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Complete Design Review

Project 5008

Christina Alzona

Benjamin Wagner (team leader)2/18/05

Project 5008Page 1of47

1.0Introduction

1.1Design Development

1.2Design Alternatives

2.0Conceptual Design

2.1Needs Assessment

2.2Feasibility Assessment

2.2.1 Aircraft Type

2.2.2 Empennage

2.2.3 Landing Gear

2.2.4 Propulsion System

2.2.5 Wings

2.2.6 Radio and Transmitter

2.2.7 Building Materials and Construction Methods

2.3Conclusions

3.0Preliminary Design

3.1Construction Methods

3.1.1 Wing

3.1.2 Empennage

3.1.3 Fuselage

3.1.4 Takeoff Mechanism

3.2Analysis Methods and Sizing

3.2.1 Aerodynamics

3.2.2 Structures

3.2.3 Payload

3.2.4 Propulsion

4.0Prototype Design

4.1Construction Methods

4.1.1 Wing

4.1.2 Empennage

4.1.3 Fuselage


4.2Aircraft Configuration


4.2.2 Weight Analysis

4.3Predicted Performance

5.0Analysis and Design Testing

5.1Wing Failure Analysis

5.2Aircraft Load Testing

5.3Center of Gravity Calculations

6.0 Bill of Materials

7.0 Time Line

7.1Fall 2004 Timeline

7.2Winter 2004 Timeline

7.2.1 Predicted Winter 2004 Timeline

7.2.2 Actual Winter 2004 Timeline

8.0 Acknowledgements

9.0 References

Appendix A: Construction Pictures

Appendix B: CAD Drawings

Appendix C: Finite Element Analysis

Appendix D: Electric Motor Calculations

Appendix E: Airfoil Data


1.0 IntroductionRochester Institute of Technology College of Imaging Arts and Sciences had

expressed an interest in the creation of an unmanned airborne sensing platform

to assist their Wildfire Airborne Sensor Program (WASP). In an effort to assist

CIAS, two design teams were assembled. One will design and build the body of

the UAV, and another team will design and integrate onboard telemetry and

stability augmentation. This UAV approach is scheduled to be an ongoing

project in coming years. Current mission objectives include flight ranges of at

least two miles from base station and an endurance of at least one hour. CIAS

also requested a fairly large payload capacity and ease of repeatability in design

and construction. CIAS has requested a UAV capable of a 3.2 km range, able to

carry a 1.5 kg payload, and adaptable to various unspecified mission

requirements.

1.1 Design Development

During the conceptual design phase, aircraft size approximations and

feasibility assessment were developed from the needs of our sponsors in

the Mechanical Engineering department and CIAS as RIT. Using the

initial size approximations, feasibility assessment and the needs of our

sponsors, the team prioritized the most critical elements of the project for

further consideration. These elements were then analyzed based on

aerodynamics, propulsion and structural integrity of the aircraft. These

categories were further divided into building materials, aerodynamic

designs, structural designs, and construction methods. In this phase,

numerous airplane configurations were considered using our design

parameters and compared against our feasibility assessment of the

different aspects of the aircraft design. The models that most closely

achieved the desired design criteria were later utilized in the preliminary

design phase for closer examination.


In the preliminary design phase, a more detailed analysis and theoretical

performance calculations were performed. Flight characteristics were

predicted and loads were analyzed for the optimal design. Payload

volumes and weights were dynamic throughout the preliminary design

phase making the fuselage difficult to conceptualize, but the final fuselage

design should meet the specifications that were given as of this

preliminary design review. Construction methods for the entire airborne

platform had to take into account the limited human labor that will be

available with the current team. It is the recommendation of the team that

more members be allocated to the construction of the airborne platform to

ensure the completion of the prototype in the time constraints imposed.

1.2 Design Alternatives

Numerous propulsion, aerodynamic, and structural configurations were

compared and analyzed. Due to existing molds and past experience the

airfoil for the main wing was chosen to be an Eppler 423. Main wing

locations that were considered were high, low, and mid fuselage mounts.

A high wing location was chosen because of increased stability over other

the designs and it would allow the integrating teams to access hardware

more easily within the fuselage. Empennage styles that were considered

were conventional, T-tail, cruciform, and canard. The conventional tail

was chosen for ease of constructability and to keep weight down.

Propulsion options that were analyzed were electric, hybrid, and

nitromethane glow engine. Electric was chosen to meet the vibrations

specifications of the needs assessment, it produces no emissions to affect

the integrity of the video equipment onboard, and it increases stability

because as fuel is consumed the weight of the airborne platform does not

shift. Motor location was chosen to be a front mount. The front mounted

position allowed more air to pass over the wing to aid lift during take-off

and kept weight down as other options would have required reinforcing the


airframe. Lithium polymer batteries were chosen as a fuel source

because currently it offers the highest energy density per unit of mass of

any commercially available batteries. The main drawback of lithium

polymer is the cost. A fiberglass and epoxy composite was chosen for

shell of the fuselage. It provides the strength necessary for the chosen

geometric configuration but is less stiff than other options to better achieve

the vibrations requirement.

2.0 Conceptual DesignThe conceptual design process was separated into two main tasks. The

first task was to determine the needs assessment to evaluate the minimum

requirements of our sponsor and define the goals of the design team. Secondly,

a feasibility assessment was performed on all the main aspects of the airborne

platform design. With the results of the feasibility assessment, a preliminary

design incorporating the winning designs from each aspect was created.

2.1 Needs Assessment

This design partnership needed to take into consideration that it would be

integrating the airborne platform design with at least two other design

teams in the winter and spring quarters. These two other projects that will

be involved in adding sensitive and expensive equipment of a volume and

weight that was yet to be determined as of this preliminary design review.

The design of the airborne platform had to incorporate these two black

boxes comfortably and allow their components to be accessed easily. The

airborne platform also had to provide some measure of impact protection

of its more expensive components in the unfortunate event of a crash.

In addition to the above concerns, CIAS has specified that the airborne

sensing platform must meet or exceed these specifications:


Carries a 3 lb payload (CIAS sensing equipment)

Has a cruise speed of 15-30MPH

Has a 1 hour endurance

Provides view angles both upward and downward for sensors

Can climb to 1,000 ft

2.2 Feasibility Assessment

With the needs assessment completed. The team split the design into

seven main categories to brainstorm ideas. Once the team exhausted

itself of the best options for each category, a feasibility assessment was

used to rate them against one another. The best design from each

category was then incorporated into the preliminary design.

2.2.1 Aircraft Type

Several aircraft configurations were analyzed during the conceptual

design phase of this project. These configurations were conventional,

flying wing, canard, and biplane as illustrated in figure 1.

Figure 1 Aircraft Type

Aircraft configurations were chosen based on general aircraft knowledge

and experience. The conventional configuration was chosen because of

stability and its ability to contain the unknown volumes of the payloads

being carried. Although a flying wing design may have been more efficient


to reduce drag, the unknown nature of the payload made it difficult to

design a volume to contain it. The canard design was dismissed because

the downwash of the horizontal stabilizer of this design has been proven

to disrupt the lift distribution on the wing, thereby increasing the induced

drag and shed vorticity.

2.2.2 Empennage

Several empennage configurations were considered for the design of this

airborne platform. A conventional style tail design was chosen because of

ease of constructability and weight concerns. A T-tail and cruciform were

considered because they would keep the horizontal stabilizer out of the

downwash of the wing, but it was thought the airborne platform would be

moving too slow for this advantage to be noticed considerably. The T-tail

and cruciform designs would also require reinforcement of the vertical

stabilizer and increased the weight of the airplane. The V-tail would

decrease weight, but it was discounted because it would require using a

radio transmitter capable of mixing control surface functions. V-tails also

produce a counteracting lift which would have decreased the lift

distribution of the airborne platform. A visual representation of the

empennage choices are found in Figure 2 and the feasibility analysis is

found in Figure 3.

