1. PUPR Aero Design
Polytechnic University of Puerto Rico
Team Number: 326
January 25, 2016

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Table of Contents
Contents
Executive Summary .................................................................................................................................................. 3
Schedule Summary ................................................................................................................................................... 4
1. Loads and Environments, Assumptions ............................................................................................................. 6
i. Design Loads Derivations ................................................................................................................................ 6
ii. Environmental Considerations ...................................................................................................................... 7
2. Design Layout & Trades ..................................................................................................................................... 7
i. Overall Design Layout and Size ....................................................................................................................... 7
ii. Optimization (Sensitivities, System of systems: planform, layout, power plant, etc.) ................................ 10
a) Competitive Scoring and Strategy Analysis ............................................................................................. 10
3. Analysis ............................................................................................................................................................ 11
i. Analysis Techniques ...................................................................................................................................... 11
a) Analytical Tools ....................................................................................................................................... 11
4. Performance Analysis ...................................................................................................................................... 12
i. Runway/Launch/Landing Performance .................................................................................................... 12
ii. Flight and Maneuver Performance ............................................................................................................. 13
ii. Downwash ....................................................................................................................................................... 13
ii. Dynamic & Static Stability ........................................................................................................................ 14
iii. Lifting Performance, and Margin ............................................................................................................ 14
5. Mechanical Analysis ........................................................................................................................................ 15
i. Applied Loads and Critical Margins Discussion ......................................................................................... 15
ii. Mass Properties & Balance ...................................................................................................................... 15
6. Manufacturing ................................................................................................................................................. 19
7. Assembly and Subassembly, Test and Integration .............................................................................................. 20
8. Conclusion .......................................................................................................................................................... 21
List of Symbols and Acronyms ................................................................................................................................ 22
Appendix A – Supporting Documentation and Backup Calculations ...................................................................... 23
Additional Material ................................................................................................................................................. 24

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Executive Summary
The Micro Class category requires an aircraft weighing less than 10 pounds that fits within
container of a 6” outer diameter. The goal is to have the highest payload fraction possible with the
lowest empty weight that a design will allow. This type of electric aircraft has to be hand-launched. For
this purpose, a conventional aircraft fitting those parameters was designed and manufactured. The
design consists of only three materials: balsa wood, carbon fiber, and monokote.
Our team goal for this competition was to reach a high payload fraction: an approximate value
of 90%.

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Schedule Summary
September
● Team Organization
● High lift capacity aircraft research
October
● Preliminary Design
● Airfoil Selection for Wing
November
● Fuselage Sizing
● Engineering Analysis
December ● First Prototype Flight Test
January
● Container Design
● Design Report Conclusion & Submission
February ● Aircraft Assembly Strategy
March ● Final Competition Preparations
Table 1: Schedule Summary

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Referenced Documents References Specifications
Estimating R/C Model
Aerodynamics and
Performance; Nicolai
Aircraft Design: A
Conceptual Approach;
Raymer
Payload dimensions: 1.5” x 1.5” x 5”
SAE Aero Design East and West
Rules
Mechanics of Flight:
Second Edition; Warren
Phillips
Desired high payload fraction
Tail Design; Mohammad
Sadraey
Aircraft Performance and
Design; John D. Anderson
Aircraft must be assembled in less
than 150 seconds
Propeller Static & Dynamic
Thrust Calculator; Gabriel
Staples
Introduction to Flight;
John D. Anderson
The fully packed aircraft system
container shall weigh no more than
10 pounds
Shigley’s Mechanical
Engineering Design
Aircraft container must have a
maximum diameter of 6”
Table 2: Referenced Documents, References, and Specifications

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1. Loads and Environments, Assumptions
i. Design Loads Derivations
The aircraft should be capable of landing over grassy areas to reduce the risk of breaking. The
aircraft will experience accelerations and decelerations during the flight course, such as when it is
clearing the 180° turns, in addition to centripetal forces, shown in the figure below.
Figure 1: Aircraft Forces in a Level Turn
Here, the aircraft is performing a level turn. As seen in Figure 1, the lift is inversely proportional
to the bank (roll) angle. In manned flight applications, this is the orthogonal force that the pilot will
experience when he is pulling up on the aircraft. For the flight course, operational precautions must be
taken into account to reduce this force so as to avoid any structural failures to the aircraft.

