rc airplane designii executive summary this report outlines the project scope and definition of a...

RC AIRPLANE DESIGN TEAM SKYTANIC

TEAM #3

SUBMITTED TO:

DR. J. YANG

25TH JULY, 2012

Akposeiyifa J. Ebufegha Bongai Simango

Caitlin Thompson David Lee

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ABSTRACT

A RC model airplane is to be designed under various constraints as determined by the SAE micro-class

RC airplane contest guidelines. The purpose of this competition was to produce plane with the highest

payload to total weight fraction possible. This RC airplane should employ a unique and creative team-

created design whilst being based on sound research and theory. Upon completing the research and

design phase, rigorous testing was performed to estimate the performance of the rc airplane from which

it was determined that under the optimal conditions the rc airplane should have a payload to total

weight fraction of 0.85.

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EXECUTIVE SUMMARY

This report outlines the project scope and definition of a micro-class radio controlled (RC)

airplane as well as the design process and fabrication considerations taken into account in the creation

of RC airplane. Students form groups of four, from which each team is to develop a project management

plan for the completion of the project. These plans include subtasks, as well as the time required from

completion, along with the designated person assigned to the task. This also includes the allocation and

management of the resources necessary for project’s realization.

A RC model airplane is to be designed under various constraints as determined by the SAE

micro-class RC airplane contest guidelines. This RC airplane should employ a team-created design that is

both unique and creative whilst being based on sound research and theory. The congregation of this

research should allow the project team to determine a design configuration best suited for the task -

traversing a distance of 20 meters, from which evaluation is determined from two main factors; the

model’s ability to carry the highest payload fraction possible and adhering to the project design

constraints. The imposed design constraints are organized into three groups; general, technical, and

performance.

This project is to be completed within a period of three months whilst maintaining a budget cap

of $600. The budget not only covers the fabrication of the RC airplane, but the any form of testing and

analysis to be performed over the course of executing the project.

The project team initiated the project by outlining a set of design objectives to which any

subsequent design analysis results will be compared to. These objectives were based on the following

premises; maximizing lift, minimizing drag and minimizing airplane weight. The major team objective is

to achieve a payload to total weight fraction of 0.9 as this is the main factor that is being measured for

competition purposes. Needless to say, the early phase of design required that background research be

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performed in order to obtain a sound knowledge base in regards to RC airplane design principles as well

as any other aerodynamic principles that apply to scaled-down airplanes that experience low Reynolds

number flight conditions.

Upon completing the research phase, it was determined that the rc airplane design should

implement a low taper ratio, large wingspan and a large wing area. Using the preliminary research and

objectives, it was decided that the group would proceed with flying wing or delta style design due to its

excellent lift to drag ratio and general benefits with respect to generating lift in comparison to other

airplane designs. Further research suggests that the plane’s aspect ratio should fall within the range of

6-8 and the plane frame should be primarily constructed using balsa wood or basswood.

The team proceeded to perform bending stress and deflection estimations for the fuselage and

wing utilizing the Bernoulli Euler Beam and simple beam models. These results were confirmed with

simple experiments. A vibration analysis was also performed from which it was determined that the

effects of vibration were insignificant. The project team then proceeded to perform computational fluid

dynamics analysis and aerodynamic analysis using Solidworks 3D and XFLR5. This analysis allowed the

team to conclude that the design being utilized is mostly stable both laterally and longitudinally. At this

point, it was concluded that the design was suitable for manufacture.

At this point a prototype was created to determine potential areas of concern before actual

fabrication began. After the prototyping phase the actual model was constructed and balancing tests

were performed optimize the positioning of the electronic components and confirm static stability. To

conclude the manufacturing phase the components were tested to ensure that the propulsion system

and control surfaces were functioning.

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Nomenclature

AWING Wing Area

b Wing Span

AR Aspect Ratio

λ Taper Ratio

CLIFT Coefficient of Lift

CDRAG Coefficient of Drag

ρ Density of Air

V Craft Airspeed

w Load

l Lift

L Length

M Moment

CG Center of Gravity

E Young's Modulus

I Area moment of Inertia

θ Deflection Angle

v Deflection

σ Bending Stress

Y Amplitude of Vibration in Y axis

LB Rolling Moment

NB Yawing Moment


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Table of Contents

ABSTRACT ............................................................................................................................................. i

EXECUTIVE SUMMARY ......................................................................................................................... ii

Nomenclature ..................................................................................................................................... iv

1. INTRODUCTION ............................................................................................................................ 1

1.1 PROJECT SCOPE AND DEFINITION .......................................................................................... 1

1.2 DESIGN CONSTRAINTS ........................................................................................................... 1

1.3 DESIGN OBJECTIVES ............................................................................................................... 2

2. DESIGN PROCESS .......................................................................................................................... 2

2.1 BACKGROUNG RESEARCH ...................................................................................................... 2

2.1.1 LIFT& LIFTING BODIES .............................................................................................. 2

2.1.2 WINGS ...................................................................................................................... 2

2.1.3 PLANE CONTROL SURFACES ..................................................................................... 4

2.1.4 WING DESIGN PARAMETERS .................................................................................... 5

2.1.5 THRUST& THE PROPELLER ........................................................................................ 7

2.1.6 MATERIALS ............................................................................................................... 8

2.2 DESIGN CONCEPT GENERATION ............................................................................................ 9

2.2.1 CONCEPTS ................................................................................................................. 9

2.3 CONCEPT SELECTION AND EVALUATION ............................................................................. 11

2.3.1 DESIGN CHANGES & JUSTIFICATION TO CHANGES ................................................ 13

3. DETAILED DESIGN CALCULATION & ANALYSIS ............................................................................. 13

3.1 AERODYNAMIC ANALYSIS .................................................................................................... 13

3.2 STRUCTURAL ANALYSIS ....................................................................................................... 17

3.2.1 WING DEFLECTION AND BENDING STRESS ............................................................ 17

3.2.2 FUSELAGE DEFLECTION AND BENDING STRESS ..................................................... 19

3.3 VIBRATION ANALYSIS ........................................................................................................... 21

3.4 STABILITY AND CONTROL .................................................................................................... 22

4. TESTING AND EXPERIMENTAL RESULTS ...................................................................................... 26

4.1 WING DEFLECTION TESTING ................................................................................................ 26

4.2 FUSELAGE DEFLECTION TESTING ......................................................................................... 27

4.3 DETERMINING CENTER OF GRAVITY (BALANCING TEST) .................................................... 27


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5. MANUFACTURING AND FABRICATION ........................................................................................ 28

6. PERFORMANCE PREDICTION ...................................................................................................... 29

6.1 PREDICTED FLIGHT CHARACTERISTICS ................................................................................ 29

7. CONCLUSION ...................................................................................................................................... 29

8. RECOMMENDATIONS ......................................................................................................................... 30

9. PROJECT MANAGEMENT .................................................................................................................... 30

9.1 BUDGET DETAILS ................................................................................................................. 30

9.2 PROJECT MANAGEMENT PLAN ............................................................................................ 32

10. REFERENCES ........................................................................................................................... 35

APPENDICES ....................................................................................................................................... 36

APPENDIX A ....................................................................................................................................... 37

APPENDIX B ....................................................................................................................................... 40

APPENDIX C ....................................................................................................................................... 45

APPENDIX D ....................................................................................................................................... 55

APPENDIX E ....................................................................................................................................... 57

APPENDIX F........................................................................................................................................ 88


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1. INTRODUCTION

This report outlines the project scope and definition of a micro-class radio controlled (RC) airplane as

well as the design process and fabrication considerations taken into account in the creation of RC

airplane. Students form groups of four, from which each team is to develop a project management plan

for the completion of the project. These plans include subtasks, as well as the time required from

completion, along with the designated person assigned to the task. This also includes the allocation and

management of the resources necessary for project’s realization.

1.1 PROJECT SCOPE AND DEFINITION

A RC model airplane is to be designed under various constraints as determined by the SAE micro-class

RC airplane contest guidelines. This RC airplane should employ a team-created design that is both

unique and creative whilst being based on sound research and theory. The congregation of this research

should allow the project team to determine a design configuration best suited for the task - traversing a

distance of 20 meters, from which evaluation is determined from two main factors; the model’s ability

to carry the highest payload fraction possible and adhering to the project design constraints. The

imposed design constraints are organized into three groups; general, technical, and performance.

This project is to be completed within a period of three months whilst maintaining a budget cap of $600.

The budget not only covers the fabrication of the RC airplane, but the any form of testing and analysis to

be performed over the course of executing the project.

1.2 DESIGN CONSTRAINTS

The criteria for the design are outlined in the SAE International rules. Requirements, such as the payload dimensions, geometry, types of wings, for example, were used as a guideline for the selection process. Outlined and organized in the SAE rules, the restrictions were, likewise, categorized. The restrictions were categorized into four groups; general, performance, and geometric.

General restrictions are broad, but include factors such as;

I. The required amount of time to assemble the RC airplane, the plane is to be assembled in three minutes by two people;

II. The battery type. For the sake of this project, the power supply and motor will be supplied by the Faculty of Engineering and Applied Science (See Appendix A for part specifications); and

III. Payload materials that are prohibited, the payload cannot be made of lead; IV. The use of breakable propellers.

Performance restrictions deal with the model planes phases of flight; launch, flight, and landing.

I. Launch phase specifications involve such things as launching mechanism, in this case the plane is to be hand launched;

II. The in-flight requirements include the ability for the plane to stabilize its path; and III. The landing requirements include factors such as bouncing out of the designated landing area

and landing in the same direction as launch.