Figure 2 Empennage Configurations


Tail Design Feasibility Assessment(On a scale from 1-10) Weighting Low T tail Cruciform

R1: Sufficient Skills 0.1 9 7 6R2: Sufficient Equipment 0.1 9 9 9R3: Sufficient # of people 0.13 2 2 2E1: Economically Feasible 0.07 8 7 7S1: Meeting Intermediate Milestones 0.1 8 8 8S2: Meeting PDR Requirements 0.1 8 7 7S3: Meeting CDR Requirements 0.15 8 7 6T1: Has similar technology been used before 0.12 8 7 6T2: Plane stability 0.08 6 8 7T3: Drag reducing 0.05 7 7 7

Total: 1 7.21 6.73 6.28Figure 3: Tail design feasibility

2.2.3 Landing Gear

The landing gear configurations that were considered were tail-dragger,

tricycle gear, and no landing gear. A tail-dragger configuration was initially

considered optimal for this airborne platform because it provided an

decent trade-off between aerodynamic drag and ground stability. While a

tricycle landing gear would have been the most stable in ground handling

characteristics, the large cross section of the nose gear would have added

and unacceptable amount of drag to the flight characteristic predictions.

No landing gear was also considered and was eventually chosen. By

choosing no landing gear, the design would be save weight and be more

aerodynamic. Onboard fuel would also be conserved because an external

source of power would have to be used to launch the aircraft. A method of

skids or other devices would be need, however, to protect the fuselage


and other components against a rough belly landing. A diagram of the

feasibility analysis is found in Figure 4.

Landing Gear(On a scale from 1-

10) Weighting None TrikeTail-

draggerAerodynamics 0.2 9 7 8Ground Handling 0.2 0 9 6Weight 0.3 10 6 8Ease of Construction 0.3 10 6 7

Total: 1 7.8 6.8 7.3Figure 4: Landing gear feasibility analysis


Several propulsion systems were examined. These included

nitromethane glow engines, electric motors, and a hybrid of the previous

two options. An all electric configuration was chosen because it would

produce the least vibrations of all options. It is also the cleanest and does

not produce any exhaust that may affect the onboard sensory payload.

The main disadvantages of electric power are the added weight of

batteries and electric motors generally have a lower energy per unit mass

than the glow engines. Glow engines have an advantage of a outputting

more power per weight, but create an unacceptable level of vibrations.

Glow engines also create an oily exhaust that may stick to the aircraft and

could affect the sensing ability of the payload. As fuel is consumed with a

glow engine, the center of gravity of the aircraft will shift changing the

stability of the aircraft in mid-flight. A hybrid would combine some

advantages of the previous options. The power available advantage of

the glow engine would have been utilized in takeoff, climbing, and getting

the airborne platform to the loiter zone of the flight mission. In the loiter

zone and the return flight to the landing zone, the electric motor would be

used to decrease the vibration level for the CIAS sensory equipment. The


hybrid option would require reinforcements throughout the fuselage to

support both power plants, would still create exhaust that could interfere

with sensory equipment, and would have the stability issues associated

with mid-flight center of gravity shifts. Analysis of the propulsion feasibility

can be found in figure 5.

Propulsion(On a scale from 1-

10) Weighting Electric Hybrid GasVibration 0.3 9 7 5Weight 0.2 8 9 7Stability 0.2 9 8 7Cost 0.1 7 7 7Ease of Installation 0.2 9 7 8

Total: 1 8.6 7.6 6.6Figure 5: Propulsion feasibility analysis

An issue that developed when an all electric configuration was chosen

was the option of battery type. After researching what was commercially

available at the time of this preliminary design review, the options that

were available were nickel metal hydride, nickel cadmium, and lithium

polymer. Lithium polymer was chosen because its energy density is much

higher than the other two options. Also, the discharge curve of lithium

polymer, voltage over time, is largely flat until the battery is completely

discharged. Battery weight and volume were the main concerns

considering the largely unknown black box payload. The main

disadvantages of lithium polymer batteries are the cost and they have a

slower charge rate than other options. Nickel metal hydride and nickel

cadmium were much cheaper but weight much more than the lithium


polymer option. The voltage supplied by these batteries also steadily

declined during the discharge cycle. An analysis of the battery feasibility

is shown in figure 6.

Batteries(On a scale from 1-10) Weighting LiPoly Nimh NiCad

Energy Density (charge/mass) 0.4 9 7 8Cost 0.1 6 8 7Ease of use/installation 0.2 8 8 8Recharge Ratio 0.3 7 7 7

Total: 1 7.9 7.3 7.6Figure 6: Battery feasibility analysis

2.2.5 Wings

Due to the small size of this design team and work already completed by

the RIT Aerodesign Team in their design of a heavy lifting RC aircraft in

2002, an Eppler 423 airfoil was decided upon. Molds already exist for a

wing that is suitable for the design requirements of this project.

Construction methods used will be similar to those used by the RIT

Aerodesign Team when they originally constructed wings from these

molds. More detailed information on the Eppler 423 airfoil can be found in

Appendix D.

Wing locations that were considered were high, low, and mid fuselage. A

high wing was chosen because it is the most stable out of the three

options. A low wing location would have required less reinforcement, but

this was considered less important that stability. A mid-wing configuration

is more stable than the low wing and would help minimize moments at the

center of gravity, but it would have increased weight through fuselage

reinforcements. A feasibility analysis of the wing location can be found in

figure 7.


Wing Location Feasibility Assessment(On a scale from 1-10) Weighting High Low Mid

R1: Sufficient Skills 0.1 8 8 6R2: Sufficient Equipment 0.1 9 9 9R3: Sufficient # of people 0.13 2 2 2E1: Economically Feasible 0.07 9 9 9S1: Meeting Intermediate Milestones 0.1 8 8 8S2: Meeting PDR Requirements 0.1 7 8 6S3: Meeting CDR Requirements 0.15 7 7 7T1: Has similar technology been used before 0.12 9 8 8T2: Plane stability 0.08 9 7 8T3: Drag reducing 0.05 8 8 8

1 7.34 7.16 6.84

Figure 7: Wing Location feasibility analysis

Wing to fuselage attachment was also analyzed. Possible methods

included a two piece wing joined to the fuselage through tube and pins,

bolting the wing directly to the wing saddle with screws, or sandwiching

the wing between the fuselage and a plate bolted to the fuselage.

Sandwiching the wing between the fuselage and a plate was eventually

chosen. This would maintain the structural integrity of the wing and

prevent any stress concentrations from forming on the wing because the

stress would be spread out over the area of the wing plate. A two piece

wing would have the advantage of being more portable, but the fuselage

and wing spar reinforcements in the area they would be joined would have

added weight to the final design. Bolting a one piece wing directly to the

fuselage would weigh the least of any considered designs, but it would

degrade the integrity of the wing and create stress concentrations at the

screw holes that would need to be reinforced.


2.2.6 Radio and TransmitterA radio and transmitter will be supplied by the RIT Mechanical

Engineering Department for this phase of the project. This team has

been told that the department traditionally uses Futaba brand radio

equipment and a radio kit will be available. This consists of a transmitter

with at least 4 channels and mixing controls for the ailerons. A standard

Futaba receiver has capabilities for 7 channels, a battery to power servos

and the receiver. The batteries generally operate at 4.8V and have a

capacity around 600 mAh. The design team intends to use micro-servos

which can run in the 26 oz/in torque range, which should be more than

enough to move the control surfaces of an airplane of this type.

