

have experienced insufficient flow. Therefore, a tail in which the vertical tail does not lie in the same

vertical longitudinal plane as the fuselage was more favorable.


Figure 11: Top View of Velocity Flow Field Plane Intersecting the Fuselage Section

4.3.6 Control Surface Sizing

The control surfaces of the aircraft were sized according to historical models, empirical analysis,

and flight-testing. Historical trends were referenced from Raymer [2], which suggests that the rudder

should be 25% of the chord, the elevator 25%, and the aileron 30%. After further analysis, these values

were then altered to the values shown in Table 12.

Table 12: Control Surface Values.

Control Surface Percent of Chord

Elevator 25%

Rudder 30%

Flaperon 20%

4.4 Propulsion System

The propulsion system must be efficient and optimized to provide Ample thrust. As stated in the

conceptual design process, it was decided that two motors should be used for weight savings and the

ability to produce more power with two 15 Amp limits rather than one. It was also postulated that two

motors could be used to better turn the aircraft in the TM through the use of differential thrust. The initial

goal was to examine small motors that would output the most power for their weight.

4.4.1 Motor Selection

For the initial comparisons of potential motors, PropCalc, a motor system simulation program,

was used to examine motor data [4]. Several motors of comparable size and output were chosen, while

the parameters voltage and propeller diameter were held constant at 14.6V with an eight inches.

Outrunner motors were considered in comparison to inrunner motors. Inrunners require a gearbox to

decrease revolutions per minute (RPM) and increase torque for aircraft propulsion. However, outrunners

do not require a gearbox. Therefore, outrunner motors offered the highest output power to weight ratio


when compared to inrunner motors. Table 13 below outlines the data gathered, where overall efficiency of

the motor was the crucial factor.

Table 13: PropCalc Outrunner Motor Comparison

Model KV Rating

(rpm/V) Current Drawn

(A) Thrust

Output (lb) Efficiency

Hyperion ZS 2213 1042 14.16 2.15 87%

Hyperion GS 2213 1175 15.38 2.1 85.8%

aXi 2212 1150 14.68 2.05 81%

As seen in Table 13 above, the Hyperion ZS 2213 provides Ample thrust for the mission

requirements while maintaining the highest efficiency.

4.4.2 Propeller Selection

The next step in the initial selection of the propulsion system configuration was to determine the

optimum propellers. Given a twin motor configuration, in an attempt to reduce the induced yaw from the

inertia of the spinning propellers, it was decided that the propellers would be counter-rotating. One

propeller would be a standard propeller while the other would be a pusher propeller. A major limitation in

the selection was the availability of a standard and a pusher propeller of the same dimensions. Potential

candidates were found from major suppliers to provide a baseline of propellers. It was also decided that

due to the constraint of ground clearance on the TM, three blade propellers would better handle the

mission. Table 14 below shows the various performance characteristics determined through PropCalc

with several different sizes of propellers [4].

Table 14: Thrust Output of Various Propeller Sizes Using the Hyperion ZS 2213 [4]

Diameter (in)

Pitch (in/turn)

No. of Blades

Voltage (V)

Current (A)

Thrust (lb)

7 4 3 16.93 15.14 2.61

8 6 3 16.41 29.23 4.66

9 7 3 15.94 41.8 6.17

10 7 3 15.59 51.19 7.01

As seen from the table above, PropCalc predicted rather large static thrust values, which are very

inaccurate for quantitative analysis. The data did provide a general range for the size of propeller that the

aircraft would need when compared to static thrust tests conducted using other propeller and motor

combinations. By comparing the two data sets, the PropCalc data could be calibrated, and it was

determined that the nine inch diameter propeller would best suit the aircraft for free flight, while the ten

inch diameter propeller would best suit the aircraft for fully loaded flight during M2 [4].


4.4.3 Battery Selection

In order to determine the most efficient battery for use in the aircraft, energy density was the

determining factor. Three different batteries were chosen whose characteristics most closely matched the

previously stated needs of the aircraft. Two batteries were chosen from Tenergy: a 1600 mAh and a 1200

mAh cell. This was done in order to determine if smaller energy storage per cell caused an increase in

energy density. A third cell was ultimately chosen from Elite which had a 1500 mAh capacity. The battery

testing is shown in §8.1.1.

4.5 Mission Model

A mission model was created to analyze and optimize the aircraft’s performance for all three

missions. The model translated the ideal aerodynamic and flight performance estimates into simulations

for each phase of the flight course. Figure 12 is a three dimensional model with color-coded flight phases.

Figure 12: Flight Mission Course

The flight course, color correspondence, and phases of the mission are shown in Table 15.

Table 15: Breakdown of Flight Course

Flight Phase Description Color in Figure 12 Assumptions

Take-off Acceleration from zero velocity to lift off speed Orange Constant AOA

Climb Climb to arbitrary altitude to begin flight course Pink Constant AOA

Half Turn 180 turn, 500 ft from starting line Blue L = nW, T = D

Cruise Steady Level Flight Green L = W

VMax

360 Turn 360 turn at full speed, high wing loading Red

Constant AOA

L = nW

VMax


For initial estimates, the velocity at L/Dmax was used to compute the time required to complete the

flight course. Thus, the number of laps the design could possibly complete within the time limit was

determined for M1 and M3. In M2, the aircraft was simulated with four wooden blocks as payload.

Due to the large number of assumptions and simplifications made in the model simulations, there

were inherent uncertainties associated with the mission model. The model assumed the flight path would

be flown with exactly 180 and 360 turns and perfectly parallel 500 ft straightaways with no deviation in

altitude or wind speed. Also, the thrust and power were assumed to be held constant ignoring battery

depletion effects.

4.6 Aerodynamic Analysis

4.6.1 Drag Buildup

The lift and drag were estimated at ⁄ for M2 and for M1 and M3. These values are

displayed in Table 16.