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Geometric considerations include;

I. The RC airplane must fit into a box of 8x12x24 inches when disassembled; and II. The payload must fit into a separable cargo size at 5x2x2 inches at the minimun.

Further details on design constraints are outlined in the SAE International (abbr) contest guidelines.

1.3 DESIGN OBJECTIVES

Three main ideas were taken into consideration whilst deciding what the design objectives were;

maximizing lift, minimizing drag and minimizing airplane weight. Keeping these ideas in mind the

following objectives were set;

Table 1.1: Initial Design Objectives

Plane Weight 800g – 900g

Payload to Total Weight Fraction 0.9

Assembly Time Less than 2 minutes

Target Airspeed 10 m/s

Wing Area 500 – 600 squared inches

Wing Span 60 inches

Taper Ratio 0

2. DESIGN PROCESS

2.1 BACKGROUNG RESEARCH

2.1.1 LIFT& LIFTING BODIES

Lift is defined as a mechanical force generated by a solid object moving through a fluid. Essentially, lift is the force that opposes the weight of the airplane allowing the airplane to achieve flight. A lifting body is any part of the plane that generates lift. Although lift can be generated by every part of the airplane, the majority of the lift acting on an aircraft is generated by the aircraft’s wings [1].

2.1.2 WINGS

The wing is designed to slice the air, the flowing fluid, into two streams; one that flows over the wing, and the other flowing under it the wing. In standard design, the wing geometry and orientation are designed such that the air flowing on its topside has a greater velocity than the air flowing underneath. This is important because the higher velocity air stream exerts less pressure on the wing than the lower velocity stream beneath it, this pressure difference results in the upward acting force referred to as lift [2].

The lift acting on the RC airplane is a crucial part of ensuring that the design is capable of carrying the maximum payload with the given engine size. Thus, it is important to assess the various factors that

Fig. 2.1: Lift acting on a Body [1]


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influence lift on an airplane. Lift is affected by the angle of attack, wing loading, wing geometry, the speed at which the airplane travels through the fluid, and the properties of air. All of these can be influenced by design considerations with the exception of the air properties. The wings angle of attack has a direct correlation with the potential lift generated; however, it will also affect the amount of drag the body generates in an equivalent proportion. With a greater angle of attack the wing will generate more lift; however, it should be noted that there is a “useful” lift. Useful lift refers to a situation in which most the force generated by the wings acts on the airplane body more in the vertical plane as opposed to the horizontal plane, this is where the wing geometry and orientation come into consideration [2].

Fig. 2.2: Wing Orientation and Lift and Drag Generation [1]

The wing profiles on most aircrafts are typically in the shape of airfoils, as shown in the figure below. Airfoil shape is important as it will affect the lift generated. The lift generated is determined by the flow turning on the trailing edge of the foil – with greater flow resulting in a greater lift. Most RC airplanes employ a Clark Y type airfoil is used due to its flat bottom design, high camber, efficient lift to drag ratio and the fact that it is easily constructed. It also offers predictable and gentle stall characteristics [3].

Fig. 2.3: Clark Y Airfoil Outline

The thrust, generated by the propeller and hand launch, is important to offset the effect of drag as a result of variances in wing geometry and orientation. This allows a greater angle of attack to be utilized in the design. It will also affect the velocity at which the craft body moves through the fluid, and as such, affects the maximum amount of lift that the plane can generate. The thrust generated by the plane requires optimizing the propeller design to ensure that the motor is being used in the most efficient manner to generate the most thrust.

Wing loads are a very important design consideration. It is a ratio of the aircraft weight to wing area; units are typically in oz/ft2 [3]. It is important because;

I. Wing loading is the only indicator of how "heavy" an aircraft is. The actual weight of an aircraft has relatively little meaning in determining how well the plane will fly.


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II. The lighter the wing loading, the slower the aircraft can take-off, fly and land. It will also have a better climb.

Increasing the wing area are will reduce the wing loading and allow the aircraft the benefits of being

“lighter”. Lift is a function of the lift coefficient (CLIFT), wing area (AWING), air density () and airspeed (V) as shown in the formula below [1];

𝐿𝑖𝑓𝑡 = 𝐶𝐿𝐼𝐹𝑇 ×𝜌×𝑉2

2× 𝐴𝑊𝐼𝑁𝐺

2.1.3 PLANE CONTROL SURFACES

There are four mechanisms that control the flight of an

aircraft; flaps, ailerons, elevators, and a rudder. Airplanes

make the use of flaps or slats to change the shapes of their

wings and tails. The purpose of the flaps and slats is to alter

the amount of lift and drag, ultimately controlling how the

airplane flies.

Flaps (Fig. 2.4) are located on the back of the wing and are

used during take-off and landing to adjust the shape of the

wing. The flaps extend downward from the trailing edge to

cause more lift during take-off. The change in shape also

increases drag that assists in slowing down for landing [2].

Ailerons (Fig. 2.5) are located closer to the end of the wing

and are used in opposition so that one wing will create

more lift than the other. This allows the plane to bank or

roll either left or right [2].

The rudder (Fig. 2.6) and elevators (Fig. 2.7) are stabilizers

located on the tail of an airplane. The rudder is a vertical

flap and is used to turn the plane left or right. Elevators,

located on the horizontal part of the tail, are deployed to

make the plane go up or down [2].

Fig. 2.4: Airplane with Flaps Highlighted [2]

Fig. 2.5: Airplane with Ailerons Highlighted [2]

Fig. 2.6: Airplane with Rudder Highlighted [2]

Fig. 2.7: Airplane with Elevator Highlighted [2]

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2.1.4 WING DESIGN PARAMETERS

WINGSPAN

Wingspan refers to the end-to-end length of the aircraft from one edge of the wing to the other. Deciding the wing span is one of the most basic decisions to be made in the design of a wing. It is best to utilize the largest wing span consistent with structural dynamic constraints. This should reduce the induced drag directly and create more lift [3]. It is important to note that as the wing span is increased, the wing structural weight also increases, from which the weight increase offsets the induced reduced drag. However, it is difficult to reach this point.

ASPECT RATIO Aspect ratio has two general definitions. It can be defined as the ratio of the wing span (b) to the mean wing chord (C), but is commonly defined as the ratio of the wing area (A) to the wing span squared. The formulaic representation is shown below [3];

𝐴𝑠𝑝𝑒𝑐𝑡 𝑅𝑎𝑡𝑖𝑜 =𝑏2

𝐴=

𝑏

𝐶

Aspect ratio is a major factor determining the dimensional characteristics of the ordinary wing as well as its lift-drag ratio. There are generally two options available for selecting aspect ratios; one can design a low aspect ratio wing or a high aspect ratio wing. Increasing aspect ratio results in an increase in lift experienced by the craft at a given angle of attack. It will result in changes to the wing lift distribution by intensifying the effects of all other parameters. An increase in aspect ratio with constant velocity will also decrease the drag the aircraft experiences, an effect that becomes more apparent with higher angles of attack. This will improve the performance of the wing when in a climbing attitude. Low aspect ratio wings have nearly elliptic distributions of lift for a wide range of taper ratios and sweep angles. It takes a great deal of twist to change the distribution. Very high aspect ratio wings generate more lift but are quite sensitive to direction changes in flight. Also, a higher aspect ratio also has the effect of a higher rate of lift increase, as angle of attack increases, than lower aspect ratio wings.

Fig. 2.8: Typical Wing Design Parameters [3]

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Aspect ratio has an effect on the lift curve, that is to say, a high aspect ratio wing will have a higher maximum lift coefficient, but will also have a lower stalling angle of attack than a low aspect ratio wing that employs the same airfoil shape. In ultra-lightweight planes, a higher aspect ratio is typically employed. Most ultra-lightweight planes typically have an aspect ratio between 5.5 and 8 and light general aviation aircraft between 7 and 9, averaging around 7.5 [3]. This is the range in which the project testing will be focused on. TAPER RATIO The taper ratio can be in either planform or thickness, or a combination of the two. It is defined as a ratio expressing the decrease from wing root’s chord length (Croot) to wing tip’s chord length (Ctip). The formulaic expression is shown below [3];

𝑇𝑎𝑝𝑒𝑟 𝑅𝑎𝑡𝑖𝑜 = 𝐶𝑡𝑖𝑝

𝐶𝑟𝑜𝑜𝑡

The taper ratio is important for design consideration as it can affect the weight of the wing. A lower taper ratio will typically result in lighter wings, leading to a smaller load on the engine. Also, increasing the root chord of the plane will create more space to accommodate a landing gear. A short tip chord length; however, can result in a reduced lift coefficient and unacceptable stall characteristics hindering the reduction of the taper ratio [3]. Essentially, the design goal is to keep the taper ratio as small as possible without creating excessive variation in the lift coefficient or poor stalling characteristics. SWEEP Sweep can be defined as the slant of a wing, horizontal tail, or other airfoil surface. It is typically employed for its effect on drag as it permits a greater lift coefficient without drag divergence. However, it has negative impacts on stall, increases the wing loading and destabilizes the airplane. It is best employed with lower aspect ratio planes [3]. WING TWIST Wing twist refers to an aerodynamic feature added to aircraft wings to change lift distribution along the wing and the lift coefficient distributions. It has a positive effect on pitching moment and a little effect on trimmed drag, but also increases the structural weight of the craft [3]. This is a design element to be employed on a situational basis in RC airplane design, as it might prove more harmful than beneficial. DIHEDRAL AND ANHEDRAL Dihedral and anhedral refer to the angle that results from slanting the wing by raising (dihedral) or lowering (anhedral) the wing tip with respect to the wing root. The dihedral is primarily used to counteract rolling and make the plane more laterally stable. Anhedrals are primarily used to counteract naturally occurring dihedrals [3].