2.2.7 Building Materials and Construction Methods

Building material options were also explored for construction of the

airborne platform. Because of the high strength low weight requirements

of this project carbon fiber, fiberglass, and balsa/ply were analyzed.

Carbon fiber has superior strength to weight, however, it is the stiffest

material examined. Fiberglass has the benefit of being less stiff than

carbon and would probably be able to withstand minor crashes better, but

because of the strength to weight ratio it was not used. Balsa also has a

good strength to weight ratio and provides a decent amount of

crashworthiness, but it would be difficult to create the geometries required

in some of the designs. A breakdown of the feasibility of building

materials can be found in figure 8.


Construction Material Feasibility Assessment

(On a scale from 1-10) Weighting Carbon Fiberglass WoodR1: Sufficient Skills 0.1 6 6 7R2: Sufficient Equipment 0.1 9 9 8R3: Sufficient # of people 0.13 2 2 3E1: Economically Feasible 0.07 9 9 7S1: Meeting Intermediate Milestones 0.1 8 8 8S2: Meeting PDR Requirements 0.1 8 8 8S3: Meeting CDR Requirements 0.15 7 7 7T1: Has similar technology been used before 0.09 9 8 7T2: Durability 0.08 9 9 6T3: Weight 0.08 7 8 9

Total: 1 7.13 7.12 6.86Figure 8: Construction material feasibility assessment

Because of the material properties and our needs assessment a variety of

materials were chosen for the construction. Carbon was decided for the

fuselage and cowling construction. The wing will be a balsa rib, carbon

spar, and fiberglass skin construction. The empennage will be built up

from a balsa frame. The landing gear frame will be a carbon laminate.

Carbon and fiberglass would require the construction of a mold. This has

both advantages and disadvantages in the construction process. The

main disadvantage is this design team is too small to create both fuselage

molds and construct an aircraft; the team will need additional members to

complete the project. An advantage creating molds for parts is the ease of

repeatability and conformity of constructing additional airplanes.


2.3 Conclusions

From the conceptual design phase and feasibility assessment it was

determined that the main focus for the design team would be finding

suitable propulsion system and designing a fuselage that can meet any

other future requirements of our sponsor. Given the small size of this

senior design team it was also determined that in order to complete this

project it would require the addition of more members during the

construction phase of this project.

3.0 Preliminary DesignWith completion of the feasibility assessment, a preliminary design was

created using the best methods. The preliminary design section is

separated into construction methods, component sizing analysis, and

predicted performance. A construction method was included to assist any

new members this design team may acquire to help construct the

prototype.

3.1 Construction Methods

3.1.1 Wing

Wing construction methods will be the same as those used by the RIT

Aerodesign team in their construction of similar wings. The wing skins will

be vacuum molded. The molds available include a top half and bottom

half that are imprinted with the camber of the airfoil. The halves of the

wing are molded separately and joined later. The molds must be sanded

completely smooth and treated with an anti-sticking compound called

Partol #10 before use. A film of epoxy is put in the mold followed by a

layer of fiberglass. Another layer of epoxy is added followed by a layer of

1/32” balsa wood. The balsa prevents the vacuum bag from molding

creases into the fiberglass layer. A sheet of plastic is glued to the outside

edges of the inner mold surface with Liquid Nails or suitable substitute and


a tube connected to a pump is inserted into the bag and sealed. The

pump evacuates the air from the bag until the epoxy has cured. The

pressure of the atmosphere should be enough squeeze any excess epoxy

from the wing skins and allow a uniform consistency between all layers of

the composite. The wing skins are now molded into the final shape of the

airfoil. A finished cross section is shown below in Figure 9.

Figure 9: Finished wing cross section (courtesy of RIT Aerodesign Team)

A rib and spar frame will be constructed for the wing. Wing frames shall be

constructed in halves. A spar will be created using a 7/8” x 5/8” piece of

hex cell with a layer of unidirectional carbon laminated with epoxy to the

top and bottom. The carbon fibers will run the length of the main wing

spar. This wing spar with be wrapped with a single layer of ¾ oz

fiberglass and epoxy to decrease the possibility of delaminating. Wing

ribs will be constructed of 1/16” balsa sheeting with the grain of the balsa

running the length of the ribs. Ribs will be constructed of one piece glued

aft of the wing spar. The rib at the root of each wing half shall be 1/8”

thick balsa and in one piece. Rib spacing will be 6” apart. A trailing edge

and leading edge of molded foam will then be glued to the spar frame.

The wing skins will be glued one at a time with epoxy to the spars and ribs

structure and allowed to cure before the next skin is attached.

The wings will have a 2 degree dihedral angle. Wings will be joined at this

angle using epoxy. Wings will be reinforced with an eight inch wide strip

of fiberglass glued with epoxy around the center wing.


Ailerons will be formed from solid balsa and conform to the shape of the

airfoil. They will be attached using standard fiber hinges with epoxy to the

wing. A micro servo will be mounted to a rib near the center of the aileron.

Extended wiring for the servo will run through the center of the wing.

3.1.2 Empennage

The vertical stabilizer shall be constructed of 3/16” x ½” sticks of balsa in a

truss. This balsa shall be glued together using a CA style glue. The front

and top edges of the vertical stabilizer will be sanded round. The vertical

stabilizer shall be covered with Monokote, a thin polyester film with a heat

activated adhesive on one side.

The rudder will be constructed of 3/16” x ½” balsa sticks in a truss. The

balsa shall be glued together using a CA style glue. The front of the

rudder will be sanded to a point at a 45 degree angle. The top, bottom,

and rear of the rudder will be sanded round. The rudder will be covered

with Monokote.

The horizontal stabilizer shall be constructed of a rib and spar frame. Ribs

shall be cut from 1/16” balsa with the grain of the wood running the

lengthwise of the ribs. Ribs will be glued using a CA style glue to a ¼”

spruce spar. Ribs will be spaced 2” apart. The center 3” section of the

horizontal stabilizer will be a solid balsa construction. This solid balsa

portion will allow a large surface to glue to the fuselage with epoxy.

The elevator will be constructed of 3/16” x ½” balsa sticks in a truss. The

balsa will be glued together using a CA style adhesive. The front surface

of the elevator will be sanded to a point at a 45 degree angle. The sides

and rear of the elevator will be sanded round. The elevator will be

covered with Monokote.


A strip of Monokote slightly smaller than the width and length of the

mounting surface of the fuselage will be cut from the bottom of the

horizontal stabilizer. The horizontal stabilizer will be glued with epoxy onto

the fuselage tail section. The vertical stabilizer will then be glued into

place on top of the horizontal stabilizer with epoxy with care being taken to

ensure it is at a 90 degree angle with the horizontal stabilizer. Control

surfaces will be attached to the stabilizers using fabric hinges glued with

epoxy into notches cut into the surfaces to be mated together. Mounting

locations can be found in Appendix A.

3.1.3 Fuselage

Fuselage sections will be vacuum formed in molds. Molds will be created

with particle board and modeling compound in the shape of the finished

product. The surface of the mold will be sanded completely smooth

before a part is created. The sanded surface will be covered with Partol

#10. A layer of epoxy coated carbon cross-ply will be applied to the inner

surface of the molds until the composite of the specified thickness. A thin

piece of plastic will be glued to the surface of the mold using Liquid Nails

or a suitable substitute. A tube connected to a compressor will be sealed

into the plastic bag, and the air will be evacuated from the plastic bag until

the glue is cured. The fuselage sections will then be removed from the

molds.

