1



Complete Design Review

Project 5008

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Christina Alzona

Benjamin Wagner (team leader)

2/18/05

1. Introduction

1. Design Development

2. Design Alternatives

2. Conceptual Design

1. Needs Assessment

2. Feasibility Assessment

1. Aircraft Type

2. Empennage

3. Landing Gear

4. Propulsion System

5. Wings

6. Radio and Transmitter

7. Building Materials and Construction Methods

3. Conclusions

3. Preliminary Design

1. Construction Methods

1. Wing

2. Empennage

3. Fuselage

4. Takeoff Mechanism

2. Analysis Methods and Sizing

1. Aerodynamics

2. Structures

3. Payload

4. Propulsion

4. Prototype Design

1. Construction Methods

1. Wing

2. Empennage

3. Fuselage

2. Aircraft Configuration

1. Propulsion System

2. Weight Analysis

3. Predicted Performance

5. Analysis and Design Testing

1. Wing Failure Analysis

2. Aircraft Load Testing

3. Center of Gravity Calculations

6. Bill of Materials

7. Time Line

1. Fall 2004 Timeline

2. Winter 2004 Timeline

1. Predicted Winter 2004 Timeline

2. Actual Winter 2004 Timeline

8. Acknowledgements

9. 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 Introduction

Rochester 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 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.

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 Design

The 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.

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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.

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Figure 2 Empennage Configurations

|Tail Design Feasibility Assessment |Weighting |Low |T tail |Cruciform |

|(On a scale from 1-10) | | | | |

|R1: Sufficient Skills |0.1 |9 |7 |6 |

|R2: Sufficient Equipment |0.1 |9 |9 |9 |

|R3: Sufficient # of people |0.13 |2 |2 |2 |

|E1: Economically Feasible |0.07 |8 |7 |7 |

|S1: Meeting Intermediate Milestones |0.1 |8 |8 |8 |

|S2: Meeting PDR Requirements |0.1 |8 |7 |7 |

|S3: Meeting CDR Requirements |0.15 |8 |7 |6 |

|T1: Has similar technology been used before |0.12 |8 |7 |6 |

|T2: Plane stability |0.08 |6 |8 |7 |

|T3: Drag reducing |0.05 |7 |7 |7 |

|Total: |1 |7.21 |6.73 |6.28 |

Figure 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 |Weighting |None |Trike |Tail-dragger |

|(On a scale from 1-10) | | | | |

|Aerodynamics |0.2 |9 |7 |8 |

|Ground Handling |0.2 |0 |9 |6 |

|Weight |0.3 |10 |6 |8 |

|Ease of Construction |0.3 |10 |6 |7 |

|Total: |1 |7.8 |6.8 |7.3 |

Figure 4: Landing gear feasibility analysis

2.2.4 Propulsion System

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 |Weighting |Electric |Hybrid |Gas |

|(On a scale from 1-10) | | | | |

|Vibration |0.3 |9 |7 |5 |

|Weight |0.2 |8 |9 |7 |

|Stability |0.2 |9 |8 |7 |

|Cost |0.1 |7 |7 |7 |

|Ease of Installation |0.2 |9 |7 |8 |

|Total: |1 |8.6 |7.6 |6.6 |

Figure 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 |Weighting |LiPoly |Nimh |NiCad |

|(On a scale from 1-10) | | | | |

|Energy Density (charge/mass) |0.4 |9 |7 |8 |

|Cost |0.1 |6 |8 |7 |

|Ease of use/installation |0.2 |8 |8 |8 |

|Recharge Ratio |0.3 |7 |7 |7 |

|Total: |1 |7.9 |7.3 |7.6 |

Figure 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 |Weighting |High |Low |Mid |