(3)

http://aerodynamic/

http://aircraft/

http://wing/


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2.1.5 THRUST& THE PROPELLER

Thrust is a mechanical force that pushes or pulls the aircraft through the air. Its purpose is to overcome the drag of the airplane and cause forward motion [4]. Generally, thrust is generated by the motor of the aircraft through some sort of propulsion system, which in this case is a motor powered propeller.

A propeller generates thrust by converting the power provided by the motor into thrust in a similar manner to the means by which lift is generated by the wing. By forcing the working fluid over the blades of the propeller, it creates an imbalance of pressures on the faces of the blades resulting in forces leading to translation. Flight characteristics can be drastically affected by the propeller design and as such, it is important to understand the effects of the physical properties of the propeller and factor these effects into the design of airplane[5].

The characteristics of a propeller are determined by two main physical properties; the diameter and the pitch. The diameter refers to the tip-to-tip length across a two blade propeller or twice the length of a single blade for multi-blade propellers [5]. A larger diameter propeller is more efficient and allows more thrust to be produced by the engine, thus, the airplane will be able to pull or push a greater payload. However, there is a limit to which the propeller diameter can be extended. Overextending the propeller diameter will result in a gyro effect which will counteract the yaw and pitch of the craft, making it less responsive. A large diameter propeller will increase the load required from the engine to maintain the same rotation speed due to the increased work involved in rotating the extra length. Also, it is important to note that the increase in the size of the propeller will result in the greater load on the motor. Propeller pitch is the theoretical distance the propeller would advance in one revolution under ideal conditions. A higher pitch will result in a greater distance being travelled with each revolution and thus, a faster plane. Much like with the diameter, increasing the pitch will result in a greater load being put on the engine.

Because of the loads experienced on the motor, it is important that propeller specifications, diameter and pitch, are made with motor effects in mind. The engine load being too little or too great could result in damage to the engine, therefore, to ensure that the propeller will not overload or underload the engine, the propeller utilized in the design will be sized according to Fig. 2.9.

Fig. 2.9 was developed from empirical data and displays a range of sizes for the propeller given an engine size. It illustrates the tradeoff between speed and thrust when it comes to RC airplane design. A faster plane with lower drag will utilize a greater pitch values and smaller diameters. A plane that has to push or pull a greater payload requires longer blades. However, it is important to note that the engine sizes are for typical gas type motors, and this project utilizes an electric motor to turn the propeller. Fortunately, the motor suppler provides an equivalent size range of 0.15 – 0.40 for the motor, thus, there is an engine size range to experiment with in order to ensure optimal propeller selection.

There are two methods for which the propeller can provide thrust for the plane; pull or push. Conventional propeller design assumes that the propeller is to be placed in the front of the craft pulling it forward. Alternatively, the motor can be mounted on the back and the plane can be pushed by the propeller. However, if this path is chosen, the motor spin direction must be reversed and the front face of the propeller should be oriented opposite to the convention [5].


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MULTI-BLADE OPTIONS

Two-bladed propellers are commonly used because they are relatively efficient and easy and cheap to produce but sometimes an RC airplane may require more blades, particularly in situations utilizing multiple engines driving the plane. Generally, more blades are utilized in designs as the increased number of blades allows for smaller diameter propellers to be used. Essentially, one can get a similar amount of thrust to a larger diameter propeller by employing more blades [5]. However, it is important to note that adding more blades decreases the overall efficiency of the propeller because each blade has to cut through more turbulent air from the preceding blade, in fact a single bladed propeller is the most efficient but is more difficult to construct due to the need to create a balancing counterweight on the other side of the propeller cap to offset the weight of the blade.

2.1.6 MATERIALS

The ideal RC aircraft frame is lightweight, inexpensive, strong and durable material. Research into material used in RC aircraft construction revealed that the most common materials used are reinforced plastics; polystyrene, polyvinyl chloride, Styrofoam, and wood, the most popular being balsa and bass wood. Given that balsa wood is fairly inexpensive and lightweight, it will be the focus of this research. Please refer to Appendix B for more in-depth information on Balsa wood and its specifications.

One key advantage of balsa over hardwood for model airplanes lies in the ease with which it can be shaped with a sharp knife or a razor and also the fact that it can be used in comparatively large sizes– a most desirable condition for model planes which have to maintain their aerodynamic settings for flying and stability.

There are three cuts of Balsa wood and each is best used in construction of specific plane part [3];

● Tangent Cut: is ideal for fuselage sheet covering, wing leading edges, forming tubes and flexible spars.

● Quarter Grain Cut: Ideal for wings, tails, wing ribs, wing trailing edges, and fuselage formers. ● Random Cut: is a substitute for the other two cuts.

Fig. 2.9: Propeller Size Range for specific Engine Size [4]


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In using Balsa wood, it will be necessary to reinforce the wood depending on its application. The current consensus is that the balsa wood will be reinforced with epoxy resins; a viscous fluid composed mainly of volatile fluid terpenes, with lesser components of dissolved non-volatile solids. Two important mechanical properties of any resin system are its high tensile strength and high stiffness. The resin should also make the wood less permeable and thus prevent craft degradation and weight increase from water absorption.

Team consensus on wing design is that the wing will be composed of covered wing ribs. Research into wing lightweight coverings yielded the possibilities of using Esaki tissues, Monokote, Towercote, Econocote and Ultracote.

2.2 DESIGN CONCEPT GENERATION

To facilitate design process, the team decided on and set areas of focus for the design. These foci were at the background of each design and they are as follows;

I. Maximize Lift : a. Reduced wing loading b. Wing shape & geometry (Applying wing design aspects from background research)

II. Minimize Drag : a. Higher aspect ratio b. Plane Geometry

III. Maximize Thrust: a. Proper motor sizing & Propeller dimensions selection

IV. Minimize Weight: a. Material Selection b. Design implementation considerations. Utilizing wireframe structures, trusses and

avoiding thick solid parts as much as possible.

2.2.1 CONCEPTS

Utilizing the previously mentioned foci, four concepts were developed. Fig. 2.10 shows a delta wing based design. The concept takes advantage of the space provided as per the SAE guidelines; it has a 60” wingspan and a 15” wing root chord length. The design is simplistic and should theoretically generate primarily lift since it is essentially a wing. In terms of production, the design is fairly simple to manufacture as it won’t require fabrication of any complex parts. The most complicated aspect of the design is the taper towards the end of the plane as it requires that the airfoil profile be scaled down appropriately so as to maintain the benefits of the chosen airfoil profile.

Fig. 2.11 shows a design that combines the flying wing concept with aspects of a conventional plane in an attempt to gain the benefits of having a delta wing, the high lift to drag generation ratio, whilst minimizing the inherent instability of the airplane design by introducing stabilizing features such as a tail and rudder as well as an

Fig. 2.10: Flying Wing Concept (Top and Side Views)


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additional set of wings towards tail. Much like with the flying wing concept, this concept will be simple to construct and takes full advantage of the spatial constraints set by the SAE guidelines.

Fig. 2.12 shows a concept that utilizes a more traditional airplane design. The key feature to note with this design is the swept wing. A traditional wing position is portrayed using the dotted line. However, the chosen wing is swept back and this is to improve the airplane’s lateral stability. This concept does not present as large a lift generating surface as with the previously mentioned concepts and will prove more difficult to manufacture due to the wing shape and the fuselage profile.

Fig. 2.13 shows a concept that combines a helicopter to a typical airplane. It requires the use of two motors or sequence of gears to rotate both propellers. One propeller is providing lift and the other is providing thrust. Needless to say, this concept will be difficult to fabricate and its operation will be difficult as a result of operating the two propellers simultaneously. Also power requirements must be taken into consideration with this design as it will most likely require an additional power supply to power the extra motor in the event that a combination of gears cannot be used.

Please refer to Appendix C for description of each concept’s characteristics in detail. This section of the

report merely gives a brief introduction to the concept.

Fig. 2.11: Concept #2 (Top and Side Views)

Fig. 2.12: Concept #3 (Top and Side Views)

Equipment Mounting Platform

(Located Inside plane)

Equipment Mounting Platform

(Located Inside plane)

Fig. 2.13: Concept #4 (Side and Front Views)


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2.3 CONCEPT SELECTION AND EVALUATION

In order to make an objective decision on which concept to use, Team Skytanic employed the Kepner-Tregoe (KT) Troubleshooting method for Selecting a Fix. The decision analysis is in the form of a chart, shown in Table 2.1. On the left hand side is where the list of ‘wants’ and ‘needs’ go for the given issue to be solved. Once a comprehensive list is completed, the ‘needs’ are marked with an “M” and the ‘wants’ are ranked on a scale of 10. The most important ‘want’ is first given a rank of 10 and the following ‘wants’ are ranked accordingly. Along the top margin is each alternative concept or solution. Each concept is marked according to the ‘needs’ first and if any of the concepts do not meet a ‘need’, that concept is disqualified. The remaining concepts are then given are mark out of 10 for each of the ‘wants’ listed in the very first column. Once all of the alternatives are graded, the scores are then multiplied by the weight given to each objective and the scores are totaled. The most suitable alternative will have the highest score.