The molded sections will be cut and sanded to final shape. Bulkhead will

be epoxied into the fuselage sides. The motor mounting shelf will be

glued together and glued to the firewall of the fuselage. Hinges will be

glued onto the bulkhead sides and the fuselage will be pinned together

with hinge wires.


The molded tail portion will have a 3/16” balsa sheet epoxied into the

saddle portion the horizontal stabilizer will be attached. The tail section

will be screwed onto the main fuselage section using small machine

screws.

When the sides of the main fuselage are pinned together, the motor and

motor mount shall be mounted to the motor mounting shelf. When the

motor is mounted, the cowling will be screwed to the main fuselage over

the propulsion unit.

Plastic skids will be glued to the outside skin of the bottom, front, and tail

of the fuselage.

In the wing saddle area of the fuselage, No. 4 holes will be drilled every 1”

to allow for variable wing attachment. Hatches and windows can be

screwed into the fuselage at locations specified in drawings using small

machine screws. Sensory equipment can now be mounted into the

fuselage using the molded rails on the fuselage sides. Drawings of final

layup can be found in Appendix A.

3.1.4 Takeoff Mechanism

This team proposes creating a winch to aid in the takeoff of the airborne

platform. This can be constructed using a gasoline engine of the same

size as those found on lawn mowers and chainsaws. These should have

sufficient power to launch the airplane in a distance well below the

calculations derived for takeoff with landing gear.

There are several advantages to the winch system. By using energy from

an external source to launch the airplane rather than onboard batteries,


the mission time of the aircraft can be extended beyond the needs

assessment requirements. Also there is a drag and weight reduction with

the subtraction of landing gear.

3.2 Analysis Methods and Sizing

3.2.1 Aerodynamics

The shape of the exterior of the airborne platform was designed to be as

aerodynamic as possible while allowing for flexibility in payload weight and

volume. This was particularly challenging because the molded fuselage is

also a load bearing structure; the design agreed upon was the best option

considered to maintain an aerodynamic profile on a load bearing structure.

The drag calculations shown in Section 3.3 show that the final fuselage

will have a low parasitic and induced drag coefficient. The team

attempted to create a design without sharp vortex inducing corners. This

also helped in the structural analysis by avoiding stress concentrations.

A maximum external width and height of 6” was chosen because it should

allow a comfortably sized fuselage interior to mount all necessary

payloads. Six inches width should allow any maintenance to payloads to

be conducted easily while payloads are mounted to the aircraft structure.

A NACA 0006 airfoil was chosen for the horizontal stabilizer because it

would be more efficient stabilizer and lifting surface for control. Because

of the short overall chord of the vertical stabilizer, it was decided that a flat

plate would be better than an airfoil shape. With a short overall chord, the

benefits a formed airfoil could provide would not be observed.

The cowling was molded in an upward sweep to mount the motor slightly

above the main fuselage. This will protect the motor in the unlikely

incident of a crash landing. The cowling should take the brunt of any

impact and is easily replaced.


3.2.2 Structures

Rails were molded into the skin of the fuselage to save weight and provide

more strength to the structure of the aircraft. Bulkheads were strategically

placed to allow for predicted payload volumes while strengthening

structural integrity.

For ease of construction and disassembly, it was decided to keep the

fuselage in two halves that could be easily separated. The stresses

attempting to separate the sides of the fuselage were relatively low. The

hinges, cowling, tail cone, and motor mount will be sufficient to hold the

sides together during the flight mission.

The motor mounting shelf is notched and glued into firewall of the

fuselage. It is strong enough to withstand much more than the thrust that

the motor is able to create.

3.2.3 Payload

Care had to be taken to allow for payloads of various volumes and weights

to be mounted within the fuselage. There are two rails molded into each

side of the main fuselage to allow for mounting of payloads. CIAS

estimated a fairly heavy payload, and it was decided to place this in the

payload bay closest to the firewall. The receiver, receiver battery, and

propulsion battery will be mounted in the second payload bay near the

center of gravity under the wing. Other equipment can be mounted in the

third payload bay behind the wing as necessary.

Because of the unknown nature of the payloads being transported, the

center of gravity of the aircraft can vary by a large margin. For this

reason, the wing has a variable mounting surface. Also the propulsion


batteries will be mounted with Velcro so they can be shifted fore and aft in

the second payload bay to assist in adjusting the center of gravity.

3.2.4 Propulsion

Much research went into determining an appropriate propulsion unit for

this airborne platform. After reviewing what was currently available, it was

decided that Model Motors AXI 4120 would be chosen for a power plant.

These are brushless motors that boast an efficiency of up to 86%, some of

the highest rated efficiencies for their size. They are also some of the

largest brushless motors commercially available. These motors also do

not require gear boxes like other motors that were compared. This further

saved weight. In the configuration chosen for this application, the motor is

predicted to have a maximum mechanical output of around 500 W at the

shaft using a folding 14x9.5 propellar. Weight of the motor is 10.25 oz.

After analyzing what is available commercially, the team decided to use

lithium polymer batteries from Thunder Power Batteries. The model that

calculations were based on is the TP8000-5S4P. This is a 20 cell lithium

polymer battery pack with 5 cells in series and 4 in parallel. The energy

capacity of the batteries is 8000 mAh at 18.5V for a total of 148 Watt-hrs.

These batteries can provide a maximum average discharge of 40A, which

is far below what the flight mission requirements are. These batteries can

burst at 80A for several seconds if it is needed. Battery dimensions can

be found in the CAD drawings. Battery pack weight is 1.7lbs a pack; a

second battery pack could easily be added to the existing aircraft design if

the flight mission were to require the extra power.

Predicted power and torque of motor and battery can be found in

Appendix C.


Total energy required to meet mission requirements is shown below and

compared to the total available energy from the batteries. Notice that the

takeoff value assumes the aircraft is taking off under its own power. There

should be enough power left over in the batteries to power electric

payloads or extend the loiter portion of the flight mission. A breakdown of

available power and the power required is found in figure 10.

Power Required(W)

Time to Completion(s)

Energy Consumption (Whr)

Takeoff 180 60 3(with landing gear) Climb to 90 300 7.5cruise altitude Loiter and return 75 3240 67.5Flight Total Energy Avail: 148 Whr Total Energy Req: 78 Whr

Figure 10: Available power versus total power required

3.3 Final Aircraft and Predicted Performance

3.3.1 Aircraft Configuration

Load Stress

Refer to CAD drawings in Appendix A for preliminary aircraft configuration.

For analysis, arbitrary payload locations were chosen in the areas they

would most likely be in the prototype flight tests. These payload locations

are subject to change, but should only affect the stress calculations by a

negligible margin. A shifting wing location also makes the center of gravity

calculation arbitrary depending on the payload locations. The highest

loading stress on the fuselage was 554 psi and was calculated using finite

element Analysis in IDEAS, refer to Appendix A for appropriate diagrams.


Maximum stress at the root of the wing on the wing spar was 15.8KSI.

this within a conservative factor of safety of 45KSI in compression and

tension for a carbon/epoxy composite.

3.3.2 Predicted Performance

Wing Sizing

Optimal chord length and span of the wing was calculated from the

predicted wing loading:

Because the wing molds for this project already exist, they were largely

unneeded for this project. However, they become useful in determining

the required wingspan. It was found that an effective wingspan of 100

inches would predict desired flight characteristics for this design.


Take off and Landing

Takeoff and landing was calculated based on using landing gear. This

way the team would be assured the design is capable of having a more

versatile mission roles.

Takeoff distance was calculated at 132 feet. Calculations for ground roll

can be found in the Power Required section.