|(On a scale from 1-10) | | | | |

|R1: Sufficient Skills |0.1 |8 |8 |6 |

|R2: Sufficient Equipment |0.1 |9 |9 |9 |

|R3: Sufficient # of people |0.13 |2 |2 |2 |

|E1: Economically Feasible |0.07 |9 |9 |9 |

|S1: Meeting Intermediate Milestones |0.1 |8 |8 |8 |

|S2: Meeting PDR Requirements |0.1 |7 |8 |6 |

|S3: Meeting CDR Requirements |0.15 |7 |7 |7 |

|T1: Has similar technology been used before |0.12 |9 |8 |8 |

|T2: Plane stability |0.08 |9 |7 |8 |

|T3: 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 Transmitter

A 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 |Weighting |Carbon |Fiberglass |Wood |

|(On a scale from 1-10) | | | | |

|R1: Sufficient Skills |0.1 |6 |6 |7 |

|R2: Sufficient Equipment |0.1 |9 |9 |8 |

|R3: Sufficient # of people |0.13 |2 |2 |3 |

|E1: Economically Feasible |0.07 |9 |9 |7 |

|S1: Meeting Intermediate Milestones |0.1 |8 |8 |8 |

|S2: Meeting PDR Requirements |0.1 |8 |8 |8 |

|S3: Meeting CDR Requirements |0.15 |7 |7 |7 |

|T1: Has similar technology been used before |0.09 |9 |8 |7 |

|T2: Durability |0.08 |9 |9 |6 |

|T3: Weight |0.08 |7 |8 |9 |

|Total: |1 |7.13 |7.12 |6.86 |

Figure 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 Design

With 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.

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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 |7.5 |

|cruise altitude |  |  |  |

|Loiter and return |75 |3240 |67.5 |

|Flight |  |  |  |

|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:

[pic]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.

[pic]

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:

[pic]

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 = |0.006 |

|alpha = |4.085 |rho (cruise) = |0.002308 |CL max = |1.8 |

|Cl = |1.4 |V (ft/s) = |36.67 |mu = |3.74E-07 |

|e = |0.9 |weight (lb) = |12 |Lf (fus length) (ft) = |4 |

|AR = |10 |time to TO (s ) = |300 |dia of fuselage (ft) = |0.5 |

|T @ TO (lb)= |2.57 |mu (pavement) = |0.02 |PA @ climb (ft*lbs/s) = |110 |

Figure 12: Initial values for flight characteristic predictions

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 |HP |

|CD (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 |lbs |

|PE = |12000 |ft*lbs |PR (cruise) = |26.52614691 |ft*lbs/sec |

|PR (to) = |40.09369 |ft*lbs/sec |  |0.048229358 |HP |

|Slo = |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 Analysis |

|Wing |3 |Lbs |

|Fuselage |2 |Lbs |

|empennage |0.2 |Lbs |

|Batteries |1.7 |Lbs |

|Motor |0.65 |Lbs |

|CIAS payload |3 |Lbs |

|Misc payload |1 |Lbs |

|radio equipment |0.3 |Lbs |

|Total Weight |11.85 |Lbs |

Figure 14: Weight Analysis

4.0 Prototype Design

4.1Construction Methods

4.1.1 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 Empennage

Construction

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.

[pic]

Final empennage weight is approximately 1 lb.

4.1.3 Fuselage

The 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

4.2.1 Propulsion System

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 plane

The final structural weight of the finished airplane is

|Final Weight | | |

|Fuselage |2.67 |lbs |

|Cowling |0.13 |lbs |

|empennage |1 |lbs |

|wing |2.9 |lbs |

|total |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 Performance

With 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 = |1.8 |

|Cl = |1.4 |V (ft/s) = |36.67 |mu = |3.74E-07 |

|e = |0.9 |weight (lb) = |13.5 |Lf (fus length) (ft) = |4 |

|AR = |15 |time to TO (s ) = |300 |dia of fuselage (ft) = |0.5 |

|T @ TO (lb)= |2.57 |mu (pavement) = |0.02 |PA @ 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 |mph |TR (cruise) = |0.582709783 |lbs |

|V lo = |35.37131 |ft/s |PR (cruise) = |21.36796773 |ft*lbs/sec |

|  |24.1168 |mph |  |0.03885085 |HP |

|PE = |13500 |ft*lbs |  |29.13813781 |Watt |

|PR (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 Analysis

Wing 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 Testing

Due 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.