The following concept precipitated from the KT Troubleshooting Chart;

Fig 2.14: Flying Wing Concept (Top and Side Views)


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Table 2.1: KT Troubleshooting Chart Used In Design Selection Process


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2.3.1 DESIGN CHANGES & JUSTIFICATION TO CHANGES

Over the course of the design process the initial concept has evolved from the simple delta wing shown in Fig. 2.14. The primary idea behind the concept is still the cornerstone of the design, but the changes made were necessary to make the design more practical as well as improve certain aspects of performance. The one major change between the original concept as shown in Fig. 2.14 and the current rendition of the aircraft (Fig. 2.15) is the nose of the plane is no longer a tip. This was done to minimize the impact of crash as well as simplify the construction process. The design still utilizes a 5 ft wingspan, but the fuselage has now been extended to cover 24 inches. This was done to maximize the lift generating surfaces, it also enhances the crafts lateral stability. These are just a few of the changes made to the original concept.

A number of other changes have been made to the design, as the figures show. This will report address the reasons behind these changes, detailing how the design went from the original concept to the current design. It presents the results of the aerodynamic and structural analysis as well as, analyzing aspects of the design relating to stability and control.

3. DETAILED DESIGN CALCULATION & ANALYSIS

3.1 AERODYNAMIC ANALYSIS

3.1.1 PLANE DIMENSIONS RELEVANT TO AERODYNAMIC ANALYSIS

Table 3.1: Results Calculated for Relevant Aerodynamic Factors in Wing Design

Wingspan (inches) 60

Wing Area (inches2) 544.8

Aspect Ratio 6.61

Root Chord Length (inches) 24

Tip Chord Length (inches) 0

Taper Ratio 0

Table 3.1 shows that the airplane has a fairly large wing span and area, this takes advantage of the dimensional restraints set forth in the SAE competition guidelines. Also, the aspect ratio is 6.61 and as such, falls within the optimal range of 6-7 for generating lift. This range was the target range set in the research stage of the project. The taper ratio for the design is zero, and this should reduce the drag experienced by the plane whilst increasing the lateral stability.

3.1.2 CALCULATING THEORETICAL LIFT AND DRAG

The design being implemented uses a Clark Y type airfoil, taking advantage of its high camber and relatively flat base. This foil is chosen as it is a standard choice for RC airplanes. Assuming the plane

Fig. 2.15: Completed RC Airplane


14

takes off horizontally, the angle of attack of the airfoil is approximately 8.300, given the foil dimensions. The angle of attack and the aspect ratio are important in determining the lift and drag coefficients. For the sake of the analysis, the Fig. 3.1 is used to determine the lift and drag coefficients,

Fig 3.1: Lift and Drag Coefficient Chart for Clark Y airfoil [2]

Fig. 3.1 implies that at an angle of attack of 8.300, the lift coefficient is approximately 1.00 and the drag coefficient is approximately 0.065. These results indicate that plane should experience very little drag, but significant amounts of lift in relation to the experienced drag. Determining the lift requires the application of the formula below [1];

Determining the theoretical drag requires the application of the formula below [1];

Equations 4 and 5 are use the coefficient of lift (CLIFT) and coefficient of drag (CDrag) to develop plots that show the theoretical lift and drag at varying air speeds (V) for the given wing area (AWING). For the sake of the analysis, it is assumed that the plane is flying in conditions similar to standard room temperature

and pressure thus providing an air density () of 1.2041 kg/m3. For a list of the numerical values calculated in tabular form refer to Appendix D.

Fig. 3.2 and Fig. 3.3 show that both drag and lift generated should increase as the speed of the craft increases. However, the lift generated for the given design far outweighs the drag generated in magnitude. The lift to drag ratio at any given airspeed is 15.38 : 1, this indicates that the airplane should be able to move at great speed without risk of generating excessive drag.

(4)

(5)


15

Fig. 3.2: Theoretical Lift vs. Airspeed at angle of attack of 8.300

Fig. 3.3: Theoretical Drag vs. Airspeed at angle of attack of 8.300

3.1.3 COMPUTATIONAL FLUID DYNAMICS ANALYSIS USING SOLIDWORKS

Fig. 3.4 shows the flow trajectory juxtaposed on a

surface plot for the model. For both scales, lower

values are shown in blue while higher values are

shown in red. Based on this, it is shown that the

increase of velocity is correlated to the decrease of

pressure, and vice versa.

Fig. 3.5 shows the flow trajectories over the top view

of the plane. It can be seen that the flow leading up

to the plane has a greater velocity than that of the

flow behind the plane. This shows that as air is

passed over the surface of the plane, the air slows

down. At the nose of the plane, the flow of air is

slowing down which indicates a build-up of

0

20

40

60

80

100

0 5 10 15 20 25

Lift

(N

)

Velocity (m/s)

Lift (N) vs Velocity (m/s)

0123456

0 5 10 15 20 25

Dra

g (N

)

Velocity (m/s)

Drag (N) vs Velocity (m/s)

Fig. 3.4: Surface Plot Showing Pressure and

Velocity Distributions on Craft


16

pressure. This indicates the presence of a

boundary layer from which the flow over the

surface is laminar in the beginning and is

turbulent towards the end. Another indication

of this is the fact that the flow trajectory at the

tail end of the plane are oriented randomly

compared to the flow elsewhere on the plane.

The decreased velocity at the falling edge of the

plane indicates the influence that a sharp drop

of the surface has on the velocity, and thus the

pressure. However, not included in this analysis

was the effect of the presence of a propeller.

Due to the amount of processing power and

time constraint, the propeller was omitted from

the analysis.

Additionally provided is a pressure contour of

the top and bottom of the plane. As mentioned

before, the decrease in velocity at the plane’s

nose is demonstrated with the increase of

pressure experienced. This is shown in Fig. 3.7.

The bottom of the plane is shown in Fig. 3.6,

and indicate that the least pressure experienced

was felt at the plane’s “chin” and front edge of

the wings. The top of the plane, as shown in Fig.

3.7, indicates that the lowest pressures are

experienced over the majority of the top of the

wings, as well as over the center of the

fuselage. With this variation of pressure over

the top and bottom of the model, lift can

potentially occur. However, it is also because of

this variation that a moment about the Z-axis

can be created resulting in instabilities during

flight. The use of elevons and launching techniques should help alleviate this issue. It was also the case

that during the simulation set-up, only the x-component of fluid velocity was utilized and as such, do not

represent the intended angle of attack of approximately eight degrees from the horizontal.

Fig. 3.5: Surface Plot Showing Velocity

Distribution on Craft

Fig. 3.6: Surface Plot Showing Pressure

Distribution on Bottom of Craft


17

3.2 STRUCTURAL ANALYSIS

3.2.1 WING DEFLECTION AND BENDING STRESS

The wing deflection and bending stress was calculated under static conditions and dynamic conditions; when the plane is not in flight and when the plane is in flight.

DEFLECTION AND BENDING STRESS UNDER STATIC CONDITIONS

Before delving into the deflection calculations it is important to understand how the wing was modeled for

the purpose of the analysis. The wing is modeled as a cantilevered beam that is subject to a non-uniform distributed load (w). This load is assumed to be non uniform due to the triangular shape of the wing as a result of the taper [6].

Fig. 3.8 shows the wing loading, this load distribution suggests that the moment is greatest at the wing root, this indicates that the bending moment and thus, the bending stress is greatest at the wing root. However, the loading also indicates that the deflection and slope change is greatest at the wing tip. To compensate for these issues, the end of the beam will be made with a stiffer basswood as opposed to balsa wood. Also, the wing root is thicker than the rest of the wing and the wing tip is essentially a point, this should serve to minimize both the effect of the moment at the wing root and the deflection at the wing tip.

Solving for the structural loads and bending requires manipulation of the Bernoulli-Euler beam model provides these standard differential equations, these equations relate the beam length (L), load

distribution, (w0), moment (M), and bending stiffness (EI) to the resulting deflection (max) and bending

moments (max) [7].

𝑑𝑉

𝑑𝑦= 𝑤 𝑦

𝑑𝑀

𝑑𝑦= 𝑉(𝑦)

𝑑𝜃

𝑑𝑦=

𝑀

𝐸𝐼

24”

y

z

(6)

(7)

(8)

Fig 3.8: Mass Loading on the Wing

Fig. 3.7: Surface Plot Showing Pressure

Distribution on Top of Craft


18

Manipulating these formulas allows one to yield the equations for the maximum deflection and the maximum bending stress is derived from the flexure formula. The maximum conditions are the most important for design consideration and as such they are the primary focus in this analysis. Equations utilized in the analysis are as follows;

𝑚𝑎𝑥 = −𝑤0𝐿

4

30 𝐸𝐼

𝜎𝑚𝑎𝑥 = 𝑀 𝑐

𝐼

DEFLECTION AND BENDING STRESS UNDER DYNAMIC CONDITIONS

The wing is modeled as a cantilevered beam that is subject to a non-uniform distributed load “w(y)” as a result of the weight of the wing and a non-uniform distributed load “l(y)” as a result of the lift generated. These loads are assumed to be non-uniform due to the triangular shape of the wing as a result of the taper [6]. The analysis is based on the assumption that the RC airplane achieves an airspeed of 10 m/s and as such is subject to maximum lift of about 21.16 N at the wing root. This indicates that the effect of the lift on the wing should overcome the effect of the mass of the wing on the, leading in a resultant non uniform load distribution in the positive z-axis direction.