Empennage sizing and control surface sizing

Stabilizers and control surfaces for this model were found using common

volume predictions. These formulas were compared to coefficients that

are commonly associated with a payload carrying aircraft of this type and

were found to exceed those standards:


More detailed analysis can be performed, but are unnecessarily

complicated for a project of this nature and the scope of this design team.

Actual sizing data for the empennage can be found in the CAD drawing in

Appendix A.

Power Required Predictions

Initial flight conditions and constants are shown below in figure 12 to

predict the power required to complete the flight mission.

alpha knot = 4.775 rho (SL) = 0.002377 Cd = 0.006alpha = 4.085 rho (cruise) = 0.002308 CL max = 1.8

Cl = 1.4 V (ft/s) = 36.67 mu = 3.74E-07e = 0.9 weight (lb) = 12 Lf (fus length) (ft) = 4

AR = 10 time to TO (s ) = 300 dia of fuselage (ft) = 0.5T @ TO (lb)= 2.57 mu (pavement) = 0.02 PA @ climb (ft*lbs/s) = 110

Figure 12: Initial values for flight characteristic predictions


Symbols: horizontal tail volume

VV vertical tail volumeaileron volume coeffcient

horizontal tail areavertical tail areahorizontal distance from leading edge of wing to aerodynamic center of airfoil

With these initial conditions chosen, the power required to fly the aircraft

was calculated. The maximum power required is at takeoff if landing gear

is to be used, and this is roughly 182 Watts. This 182 Watts assumes a

power loss of 50% which is very conservative. The motor used in this

aircraft is rated to be roughly 80% efficient depending on the speed it is

running and the propeller will be rated better than 70% efficient. The rate

of climb for the aircraft will be roughly 2.5 ft/s. Other values related to

power required can be seen in the figure 13 below.

CL = 1.197696 R/C = 0.072897611 HPCD (wing) = 0.056734 54.67320817 Watt

S = 6.456365 ft Re (fus) = 9.06E+05 b = 8.035151 ft Cf = 4.76E-03

D (wing) = 0.568434 lb CD (fus) = 0.004634576 V stall = 29.47609 ft/s D (fus) = 0.106970646 lbs

20.09733 mph CD (AC) = 0.072198593 V lo = 35.37131 ft/s D (AC) = 0.794593132 lbs

24.1168 mph TR (cruise) = 0.723374609 lbsPE = 12000 ft*lbs PR (cruise) = 26.52614691 ft*lbs/sec

PR (to) = 40.09369 ft*lbs/sec 0.048229358 HPSlo = 132.0693 ft 36.17201851 Watt

R/C = 2.484386 ft/s Assuming 50% efficiency on the motor

PR = 181.6905 Watt Figure 13: Predicted flight characteristics

Weight

A weight analysis was done on the components of the aircraft to ensure

we were within the projected estimate from the conceptual design phase. Wing

weight was taken from existing wings from the same molds and construction

methods as will be used in this design. Balsa wood density was taken at 10

lb/cft. Fiberglass and epoxy density was estimated at 124 lb/cft. Carbon and

epoxy density was estimated to be at 97 lb/cft. Motor weights and off the shelf

electronics weights were taken from the manufacturer’s documentation. Figure

14 outlines the weight estimate. The total projected weight of the design is

11.85lbs.


Weight AnalysisWing 3 LbsFuselage 2 Lbsempennage 0.2 LbsBatteries 1.7 LbsMotor 0.65 LbsCIAS payload 3 LbsMisc payload 1 Lbsradio equipment 0.3 LbsTotal Weight 11.85 Lbs

Figure 14: Weight Analysis


4.0 Prototype Design

4.1Construction Methods4.1.1 Wing

The Eppler 423 wing plan form was used in the prototype design. The

wing chord was kept at 8 inches. The wing length was changed to 120 inches

which changed the aspect ratio of the wing to 15. This added nominal weight to

the plane as the final product was lighter than predicted in the preliminary design.

This will allow greater mission versatility for our sponsor. It may be possible to fly

the aircraft at slower speeds and carry heavier payloads. The outer 12 inches of

each wing tip has a 13 degree dihedral to aid in roll stability. Aileron volume

calculations suggested the control surfaces be 24 inches long by 2 inches wide.

A wing spar was constructed from 5 spliced pieces of basswood. The

final spar dimension was .25x1x110 inches. Care was taken to avoid placing a

splice at the center point of the wing, instead one of the 24 inch long pieces of

basswood was center along the centerline of the wing. Cyanoacrylic glue and

thread was used to splice the basswood pieces together.

Wings were formed from Pactiv Green Guard extruded polystyrene that

was cut with a hot wire (Picture 1,2,3). Foam cores were sanded smooth. A .25

inch diameter hole was bored out of the wing core with a hot wire for a servo

channel. Wing cores are glued together using 3M foam spray glue. Foam spray

glue was also used to attach the spar to the quarter chord point of the wings.

Wing core was sanded smooth and spackled with lightweight spackle in

preparation for vacuum bag.

The wing was covered with one layer of unidirectional carbon fiber and a

top layer of (0, 90) woven fiberglass in an epoxy matrix. The spar was

completely through the wing core and was glue directly to the carbon sheet.

Wing lay-up was a solid 8 foot section (picture 4), so there would not be a glue

joint at the center of the wings that could cause stress concentrations.


Wings were vacuum bagged. Laminating sheets were laid out onto waxed

.003 inch Mylar skins. Epoxy was impregnated into the fibers of the laminating

sheets. The wing core was sandwiched between the laminating plies and placed

in a bag with a layer of porous material. The bag was sealed and a compressor

was used to evacuate the air from the bag. The wings were place back into the

wing beds the cores were formed form and weights were placed on the wing

beds. The wing beds and weights ensured the wings would cure straight and not

warped.

A 2 inch piece of wing spar was left at each side of the wing for wingtip

placement. Wingtips were laminated in the same manner as the main wing.

Thin 1/16 inch birch plywood root wing ribs were attached to the wingtips and

wing edges and beveled to a 13 degree dihedral angle. Wingtips were glued in

place with epoxy. The joint between the wingtips and the main wing was

reinforced with a 2 inch wide strip of (45,-45) woven fiberglass strip epoxied to

the wing surface (picture 5).

The centerline of the ailerons was mounted two-thirds the distance from

the center of the wing. This maintained the effectiveness of the aileron while

keeping them from tip vortices distortion. Ailerons were mounted with clear

packing tape on the top and bottom of the surface of the wing. This tape

provided an easy and tight bond between aileron and wing surface. This tape

also prevented air from slipping between the aileron and wing to lessen

effectiveness.

Servos are mounted to the wing near the spar at the midpoint of the

ailerons (picture 6). Mounting near the midpoint will lesson moments along the

aileron from servo movements and other aerodynamic forces. Pushrods are

constructed of 2-56 wire with metal clevises. Control horns are mounted to the

ailerons with machine screws and are made of nylon. Servo to pushrod

connections uses z-bends because of superior slip resistance.

The wing is mounted onto the fuselage using a formed foam saddle that

sits between the main fuselage and the wing. This saddle provides some

damping from vibration and other shocks during flight. This foam saddle can also


be sanded to an angle to change the angle of attack of the wing. The wing is

held in place with a formed (45,-45) woven carbon angle-ply panel. This panel is

screwed into the fuselage using 4 ¼-20 nylon bolts. According to calculations

the bolts should be more than able to withstand the loading scenarios the aircraft

would sustain in flight but will shear off in the event of a serious landing mishap.