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ii. Environmental Considerations
Based on our design, several aspects of the location’s weather conditions were taken into
consideration. The aircraft is structurally made out of balsa wood that acts as the skeleton of the plane
but is almost completely covered in monokote. Balsa wood was chosen as the primary material for
the skeleton of the plane because of its high strength to weight ratio. It is suggested that this material
should not be exposed to areas of high humidity for long periods of time. Exposure to water may cause
the plane to absorb the fluid from the environment, and thus add more weight to the structure; the
monokote should protect the plane against this.
Because our design does not include any landing gear, it is preferred to only fly over grass fields
to reduce the risk of damaging the fuselage upon landing.
2. Design Layout & Trades
i. Overall Design Layout and Size
The design process is considered a critical activity. Manufacturing and cost processes are
determined by the decisions made in the initial design stages. After reviewing the requirements stated
in the competition, the project execution was made possible. After reviewing the requirements, it was
assessed that a high payload fraction is the most important variable used in calculating our final score.
To achieve a high payload fraction value, it is desired to decrease the wing loading as much as possible;
however, the wing area is constrained by the container’s diameter. To compensate for constraint, a

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combination of an airfoil capable of creating the necessary lift as well as lightweight materials were
was selected.
An extensive analysis of different airfoils was conducted at a Reynolds number of
approximately 100,000. Figure illustrates the lift curves slopes. The 3-D aerodynamic effects were
already taken into consideration in the analysis. The S1223 and flat plate airfoils have a maximum lift
coefficient of 2 and 1.10; the NACA S1223 airfoil was selected because, as it can be seen, although it
has a moderate maximum lift coefficient, it will not stall immediately at high angles of attack, unlike
the AG03. This is of great importance since low-speed flight is involved.

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+
A straight high wing with a wingspan of 36 inches was selected as the final wing configuration due to
it being more structurally and aerodynamically efficient than a constant chord wing. A wing of this type
would have produced a non-elliptical lift distribution and the bending moments would have been more
severe. Also, the addition of wing twist would have increased the volume necessary for the wing to fit
in the container. Adding sweep was not considered for many reasons: our design will not operate at
very high speeds, and it would not be structurally beneficial.
A T-tail was chosen for its reduction in vertical tail size due to the end-plate effect. For stability
reasons, a symmetrical airfoil, with an aspect ratio of 3.425 was chosen. This is due to symmetrical
airfoil behavior during both positive and negative angle of attack. The configuration will improve the
efficiency because of less downwash affecting it, due to the clearance between the wing and the tail.
The NACA 0012 airfoil was selected for structural and data availability reasons.

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ii. Optimization (Sensitivities, System of systems: planform, layout, power plant, etc.)
Wing Tail General
Airfoil: NACA S1223 Airfoil: NACA 0012
Empty Weight: Approx. 1.6
pounds
Span: 36 inches Span: 17.125 inches Taper Ratio: 0
Reference Area: 223 in2
Reference Area:
40 in2
(Horiz. Proj.)
17.5 in2
(Vert. Proj.)
Moment Arm: 19.34 in.
Aspect Ratio: 5.8 Aspect Ratio: 3.425 Aircraft Length: 21.86 in.
Taper Ratio: 0 Taper Ratio: 0 Max Fuselage Diameter: 3.74 in.
Propeller: 11” diameter x 7” pitch
Table 3: General Aircraft Layout
a) Competitive Scoring and Strategy Analysis
According to Section 9.8 in the rule guide, the Final Flight Score is mostly dependent on the
payload fraction; a high payload can result in a higher flight. A high payload fraction can be achieved by
designing an aircraft that is lightweight that can carry the highest
payload possible. Lightweight materials with high strength to weight
ratios for the structure of the plane were selected. For the payload,
Tungsten was the material of choice due to its high density. Figure 8
located below shows 3 materials that were considered, all of which
have the same weight.
FIgure 8: Payload Material comparison

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Design Features and Details
The design was heavily constrained the team’s goal of a achieving a high payload fraction and
low stall velocity. The payload faction and stall velocity that the team aspired for was one of 0.9 and
about 20 miles per hour. To attempt to achieve a high payload fraction, balsa wood, carbon fiber and
monokote were the only materials selected for the building of the aircraft. A combination of an airfoil
with a high camber with a low Reynolds number and a large area on the wing was selected in order to
achieve high lift at lower speeds.
3. Analysis
i. Analysis Techniques
a) Analytical Tools
Throughout the design process, Creo Parametric was used to simulate our ideas. Creo was
initially used to monitor the sizing in the design. It helped us make decisions based on the information
we extracted from the CAD; this information includes but is not limited to location of the center of
gravity and weight of the aircraft. Using Creo, we were able to save time and money in testing. For
example, we could see tolerance errors in the design without having to manufacture any parts and as
well as physically test if the payload, motors and other electrical components would fit in the aircraft.
Microsoft Excel was extensively used for aerodynamic and performance analysis. Using excel,
data tables and graphs were developed to predict and optimize the behavior of our design. For
example, graphs comparing lift coefficients, lift-to-drag ratios and the drag polar for each airfoil were