Table 16: Mission Aerodynamic Performance

Performance Parameter

M1 M2 M3

(ft/s) 60.8 44 60

CL 0.263 0.948 0.390

CD 0.0537 0.0982 0.0581

L/D 4.9 9.7 6.7

Using a procedure outlined in Raymer [2], a drag buildup method was done on each major

component of the aircraft in order to determine the various drag forces acting on the aircraft. Table 17

illustrates the component drag buildup of the aircraft for M1 at a cruise velocity of 60.8 ft/s. The fuselage

and wing account for most of the drag as shown in Figure 13. The coefficient of drag at the aircraft‘s

Reynolds number was calculated to be 0.05317. The lift due to drag factor was calculated from the drag

induced by the lift produced from the wing.

Table 17: Drag Components

Component CD0 Percent of Total

Wing 0.01856 34.90

Tail 0.00125 2.35

Tires (2) 0.00653 12.27

Main Gear 0.00574 10.80

Fuselage 0.01792 33.70

Drag Due to Lift 0.00318 5.98

Total 0.05317 100.00


Figure 13: Parasite Drag Buildup

4.6.2 Stability and Control

Using methods outlined in Nelson [5], stability/control derivatives and coefficients were calculated for

the aircraft in the stability axis coordinate system. The stability coefficients are shown in Table 18 and

Table 19. The negative value of shows that the aircraft has dihedral stability in the spiral mode. A

dihedral of 5° was added to the wing in the initial design to ensure roll and dihedral stability. The aircraft

also has acceptable pitch static stability as shown by having a value of -0.873. Figure 14 shows the

pole-zero maps for the eigenvalues of the dynamic stability matrices for each mission configuration. This

figure shows that for all missions the aircraft is marginally stable in all modes except the spiral mode. The

spiral mode stability is not unusual, however, and at this low magnitude is easily corrected by pilot input

during flight.

Table 18: Lateral-Directional Stability Coefficients

Lateral Stability Coefficients, (rad-1

)

Y, Side Force N, Yaw Moment L, Roll Moment

, sideslip -0.221 0.124 -0.0688

p, roll rate 0 -0.121 -0.806

r, yaw rate 0.257 -0.145 0.248

, aileron

deflection 0 -0.0526 0.182

, rudder

deflection 0.103 -0.0579 0.0027

Wing Tail Tires (2) Main Gear Fuselage Drag Due to Lift


Table 19: Longitudinal Stability Coefficients

Longitudinal Stability Coefficients (rad-1

)

X, Longitudinal

Force

Z, Normal

Force

M, Pitching

Moment

u, longitudinal

velocity -0.30 -1.93 0

, angle of attack 0.465 -4.93 -0.873

, angle of attack

rate 0 2.52 -10.09

q, pitch rate 0 -4.11 -16.44

, elevator

deflection 0 -0.231 -0.925

Figure 14: Pole-Zero Maps for Each Mission

-8

-4

0

4

8

-25 -20 -15 -10 -5 0 5

ω

η

Mission 1

-8

-4

0

4

8

-40 -35 -30 -25 -20 -15 -10 -5 0 5

ω

η

Mission 2 Short-PeriodLong-PeriodRollDutch RollSpiral

-8

-4

0

4

8

-30 -25 -20 -15 -10 -5 0 5

ω

η

Mission 3


4.7 Preliminary Mission Performance Estimates

Using the mission model outlined in §4.5, the aircraft’s performance was predicted for each

mission. The aircraft’s characteristics as determined throughout the preliminary design phase were used

to estimate the following: take-off distance, rate of climb, and lap times. For these initial estimates, the

cruise velocity was assumed to be the velocity at L/DMax. This resulted in a very conservative analysis, as

maximum velocity throughout the course would be much higher. It was through this analysis that the team

was able to identify a problem with the take-off distance on M2; this led to the implementation of flaperons

for higher lift during take-off. Table 20 outlines the estimated mission performance.

Table 20: Estimated Mission Performance

Mission 1 Mission 2 Mission 3

Laps Flown 5 Blocks Carried 4 Time Flown (s) 120

Max Laps Flown 6 Max Stores Carried 5 Min Time Flown (s) 100

M1 Score 1.6 M2 Score 3.2 M3 Score 5

5.0 Detail Design

Using the preliminary design and the mission performance estimates, the aircraft dimensions

were finalized. This section reviews the detailed design characteristics of the aircraft along with design

refinements in aerodynamics, structures, propulsion, and systems to arrive at a finalized configuration.

5.1 Dimensional Parameters

Table 21 and Table 22 list the dimensional, aerodynamic, and propulsive parameters of the

finalized aircraft configuration.

Table 21: Aircraft Dimensional Parameters

Wing Fuselage

Span 63.8 inches Length 33 inches

Chord 9 inches Width 6.5 inches

Area 3.99 ft2

Height 7.6 inches

AR 7.1 Payload Volume 0.5 ft3

Airfoil NACA 67015 Blocks Carried 4

Dihedral 5°

Incidence 0°

Horizontal Stabilizer Vertical Stabilizer

Span 15 inches Span 5 inches

Chord 6 inches Tip Chord 3 inches

Area 0.625 ft2

Root Chord 5 inches

Airfoil NACA 0015 Area 0.556 ft2

Airfoil NACA 0015


Table 22: Control Surface Dimensional Parameters

Rudder (x2) Elevator Flaperon (x2)

Span (inches) 1.25 15 25.5

% of Chord 30 25 20

Max Deflection (°) 20 20 20

5.1.1 Aerodynamic Refinements

In order to reduce the induced drag on the aircraft due to wingtip vortices, Hoerner Wingtips were

added. These devices are the most efficient way to reduce downwash effect while increasing the effective

wingspan with the lowest weight penalty. Figure 15 shows the Hoerner Wingtips.

Figure 15: Hoerner Wingtip

A flap deflection angle of 10° was necessary to meet the 40 ft take-off requirement in M2. Table

23 shows the refined design values for the wing.

Table 23: Refined Wing Design Values

Wingspan 5.32 ft Tip Chord Length 9.0 inches

Root Chord Length 9.0 inches Mean Aerodynamic Chord 9.0 inches

Wing Area 3.99 ft2

Aspect Ratio 7.1

5.2 Structural Characteristics

The Structures design goals are summarized in Table 24. It was attempted to minimize the

structural weight of the aircraft while providing adequate strength, rigidity, support for high loading in turns

and resistance to impact on landing. Analyses of the ideal strength to weight characteristics were done to

ensure each part was designed with structural integrity, aerodynamic, and stability considerations in mind.

