1. 2015 SAE AERO DESIGN - WEST COMPETITION
MICRO CLASS DESIGN REPORT: L-406 SKYCRANE
PUPR Aero Design
Polytechnic University of Puerto Rico
Team Number: 329
March 9, 2015

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Table of Contents
List of Figures and Tables ..........................................................................................................................3
Executive Summary ..................................................................................................................................4
Schedule Summary...................................................................................................................................5
1. Loads and Environments, Assumptions............................................................................................7
i. Design Loads Derivations .....................................................................................................................7
ii. Environmental Considerations..............................................................................................................8
2. Design Layout & Trades..................................................................................................................9
i. Overall Design Layout and Size.............................................................................................................9
ii. Optimization (Sensitivities, System of systems: planform, layout, power plant,etc.)..............................11
a) Competitive Scoring and StrategyAnalysis ......................................................................................12
iii. Design Features and Details ...............................................................................................................13
iv. Interfaces and Attachments ...............................................................................................................13
3. Analysis.......................................................................................................................................14
i. Analysis Techniques...........................................................................................................................14
a) Analytical Tools..............................................................................................................................14
b) Developed Models.........................................................................................................................14
6.2. Performance Analysis.......................................................................................................................15
i. Runway/Launch/Landing Performance............................................................................................15
ii. Flight and Maneuver Performance..................................................................................................15
iii. Downwash ....................................................................................................................................16
iv. Dynamic & Static Stability...............................................................................................................17
v. Lifting Performance, PayloadPrediction, and Margin.......................................................................17
6.3. Mechanical Analysis.....................................................................................................................18
i. Applied Loads and Critical Margins Discussion.................................................................................18
ii. Mass Properties & Balance.............................................................................................................18
7. Assembly and Subassembly, Test and Integration ..........................................................................19
8. Manufacturing.............................................................................................................................21
9. Conclusion...................................................................................................................................23
List of Symbols and Acronyms .................................................................................................................23
Appendix A – Supporting Documentation and Backup Calculations............................................................24
Appendix B – Payload Prediction Graph ...................................................................................................26
Additional Material.................................................................................................................................28

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List of Figures and Tables
Figure 1: Aircraft Forces in a Level Turn ..................................................................................................... 7
Figure 2: Selected Airfoil ............................................................................................................................. 9
Figure 3: 3-D Lift Curve Slopes .................................................................................................................. 10
Figure 4: Selected Tail Airfoil .................................................................................................................... 11
Figure 5: Flight Score vs. Payload Fraction ............................................................................................... 12
Figure 6: Downwash vs. Angle of Attack................................................................................................... 16
Figure 7: Exploded View of Aircraft .......................................................................................................... 20
Figure 8: 3-D Printed Prototype Fuselage................................................................................................. 21
Figure 9: Assembled Prototype Aircraft.................................................................................................... 22
Figure 10: Lift-to-Drag Ratio vs. Lift & Drag Coefficients (NACA 6409).................................................... 25
Figure 11: Dynamic Thrust vs. Aircraft Speed........................................................................................... 25
Figure 12: Dynamic Thrust Equation......................................................................................................... 26
Figure 13: Payload Prediction ................................................................................................................... 27
Figure 14: Cubic Loading vs. Aircraft Empty Weight................................................................................. 28
Table 1: Schedule Summary........................................................................................................................ 5
Table 2: Referenced Documents, References, and Specifications ............................................................. 6
Table 3: General Aircraft Layout............................................................................................................... 11
Table 4: Performance Margins.................................................................................................................. 15
Table 5: Critical Structural Margins........................................................................................................... 18
Table 6: Level Turn Performance .............................................................................................................. 26
Table 7: Landing Performance .................................................................................................................. 26

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ExecutiveSummary
The Micro Class category requires an aircraft weighing less than 10 pounds that fits within a 6”
diameter container. 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,
an aircraft fitting those parameters was designed and manufactured using additive manufacturing.
Our team goal for this competition was to reach a high payload fraction: an approximate value
of 80%. The innovation that the team developed for this Micro Class Competition was a totally 3-D
Printed aircraft. This was done to achieve a better payload fraction by reducing the airplane’s weight. In
addition, it accelerated the manufacturing process.