The load distribution shown in Fig. 3.9 suggests that the greatest moment is still at the wing root. However, the diagram also shows that the moment in now positive and the magnitude of the loads suggest that the magnitude of moment is greater than in the static case. The load distribution still indicates that the deflection and slope change is greatest at the wing tip. The wing root is thicker than the rest of the wing and the wing tip is essentially a point, this should serve to minimize both the effect of the moment at the wing root and the deflection at the wing tip. The end of the beam will be made with basswood as opposed to balsa wood because basswood is stiffer than balsa wood.

The Bernoulli-Euler beam model provides equations that relate the length (L), the weight distribution (w0), lift distribution (l0), moment due to weight (MWEIGHT), moment due to lift (MLIFT), and the bending

stiffness (EI) to the resulting deflection (max) and bending moments (max) [7]. The equations are equations 6, 7 and 8. Manipulating the formulas whilst accounting for the opposing load distributions allows one to yield the equations for the maximum deflection under dynamic conditions. The maximum bending stress is derived from the flexure formula using the resultant moment. As previously mentioned, the maximum conditions are the most important for design consideration and as such they are the primary focus in this analysis. Equations utilized in the analysis are as follows [7];

𝑚𝑎𝑥 = −𝑤0𝐿

4

30 𝐸𝐼+

𝑙0𝐿4

30 𝐸𝐼=

𝐿4(𝑙0 − 𝑤0)

30 𝐸𝐼

(10)

(9)

y

z

(11)

24”

Fig 3.9: Mass Loading on the Wing during Flight


19

𝜎𝑚𝑎𝑥 = 𝑀𝐿𝐼𝐹𝑇 −𝑀𝑊𝐸𝐼𝐺𝐻𝑇 𝑐

𝐼

CALCULATION RESULTS

Table 3.1: Relevant Wing Dimensions in Metric Units

Projected Wing Mass

(kg)

Maximum Load

“w0”

(N)

Maximum Lift

“l0”

(N)

Wing Span

(m)

Wing Thickness at Root

(m)

Area Moment of Inertia

“I”

(m4)

Young Modulus

“E”

(Pa)

0.047 0.46107 21.16 0.6096 0.0356616 6.91210-6 1.28109

Table 3.2: Theoretical Static and Dynamic Maximum Deflection and Bending Stress

Condition

Maximum Deflection

(m)

Maximum Bending Stress

(Pa)

Static -2.39910-7 147.34

Dynamic 1.07710-5 6615.18

DISCUSSION OF WING STRUCTURAL ANALYSIS RESULTS

The results show that the maximum deflection, which is at the wing tip, should be negligible in either case. Similarly, the maximum bending stress calculated for both static and dynamic conditions suggests that the design has a large factor of safety given that the tensile strength of Balsa wood which ranges between 7.6 MPa – 32.2 MPa.

This suggests that the wing design is structurally sound and should not be subject to flutter. However, these calculations assume steady flight under controlled environmental conditions, and as such, the actual values may be greater than the theoretical calculations indicate.

3.2.2 FUSELAGE DEFLECTION AND BENDING STRESS

The fuselage deflection and bending stress was calculated under static conditions; when the plane is not in flight, and under dynamic conditions; when the plane is in flight.

DEFLECTION AND BENDING STRESS UNDER STATIC CONDITIONS

The fuselage is modeled as a simple supported beam that is fixed on both ends [6]. This model is based on the assumption that the airplane is being held in place and as such the ends are not free to move up or down. This beam is subject to a concentrated load “w” as a result of the weight of the components in

(12)

24”

y

z

Fig 3.10: Mass Loading on the Fuselage


20

the mounting space. The components mounted on the fuselage are modeled as point mass acting at the center of the fuselage under the assumption that the weight of each component acts through its center of mass and that the components are mounted in such a way as to ensure that its center of mass is in-line with the center of the fuselage.

The loading shown in Fig. 3.10 suggests that moments are greatest at the center of the fuselage indicating that the bending stress is greatest there. The loading also indicates that the deflection and slope change is greatest at the center as well. It is important to note, however, that the shear is constant throughout the body; this suggests that the center is the most important point to analyze as the basis of structural integrity.

Manipulating equations 6, 7 and 8 allows one to yield the equations for the maximum deflection. The maximum bending stress is derived from the flexure formula. It is important to note that the maximum conditions are the most important for design consideration and as such, they are the primary focus in this analysis. Equations utilized in the analysis are as follows [7];

𝑚𝑎𝑥 = −𝑤𝐿3

48 𝐸𝐼

𝜎𝑚𝑎𝑥 = 𝑀𝑊𝐸𝐼𝐺𝐻𝑇 𝑐

𝐼

DEFLECTION AND BENDING STRESS UNDER DYNAMIC CONDITIONS

The wing is modeled as a supported beam that is subject to a concentrated load “w” as a result of the weight of the components and a uniform distributed load “l(y)” as a result of the lift generated. The analysis is based on the assumption that the rc airplane achieves an airspeed of 10 m/s and as such is subject to maximum lift of about 21.16 N at the wing root. This indicates that the effect of the lift on the wing should overcome the effect of the mass of the components on the fuselage.

The loading shown in Fig. 3.11 suggests that center is still the primary focus of the analysis as it still experiences the greatest moments and deflection. However, it is important to note that shear is greatest towards the ends and zero in the center due to the lift distribution. Shear is not as significant as bending in this scenario and as such, the more weight is given to the results pertaining to the center.

Manipulating equations 6, 7 and 8 allows one to yield the equations for the maximum deflection. The maximum bending stress is derived from the flexure formula using the resultant moment. As previously mentioned, the maximum conditions are the most important for design consideration and as such they are the primary focus in this analysis. Equations utilized in the analysis are as follows [7];

𝑣𝑚𝑎𝑥 = −𝑤𝐿3

48 𝐸𝐼+

5(𝑙0𝐿4)

384 𝐸𝐼

(14)

(13)

y

z

(15)

Fig. 3.11: Mass Loading on the Fuselage during Flight

24”


21

𝜎𝑚𝑎𝑥 = 𝑀𝐿𝐼𝐹𝑇 −𝑀𝑊𝐸𝐼𝐺𝐻𝑇 𝑐

𝐼

CALCULATION RESULTS

Table 3.3: Relevant Fuselage Dimensions in Metric Units

Component Mass

(kg)

Maximum Load

“w0”

(N)

Maximum Lift

“l0”

(N)

Fuselage Length

(m)

Area Moment of Inertia

“I”

(m4)

Young Modulus

“E”

(Pa)

0.359 3.52179 21.16 0.3048 3.43010-10 1.28109

Table 3.4: Theoretical Static and Dynamic Maximum Deflection and Bending Stress

Condition

Maximum Deflection

(m)

Maximum Bending Stress

(MPa)

Static -4.7310-3 1.86

Dynamic 6.8510-4 1.55

DISCUSSION OF RESULTS OF FUSELAGE STRUCTURAL ANALYSIS

The results show that the maximum deflection should be negligible in either case. Similarly, the maximum bending stress calculated suggests that the bending stress is less that the tensile strength of Balsa wood which ranges between 7.6 MPa – 32.2 MPa .

This suggests that the design is structurally sound. However, these calculations assume steady flight under controlled environmental conditions, and as such, the actual values may be greater than the theoretical calculations indicate.

3. 3 VIBRATION ANALYSIS

The primary source of vibration is the motor; all other sources would be as a result of the environment and thus cannot be accounted for. For the sake of the analysis, the system is modeled as a motor balanced at the center of a simply supported beam with fixed ends. This model assumes that the plane is held in place and as such the ends of the fuselage are not free to move in the vertical direction.

(16)


22

Fig. 3.12: The model used for the vibration analysis

The model indicates that the system response is an undamped free vibration as a result of a harmonic force. The maximum magnitude of the applied harmonic force (Fo) is assumed to be the weight of the motor and propeller. As such, the maximum amplitude of the vibration (Y) can be determined as a function of maximum magnitude of the applied harmonic force (Fo), the material stiffness (k), the mass of the beam (m) and the angular velocity (ω) as shown in the equation below [8];

CALCULATION RESULTS

Table 3.5: Calculation Results and Relevant Measurements

Motor Mass w/ Propeller

(kg)

Maximum Load “F0”

(N)

Angular Velocity

(RPM)

Maximum Vibration Amplitude

(mm)

0.202 1.982 15200 4.0310-3

DISCUSSION OF RESULTS OF VIBRATIONAL ANALYSIS

The results show that the effect of vibration should be negligible and as such vibration should not play a major role in design considerations. This being said, further examination of vibration and its effects ceased. However, these calculations assume ideal conditions, and as such, the actual values may be greater than the theoretical calculations indicate. Testing and empirical observations after constructing the plane should aid in refining the design if vibration proves to be an issue.