Moments of inertia were heavily considered in part design due to the large fuselage and payload.


Table 24: Design Requirement for Structures

Overall Design Requirements Structures Sub-team Design

Requirements

Carry four 6x6x6 inch (±0.125) blocks

Design a fuselage around the blocks

Minimize weight, minimize parts

Include room for tolerances and restraints for internal payloads

Low drag for M1 and M2 Streamline the large, square fuselage

Durability Endure high loading during turns

Survive impact on landing

5.2.1 Load Paths

To handle compression forces, a two ply 3K carbon fiber C-channel main wing spar connects to

the 6.5x2.25 inch XPS foam-carbon fiber composite wing box. Aiding in torsional rigidity, a secondary

0.125 inch balsa wing spar spans the wing at 75% of the chord. 1/32 inch balsa sheeting spanning the

length of the wing on leading edge provides support for tension and torsion. A 0.75 inch diameter carbon

fiber tube spans the central spine of the aircraft bearing the static and dynamic loads of the M2 and M3.

The spine transmits its loads through the spar and the wing box.

5.2.2 Structural Analysis

Two failure modes are assumed as most critical for the structural design of the aircraft: turning

and landing. This analysis assumed that the wing spars, spine, leading edge, and wing box carry-through

would bear all the structural loading. The fully loaded aircraft (M2) undergoes an approximate load factor

of five in turning and landing.

5.2.3 Main Spar Structure

The aircraft’s wing is designed to withstand a 5 g load at MGTOW. In order to minimize weight

without sacrificing structural integrity, the C-channel main spar is designed with two ply 3K carbon fiber

with vertical shear web fibers running at -45° and +45° to the horizontal (Figure 16). The cross-sectional

shape consist of a C-channel with a shear web of height 1.32 inches, driven by the thickness of the airfoil

section at the quarter chord with a flange width of 0.33 inches. Uniaxial carbon fiber spans parallel to the

flanges for added rigidity in tension and compression.

Figure 16: Varying Layers of Carbon Fiber on the Wing Spar

The spine of the aircraft is aligned orthogonal to the spar web through the hole shown in Figure

16. In order to strengthen the connection between the spar and the spine of the aircraft, a XPS foam-


carbon composite wing box carry through is laid up in the central section of the main spar between the

flanges (Figure 17).

Figure 17: Cross-section of the Foam-Carbon Composite Box

5.2.4 Wing Rib Structure

The detailed shape of the wing ribs is shown in Figure 18. The D-shape leading edge ribs are

attached to the forward face of the main spar web. Aft ribs are mounted in to the C-channel of the spar

between the flanges. Because the main spar, secondary spar, and leading edge handle the critical loads,

the ribs are constructed from XPS foam because of the foam’s low density and ability to handle loads

uniformly in all directions. Balsa, in comparison, is weak at handling loads not along the grain of the

wood. At 0.5 inches wide, the XPS foam ribs offer a large surface for Monokote attachment while

reducing the number of members required along the wingspan. In addition, the wide surface area of the

XPS foam ribs prevents any undesired curvature or bows in the wing between ribs.

Figure 18: Foam Rib Structure Around the C-channel Main Spar

The ribs were spaced evenly across the wing span with a full rib structure at the wing root

providing additional loading strength as shown in Figure 19.


Figure 19: Wing Rib Structural Arrangement

5.2.5 U-Tail Structure

A rigid tail was required due to the large wake from the fuselage and buffeting from the propellers.

The main structures of the vertical and horizontal stabilizers were solid pieces of XPS foam. Torsional

stiffness of the horizontal stabilizer was increased by the addition of a leading balsa spar, a main balsa

spar, and an aft balsa spar. The control surfaces are made of balsa for increased rigidity (Figure 20).

Figure 20: U-tail Assembly

5.2.6 Fuselage Structure

The design of the fuselage structure was based on the idea of reducing assembled parts, thereby

reducing stress concentrations and glue weight. The size of the bottom section of the fuselage (Figure 21)

was designed to hold four internal payloads for M2 at 24.5 inches long, 6.125 inches wide, and 6.125

inches tall to account for the maximum tolerance of the blocks. The blocks were designed to rest on flat-

topped foam inserts that sit on top of the spine as it runs through the fuselage. The payload restraints

were incorporated around the spine into the bottom of the fuselage. To reduce weight while and still

maintain structural integrity, circular sections were cut from the side of the fuselage. The wing box carry-

through fits in the central cut-out of the bottom fuselage.


Figure 21: Bottom Fuselage Section Cut from a Block of XPS Foam

The nose cone and tail cone were designed to make the aircraft as aerodynamic as possible

while minimizing weight. Since the nose cone and tail cone handle minimal loads, each was designed as

a single layer of 3K carbon fiber. Foam inserts were added to the nose to maintain its shape and provide

a shelf for the receiver and receiver battery pack in the nose cone (Figure 22). A single layer of carbon

fiber (not shown) was added to the shelf containing the receiver battery to protect the foam from the heat

produced by the battery pack while in flight. Since carbon fiber limits the radio reception of the receiver,

the antenna was placed in the forward section of the bottom fuselage.

Figure 22: Carbon Fiber Nose Cone with Foam Shelves (left) and Tail Cone (right)

5.2.7 V-n Diagram

An important aspect of the structural design is the development of a V-n diagram that details the

operational limits of the aircraft. The V-n diagram for this aircraft is presented in Figure 23. The curve in

the figure shows the aerodynamic limits of the aircraft, indicating the conditions at which the aircraft will

stall. The horizontal line is known as the maximum positive load factor, representing the maximum flight

loads allowed by the aircraft’s wing structure. The positive load factor is determined from the destructive

wing loading test. The vertical line in Figure 23 represents the aircraft’s maximum speed with the

designed propulsion system. The V-n diagram does not depict the negative loading effects which can

occur during flight testing [6].


