2015 SAE AERO DESIGN - WEST COMPETITION

MICRO CLASS DESIGN REPORT: L-406 SKYCRANE

PUPR Aero DesignPolytechnic University of Puerto Rico

Team Number: 329March 9, 2015

Table of Contents1 | P a g e

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 Strategy Analysis.............................................................................................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, Payload Prediction, 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.........................................................................................................................24

Appendix A – Supporting Documentation and Backup Calculations................................................................25

Appendix B – Payload Prediction Graph...........................................................................................................27

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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Executive Summary

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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Schedule Summary

October Conceptual Design

Novembe

r

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

Referenced Documents References Specifications

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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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-15.00 -5.00 5.00 15.00 25.000

0.5

1

1.5

2

2.5

3 - D Li ft Curve Sl ope s [Al l A i rfoi l s ]

NACA 6409 CH10 E423

Angle of Attack (deg)

Lift C

oeffi

cient

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

selected for the tail. This is desired because symmetrical airfoils have identical upper and lower

surfaces, and find applications in V-tail designs, 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.

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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 GeneralAirfoil: NACA 6409 Airfoil: NACA 0012 Empty Weight: Approx. 3 pounds

Span: 45 inches Span: 10.4 inches Taper Ratio: 0.4

Reference Area: 259 in2Reference 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.

0.50 0.55 0.60 0.65 0.70 0.75 0.8050.00

55.00

60.00

65.00

70.00

75.00

80.00

85.00

90.00

95.00Fl i ght Score vs Payl oad Fracti on

10 in Con-tainer15 in Con-tainer18 in Con-tainer20 in Con-tainer

Figure 5: Flight Score vs. Payload Fraction

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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. 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 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.15 | P a g e

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=2CL, wπ ARw

-10.00 -5.00 0.00 5.00 10.00 15.00 20.00 25.00 30.000.00

2.00

4.00

6.00

8.00

10.00

12.00

f(x) = 0.332648817559747 x + 2.16221731413836

Downwash vs. Angle of Attack

Angle of Attack [degrees]

Dow

nwas

h [d

egre

es]

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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iv. Dynamic & Static Stability

An important measure of the tail effectiveness is the horizontal tail volume coefficient, shown

in the following equation.

V H=SH lHSc

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. Aircraft Design: A Conceptual Approach, Fifth Edition.3 Graph shown in Additional Material (page 28)

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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.18 | P a g e

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 will be joined with nylon screws. The payload will be carried inside the 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 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 West competition.

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.

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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

0.000 0.500 1.000 1.5000.00

2.00

4.00

6.00

8.00

10.00

12.00

0.000

2.000

4.000

6.000

8.000

10.000

12.000

Lift-to-Drag Ratio Drag Polar

CL

L/D

Ratio

CD

Figure 10: Lift-to-Drag Ratio vs. Lift & Drag Coefficients (NACA 6409)

0.00 5.00 10.00 15.00 20.00 25.000.00

1.00

2.00

3.00

4.00

5.00

6.00

7.00

8.00

9.00

f(x) = − 0.365550595621114 x + 8.24253452380857

Aircraft Airspeed, V0 (mph)

Thru

st, F

(lbf

)

Figure 11: Dynamic Thrust vs. Aircraft Speed

V Load Roll Angle φ Turn Turn Rate

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[mph]

Factor nmax

(degrees) Radius (feet) (degrees/s)

40 1.08 22 259 13.045 1.37 43 145 26.150 1.69 54 123 34.355 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 6390.64 14.04 0.071 4.08 5.21 7020.58 15.40 0.065 3.72 4.75 7700.54 16.79 0.060 3.41 4.36 8400.56 16.01 0.062 3.57 4.57 8010.57 15.80 0.063 3.62 4.63 7900.84 10.66 0.094 5.36 6.85 5331.06 8.52 0.117 6.70 8.55 4261.27 7.08 0.141 8.04 10.25 354

Table 7: Landing Performance

Figure 12: Dynamic Thrust Equation

Appendix B – Payload Prediction Graph

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300 700 1100 1500 1900 2300 2700 31005.40

5.60

5.80

6.00

6.20

6.40

6.60

6.80

f(x) = − 0.000445277478250225 x + 6.77441787691382

Payload Prediction Graph at Maximum Velocity

Density Altitude [slug/ft3]

Payl

oad

Wei

ght [

lbf]

Figure 13: Payload Prediction

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Additional Material

Figure 14: Cubic Loading vs. Aircraft Empty Weight

30.00 40.00 50.00 60.00 70.00 80.00 90.0010.00

11.00

12.00

13.00

14.00

15.00

16.00

17.00

Stall Speed [mph]

Lift

-to-

Dra

g Ra

tio

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