4.1.2 EmpennageConstruction

The tail cone of the empennage was constructed using 1/16 inch thick

balsa sheets reinforced with ¼ inch square balsa sticks (picture 13,14,15). The

balsa sticks created a truss underneath the balsa sheets. The tail cone tapers up

from the 6x6 inch dimension of the fuselage bulkhead to a 1x1 inch square at the

rear of the tail cone. The tail cone is hollow and has a hatch screwed onto the

bottom front to allow easy access to servo lines and nylon bolts securing the tail

cone to the fuselage. The front bulkhead of the tail cone is constructed of 1/8

inch thick birch plywood.

The horizontal stabilizer is 30 inches wide and has a mean chord of 8

inches. It is 11 inches at the largest chord at the root. The horizontal stabilizer

utilizes a NACA 0008 airfoil. It was cut using a hot wire passed through Pactiv

Green Guard extruded polystyrene foam blocks. These foam cores were sanded

and laminated with a single layer of (0, 90) woven fiberglass cross-ply

impregnated with epoxy and black pigment. The foam cores and fiberglass were

vacuum bagged to eliminate air bubbles and delaminating points.

The vertical stabilizer was constructed in the same manner as the

horizontal stabilizer. The vertical stabilizer was constructed using a NACA 0006

wing plan form. The mean chord is 6 inches and the height is 15 inches.

The vertical stabilizer was formed to the horizontal stabilizer (picture 12).

Both the horizontal and vertical stabilizers were epoxied to the tail cone. The tail

cone was primered and painted to provide limited protection from moisture and

punctures.


Control surfaces were cut from the trailing edge of the horizontal and

vertical stabilizers. Control surface on the vertical stabilizer is 2.5 inch wide and

runs the length of the trailing edge. The elevator is 2 inches wide and runs the

length of the trailing edge of the horizontal stabilizer. The leading edge of the

control surfaces was beveled with a razor to allow freedom of movement when

hinged. Hinging of the control surfaces was completed using clear packaging

tape on the top and bottom of the mated surfaces of the stabilizers and control

surface. This provided a more simple design and prevents air from passing

through the hinge line and limiting the effectiveness of the control surfaces.

The empennage section was bolted to the main fuselage using ¼-20

nylons bolts through holes in the bulkhead. Calculations show in the event of a

mishap, the nylon bolts should fail before tail cone failure. If this proves to be

incorrect, fewer bolts may be used without adversely affecting aerodynamic loads

from the empennage through the fuselage.

Servos for the rudder and elevator were mounted on the horizontal

stabilizer. Servo lines were spliced longer and extend back into the main

fuselage. Pushrods are 2-56 gauge wire and have threaded metal clevis at the

control horn attachment point. Control horns are nylon and are screwed onto the

control surfaces. By mounting servos on the horizontal stabilizer instead of

within the main fuselage, there is more useable payload volume and pushrod

weight is saved.

Tail Volume Coefficients

There was initially concern that the aerodynamic center of the wing was

too close to the aerodynamic center of the empennage for effective control. The

horizontal stabilizer is located above the wing and flow over the stabilizer should

not be affected by air circulation over the wing. There are formulas for tail

volume coefficients that are statistically based off of empirical data. These

formulas are used to determine the theoretical effectiveness of tail surfaces.


Aircraft stabilizer volumes generally range between .3-.8, with .8 being more

stable. Our design volume coefficient for the horizontal stabilizer was .78, which

means it should theoretically be stable. The vertical stabilizer was calculated to

have a volume coefficient of .195. Average vertical stabilizer coefficients range

from .005 to .015, with .015 being more stable. Tail dimensions can be verified

from the CAD drawings.

Final empennage weight is approximately 1 lb.

4.1.3 FuselageThe fuselage was design and constructed around the payload

requirements of the sponsor. The payload objective was to fully enclose a

6x6x12 inch black box with a mass of 3 lbs. It was also uncertain as to the size

of the telemetry and flight control system to be integrated later. A relative shape

of 6x6x33.5 inches was chosen for the main section of the fuselage. Due to the

simple geometry of the rectangular box, two .5 inch rails were molded into each

side of the fuselage. These railings offered increased bending and torsional

stiffness compared to a standard box.

Composite Lay-up

Composite construction techniques were employed in lay-up of the

fuselage. The fuselage skin was constructed from a layer of (0, 90) weaved

carbon cross-ply sandwiched between two layers of (0, 90) weaved fiberglass

cross-ply in an epoxy matrix. The carbon weave was chosen to provide stiffness


in the event of a landing mishap to protect sensitive onboard equipment as well

as to limit deflection in the fuselage due to aerodynamic forces. The laminated

fiberglass provided a smoother outer finish to decrease parasite drag as well as

to negotiate the geometry of the molded railings during composite lay-up.

The mold for the fuselage was a female mold. Because of two axes

symmetry only one half of the fuselage had to be represented in the mold. It was

found early on that foam would not be an adequate material to construct the mold

(picture 16). The mold was constructed from .75 and .5 inch thick MDF (picture

17). The railing and corner contours were formed using a router directly into the

MDF (picture 18). Imperfections in the MDF mold were smoothed out with Bondo

and wet sanded (picture 19,20). When the desired geometries were achieved,

mold smoothness was achieved using primer and spray paint (picture 21). A thin

layer of epoxy was painted over the layer of spray paint. To prevent the epoxy

from the lay-up from bonding to the sides of the mold, several layers of Partall #9

mold release were buffed onto the mold surface and a layer of PVA was painted

over the mold release. After initial tests were conducted with limited success,

this method of mold conditioning proved to work best for this purpose.

Fuselage sides were vacuum bagged in the mold. A (0,90) woven piece

of carbon fiber was sandwiched between two layers of (0,90) woven fiberglass

(picture 22, 23, 24, 25). Initial tests showed that using only fiberglass would not

give the desired strength or stiffness needed for the loads the aircraft was

expected to see in flight.

Completed fuselage halves were trimmed leaving 1.5 inch excess material

on the top and bottom. This 1.5 inch overlap was roughed up with sand paper

and provided extra gluing surface to reinforce the fuselage construction (picture

29, 30, 31). 2 inch wide strips of (45,-45) woven carbon angle ply were epoxied

over the exterior glue joint of the fuselage to further reinforce this mating surface

as well as to provide additional torsional stability.

Bulkhead Placement


Interior of the fuselage contains 3 payload compartments separated by 5

bulkheads. Bulkheads were constructed of 1/8 inch thick birch plywood. A 1.5

inch diameter hole was centered in each bulkhead to decrease weight as well to

allow cables to be routed throughout the fuselage (picture 26, 27, 28). Small .25

inch holes for servo line routing were drilled into the bottoms of the bulkheads.

Bulkheads were epoxied into the fuselage before the fuselage skins were glued

together.

Bulkheads were constructed to maximize the modular capability of

fuselage. The rear most bulkhead is flush with the edge of the fuselage. Four ¼-

20 threaded holes were tapped into the corners of the bulkhead and reinforced

with basswood to provide at least 3/8 inch of effective threading. Threads were

reinforced using a cyanoacrylic style super glue. These bolt locations provide

best placement to distribute the aerodynamic forces from the empennage along

the load lines of the fuselage. These bolts allow for the easy removal and

replacement of the empennage of the aircraft. Removal of the tail also provides

an additional access point for the rear most payload bay.

Two bulkheads were placed under the wing saddle to support the wing

loads. Wing chord is 8 inches, and bulkheads were placed at the leading and

trailing edges of the wing. The top of the fuselage at the wing saddle was cut

away 1 inch to create wing mounting surface that was flush with the top of the

fuselage. A 6x12 inch sheet of 1/8 inch thick birch plywood was mounted to the

fuselage railings and the tops of the bulkheads to strengthen the wing opening.