Figure 23: V-n Diagram

5.3 Aircraft Systems Design, Component Selection, and Integration

5.3.1 Fuselage Hatch

The large size of the mission payloads in relation to the aircraft necessitated a large opening in

the fuselage to allow for easy loading and arranging of the various payloads. To accommodate this, a

large opening with a lid was designed so that the entire cargo area of the plane could be accessed at the

same time. The lid was designed to fit flush with the fuselage, and was secured in place by eight sets of

rare-earth magnets running along the edges of the opening in the fuselage.

5.3.2 Battery Placement

The battery packs needed to be carefully placed both to balance the plane and to ensure that the

batteries are cooled during flight to eliminate the risk of overheating and damaging the XPS structure of

the plane. The battery packs are therefore placed in pods that are integrated into both sides of the

fuselage forward of the wings. A removable panel allowed the batteries to be accessed easily and slits in

the paneling allow for airflow over the batteries during flight.

5.3.3 Wheels

Through testing of various aircraft on the rough taxi course, the systems team determined that it

would be important for the aircraft to have large wheels for better performance on the corrugated surface.

Large wheels available from common retailers, however, were excessively heavy, so the team

determined custom wheels should be designed and fabricated custom wheels for the aircraft.

As shown in Figure 24, the wheels were designed with a 4.5 inch radius to mitigate the bouncing

motion of the aircraft travelling over the ridges. A 0.0625 inch thick circle of plywood was glued to either

side of a 0.875 inch thick piece of XPS, and the assembly was cut out using a hot wire. A brass tube was

pushed through the center of the assembly and glued into place to act as an axle. A strip of 1 inch wide

0

1

2

3

4

5

6

0 20 40 60 80

Load

ing

(g's

)

Velocity (ft/s)


by 0.25 inch thick foam insulation strip was glued in place around the wheel for traction. The resulting

wheels were considerably lighter than the retail-bought wheels.

Figure 24: Wheel Assembly

5.3.4 Propulsion Mounting

Since the motors for this aircraft were to be located above the surface of each wing, it was

necessary to design a mount that could be secured to the front wing spar that would accommodate the

particular geometry of the motors, their wiring, and that would also hold the motors in place during

operation. The two halves of the mount were molded out of 3K carbon fiber and were then mounted back-

to-back to the spar. The wing structure was then built up around this mount. Figure 25 shows the motor

mounted through the leading edge, attached to the spar.

By examining twin engine concepts on production aircraft, it was found that the propeller most

generally sits in front of the leading edge by approximately 15% of the chord. For a chord length of nine

inches on this aircraft, the propeller needed to be about 1.35 inches in front of the leading edge. It was

also found that the rotational axis of the propeller should be offset from the chord line, allowing the

majority of the washout travel over or under the wing. Most production aircraft have engines mounted

under the wing due to the noise considerations in the fuselage. It is, however, more beneficial for the

washout to travel over the upper wing surface, further accelerating the flow and in turn providing more lift

on that section of the wing.

The location along the span where the motors should be mounted was determined through

propeller clearance and through the desire to have a sufficient moment arm for the propeller to allow for

effective turning through differential thrust. Each motor is mounted at 12.24 inches from the centerline

giving a turning moment of 1.63 ft·lb at maximum thrust.


Figure 25: Motor Mount Wing Assembly

5.3.5 Electronics Selection

The control system for the aircraft uses a transmitter to communicate with a receiver, which

sends commands to the servos. The control systems are powered by a separate battery on board the

aircraft. The JRX9503 transmitter was selected along with the JR R921X receiver, since the team already

owned the transmitter and receiver set and they have performed well in the past. The servo selection for

this year’s aircraft started with a survey of the servos already owned by the team plus a few additional

servos. The four servos with the highest torque to weight ratio are shown in Table 25.

Table 25: Servo Analysis

Model 6V Torque (oz·in) Weight (oz) Torque to Weight Ratio (oz·in/oz)

JR Servo DS398 75.0 0.75 100.0

Futaba S3102 64.3 0.74 86.9

Hyperion DS13TMB 58.3 0.68 85.7

Futaba S3114 23.6 0.28 84.3

After analysis of the moments required on the aircraft’s control surfaces during flight, a factor of

safety was added to the required torques to account for wind effects. It was decided that the JR DS398

would be used for all control surfaces, since it provided more than enough torque, had the best torque to

weight ratio, and had a convenient, slim profile that fit into the aircraft well. The electronics selections for

the aircraft are summarized in Table 26 below.

Table 26: Electronics Selection Summary

Transmitter Receiver Servo

JR X9503 JR R921X JR DS398

5.4 Payload Systems Design

It was important that the payloads did not shift during flight to ensure that the plane remained

balanced and undamaged. The lightweight nature of the fuselage precluded restraints being anchored

there, so straps were threaded through the floor of the fuselage and around the spine of the aircraft to


ensure that the loads associated with holding the payloads would be transferred to a structurally sound

member. Velcro was chosen as the material for the straps because of its ease of use and its slim profile.

5.5 Weight and Balance

The component breakdown of the aircraft’s empty weight by system and subsystem is shown in

Table 27. Approximately one half of the aircraft’s empty weight is contributed by the propulsion systems,

and the other half is contributed by the structural airframe and the control systems.

Table 27: Weight and Moment Balance of the Aircraft

Component Qty. Weight

(lb)

Moment Balance (ft·lb) Total (lb) X Y

Structures

Empennage 1 0.092 -0.304 0

1.687

Fuselage Bottom 1 0.165

-0.015 0

Fuselage Lid 1 0.068 -0.011 0

Landing Gear 1 0.325 0.001 0

Nose Cone 1 0.070 0.070 0

Spine 1 0.086 -0.100 0

Tail Cone 1 0.081 -0.115 0

Tail Skid 1 0.019 -0.058 0

Wing 1 0.781 -0.185 0

Controls

Receiver 1 0.033 0.031 0

0.562 Receiver Battery 1 0.335

0.314 0

Servos 4 0.194 -0.359 0.001

Propulsions

Propeller 2 0.088 0.014 0

1.932

Motor Assembly 2 0.263

0.017 0

Battery Pack 2 1.460 0.044 0

Speed Controller 2 0.120

-0.010 0

The origin of reference for the moment balance calculation of the aircraft’s empty weight is the

aircraft body reference illustrated in Figure 26.