3.4 STABILITY AND CONTROL

3.4.1 FEATURES THAT PROMOTE STABILITY

LONGITUDINAL STABILITY

Longitudinal stability, which is also referred to as pitch stability is dependent on maintaining the

moment about the center of gravity. Basically, the craft is said to be stable if the rate of change in

moments about the center of gravity (MCG) with respect to the lift (L) is negative. However, this rate is

best analyzed with respect to the discretized Moment (CMCG) and the lift coefficient [6];

2mk

FY o

(17)


23

𝑑𝐶𝑀𝐶𝐺

𝑑𝐶𝐿𝐼𝐹𝑇< 0

Where CMCG is defined as;

𝐶𝑀𝐶𝐺 = 𝑀𝐶𝐺

12

× 𝜌 × 𝑉2 × 𝐴 × 𝑙𝐶𝐺

Typically, pitch stability is dependent on the tail

configuration and the plane center of gravity [2],

however, the chosen design does not have a tail and as

such the stability is dependent on the center of gravity

alone. The center of gravity of the chosen design is

located at 10.4" from the tip of the plane as determined

by SolidWorks software package and the CAD model of

the designed craft.

As the Fig. 3.13 shows, the wing center of gravity is located towards the front of the plane, which is ideal

for maintaining stability as it will cause the craft to drop its nose if faced with an increasing gust of wind

or raise the nose in the opposite scenario. This action will counteract the effect of changing lift and thus,

return the craft to its static equilibrium position.

LATERAL (ROLL) & DIRECTIONAL (YAWING) STABILITY

The factors affecting roll and yaw stability are virtually the same and as such the two are addressed together. Here, stability is concerned with balancing the rolling moment (LB) and yawing moment (NB) resulting from the two wings. The moment generated by each wing can be determined using the following equation [6];

These particular aspects of the design are critical to lateral and directional stability;

I. The current design employs a low wing position which is laterally destabilizing. This could however be rectified with a dihedral should it prove to be an issue [2].

II. The current design implements a taper, resulting in the leading edge of the wing appearing to be swept. Wing sweep is laterally stabilizing as the wing toward the sideslip will experience higher velocities and thus more lift than the other wing, resulting in a correcting moment that returns the plane to the equilibrium condition [2].

0 CGM(18)

(20)

WINGLBB AVCL 2

2

1 (21)

WINGNBB AVCN 2

2

1 (22)

(19)

Fig 3.13: Wing Root Chord Length showing

distance to Center of Gravity


24

3.4.2 STABILITY ANALYSIS USING XFRL5 AERODYNAMIC STABILITY SOFTWARE

A study of the airplane’s aerodynamic stability was conducted using the XFLR5 software package. The

following root locus plots shown in Fig. 3.14 and Fig. 3.15 obtained using this software package.

LONGITUDINAL STABILITY

Fig. 3.15 shows that most the points are on the left hand side of the imaginary axis. This indicates that

the system is relatively longitudinally stable. The plot shows that any disturbances in the system will

depreciate with time; this is because the imaginary terms are located to the left of the imaginary axis.

There is one point on the real axis that is positive, this indicates that there is a small range of values for

which the flight characteristics are unstable, but, this point is close to the imaginary axis suggesting that

the craft can be considered to be longitudinally stable in all conditions.

LATERAL STABILITY

Fig. 3.14 shows that most the points are on the left hand side of the imaginary axis. This indicates that

the system is relatively laterally stable. The plot shows that any disturbances in the system will

depreciate with time; this is because the imaginary terms are located to the left of the imaginary axis.

There is one point on the real axis that is positive, this indicates that there is a small range of values for

which the flight characteristics are unstable, but, this point is close to the imaginary axis suggesting that

the craft can be considered to be laterally stable in all conditions.

3.4.3 AIRPLANE CONTROL

The current design will need to implement 3 channels for the purpose controlling flight. These are the channels for the two servomotors and throttle which are related to the elevons and propulsion system.

PROPULSION SYSTEM The main source of the thrust propelling the aircraft forward is the motor and its propeller. The motor for the system is restricted to being a Turnigy Aerodrive SK3 - 3542-800kv Brushless Outrunner Motor.

Fig 3.14: Lateral Stability Root Locus Plot Fig 3.15: Longitudinal Stability Root Locus Plot


25

PROPELLER SELECTION AND TESTING

The supplier indicates that the motor provided is equivalent to a 0.15 - 0.4 size gas powered motor, as such, the propeller size for testing was determined to be between 9" x 4" to an 11" x 6" size propeller. This range is based on values specified in Fig. 2.9. For the testing process 4 propeller sizes where chosen; 9”x 4.7”, 10”x 4.7”, 10”x 7SF”and 11”x 5.5".

The chosen propellers where tested using the apparatus shown in Fig. 3.14. It records the current drawn by the motor whilst in operation. The setup also allows the user to measure the thrust generated by the propeller. It does this by suspending the spinning propeller over a load cell and measuring the force of the displaced air pushing against it. The load cell gives the results in terms of a mass equivalent as shown in Fig. 3.16. RESULTS OF PROPELLER TESTING

Fig. 3.15: Chart showing Current Drawn against Power for each Propeller

Fig. 3.16: Chart showing Thrust against Power for each Propeller

0

0.5

1

1.5

2

2.5

0 5 10 15 20 25 30

Cu

rre

nt

(Am

ps)

Power (Watts)

Current Drawn Vs. Power Supplied

11X5.5

9X4.7

10X7SF

10X4.7

050

100150200250

0 5 10 15 20 25 30

We

igh

t (g

)

Power (Watts)

Thrust vs Power

11X5.5

9X4.7

10X7SF

10X4.7

Fig. 3.14: Apparatus Used Preliminary Propeller Testing


26

Fig. 3.15 shows that there is little difference in terms of current being drawn by the motor for each propeller. It also shows that the current being drawn by the motor will never reach 32A, the limit the speed controller can handle, at least given the power the battery can supply. As such there is no danger of burning out the speed controller. That being said, Fig. 3.16 suggests that the most thrust is generated whilst using the 10”x 4.7" and 11”x 5.5" propellers. However, the performance of the 11”x 5.5" propeller eliminated it from further consideration. As such, the 10”x 4.7" was selected for the final product.

CONTROL SURFACE (ELEVONS)

The current design utilizes a tapered elevon to perform both roll and pitch control. Elevons are flaps that act as both the ailerons the elevators, essentially, the elevons are the only control surface for the design being used. The elevons will work by causing the nose of the plane to go upwards when both elevons are lifted up and vice versa. This is the primary means of causing the plane to climb or dive. To turn, one elevon is tilted up and the other is tilted downwards. This will result in the craft rolling towards the desired direction. They will be controlled using two servomotors and a mixer to mix the channels that send signals to these servomotors such that the elevons function properly.

LANDING GEAR

The chosen design does not utilize landing gear as this would add weight and drag [2], which is contrary to the objectives outlined earlier in the report. The craft is expected to glide to a landing base is made flat to accommodate a safe landing.

4. TESTING AND EXPERIMENTAL RESULTS

4.1 WING DEFLECTION TESTING

Wing deflection testing was performed

by measuring the difference in elevation

at the wing root and the wing tip. The

initial experimental setup utilized a

measuring stick to measure this

difference, but, the deflections were so

small that a proximity sensor was used to

obtain more accurate results. Testing of

the wing as initially designed showed an

unacceptable amount of deflection, 2.54

cm. This was rectified by adding supports

as shown in Fig. 4.2. To estimate the

effects of lift on the wing, a 1 kg mass was

hung 8 inches from the wing root as the

model used in Fig. 3.9 suggests. The

deflection was then measured and the

deflection without the mass is subtracted

from this value to obtain the deflection due to a lift of 9.81 N. The results in Table 4.1 indicate that the

actual deflections are more significant than the theoretical analysis suggested, but, the deflection is still

relatively imperceptible and thus, still negligible.

Fig. 4.1: Wing Deflection Testing

Fig. 4.2: Wing Modification


27

Table 4.1: Results of Wing Deflection Testing

Initial Wing Design Modified Wing Wing with 1kg Mass

Deflection (cm) 2.54 0.0015 0.128

4.2 FUSELAGE DEFLECTION TESTING

Fuselage deflection was tested by

measuring the difference in elevation at

the fuselage ends and its center. The

initial experimental setup utilized a

measuring stick to measure this

difference, but, the deflections were so

small that a proximity sensor was used to

obtain more accurate results. To estimate

the effects of lift on the fuselage, a 1 kg

mass was hung 8 inches from the wing root as the model used in Fig. 3.11 suggests. The deflection was

then measured and the deflection without the mass is subtracted from this value to obtain the

deflection due to a lift of 9.81 N. The results in Table 4.2 indicate that the actual deflections are more

significant than the theoretical analysis suggested, but, the deflection is still relatively imperceptible and

thus, still negligible.

Table 4.2: Results of Fuselage Deflection Testing

Fuselage Fuselage with 1kg Mass

Deflection (cm) 0.0001 0.0027

4.3 DETERMINING CENTER OF GRAVITY (BALANCING TEST)

The center of gravity was determined by

balancing the plane on a rod and

measuring the distance from the nose of

the plane to the point at which the plane

is balanced.

The purpose of this test is as follows;

I. Determine Pitch stability

II. Determine Lateral stability

III. Determine positions for electronic components

The results show that the airplane has lateral stability as it balances at the center of its 60 inch wingspan

and the craft has pitch stability as it balances more towards the front of the plane as opposed to the

back of the plane. This supports the data from the XFLR5 software package in regards to stability. The

center of gravity is located at 10.6 inches from the tip of the plane’s nose when all the components are

Fig. 4.3: Fuselage Deflection Testing

Fig. 4.4: Balancing Test


28

added into the plane. This piece of information indicates that the payload bay should be positioned such

that the center of mass of the payload acts through the center of gravity of the plane.