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developed to analyze how well one performs compared to the other. Microsoft Excel was also
extensively used for the stability calculations of the aircraft.
4. Performance Analysis
Aircraft must be hand-launched.
Aircraft is required to remain airborne and fly past the designated turn points, perform the two 180°
turns in heading, and arrive at the landing zone.
The aircraft must take off and land intact to receive points for the flight.
All parts must remain attached to the aircraft during flight and during the landing maneuver.
Aircraft must land in a designated landing zone measuring 200 feet in length.
Table 4: Performance Margins
i. Runway/Launch/Landing Performance
The aircraft will be hand-launched, according to the stated requirements by SAE. An estimated
launch speed of 35 miles per hour is needed for the aircraft to fly. The installed motor will provide
approximately 11,160 revolutions per minute (RPM) to the propeller, with dimensions of 11” diameter
and 8” pitch. This, in turn, will operate the aircraft at a range of speeds between 40 and 55 miles per
hour (MPH). Since one of the competition objectives is to clear two 180° turns, the turn rate needs to
be compensated for the load factor so as to avoid wing support failure. The range for unpowered flight
was determined assuming that the maximum flying altitude is 50 feet.1
1
See Appendix A for calculations.

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ii. Flight and Maneuver Performance
The flight characteristic for the aircraft allows the plane to fly at maximum weight capacity at stall
speeds and maneuver at a turn rate of 27 degrees per second. Maneuvering for the craft is done using flaperons
at .35 chord for the full wingspan and a horizontal stabilator for a reduction in weight.
ii. Downwash
The downwash angle of a typical wing is a function of its sectional lift coefficient and aspect
ratio, and can be approximated by the following equation.
𝜖 =
2𝐶!,!
𝜋𝐴𝑅!
Figure 6: Downwash vs. Angle of Attack
It is observed from the above figure that the downwash experienced by the wing is directly
proportional to its angle of attack. This is a consequence of the increasing lift in the wing. Too much
downwash can create a turbulent airflow over the tail, negatively impacting its performance.

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ii. Dynamic & Static Stability
An important measure of the tail effectiveness is the horizontal tail volume coefficient, shown
in the following equation.
𝑉! =
𝑆! 𝑙!
𝑆𝑐
SH is the horizontal stabilizer planform area, lH is the horizontal stabilizer moment arm, S is the
wing planform, and c is the wing chord. For this aircraft, the chosen tail volume coefficients for the
horizontal and vertical tails were 0.63 and 0.036, respectively. These values were picked for a
homebuilt aircraft.2
Another important parameter required for stability is the location of the aircraft’s center of
gravity. The wing was placed on a location that would provide a static margin of approximate value of
14.76% was obtained.
iii. Lifting Performance, and Margin
We used our previous research on thirteen heavy-lift airplanes which demonstrates that none
of them would exceed a cubic loading of 3.20, and in fact, a payload fraction above 80% was obtained
with an airplane with a cubic loading of 2.76. Therefore, we used 90% payload fraction and a maximum
cubic loading of 3.0 as our goal using the largest wing area we could fit in the container, that we could
2
Table 6.4, Page 160. Aircraft Design: A Conceptual Approach, Fifth Edition.

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add flaperons to during the assembly. Figure 133
illustrates the sensitivity of empty weight to cubic
loading and payload fraction.
5. Mechanical Analysis
Aircraft must stay intact during flight and support all dynamic loads.
Aircraft must support Fixed payloads according to SAE requirements.
The aircraft must take off and land intact to receive points for the flight.
Broken propellers are allowed.
Table 5: Critical Structural Margins
i. Applied Loads and Critical Margins Discussion
During the Four rounds of the competition, the aircraft will have to carry a fixed payload in the
rounds chosen to do so; this will be done to test how well the aircraft is designed. In addition, as
mentioned in Section 3.1 of this report, the aircraft needs to support the forces encountered when
executing level turns, such as the G-forces. Table 54
shows the calculated level turn parameters for the
installed motor’s speed limits. Our airplane was designed to routinely sustain a G-force of 2.0 by
assuming a 3.0 ultimate load factor limit.
ii. Mass Properties & Balance
The weight and balance was determined applying a simple exercise. We plan to weigh every
component including the electrical, and the airplane’s main. The moment arms were also measured
with respect to the airplanes datum or nose. With that we determine the center of gravity of the
3
Graph shown in Additional Material (page 28)
4
See Appendix B.