Figure 26: Reference Axis for the Moment and Balance Calculations

Table 28 presents the overall weight and balance numbers for the aircraft for each mission’s

configuration. This table confirms the aircraft is balanced through all missions. The sum of the moments

about the X-axis equals zero slightly forward of the center of lift.

Table 28: Weight and Moment Balance for Each Mission

Mission M1 M2 M3

Weight (lb) 4.18 8.18 6.18

X (ft·lb) 0.00 0.00 0.00

Y (ft·lb) 0.00 0.00 0.00

5.6 Flight and Mission Performance

After further refining the aircraft, the design was put through the mission model as seen in §4.5.

Each phase of the flight course was simulated and an estimate of a flight score was produced. For

scoring purposes, it was assumed that the aircraft successfully completed the TM and received a score of

1. Table 29 shows relevant flight characteristics used in the model simulation of the aircraft. These values

reflect the detailed aircraft design that has been slightly refined since the preliminary design phase.

Table 29: Aircraft Flight Characteristics

Mission 1 Mission 2 Mission 3

Mission Weight (lb) 4.18 Mission Weight (lb) 8.18 Mission Weight (lb) 6.18

Take-off Distance (ft) 13 Take-off Distance (lb) 39 Take-off Distance (ft) 32

Cruising Speed (ft/s) 70 Cargo Blocks Carried 4 Cruising Speed (ft/s) 65

Turning Speed (ft/s) 69

Turning Speed (ft/s) 55

Turning Load Factor 6

Turning Load Factor 3

Lap 1 Time (s) 41

Lap 1 Time (s) 53

Subsequent Lap Times (s) 36

Subsequent Lap Times (s) 45

Number of Laps 6

Mission Time (s) 142


5.6.1 M1 Performance

M1 was modeled with the aircraft using its maximum cruise speed during the straight portions and

making its turns at a load factor of six. To account for uncertainty in the lap times due to wind and flight

path deviation an extra ten seconds were added to each lap time. Estimated results from M1 are shown in

Table 30.

Table 30: M1 Performance

Mission 1

Lap 1

Mission Phase Time (s) Distance (ft) Velocity (ft/s)

Take-off 0.90 13.5 0-27.8

Initial Climb 8.40 504 42.1-69.4

180o Turns (2) 3.88 218 57.00

Cruise 20.00 2000 48.42

360o

Turn 3.88 177 57.00

Subsequent Laps

Cruise 26.67 2000 60.80

180o Turns (2) 3.88 177 57.00

360o

Turn 3.88 177 57.00

Totals - 240.00 14667 -

Number of Laps 6 M1 Score 1.71

The highest number of laps for M1 was estimated to be seven based on historical DBF

competition data. The estimated score for M1 was 1.71.


M2 was modeled in the same fashion as in the preliminary design. It was assumed that the

highest number of blocks carried by the competition would be six, a generous estimate primarily driven by

propulsive constraints. The take-off distance of the aircraft was estimated to be just under 40 ft. This

determined that the aircraft would successfully fly the mission and yield an M2 score of 2.67.


M3 assumed that the aircraft would turn with a load factor of three. The aircraft was also assumed

to fly at cruise speed during the straight portions. The score assumed that the best aircraft in the

competition would be able to complete the mission in two minutes.

Table 31 below shows the predictions for M3. Estimated time to complete three laps is 140

seconds, which yields a flight score of 4.9 for M3.


Table 31: M3 Performance

Mission 3

Lap 1

Mission Phase Time (s) Distance (ft) Velocity (ft/s)

Take-off 1.8 32 0-33

Initial Climb 14.2 502 38.3-56.25

180o Turns (2) 3.88 238 56.25

Cruise 20 2000 60.0

360o Turn 3.8 238 56.25

Laps 2 and 3

Cruise 26.64 2000 60.0

180o Turns (2) 3.88 238 56.25

360o Turn 3.88 238 56.25

Totals - 146.88 7986 -

Number of Laps 3 M3 Score 4.90

5.6.4 Total Mission Score

Using the projected data from this section, as well as assuming that the aircraft could complete

the 40 ft take-off requirement in all missions, and successfully complete the TM, the TFS was calculated.

This aircraft was predicted to complete the competition with a TFS of 9.82.

5.7 Drawing Package

This section of the report contains the drawing package for the aircraft. The drawing package

includes an annotated three view drawing of the aircraft, its structural arrangement, and the location of

payloads for different mission configurations.




3.76

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University of Oklahoma

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4.50 8.50

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6.0 Manufacturing Plan and Processes

The aircraft’s design was an ever evolving process that revolved around weight, strength, and the

manufacturing of the components. Beginning with prototype design, the feasibility of manufacturing

individual components was analyzed. If it was determined that strength and/or weight could not be

sacrificed for ease of manufacturability, a detailed process was laid out for how to develop a component.

If a manufacturing process was not deemed feasible, because of a lack of tool availability and team

experience, then consideration of alternate manufacturing processes became a key role in the overall

design process.

6.1 Manufacturing Process Selection

Materials and manufacturing processes selected for construction of the aircraft play a vital role in

the overall weight and flight scores. A major focus was to reduce the weight of the aircraft by reducing the

number of assembled parts. The primary materials analyzed were:

Balsa - With a high strength-to-weight ratio and ease of construction, balsa manufactured parts

result in quick, rigid structures. However, balsa is limited by its grain direction and limited

durability. Balsa buildups have been the preferred method in the past and the team is most

familiar with balsa construction.

Extruded polystyrene (XPS) foam - XPS construction is lightweight and is strong enough in axial

loads for RC construction. Parts can be produced quickly, easily, and economically. However, the

team is not as experienced in producing parts from XPS than from other materials.

Carbon fiber – A strong and rigid material for its weight, carbon fiber can strengthen any load

bearing structure and complex shapes can easily be made from molds. The team is experienced

in creating carbon fiber lay-ups.