Table 4.3: Results of Balancing Test

Plane Plane with Components

Distance from Plane Nose Tip (inches) 10.4 10.6

Deviation from the center (inches) 0 0

5. MANUFACTURING AND FABRICATION

Table 5.1: Fabrication Decisions

Part Fabrication Process

Wing Ribs - These were cut using the Laser cutter as precision is required. The ribs will be attached to the support spar and dowel spar using cyanoacrylate glue

Support Spar - Cut with the laser cutter and attached using cyanoacrylate glue. This was made of basswood for extra stiffness

Spar (Dowel) - Cut with the laser cutter and attached using cyanoacrylate glue

Covering - The covering was Monokote and was done by hand using heating iron and knife for trimming

Elevon - Cut with the laser cutter and attached to the wing using adhesive tape

Servomotors - Attached using 2 sided tape, and connected to the elevons using wire and an aluminium tube

Magnets – Magnets will be glued to the end of the wings and used to the attach the wings to the fuselage

Fuselage Airfoil - These were cut using the Laser cutter as precision is required. They were glued into position on the mounting space using cyanoacrylate glue

Mounting space - Can be cut with the laser cutter, but band saw will be just as effective and more time efficient. This will be made of basswood for additional stiffness

Nose - the nose of the plane is a foam block that consists of 1.5 inch segments that are glued together. Each segment is cut using the laser cutter

Covering - The covering is Monokote and was done by hand using heating iron and knife for trimming

Magnets – Magnets will be glued to the end of the fuselage and used to the attach the fuselage to the wings

Fuselage Case

Sides - These were cut using the Laser cutter as precision is required

Spars - These needed little precision and as such were cut with a band saw

Covering - The covering is Monokote and was done by hand using heating iron and knife for trimming

Mounts Battery & Speed Controller - Attached using Velcro

Motor - mounted in specially designed basswood mount

Payload - Mount was constructed using aluminium sheets. The sheets were bent into shape and holes were tapped into the side of the case to insert dowels to hold the payload in position

For drawings of each specific part refer to Technical Drawing Package in Appendix E.


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5.1 PROTOTYPING

To aid in fabrication decisions, a prototype (seen in Fig 5.1) of the design was created using corrugated cardboard and masking tape. This was done in order to get an idea of the possible issues that will be encountered in the process of creating the frame of the model. It also helped the team to gain an idea of what process would best be employed in order to fabricate each part. The Part Fabrication considerations are listed in Table 5.1

6. PERFORMANCE PREDICTION

Utilizing the information from the design analysis phase and Newtons force equation as well as data

from the SolidWorks 3D and XFLR5 software packages, the following performance specifications were

estimated;

Table 6.1: Predicted Performance Specifications

Plane Weight 854 g

Plane Airspeed 16.42 m/s

Maximum Lift Generated 57.04 N

Maximum Payload that can be lifted 4.96 Kg

Payload to Total Weight Fraction 0.85

Assembly Time 30 – 60 seconds

6.1 PREDICTED FLIGHT CHARACTERISTICS

The following are the expected flight characteristics based on research, the design features and the

results of the calculations and testing phase;

I. Stable lateral and directional flight conditions

II. Slight instability in terms of pitch stability

a. Care must be taken to maintain speed above 12 m/s

III. Plane should be self correcting in most conditions

IV. Time required to complete course should be 1.22 s

7. CONCLUSION

Upon completing the research phase of the project, it was decided that the best design route for

maximizing lift would be utilizing a flying wing or delta wing design. To minimize weight and still

maintain structural integrity, the rc airplane was constructed from balsa wood and basswood, with

basswood being used in areas that require greater stiffness. Upon deciding on a design direction,

vibration, deflection and bending stress tests performed both analytically and experimentally to refine

the design parameters. From these tests it was determined that the craft should be able to withstand

the effects of the loads that will be imposed on the craft. Further analysis proved that the craft should

Fig. 5.1: Prototype (See Appendix F)


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be able to lift a 4.96 kg weight at the maximum whilst weighing approximately 854g. This yields a

payload to total weight fraction of 0.85 which is fairly close to the 0.9 fraction set as a design objective.

8. RECOMMENDATIONS

If this project were to be undertaken again the following are areas that should be further examined for

the sake of improving the current design;

I. Examine the effect of using a swept wing type delta wing as opposed to a low taper ratio

II. Examine the effect of utilizing vertical stabilizers at the wing tips

III. Examine the effect of implementing a dihedral

IV. Perform a more in-depth study of airfoil profiles to determine if the Clark Y airfoil is the optimal

choice for the design

9. PROJECT MANAGEMENT

9.1 BUDGET DETAILS

Table 9.1: Items Purchased for Construction

Product Price Quantity Total ($) 36" 1/8" Dowel 0.49 4 2.21 3/16" Dowel 1.49 2 3.37 1/16 Balsa 2.99 3 10.14 3/32 Basswood 2.99 1 3.38 MonoKote RED 18.99 1 21.46 Batteries 1.00 6 6.78 Wood

26.51

Sand Paper 2.89 1 3.27 Industrial Velcro 5.99 1 6.77 Magnetic Strips 1.69 1 1.91 Pink Insulation 8.00 1 9.04 Box Knife 1.99 1 2.25 Scissors 3.99 1 4.51 3/16" Dowel 1.49 3 5.05 Balsa 3.69 2 8.34 Basswood 2.99 2 6.76 1/8" Basswood 3.29 2 7.44 3/32 Basswood 3.49 3 11.83 Strip AIL Horn Wires W/B 2.99 1 3.38 Medium CA Glue 2 Oz 13.99 1 15.81 3/32 Balsa 3.49 2 7.89 MonoKote Red 18.99 1 21.46 SUB-TOTAL

189.54


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Table 9.2: Items Gifted to the Team

Product Price Quantity Total ($) Magnetic Clasps 2.00 5 11.30

Payload Material

(Aluminium)

5.00

SUB-TOTAL

16.30

Table 9.3: Common Parts Provided to the Team

Product Price Quantity Total ($)

TGY-D1290P High Speed Micro Servo 6.57 1 7.42

3542-800kv Brushless Outrunner Motor 32.81 1 37.08

35A Fixed Wing Brushless Speed

Controller 12.90 1 14.58

Turnigy 2200mAh 3S 20C Lipo Pack 7.89 1 8.92

Turnigy 2S 3S Balance Charger. Direct

110/240v Input 11.44 1 12.93

Hobby King 2.4Ghz 4Ch Tx & Rx V2

(Mode 2) 22.99 1 25.98

SUB-TOTAL

106.90

Table 9.4: Expenses at Each Phase

Prototyping $11.30

Testing $15.00

Fabrication $220.85

Repairs and Modifications $65.58

TOTAL $312.73


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9.2 PROJECT MANAGEMENT PLAN

Devising a project management plan began by gathering the project information and defining the project requirements. This was completed in compliance with the 2012 SAE Aero Design competition rules as defined by SAE International. The goal is to administer a series of chronological tasks whilst maintaining a record of each task to ensure successful completion and compliance. This has been achieved by the use and implementation of a Gantt chart as shown on the next page. All associated tasks are addressed in the Gantt chart whose structure will change over the course of the project until completion. A Gantt chart was used throughout the scope of this project in order to keep track of activities, tasks and

the persons responsible with completing them. The Gantt chart shows the activities in sequential order

according to completion dates and the order of linearity in project organization. Each activity was logged

in the log books and scheduled on the Gantt chart and this integration of two systematic scheduling

tools ensured that tasks were completed on time.

Although a critical path method was suggested in the first report, the project team decided that this was

not necessary because the Gantt chart was sufficient to monitor whether the project was on schedule

and also to ensure that all tasks were completed and executed within the given time frame.


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34


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10. REFERENCES

[1] Glenn Research Center, What is Lift? 2010. Web. May 16th 2012

<<http://www.grc.nasa.gov/WWW/K-12/airplane/lift1.html>>

[2] U.S. Centennial of Flight Commission, Theories of Flight, 2011. Web. May 15th 2012 <<http://www.centennialofflight.gov/essay/Theories_of_Flight/Stability/TH26.htm>>

[3] Peter Carpenter, RC Airplane World, 2011. Web. May 15th 2012

<<http://www.rc-airplane-world.com>>

[4] Glenn Research Center, What is Thrust? 2010. Web. May 16th 2012

<<http://www.grc.nasa.gov/WWW/K-12/airplane/thrust1.html>>

[5] Buczynski James. Propeller Selection Guide. Johnson County, N.Y.: Academic Press, 2007, pp. 252-270

[6] Megson, T.H.G. An Introduction to Aircraft Structural Analysis. Kidlington Oxford: Elsevier Ltd, 2007. Print

[7] Hibbeler, R.C. Mechanics of Materials. New York: Prentice Hall, 2004. Print

[8] Rao, S. Mechanical Vibrations. New York: Prentice Hall, 2010. Print

http://www.grc.nasa.gov/WWW/K-12/airplane/lift1.html

http://www.centennialofflight.gov/essay/Evolution_of_Technology/flying_wing/Tech8.htm

http://www.rc-airplane-world.com/

http://www.grc.nasa.gov/WWW/K-12/airplane/thrust1.html


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APPENDICES


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APPENDIX A COMMON PARTS PROVIDED


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APPENDIX B MATERIALS RESEARCH


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APPENDIX C DISCUSSION OF CONCEPTS


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FLYING WING CONCEPT


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FLYING WING CONCEPT DISCUSSION

The concept has the following characteristics;

Design Optional Add-on : 9” long Tail to address stability issues if testing highlights any;

Large wing span and large wing area to maximize lift;

Low wing loading;

Structurally stable design. There is a current debate between proceeding with flat design or

utilize an airfoil. Debate to be resolved in initial testing phase if this is a chosen concept;

Simple design, to aid construction and repairs;

o Provides ample space for modifications to the current design to be made to address any

issues discovered through testing;

o The design provides the potential to test an alternative design as well. DESIGN CONCEPT

# 3 has features that are interchangeable with the FLYING WING CONCEPT and as such

both concepts can be tested to determine which yields the best result;

Parts are easily detachable and replaceable in the event that accidents occur during testing.