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airplane. This lead our plane to comply with a new challenge, match the MAC of the wing to and find a
negative static margin to have a balanced and stable aircraft
Assembly and Subassembly
There are 3 assembly points each: fuselage, wings and tail. Each of the assembly points will be
wrapped completely in monokote.
Fuselage
The fuselage will be assembled using eleven laser cutted balsa wood parts and wrapped in
monokote. All eleven of the pieces will be permanently joined using glue. The payload will be carried
inside of the fuselage, on a payload bay installed on the center of gravity of the plane, exactly 5.26
inches from tip of the fuselage and 2.78 inches from the leading edge of the wing . There will be spinal
ribs in the fuselage that will support the payload bay, keeping it vertically centered.

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Wings
The wings will have two parts that will be divided into a total of seventeen pieces that will be
glued together. The pieces include twelve laser cutted balsa wood ribs for one wing; each wing will
carry one “flaperon”, and two sections will be permanently joined together. The wings will be
structurally held together using 0.03mm OD carbon fiber road going through the ribs and a spar made
from balsa wood on the trailing edge of the wing.
Tail
The tail will be divided into eleven pieces. Eight of the eleven pieces are ribs made out of balsa
wood. The Vertical stabilizer will also be made of balsa wood but will also have several holes to
decrease the weight. Made one carbon fiber rod to hold the structure together.

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Electronic components
Three servos will be installed for each control surface: two for the “flaperons” on the wing and
one elevator for the T-tail. A receiver, antenna, 11.1-volt lithium polymer battery, a Castle creations
Phoenix edge 75 34V 75A ESC w/5A BEC, and a AXI Gold 2212/34 Outrunner brushless motor will be
the primary electronic components used.
Figure 7: Exploded View of Aircraft

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6. Manufacturing
The aircraft will be fabricated using subtractive manufacturing. The wings will be manufactured
in 2 sections. The rib was laser cut making 10 ribs, 5 on each section of the wing. Each rib will have a
root chord of 6.2 inches. The wings contain 1 spar measuring 36” long that will cross the leading edge
of the wing. To manufacture the spar, the team will use a balsa wood block and will mold the block to
the curve of the leading edge of a NACA S1223 airfoil. Each section of the wings will be glued to hold
the joints together and clamps will be used whenever needed in the assembly process and also to
resist axial loads that might be applied to the wings. These attachments were primarily made to be
able to connect each section of the wing to each other and the fuselage.
8 identical ribs made from balsa wood will be laser cut for the T-tail. The carbon fiber rod will
be purchased having a diameter of 0.125 inches. The vertical stabilizer will also be laser cut and
molded to have the curve of a NACA 0012. The base of the tail and fuselage will be connected through
a boom made out of 0.125 inches of carbon.

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7. Assembly and Subassembly, Test and Integration
Assembly and Subassembly
We are currently in the process of designing our aircraft’s assembly and subassembly strategy.
Since we have not manufactured the airplane yet, we have not being able to do flight tests, but we will
do so as soon as we complete the manufacturing phase of the aircraft.

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8. Conclusion
The PUPR Aero Design team has conducted a complete conceptual design, performed a
thorough engineering analysis, and completed the construction of a final design that will meet the
requirements laid out by the Society of Automotive Engineers for the Aero Design East competition.
With a low empty weight and a smooth, streamlined body, the “Skycranes” is more than prepared to
take to the skies in the March competition. The aircraft is extremely lightweight, aerodynamically
efficient, and stable.

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List of Symbols and Acronyms
AR Aspect ratio
W Aircraft weight
α Angle of attack
CL Lift coefficient
MAC Mean aerodynamic chord
λ Taper ratio
D Total drag
L Total lift
V Velocity
S Wing area
c Wing chord
b Wingspan
α0 Zero-lift angle of attack
CD 3D Polar Drag

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Appendix A – Supporting Documentation and Backup Calculations
Figure 10: Lift-to-Drag Ratio vs. Lift & Drag Coefficients (NACA 6409)
Figure 11: Dynamic Thrust vs. Aircraft Speed

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Additional Material