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TRANSCRIPT
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
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4.2Aircraft Configuration
4.2.1 Propulsion System
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
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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.
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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
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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:
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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
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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
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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
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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
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
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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
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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.
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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.
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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.
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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.
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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
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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.
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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.
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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.
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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,
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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.
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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
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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.
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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.
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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.
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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:
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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
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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.
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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
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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.
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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
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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.
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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.
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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
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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
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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.
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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
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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 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
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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
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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.
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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
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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
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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
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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
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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
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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
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could not be completed due to inclement Rochester weather. Project did,
however, get completed by the time of this report.
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