Design consists of three parts; left wing, right wing and the component that is effectively a

fuselage;

Ample space for mounting equipment and payload case;

High aspect ratio;

Zero taper ratio;

Fuselage is incorporated into the wing and there are no stabilizing surfaces. This reduces the

drag experienced by the craft but makes the flight characteristics, particularly turning, more

unstable. It also reduces the overall weight of the craft;

Uses ailerons/elevators as one and the same flap;

Propeller is positioned in the back;

o Requires motor to be wired in reverse;

o There is the option is mount a vertical stabilizer over the propeller;

o There is the option to add diffusers to direct airflow and thus stabilize flight conditions

further;

No landing gear, as it would add more drag. Competition will be occurring in the grass, as such,

landing gear might proved problematic. Plane is designed to glide to a landing;


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CONCEPT #2


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DESIGN CONCEPT # 2 DISCUSSSION

The concept has the following characteristics;

Large wing span and large wing area to maximize lift;

Low wing loading;

Structurally stable design. There is a current debate between proceeding with flat design or

utilize an airfoil. Debate to be resolved in initial testing phase if this is a chosen concept;

Simple design, to aid construction and repairs;

o Provides ample space for modifications to the current design to be made to address any

issues discovered through testing;

o There is space in the front to add additional lifting surfaces;

o The design provides the potential to test an alternative design as well. The FLYING WING

CONCEPT has features that are interchangeable with DESIGN CONCEPT # 3 and as such

both concepts can be tested to determine which yields the best result;

Parts are easily detachable and replaceable in the event that accidents occur during testing.

Design consists of three parts; left wing, right wing and the component that is effectively a

fuselage;

Ample space for mounting equipment and payload case;

High aspect ratio;

Zero taper ratio;

The shape of the craft is a hybrid of the conventional airplane design and a flying wing design, as

such the effect of drag on the body should be increased in comparison to that of a flying wing;

Fuselage is incorporated into the wing, but there are vertical and horizontal stabilizing surfaces.

This adds to the drag experienced by the craft but makes the flight characteristics, particularly

turning, more stable. It also increases the overall weight of the craft, but this increase may not

be significant;

Potential for separate ailerons and elevators or for both to be one and the same flap;

Propeller is positioned in the back;

o Requires motor to be wired in reverse;

o There is the option is mount a vertical stabilizer over the propeller;

o There is the option to add diffusers to direct airflow and thus stabilize flight conditions

further;

No landing gear, as it would add more drag. Competition will be occurring in the grass, as such,

landing gear might proved problematic. Plane is designed to glide to a landing;


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Equipment Mounting Platform (Located

Inside plane)

DESIGN CONCEPT #3 (HELI-PLANE)


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DESIGN CONCEPT #3 (HELI-PLANE)

Aside from the conventional airplane, with the propeller located at the front of the craft, another design could incorporate propellers on top of the aircraft, or the bottom (hovercraft), tilted to provide both an upward and forward thrust (helicopter). There have been crafts, as well, that incorporate the forward propellers with an upward propeller on the craft's wings, providing both benefits of an airplane and a helicopter. These designs, although not as fast as a plane, offer increased stability and control. As an aide for control, similar to the insect counterpart - the dragonfly - helicopters utilise an extended tail designed for yaw correction, and is demonstrated in a study investigating the abdomen posture of locusts under varying winds (Camhi 1970). From the study, it is observed that abdomen (tail) flexes downwards under low wind velocities resulting in increased lift, as shown in the figure below. As wind speeds increase, the tail levels horizontally. This indicates that wind, and therefore thrust, has a dominant affect on the required lift for the helicopter. The difference between the helicopter and the dragonfly, besides being mechanical, is the flexing tail. To account for this; however, employed on the tail boom is a rotor.

However, because it is stated in the SAE International rules that rotating wings, such as in an helicopter, are prohibited. Although the helicopter is prohibited, at the very least, the control mechanism (gyroscopic assists are allowed) located on the tailboom of a helicopter may be utilised in the final design.

Camhi, J., 1970. Sensory Control of Adbdomen Posture in Flying Locusts. Journal of Exp. Biol. Volume 52, pages 533-537. Section of Neurbiology and Behaviour, Cornell University, Ithaca, N.Y. 14850.


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CONCEPT #4


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CONCEPT DESIGN #4 DISCUSSION

This airplane model would involve two wing profiles that can be activated depending on operating conditions and flight requirements. A variable geometry wing profile would be employed, which would include a straight wing which can then convert to a swept back wing profile. This is made possible by actuators or manually adjusting the wings before take off. When in straight wing profile mode, the plane operates in the most efficient structural mode and this is most common for low speed designs. A swept wing profile reduces the lift coefficient for any given angle of attack and this is noticeable at low speeds and results in longer takeoff and landing distances. Lift coefficient rlates the lift generated when lifting a body, the dynamic pressure of airflow over and around the airplane and the localized wing area associated with the body. The wings would include right and left ailerons to control roll, which results in a change in planform heading due to the tilting of the lift vector. The wings would have an airfoil and the the fuselage would take the same profile as normal airplanes. A tail would be attached with a rudder, included primarily to counter adverse yaw and the p-factor which is an aerodynamic phenomenon experienced by a moving propeller. The propeller would be attached at the front of the fuselage thus generating thrust. The benefits for wing sweep are that it provides lateral stability and reduces wave drag on the aircraft. The inherent disadvantages to this design are that when a swept wing travels at lower speeds, for which the designed plane would operate in, the airflow is pushed spanwise by the angled leading edge, towards the wing tip. The main stability problem is the tendency to flat spin when thermalling (rising air due to heating which occurs abruptly) or steep turning at low speeds. Severe sweep angles, however, adversely affect spanwise flow and wing efficiency, and make launching via hand more difficult.


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The complicated flight dynamics include dutch roll shown above and below, while not inherently dangerous, it produces a tremendous amount of drag and is caused by insufficient vertical stabilizer area, insufficient directional stability and too much spiral stability and spiral divergence as shown below. The plane would be thrown by hand to ensure takeoff and when considering the adverse flight dynamics that the plane would not only need to "intuit" but also overcome and this requires controls and applications that are not within the defined rules and competition context as ascribed by SAE International. Therefore a plane which employs this profile would be feasible only if it engaged in manned flight.


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APPENDIX D THEORETICAL LIFT AND DRAG FOR VARIOUS AIRSPEEDS


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V (m/s) Lift (N) Drag (N)

1 0.21161 0.013755

2 0.846442 0.055019

3 1.904494 0.123792

4 3.385767 0.220075

5 5.290261 0.343867

6 7.617976 0.495168

7 10.36891 0.673979

8 13.54307 0.880299

9 17.14045 1.114129

10 21.16104 1.375468

11 25.60486 1.664316

12 30.4719 1.980674

13 35.76216 2.324541

14 41.47565 2.695917

15 47.61235 3.094803

16 54.17227 3.521198

17 61.15542 3.975102

18 68.56178 4.456516

19 76.39137 4.965439

20 84.64418 5.501871


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APPENDIX E TECHNICAL DRAWING PACKAGE


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BILL OF MATERIALS

ITEM NO. PART NAME DESCRIPTION MATERIAL WEIGHT (lbs) QTY. 1 Fuselage - Bottom Bottem of Fuselage Basswood 0.1760733 1

2 Fuselage - Body Sides of Fuselage Body (Attached to

Bottom) Balsa 0.0208257 1

3 Fuselage - Nose Front of Fuselage Celfort Extruded Polystyrene 0.0577354 1 4 Fuselage - Top Frame for Removable Plane Cover Balsa 0.0108469 1 5 Tail Tail Frame Balsa 0.0063313 1 6 Reinforce Wing Reinforcment Basswood 0.0320321 2 7 Elevon Elevator/Aileron Balsa 0.0153459 2 8 Dowel Wing Support Beech 0.0211373 2 9 19.2 19.2" Clark Y Foil Basswood 0.0420362 2

10 12 12" Clark Y Foil Basswood 0.0108855 2 11 10 10" Clark Y Foil Balsa 0.0020003 2 12 8 8" Clark Y Foil Balsa 0.0013066 2 13 6 6" Clark Y Foil Balsa 0.0007421 2 14 4 4" Clark Y Foil Balsa 0.0003373 2

15 High Speed Micro

Servomotor Elevon Control

0.063934 2

16 Brushless Outrunner Motor Propulsion System 0.3130564 1

17 Fixed Wing Brushless Speed

Controller 0.0705479 1

18 Turnigy 2200mAh 3S 20C

Lipo Pack Battery Pack

0.407855 1

19 Propeller 0.1366866 1 20 Magnets Used to Attach Wing to Fuselage 0.0061729 20


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APPENDIX F PROTOTYPE CREATION


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rc airplane designii executive summary this report outlines the project scope and definition of a...

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