[pic]

[pic]

where:

rho = .00238 slugs/ft^3

V = 37 ft/s (cruise velocity)

Cla = 5 /rad

S = 6.7 ft^2 (wing area)

W = 13.5 lb

KU = 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.

[pic]

[pic]

Where;

Figure 17[pic]

Figure 18[pic]

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

| | |

| | | | | |

|Cg*Wtotal05009=Wtelemetry*htelemetry+Wcamera*hcamera |

| | | | | |

|Cg |5009 |= |20.4857 |in |

Figure 19

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

[pic]

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 Materials

Materials 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 |Quantity |Total Cost |Vendor |

|Hitec Mighty Mini BB MG Servo J |$27.99 |ea |4 |$111.96 |RIT |

|AXI 4120/18 |$159 |ea |1 |$159.00 |HobbyLobby |

|radial motor mount |$18.50 |ea |1 |$18.50 |HobbyLobby |

|TP8000-5S4P battery |$399.00 |ea |1 |$399.00 |ThunderPower |

|13.5V Power Supply |$74.95 |ea |1 |$74.95 |Tower Hobby |

|AstroFlight 1-9 Cell Lithium Charger |$114.95 |ea |1 |$114.95 |Tower Hobby |

|JESA70OP Jeti 70 Amp |$139.00 |ea |1 |$139.00 |HobbyLobby |

|.5" foam rubber sheets |NA |ea |2 |NA |Tower Hobby |

|APC 13x10 folding propellar |$7.99 |ea |2 |$15.98 |Tower Hobby |

|APC propellar hub |$4.79 |ea |2 |$9.58 |Tower Hobby |

|CA glue (2oz) thin |$6 |bottle |1 |$6.00 |Tower Hobby |

|CA glue (2oz) medium |$6 |bottle |1 |$6.00 |Tower Hobby |

|Great Planes fiber hinges (24) |NA |ea |1 |NA |Tower Hobby |

|Great Planes control horns (2) |NA |ea |2 |NA |Tower Hobby |

|Great Planes metal clevis (12) |NA |ea |1 |NA |Tower Hobby |

|pushrods, 2-56, threaded (6) |NA |ea |1 |NA |Tower Hobby |

|3M 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 |RIT |

|Glinks 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 |RIT |

|Epoxy (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 Crafts |

|Monokote, (6') |$11 |roll |1 |$11.00 |Tower Hobby |

|Derek Miller contracted for empennage |$200 | |1 |$200 |RIT |

|  |  | | | |  |

|Wing |  | | | |  |

|Pactiv Green Guard |na |yd |6 |$0.00 |RIT |

|Foamular 250 |NA |sheet |1 |NA |  |

|fiberglass cloth |NA |yd |2 |$0.00 |RIT |

|HexCell (7/8x5/8x100") |  |sheet |1 |$0.00 |RIT |

|uni-directional carbon fiber (24x120x.025") |  |yd |3 |$0.00 |RIT |

|Cumulative Purchase Costs | |

|Item |Price |

|Motor |$334.89 |

|Battery Packs |$406.95 |

|Tower Hobbies Order |$274.34 |

|Hardware Store run #1 |$31.23 |

|Hardware Store run #2 |$16.58 |

|Dan'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 Timeline

Project 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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Symbols:

[pic] horizontal tail volume

VV vertical tail volume

[pic] aileron volume coeffcient

[pic] horizontal tail area

[pic] vertical tail area

[pic] horizontal distance from leading edge of wing to aerodynamic center of airfoil

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