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ScheduleSummary
October  Conceptual Design
November
 Preliminary Design
 Airfoil and Wing/Tail Geometry Selection
December
 Fuselage Geometry Selection
 Engineering Analysis
January
 Prototype Manufacturing
 Engineering Analysis
February
 Prototype Manufacturing
 Final Aircraft Design Settled
March
 First Prototype Flight Test
 Design Report Conclusion & Submission
April
 Aircraft Assembly Strategy
 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
Given that our aircraft has a non-retractable propeller, it should be landed 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. It can be seen that 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 was manufactured completely out of PLA using a 3-D printer, and it is
suggested that this material should not be exposed to areas of high humidity for long periods of time,
since it can absorb the water in the environment, and thus adding more weight to the structure.
Due to the mountains that surround the field, lack of air pressure is also being taken into
consideration, something our pilot is aware of. The temperature during the time of the event is said to
be in an average of 23°C and the modest elevations, there will be no problems with the flight path or the
aircraft’s performance.
Due to the limited wind information we had available, we decided to test the prototype in the
harshest wind conditions in the PR metropolitan area (approximately 20 knots).

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2. Design Layout & Trades
i. Overall Design Layout and Size
The design process is considered a critical activity, because it becomes clear that the
manufacturing and cost processes are determined by the decisions made in the initial design stages. By
pointing out the stated requirements, the project execution was made possible.
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 this, a
combination of an airfoil capable of creating the necessary lift with high-lift devices was decided upon.
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 CH10 and E423 airfoils both have a maximum lift coefficient of 2. The
NACA 6409 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 other airfoils. This is of great
importance since low-speed flight is involved.
Figure 2: Selected Airfoil

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Figure 3: 3-D Lift Curve Slopes
A tapered high wing with an aspect ratio of 8.05 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. Finally, adding sweep was not considered for many reasons: our design will not
operate at very high speeds, and it would not be structurally beneficial.
For stability reasons, a symmetrical airfoil with a projected horizontal aspect ratio of 4.68 is
selectedfor the tail. This is desired because symmetrical airfoils have identical upper and lower surfaces,
and find applications in V-taildesigns,which is the chosen configuration for our aircraft’s tail. To account
for stability, a tail sweep of 30° was incorporated to ensure longitudinal control at the high angles of
attack that this aircraft will be expected to operate at. The NACA 0012 airfoil was selected for structural
and data availability reasons.
0
0.5
1
1.5
2
2.5
-15.00 -5.00 5.00 15.00 25.00
LIFTCOEFFICIENT
ANGLE OF ATTACK (DEG)
3-D Lift Curve Slopes [All Airfoils]
NACA 6409 CH10 E423

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Figure 4: Selected Tail Airfoil
The V-Tail configuration was selected for three reasons:
 Less wetted area, which in turn produce less drag.
 Less material used due to vertical tail elimination.
 Less servos and linkages are required for control surface operations.
ii. Optimization (Sensitivities, System of systems: planform, layout, power plant, etc.)
Wing Tail General
Airfoil: NACA 6409 Airfoil: NACA 0012 Empty Weight: Approx. 3 pounds
Span: 45 inches Span: 10.4 inches Taper Ratio: 0.4
Reference Area: 259 in2
Reference Area:
20.33 in2 (Horiz. Proj.)
6.61 in2 (Vert. Proj.)
Moment Arm: 13.85 in.
Aspect Ratio: 8.05 Aspect Ratio: 4.2 Aircraft Length: 24.04 in.
Taper Ratio: 0.4 Taper Ratio: 0.4 Fuselage Diameter: 5.3 in.
Propeller: 12” diameter x 7” pitch
Table 3: General Aircraft Layout

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a) Competitive Scoring and Strategy Analysis
According to Section 6.5 in the rule guide, the Final Flight Score is mostly dependent on the
payload fraction. Using the Flight Round formula, an interpolation of payload fractions and 4 different
container lengths was achieved. The resulting plots were linear in nature and the equations for each
container size were obtained. Figure shows the Flight Score versus the Payload Fraction for each of these
lengths.
Figure 5: Flight Score vs. Payload Fraction
50.00
55.00
60.00
65.00
70.00
75.00
80.00
85.00
90.00
95.00
0.50 0.55 0.60 0.65 0.70 0.75 0.80
Flight Score vs Payload Fraction
10 in
Container
15 in
Container
18 in
Container
20 in
Container

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iii. Design Features and Details
The design was heavily constrained by the container’s dimensions, so compact, modular design
for fast and easy assembly was implemented. For example, ailerons are removable and the airframe was
constructed by additive manufacturing; the largest part measuring 14.28 inches.
The transmitter was programmed so that two control surfaces have the function as rudders and
elevators for the tail, and ailerons and flaps for the wing. These configurations were decided upon in
order to maximize the wing and tail area that can be fit inside the container.
iv. Interfaces and Attachments
 Custom-made fittings were designed and 3-D printed for junctures of the V-Tail and wings.
 Some of the fittings for the 3-D printed parts were constructed with other materials like wood.
 Tie wraps and nylon screws are considered for use to join the fuselage parts together.