A 3x5 inch rounded rectangle hole was cut into the wing saddle to provide

access to the payload bay below. Four 1 inch cube blocks of basswood for wing

mounting were tapped to a ¼-20 threading and thread reinforced with thin

cyanoacrylic glue. These tapped blocks were epoxied to the plywood wing

saddle and bottom surface of the top of the fuselage. Mounting blocks in this

manner allows the fuselage skin to receive the majority of the wing loading

forces.


The front two bulkheads in the fuselage provide a mounting point for the

motor. The two bulkheads were cut from 1/8 inch thick birch plywood and have a

2.5 inch square mounting point centered along the tip surface. One bulkhead is

mounted flush to the front of the fuselage, and the second is mounted 3 inches

aft of the first bulkhead. The centerline of the motor is mounted 1.5 inches from

the top of the fuselage. Minimal moment force was incurred by mounting the

motor above the aerodynamic center of the aircraft. A .25x1 inch balsa stick was

glued between the mounting points of the bulkheads to provide compression

strength. The screws in the motor mount are mounted through both bulkheads to

distribute thrust loading over a greater area of the fuselage.

The motor cowling was vacuum formed from .02 inch thick styrene plastic.

A balsa plug was carved to represent half of the motor cowling. The cowling was

created large enough to cover the entire motor and most of the mounting portion

of the bulkheads. Because the main purpose of the motor cowling was to

provide assistance to minimize drag from aerodynamic forces and little actual

structural support, the strength of the styrene was not particularly important. The

motor cowling did stiffen considerably when it was mounted to the motor

bulkheads and because it was molded in halves, still provided easy access to the

motor. Other benefits of the motor cowling were to protect UAV operators from

the spinning casing of the outrunner style motor and it was determined that the

motor cowling provided some aesthetic improvements to the overall look of the

UAV.

The fuselage cowling was vacuum formed using .03 inch thick clear PVC

plastic. A balsa plug was carved and used to vacuum form the plastic over.

These vacuum forming plugs aid in the module design of the aircraft and the

ability to easily construct new parts and assembly. It was found that when the

PVC was stretched over the balsa plug, it was too thin to provide structural

support. The nose of the fuselage is a particularly vulnerable portion of the UAV.

The fuselage cowling could not extend past 3 inches from the front of the

fuselage because of the propeller clearance. It must not deform under the prop


wash and must be able to withstand minor landing mishaps. A video camera to

stream real-time data to the aircraft pilot is planned to be mounted under the

fuselage cowling and it would need protection from these forces. As a

consequence of these concerns it was decided that the fuselage cowling would

be reinforced. Several layers of fiberglass in varying directions and layer of

unidirectional carbon fiber were used to stiffen the fuselage cowling. It is

anticipated that a hole will have to be drilled into the fuselage cowling to provide

field of view for the telemetry camera. The fuselage cowling is mounted to the

front bulkhead using small wood screws and .25x1 inch balsa blocks.

The weight of the finished fuselage was approximately 2.67 lbs.

4.2 Aircraft Configuration


The Model Motors AXI 4120/18 was chosen as a propulsion source for

this aircraft. Preliminary testing done at the time of this report suggested the

engine could average 4.2 lbs of thrust (picture 34, 35). This is in line with what

was predicted in the preliminary design. Testing was completed by Senior

Design Team 05009. Not much testing had been done with the battery at the

time of the report. The battery that was purchased is a ThunderPower 8000mAh

lithium polymer battery (picture 36). Battery chosen has the same specifications

as the one outlined in the PDR, but it is shorter and wider than specified battery.

4.2.2 Weight of planeThe final structural weight of the finished airplane is

Final WeightFuselage 2.67 lbsCowling 0.13 lbsempennage 1 lbswing 2.9 lbstotal 6.7

Figure 15


Total final weight was 6.7 lbs, which makes the final weight of the aircraft 13.5 lbs after telemetry, propulsion, and payload are added. This is one half pound lighter than was specified in the preliminary design.

4.3 Predicted PerformanceWith the changes in the wing sizing and weight of the aircraft, it became

apparent that there would be changes in the predicted flight characteristics of the

aircraft. Most noteworthy of these changes is the power required to maintain

level flight dropped to 29 Watts. Power to take off also dropped to 61 Watts.

Other changes are noted in the table below.

alpha knot = 4.775 rho (SL) = 0.0023769 Cd = 0.006

alpha = 4.085 rho (cruise) = 0.0023081 CL max = 1.8Cl = 1.4 V (ft/s) = 36.67 mu = 3.74E-07e = 0.9 weight (lb) = 13.5 Lf (fus length) (ft) = 4

AR = 15time to TO (s )

= 300 dia of fuselage (ft) = 0.5T @ TO

(lb)= 2.57mu

(pavement) = 0.02PA @ climb (ft*lbs/s)

= 110

CL = 1.197696 Re (fus) = 9.06E+05 CD (wing)

= 0.039823 Cf = 4.76E-03 S = 7.263411 ft CD (fus) = 0.004119623 b = 10.43797 ft D (fus) = 0.095085018 lbs

D (wing) = 0.448868 lb CD (AC) = 0.05169699 V stall = 29.47609 ft/s D (AC) = 0.639945215 lbs

20.09733 mphTR (cruise)

= 0.582709783 lbs

V lo = 35.37131 ft/sPR (cruise)

= 21.36796773 ft*lbs/sec 24.1168 mph 0.03885085 HP

PE = 13500 ft*lbs 29.13813781 WattPR (to) = 45.07545 ft*lbs/sec Slo = 138.7783451 ft

0.081955 HP R/C = 1.654091278 ft/s 61.46653 Watt

Figure 16


5.0 Analysis and Design Testing

5.1 Wing Failure AnalysisWing failure testing was done using the Mechanical Engineering

Department’s Tinius-Olsen Tension Tester. The test section was supported as a

beam on two ends by cinder blocks and foam pieces of the foam core bed. A 4

inch piece of foam wing bed was used to support the pressure from the Tinius-

Olsen along the top center of the wing. A two foot test section (picture 7,8) with

basswood spar was mocked up for this purpose.

It was determined that a failure would be defined as a de-lamination of the

wing covering, because it was presumed that may indicate a failure of the wing

spar. When the wing was loaded to approximately 240 ft-lbs a compression

failure measuring .25” (picture 9) appeared at the center of the wing at the spar

line.

The test specimen was dissected to observe the extent of the de-

lamination. The lamination containing the failure was removed for any visible

signs of failure along the spar (picture 10). Dissection continued with the

eventual removal of all foam along spar to attempt to discern any failures (picture

11). After no visual signs of spar failure were observed, the spar was hand

tested by members of the team to determine if any weak points had occurred

(picture 12). The spar and foam were found to be in good shape, and the

observed failure had only occurred in the lamination in compression at the top of

the spar.

240 ft-lbs translates to a load factor of three for a ten foot wing section and

a total fuselage weight of fourteen pounds. A wing load factor of three would be

considerably more than the aircraft would sustain in normal flight with a nominal

factor of safety. It appeared that the test section could sustain much higher

loads. Because the integrity of the foam and spar were maintained at these

loads, if this failure happened in flight the aircraft would maintain airworthiness.


5.2 Aircraft Load TestingDue to inclement weather that normally occurs in Rochester this time of

year, flight and glide testing was unable to be completed before this report was

written. The aircraft was statically loaded to test whether the finished aircraft

could withstand forces equal to at least 3 times the gross weight of the aircraft.