The manufacturing methods used were a combination of processes that produce the highest

quality components while also being within the scope of the team’s resources. The selection of materials

was driven by the required weight and strength of each component, as well as by the team’s experience

with manufacturing that material (Table 32). It was decided to construct the aircraft as a composite of all

three analyzed materials. The aircraft's overall structure is primarily XPS and carbon fiber. XPS is favored

due to its low density (a fifth of the density of balsa), ease of manufacturing, and dynamic load durability

critical to a bush plane design. Carbon fiber parts primarily handle the major static and dynamic loads of

the aircraft. XPS makes up the profile and majority of the structure of the aircraft. Balsa is used for less

heavily loaded components and for sheeting along the leading edge of the wing.


Table 32: Material Selection FOM

Category Weight Balsa XPS Carbon Fiber

Strength 0.35 3 2 4

Weight 0.35 3 5 3


Experience 0.15 4 3 4

Total 1 3.3 3.5 3.5

It was imperative that the simplest, most effective manufacturing process be selected for each

component on the aircraft. The highest available level of technology and industry manufacturing

standards were heavily considered in the design of each component.

A major contributor to the high cost of building an aircraft of this type is molds for carbon fiber lay-

ups. In order to alleviate this rising manufacturing cost, the team researched the best way to construct

molds using existing university facilities. It was determined that medium density fiberboard (MDF) could

be milled in a CNC machine and then painted to provide an excellent surface for molding. This allowed

the team to rapidly and inexpensively prepare molds for carbon fiber and fiberglass layups. Components

made from carbon fiber composites were those in the aircraft where high loading was expected. This

usually included parts with complex geometry, such as the spar and landing gear.

The selection of XPS as an alternative to balsa made it necessary to determine the most efficient

way of cutting the material to the proper dimensions. XPS has been used widely in R/C aircraft, and

proved to be the best type of foam to use. It was determined that a hot wire cutting tool was the best way

to do this. The team designed and built a custom hot wire cutting tool capable of manufacturing large

pieces such as the fuselage (Figure 27). By using two dimensional CAD projections, wire guides could be

precision laser cut and used to cut XPS blocks with a hot wire to the desired shapes and sizes.

Figure 27: Custom-Built Hot Wire Cutter with Power Supply


6.2 Manufacturing of Components

6.2.1 Wing Assembly

The wing is composed of a C-channel carbon fiber main spar, balsa rear spar, balsa sheeted

leading edge, XPS ribs, and a carry-through XPS-carbon fiber composite wing box. The wing spar, the

main structural component in the wing assembly, was made of two ply of 3K carbon fiber. A mold was

designed in SolidWorks, translated to a Mastercam code for CNC, and machined from MDF. The mold

was sanded and painted into a smooth, non-porous surface for easier part removal after carbon fiber

layup. 3K carbon fiber weave was used on all surfaces of the spar, and in order to increase bending

stiffness, uniaxial carbon fiber was placed at the front section of the spar.

After the spar finished curing, the wing box carry through structure, composed of a rectangular

piece of XPS wrapped in one ply of 3k carbon fiber, was added to the assembly. The motor mounts were

then attached to the front of the spar in the proper positions. Finally, the spine was inserted through the

wing box and the spar, and secured in place.

Wing ribs were cut from XPS and attached to the spar. The ribs were spaced approximately 5

inches apart, starting from the wingtip. A rear spar made of 0.125 inch balsa was added at the end of the

trailing edge ribs. This helped keep the ribs in place and created a flat surface where the ailerons could

be attached. The leading edge of the wing was sheeted with 1/32 inch balsa. This gives the wing torsional

stiffness and creates a smooth surface around the leading edge.

Figure 28: Wing Assembly with Spine

6.2.2 Fuselage

As a new method to fabricate lightweight structural components, the fuselage bottom and top

were fabricated as solid XPS pieces using a hot wire cutter, rather than being constructed from balsa or

hardwood. A rectangular section was cut from the fuselage bottom where the fuselage attaches to the


wing box (Figure 29). The body pieces were sanded to improve aerodynamic performance and to remove

excess material not critical to structural needs. Along the fuselage top, circles were cut to reduce weight.

A circular notch was cut down the center of the fuselage where the spine would lie. An XPS floor was

epoxied to the inside of the fuselage bottom.

Figure 29: Fuselage Top (left) and Fuselage Bottom with Floor Insert

6.2.3 Nose and Tail Cone

A negative mold was created for the nose cone using the same method as for the wing spar. A

single layer of carbon fiber was then applied the mold, and the resulting shell was reinforced with XPS

inserts. The tail cone was cut from XPS and sanded smooth to serve as a positive mold. A carbon fiber

shell was laid up on to the tail cone, cured, and removed.

Figure 30: Nose (left) and Tail Cones

6.2.4 Tail Assembly

Similar to the fuselage, the horizontal and vertical stabilizers of the tail assembly were cut from

solid XPS pieces. To strengthen the horizontal stabilizer in bending and torsion, a 0.125” balsa spar was

added at the quarter chord. A central cut was made for the spine to be inserted, which was then epoxied

to the horizontal stabilizer and the balsa spar. A box cut was made in the horizontal stabilizer to place the


two servos required for the rudders and the elevator. The rudders and elevators were constructed from

1/32” balsa sheeting and 0.125” balsa ribs.

Figure 31: Tail Assembly

6.2.5 Landing Gear Assembly

The landing gear was manufactured from a thin XPS core wrapped in carbon fiber. A two

dimensional profile of the landing gear was used to cut the foam core from a 0.5 inch sheet of XPS. The

XPS core was trimmed to fit and was wrapped in two ply 3K carbon fiber and placed into the negative

mold. The negative mold was used to maintain the resulting shape of the landing gear. After curing, holes

were drilled in the lower legs and the wheels were attached. Holes were then drilled in the top of the

landing gear and the assembly was bolted to the wing box within the fuselage with Nylon bolts.

Figure 32: Landing Gear Mold and Assembly

6.3 Manufacturing Schedule

To keep the project on track during the fabrication period, a manufacturing schedule was


established. This schedule also ensured the build of another prototype, and finally of the competition

aircraft. The Gantt chart below shows the estimated and actual manufacturing schedules for the aircraft.

Figure 33: Manufacturing Milestone Chart

7.0 Test Plan

The aircraft test plan is outlined in Table 33.