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3. Analysis
i. Analysis Techniques
a) Analytical Tools
Throughout the design process, Creo Parametric was used to simulate our ideas. This helped us
to make decisions based on the information we extract from the CAD. This is also an advantage in terms
of time and budget regarding the design. For example, we could see tolerance errors with the design
without building the parts, and procure whether or not the aircraft would fit into the designed container.
Microsoft Excel was extensively used for aerodynamic and performance analyses. This permitted
the development of data tables and graphs 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
developed to analyze how well one performs compared to the other.
b) Developed Models
A prototype was built to test the flying qualities of the planform chosen for the wings and V-tail.
Also, the programming of the control system (transmitter) was developed using this prototype, thus
avoiding the risk of damaging the final aircraft.

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4. PerformanceAnalysis
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 30 feet per second was assumed. This will give the aircraft the extra push it needs to
achieve the pre-analyzed flight performance. Using Anderson’s text, the landing performance was
calculated. The ground roll was not taken into consideration since the aircraft does not have a landing
gear, and our runway in this case will be grass. Using approximations stated by the book, such as the
approach angle, the estimated landing distance from a 50-foot obstacle was determined to be 974 feet.
ii. Flight and Maneuver Performance
The installed motor will provide approximately 11,160 revolutions per minute (RPM) to the
propeller, with dimensions of 12” diameter and 7” 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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iii. 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 𝐿,𝑤
𝜋𝐴𝑅 𝑤
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.
y = 0.3326x + 2.1622
-2.00
0.00
2.00
4.00
6.00
8.00
10.00
12.00
-10.00 0.00 10.00 20.00 30.00
DOWNWASH[DEGREES]
ANGLE OF ATTACK [DEGREES]
Downwash vs. Angle of Attack

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iv. Dynamic & Static Stability
An important measure of the tail effectiveness is the horizontal tail volume coefficient, shown in
the following equation.
𝑉𝐻 =
𝑆 𝐻 𝑙 𝐻
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 0.5 and 0.04, 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 positive static margin: an approximate
value of 10% was obtained.
v. Lifting Performance, Payload Prediction, and Margin
We researched several heavy-lift airplanes and saw that none of them would exceed a cubic
loading of 3.0, and in fact, a payload fraction above 80% was obtained with an airplane with a cubic
loading of 2.76. Therefore, we used 80% 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 add flaperons to during the
assembly. Figure 133 illustrates the sensitivity of empty weight to cubic loading and payload fraction.
2 Table 6.4, Page 160. AircraftDesign:A Conceptual Approach,Fifth Edition.
3 Graph shown in Additional Material (page28)

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5. Mechanical Analysis
Aircraft must stay intact during flight and support all dynamic loads.
Aircraft must support variable 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 three rounds of the competition, the aircraft will have to carry an increasing payload
every round, 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 prediction of the airframe was performed using Creo Parametric. Inputting the values
of the material properties of the PLA, a full report was obtained.
4 See Appendix B.

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6. Assembly and Subassembly, Test and Integration
The aircraft will be divided into 3 assembly points each: fuselage, wings and tail.
Fuselage
The fuselage was printed in 5 different sections; two of them will be permanently joined. Each
section willbe joined with nylon screws.The payload willbe carried insidethe center section. The frontal
section will contain the motor and propeller. The rear section will hold the tail assembly piece and will
hold them together with 2 nylon screws. Also the wiring for all the electrical components will mostly be
inside the fuselage.
Wings
The wings will be divided into a total of 12 pieces: 3 for one wing and 3 for the “flaperons”, and
two sections will be permanently joined together. These will be attached to the fuselage using
rectangular spars. Each wing has two channels in which the spars will be passed from one wing to the
other, passing through the top part of the middle section of the fuselage. The “flaperons” for each wing
will be attached with hinges to improve the stability.
Tail
The tail will be divided into 6 pieces. The control surfaces on the tail wings will be attached the
same way as for the wings. When these are attached, the tail wings will be placed in between 2 pieces
that will hold the fuselage together with 2 nylon screws.