A load factor of three was chosen considering the maximum gusting

affects expected to be experienced by the aircraft.

where: rho = .00238 slugs/ft^3V = 37 ft/s (cruise velocity)Cla = 5 /radS = 6.7 ft^2 (wing area)W = 13.5 lbKU = 30 ft/s (K = 1, gusting coefficient)

Commercial sailplanes are often tested to a load factor of 3 to assure they

can withstand wind gusting. With a gross weight of 13 lbs and a wing weight of 3

lbs the load to be distributed by the wings is estimated to be 10lbs. With a load

factor of 3, it was determined that the aircraft needed to be statically loaded with

30 lbs of weight for an accurate simulation. The aircraft was successfully loaded

to 33.5 lbs during the static test (picture 32, 33). This means the aircraft achieved

a load factor of 3 with a factor of safety of 1.2. This is common in the aerospace

community.

After the static loading we were able to compare the stresses experienced

in the wing section with the stresses experienced by the test wing section.

Several assumptions were first made. First, the wing is assumed to be an I

beam with the flanges made of the carbon and fiberglass laminations and the

center made of basswood. The effective surface of the flange is assumed to be

4 inches wide. The modulus of elasticity is assumed to be constant throughout


the entire I beam. Because of the variety of composite materials involved, it was

impractical to attempt to determine actual modulus of elasticity.

Where;

Figure 17


Figure 18

Even with a load factor of 3 and a factor of safety of 1.2 tested in the static

load test, the stress calculated is only 1.9 ksi. The failure point of the test section

was measured at 6 ksi. This indicates the aircraft may be able to carry larger

than anticipated payloads without mishap.

5.3 Center of Gravity Calculations

W total = 10.458 lbs. L total = 48.5 in h empennage = 6 inW empennage = 1.000 lbs. L fuselage = 33.5 in h telemetry = 18 inW telemetry = 1.000 lbs. L empennage = 12 in h battery = 25 inW battery = 1.713 lbs. L cowling = 3 in h fuselage = 28.75 inW fuselage = 2.670 lbs. h cowling = 47 inW cowling = 0.130 lbs. h motor = 47 inW motor = 0.851 lbs. h payload = 39.5 inW payload = 3.000 lbs. h camera = 47 inW camera = 0.094 lbs.W total05009 = 1.094 lbs.

#05009 Components

Cg*Wtotal05009=Wtelemetry*htelemetry+Wcamera*hcamera

Cg 5009 = 20.4857 inFigure 19Cg*Wtotal=Wempennage*hempennage+Wtelemetry*htelemetry+Wbattery*hbattery+Wfuselage*hfuselage+Wcowling*hcowling+Wmotor*hmotor+Wcamera*hcamera


Weight calculations pushed the center of gravity to roughly 20.5 inches from the rear of the aircraft. This is in line with the aerodynamic center of the aircraft. Assumptions were made that a 3 pound payload would always accompany the aircraft in flight and the weight of the telemetry systems are accurate.

6.0 Bill of MaterialsMaterials are readily available at no cost to the team unless otherwise

noted. The Bill of Materials is complete as to the knowledge of the team.

A purchaser from CIAS was used to procure items on the list, and

because of this an itemized list of prices was not available at the time this

report was written. The purchaser did provide a cumulative list of

purchase prices, and the values included in this report should be within

10% of the actual cost incurred in this project.

Price UOP QuantityTotal Cost Vendor

Hitec Mighty Mini BB MG Servo J $27.99 ea 4 $111.96 RITAXI 4120/18 $159 ea 1 $159.00 HobbyLobbyradial motor mount $18.50 ea 1 $18.50 HobbyLobbyTP8000-5S4P battery $399.00 ea 1 $399.00 ThunderPower13.5V Power Supply $74.95 ea 1 $74.95 Tower HobbyAstroFlight 1-9 Cell Lithium Charger $114.95 ea 1 $114.95 Tower HobbyJESA70OP Jeti 70 Amp $139.00 ea 1 $139.00 HobbyLobby.5" foam rubber sheets NA ea 2 NA Tower HobbyAPC 13x10 folding propellar $7.99 ea 2 $15.98 Tower HobbyAPC propellar hub $4.79 ea 2 $9.58 Tower HobbyCA glue (2oz) thin $6 bottle 1 $6.00 Tower Hobby


CA glue (2oz) medium $6 bottle 1 $6.00 Tower HobbyGreat Planes fiber hinges (24) NA ea 1 NA Tower HobbyGreat Planes control horns (2) NA ea 2 NA Tower HobbyGreat Planes metal clevis (12) NA ea 1 NA Tower Hobbypushrods, 2-56, threaded (6) NA ea 1 NA Tower Hobby3M Foam spray glue NA bottle 1 NA Sprayway 66 spray glue NA can 1 NA Krylon Primer NA can 2 NA vacuum forming pvc $0 sheet 2 $0.00 RITGlinks epoxy NA qt 1 NA MDF, .5x24x60" NA sheet 2 NA MDF, .75x24x60" NA sheet 1 NA Custom, Lightweight Spackle NA qt 1 NA Bondo NA gal 1 NA Painter's Plastic, 200yd NA roll 1 NA .003 Mylar NA roll 1 NA Sandpaper NA pack 1 NA Elmers Polyurethane glue NA bottle 1 NA X-acto blades (15) NA ea 1 NA X-acto knife NA ea 1 NA Krylon spray paint NA can 3 NA servo wire NA spool 1 NA Dan's Crafts Fuselage cross ply carbon laminate NA yd 3 $0.00 RITEpoxy (5min) $3.00 oz 4 $12.00 1/8" birch plywood (6x6x12") NA sheet 4 NA 1/16" birch plywood (6x6x12") NA sheet 2 NA epoxy (2hr) $75.00 gal 1 $75.00 Fiberglast Empennage balsa, 3/16x1/2x36"(5) $5 bundle 1 $5.00 Dan's CraftsMonokote, (6') $11 roll 1 $11.00 Tower HobbyDerek Miller contracted for empennage $200 1 $200 RIT Wing Pactiv Green Guard na yd 6 $0.00 RITFoamular 250 NA sheet 1 NA fiberglass cloth NA yd 2 $0.00 RITHexCell (7/8x5/8x100") sheet 1 $0.00 RITuni-directional carbon fiber (24x120x.025") yd 3 $0.00 RITCumulative Purchase CostsItem PriceMotor $334.89Battery Packs $406.95 Tower Hobbies Order $274.34Hardware Store run #1 $31.23Hardware Store run #2 $16.58Dan's Crafts $51.23


Empennage $200 Hardware run #3 $61 Total $1376.22

Figure 20

Total cost for this project is estimated to be around $1376.

7.0 Project TimelineProject timeline was separated into work completed in the fall of 2004 and the

projected timeline for the upcoming winter quarter of 2004. The majority of the

tasks to be completed in the Winter quarter timeline were behind schedule. It

was stated in the preliminary design report in order to maintain such an ambitious

schedule additional assistance would have been required. Assistance to

complete the project was not offered until week 9 of the project.

7.1 Fall Timeline

The projected timeline for fall is shown in Figure 21. The majority of the

time spent on this project was spent in the conceptual design phase. A

majority of the time was spent in conceptual design because the needs of

our sponsor have been very dynamic over the course of Senior Design I.

After losing two team members and being denied additional members, the

work load for the remaining members increased in all tasks following the

conceptual design phase.

7.2 Winter Timeline

The projected timeline for winter is shown in Figure 22. The actual

timeline for winter is shown in Figure 23. Many delays were experienced

during the project. Early in the quarter purchasing was proving to put days

of lag time into the project. Initial mold creation failed and the molds had

to be completed after the winter break. Team also did not receive

additional man hours for project until very late in the quarter. Flight testing


could not be completed due to inclement Rochester weather. Project did,

however, get completed by the time of this report.


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