Table 33: Test Procedures

System Test Dates

Propulsions

Building static thrust stand Setting up battery testing system

8/28-9/6

Battery type selection Motor selection using PropCalc

9/8-9/14

Static Thrust Propeller selection using static thrust

9/15-19/30

Structures Wing Spar Loading Test 12/28-2/14

Landing Gear

Composite Landing gear static load test Composite Landing gear drop test

Foam wheel static load test Foam wheel dynamic load test

2/4-2/5

Flight Mission Requirements Testing 1/8-4/2

Mold Prep

Composite Layup

Foam Cutting

Assembly

Feb.

Foam Cutting

Assembly

Mar. Apr.

Prototype 2

Sept. Oct. Nov. Dec. Jan.

Wing Loading Test

Mold Prep

Composite Layups

Foam Cutting

Assembly

Prototype 1

Mold Manufacturing

Composite Layups


Mold Prep

Composite Layups

Foam Cutting

Assembly

Estimated Actual


7.1 Propulsion Testing

7.1.2 Static Thrust Test

In order to gain an understanding of the propulsion systems for the aircraft, static thrust testing

was conducted to measure the performance of various battery, motor, and propeller combinations to

optimize thrust and flight time. A static thrust test apparatus was designed and built to directly measure

thrust by opposing gravitational force rather than indirectly through use of a strain gage. A metal

pyramidal structure was constructed so that the motor’s thrust vector would act completely in the vertical

Z-direction. The scale was zeroed before each test to remove the motor, battery, and stand weight. As

the motor was operating, the scale reading in pounds was recorded as the net static thrust.

Because the motor and propeller pull up in the Z-direction, the propeller wash influences the

scale through dynamic pressure. To account for this force a Pitot-static tube was placed directly under the

propeller so that the dynamic pressure could be measured by a pressure transducer connected to a volt

meter. After several tests it was found that the dynamic pressure induced on the scale was minimal and

could be ignored.

Figure 34: Static Thrust Test Apparatus

7.2 Structural Testing

7.2.1 Wing Loading Test

The spar is the most critical structural component in the wing assembly. This member carries the

load of the aircraft in flight. The wing assembly including the composite spar, the XPS ribs, the trailing-

edge balsa spar, and the balsa-sheeted leading edge was built up, excluding the Hoerner wing tips and

the flaperons. A carbon fiber rod simulating the spine ran through the wing box carry through structure.

The carbon rod was clamped to a test jig that levelly suspended the test wing 1.25 ft above the table.

Incremental loading was accomplished using 1 lb bags of sand up to a load of 32 lb, then with 0.5 lb

bags. To ensure that the loading was properly simulated on the test wing, as would be encountered in


flight, the bags of sand were added simultaneously and symmetrically onto the wing in an elliptical shape.

Bags of sand were continually added until the wing assembly failed. Figure 35 shows the test set up.

Figure 35: Wing Loading Test Setup

7.2.2 Landing Gear Test

Initial, but not exhaustive, tests of this aircraft’s landing gear looked promising, but after the first

landing gear broke on a hard landing during a test flight, testing methods were reevaluated. After

repairing this landing gear, which had a balsa core, a new one was built with an XPS core. Each landing

gear was tested by bolting it upside down to a board clamped to a workbench, then suspending weights

from the axles, being sure to distribute the weight evenly between both sides of the gear.

The balsa landing gear broke under a load of only twenty pounds by failing in the same place as it

had been repaired. The XPS core landing gear performed much better: though it deformed greatly, it held

a total load of 30 lb without any signs of damage. The board was then unclamped from the table and

wheels were added to the landing gear. To simulate the dynamic load of a rough landing, the board was

loaded with a weight of 8 lb, approximately the maximum weight of the fully loaded aircraft, and dropped

from a height of six inches. The landing gear did not survive this test, but the design still seemed viable. A

new landing gear was then manufactured with more attention paid to the composite layup, and this

landing gear was able to pass all the tests performed on the first model.

7.2.3 Wheel Test

The wheels for this aircraft were tested both statically and dynamically using a setup similar to the

dynamic test for the landing gear. The landing gear was again bolted to a wooden board and two wheels

were attached, one with a solid piece of XPS and one with the XPS between the wooden spokes cut out.

The wheels were tested and held up to 31 lb in this manner. The board was then tested dynamically as in

the landing gear test. Both wheels were unharmed in this test.


Figure 36: Landing Gear/Wheel Static Load Test

7.3 Flight Testing

After the completion of the first prototype, the aircraft entered an extensive flight test program to

examine and validate its flying characteristics and operating envelope. Goals and milestones for the

aircraft flight test program included:

First Flight

o Proof of Concept

Pilot Familiarization/ Handling Qualities

o Flight Controls Response (trimming)

Performance Testing

o 40 ft Take-Off

o Execute Flight Maneuvers with Maximum Payload

o Maximum Speed for all Configurations

These goals helped the team understand the performance of the aircraft and find improvements

that needed to be made on the path toward optimizing the design of the aircraft. Before every flight test, a

preflight inspection of the aircraft was carried out using the checklist in Table 34.


Table 34: Preflight Checklist

Flight-Line Checklist

Structural

Visual Inspection of Each Component For Damage

□ Wing □ Spar

□ Fuselage □ Landing Gear

□ Tail □ Empty Weight

□ Control Surfaces □ CG Locations

□ Spine

Payload

Payload Check

□ Check Payload Configuration □ Top Lid Secure

□ Payload Secure

Controls

Aircraft Controls Check

□ Wires Connected □ Radio Range Check

□ Control Surfaces Check □ Throttle Check

Propulsions

Propulsions Systems Connected and Secure Check

□ Batteries Charged □ Batteries Secure

□ Propellers Secure □ Batteries Connected

□ Motors & ESC Secure □ RPM Calibration

Ground Check

Alert All at Airfield Before Flight

□ Clear Runway □ Aircraft Total Weight

□ Pilot/Spotter Check □ CG Location

□ Flight Visual Check

The RPM calibration consisted of using a Tachometer to check that both propellers were spinning

within 50 to 100 RPMs of each other for safe flight. This was done to prevent any inadvertent yawing

during flight. For flight tests an ArduPilot was placed onboard the aircraft and used exclusively for data

acquisition; data were transmitted through telemetry to a ground station. It was utilized on several test

flights to record data for further examination of the flight dynamics of the aircraft. This technology allowed

the pilot to replay the flight on the Primary Flight Display (PFD) to debrief after every flight. This helped

him optimize his control input to the aircraft during flight.