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Electronic components
4 servos will be installed for each control surface: two for the “flaperons” and two for the V-tail.
A receiver, antenna, 11.1 volt lithium polymer battery, controller with BEC system, outrunner brushless
motor will be the primary electronic components used.
Figure 7: Exploded View of Aircraft

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7. Manufacturing
The aircraft was fabricated using additive manufacturing. This was decided because when using
wood, the manufacturing of each piece would have required 2 to 3 weeks. When using 3-D printing, the
manufacturing of the aircraft took nearly 23 hours. The material the team chose was PLA because it is
cost efficient, easier to manufacture, and lighter than wood. This also gives us the advantage of lighter
structures throughout the airframe.
Figure 8: 3-D Printed Prototype Fuselage
The wings were manufactured in 6 sections per wing. Each section of the wing was printed at a
length of 7.15”. The wings contain 2 spars, one measuring 45” long and the other measuring 27” long,
each crossing from one wing to the other. Each section of the wings will have an interlocking attachment
to help 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.

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The base of the tail, which has a 0.50” diameter and 1” length tube on one of its faces, was
inserted into the fuselage.The base’s dimensions are 2”diameter and a 2.40” length. The V-tail with each
tail wing have dimensions of 5.50” of width, and 1.95” in depth.
Figure 9: Assembled Prototype Aircraft

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8. Conclusion
The PUPR Aero Designteam has conducted acomplete conceptual design,performed a thorough
engineering analysis, and completed the construction of a final design that will meet the requirements
laidout by the Society of Automotive Engineers for the Aero DesignWestcompetition. With a low empty
weight and a smooth, streamlined body, the “L-406 Skycrane” is more than prepared to take to the skies
in the April competition. The aircraft is extremely lightweight, aerodynamically efficient, and stable.
List of Symbols and Acronyms

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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
AppendixA – Supporting Documentation and Backup Calculations

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Figure 10: Lift-to-Drag Ratio vs. Lift & Drag Coefficients (NACA 6409)
Figure 11: Dynamic Thrust vs. Aircraft Speed
V
[mph]
Load
Factor
nmax
Roll Angle φ
(degrees)
Turn
Radius
(feet)
Turn Rate
(degrees/s)
0.000 0.500 1.000 1.500
0.000
0.020
0.040
0.060
0.080
0.100
0.120
0.140
0.160
0.00
2.00
4.00
6.00
8.00
10.00
12.00
14.00
16.00
18.00
CD
L/DRATIO
CL
Lift-to-Drag Ratio Drag Polar
F = -0.3656*V0 + 8.2425
0.00
1.00
2.00
3.00
4.00
5.00
6.00
7.00
8.00
9.00
0.00 5.00 10.00 15.00 20.00 25.00
Thrust,F(lbf)
Aircraft Airspeed, V0 (mph)

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40 1.08 22 259 13.0
45 1.37 43 145 26.1
50 1.69 54 123 34.3
55 2.05 61 113 40.8
Table 6: Level Turn Performance
Required
Thrust (Drag)
[pounds]
Lift-to-
Drag
Ratio
Thrust-to-
Weight
Ratio
Glide
Angle
(degrees)
Sink/Climb
Rate @ 50
mph [feet/s]
Range
(50 foot
obstacle)
[feet]
0.70 12.78 0.078 4.48 5.72 639
0.64 14.04 0.071 4.08 5.21 702
0.58 15.40 0.065 3.72 4.75 770
0.54 16.79 0.060 3.41 4.36 840
0.56 16.01 0.062 3.57 4.57 801
0.57 15.80 0.063 3.62 4.63 790
0.84 10.66 0.094 5.36 6.85 533
1.06 8.52 0.117 6.70 8.55 426
1.27 7.08 0.141 8.04 10.25 354
Table 7: Landing Performance
Figure 12: Dynamic Thrust Equation
AppendixB – Payload Prediction Graph

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Figure 13: Payload Prediction
PW = -0.0004hdensity+ 6.7744
5.40
5.60
5.80
6.00
6.20
6.40
6.60
6.80
300 700 1100 1500 1900 2300 2700 3100
PayloadWeight[lbf]
Density Altitude [slug/ft3]
Payload Prediction Graph at Maximum Velocity

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AdditionalMaterial
Figure 14: Cubic Loading vs. Aircraft Empty Weight
10.00
11.00
12.00
13.00
14.00
15.00
16.00
17.00
30.00 40.00 50.00 60.00 70.00 80.00 90.00
Lift-to-DragRatio
Stall Speed [mph]