Figure 37: PFD Display

7.4 Rough Field Taxi Test

The aircraft’s ground handling performance for the Rough Field Taxi Mission was tested. The

roofing panels to be used for the competition were purchased and a miniature version of the taxi course

was constructed. A switch was set up on the transmitter to change from flight mode to taxi mode, allowing

the pilot to use the rudder controls on the transmitter for differential thrust on the ground. The first goal of

the test was to get the pilot acclimated to the motor control for differential thrust. Once the pilot felt

comfortable with this procedure, the aircraft’s turning capabilities were tested. Finally, taxi tests were

performed over the corrugated roofing panels with turning between the 2x4 inch obstacles.

Figure 38: Aircraft Taxi Test over Corrugated Surface


8.0 Performance Results

Results from the tests outlined in Section 7.0 were used to determine if redesigns would be

necessary.

8.1 Propulsion System Results

8.1.1 Battery Pack Performance Testing

To determine the energy density of each cell, a test was designed in which a large variable

carbon plate resistor discharged the assembled 13 cell series battery pack. The voltage was logged every

0.5 seconds using a data acquisition unit. Five trials were conducted with a 10 A, 13 A, and 15 A current

draw. Table 35 below compares each cell tested where the Elite 1500 mAh was clearly the most energy

dense cell for the range of current draws.

Table 35: Energy Density of Three Battery Cells Tested at Varying Current Draws

Cell Type Mass/Cell 10 A 13 A 15 A

Tenergy 1200 20 grams 121.72 138.84 163.51

Tenergy 1600 25 grams 143.46 173.99 196.67

Elite 1500 23 grams 148.81 188.14 208.26

Figure 39 below shows the data collected for the Elite 1500 mAh battery pack where the energy

density for each test was determined by the area under the curve of the voltage over time using the

following equation:

∑

(7)

Figure 39: Elite 1500 mAh Battery Pack Characterization of Voltage Over Time

0.00

2.00

4.00

6.00

8.00

10.00

12.00

14.00

16.00

18.00

20.00

0:00:00 0:01:26 0:02:53 0:04:19 0:05:46 0:07:12 0:08:38

Vo

tla

ge

(V

)

Time (hh:mm:ss)

10A Test 1

10A Test 2

15A Test 1

15A Test 2


8.1.2 Static Thrust Testing

Static thrust testing was performed to measure the amount of thrust produced under the take-off

conditions. Using Elite 1500 mAh battery packs and the three-blade propeller, a static thrust of 3.2 lb was

found to be available at full throttle for take-off.

8.2 Structural Performance

The spar was structurally intact until a load of 40 lb was applied. At that point, the main spar

fractured on the left wing one inch from the wing root as seen in Figure 40.

Figure 40: Wing Test Main Spar Fracture at a 40 lb Load

Deflections at the wingtips were measured after every 2 lb added, up to 32 lb, and then measured

after every 1 lb was added, up to failure (Figure 41). By the end of the test, the total wing deflection was

found to be1.5”on each side of the wing.

Figure 41: Wingtip Deflection during Wing Loading Test

y = 0.0008x2 + 0.0055x R² = 0.9855

0.00

0.25

0.50

0.75

1.00

1.25

1.50

0 5 10 15 20 25 30 35 40 45

De

fle

cti

on

(in

)

Wing Load (lb)


By using the equation shown in Figure 41, the estimated wingtip deflection was used to optimize

the flight performance for each mission. The maximum load in the test was much greater than the

maximum load of 30 lb expected to act on the wing during flight. Because the test proved the structure

could handle a larger load factor than necessary, the overall design of the wing structure was confirmed.

8.3 Taxi Performance

Testing proved that differential thrust control of the twin motors has the capability of steering the

aircraft through the taxi course. With the large wheel design, the aircraft was able to navigate through the

course without too much difficulty. It was noted from testing that redesigning the landing gear attachment

by locating it slightly farther forward would lessen the plane’s tendency to tip forward. Also, fine tuning the

transmitter’s left and right motor outputs to give the pilot better control over the differential thrust improved

the ground handling control over the rough surface.

8.4 Flight Performance

Numerous flights of the first prototype have been completed and documented. The aircraft

exhibited exceptional handling characteristics, especially when fully loaded with the M2 payload. The

initial prototype was flown heavier than subsequent aircraft will be, primarily due to initial construction

problems and weight and balance issues with the first prototype. Initially, the aircraft marginally satisfied

the 40ft take-off requirement at MGTOW. This was resolved using Hoerner wing-tips and the use of flaps.

The second aircraft, with performance and design enhancements, will be built and rigorously test flown

following the completion of this report. Figure 42 shows the aircraft in flight.

Figure 42: Picture of aircraft in flight.


References

[1] AIAA. 2013/14 Rules and Vehicle Design. [http://www.aiaadbf.org/2014_files/2014_rules_31Oct.html]

[2] Raymer, Daniel P., Aircraft Design: A Conceptual Approach. Fifth Edition. AIAA Education Series.

Reston, VA, 2012.

[3] Hepperle, Martin, “Javafoil v2.0,” 1996-2008[http://www.mh-aerotools.de/airfoils/jf_applet.htm] pg 18

[4] Müller, Markus, “ecalc v P5.17 06Sept12 [http://www.ecalc.ch/motorcalc_e.asp?ecalc.

[5] Nelson, Robert C., Flight Stability and Automatic Control, Second Edition. McGraw-Hill International

Editions Aerospace Science and Technology Series. Boston, MA, 1998.

[6] Bertin, John J., Brandt, Steven A., Stiles, Randall J., Whitford, Ray, Introduction to Aeronautics: A

Design Perspective, Second Edition. AIAA Education Series. Reston, VA, 2004.

Documents

2014DBF_University_of_Oklahoma