fornia, davismae.engr.ucdavis.edu/chattot/sae_aero_design/amat13/amat... · 2014-05-28 · system...

Team Captains:

Ilya Anishchenko

Alex Beckerman

Logan Halstrom

Team Members:

Michael Wachenschwanz (DAS lead)

Gene Ang Chris Lorenzen

Joshua Barram Kelley Lundquist

Max Bern Arlene Macias

James Dionisopoulos Robyn Murray

Louis Edelman Nohtal Partansky

Robert Edwards Jason Petersen

Hashmatullah Hasseeb Patricia Revolinsky

S. Sheida Hosseini Adam Simko

Steven Hung Stefan Turkowski

Sara Langberg

Faculty Advisors:

Jean-Jacques Chattot, PhD

Stephen K. Robinson, PhD

Advanced Modeling Aeronautics Team Advanced Class, Team 215

Un

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of

Cali

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AM

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20

13

Advanced Modeling Aeronautics Team 2013

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Table of Contents________________________________________________________

Statement of Compliance …………………………………………………………………... 2

List of Figures and Tables ………………………………………………………………….. 4

Nomenclature ……………………………………………………………………………..… 4

1.0 Introduction ….………………………………………………………………………….. 5

2.0 Design Process ………………………………………………………………………...… 6

2.1 Design Considerations ...………………………………………………………………... 6

2.1.1 Aircraft Design ………………. ……...………………………………………. 6

2.1.2 Avionics Design ……………... ……...………………………………………. 8

2.2 Previous Design ………………………………………………………………………..... 8

2.3 Experiments and Iterations …………………………………………………………….. 9

2.3.1 Prototyping and Material Testing ……...………………………………………. 9

2.3.2 Engine Performance and Propeller Sizing …………………………………....... 11

2.3.3 Model Testing ………………………………………………………………….. 11

2.4 Final Selection …………………………………………………………………………... 12

3.0 Aerodynamics and Performance Analysis ………………………………………..…… 13

3.1 Aircraft Sizing ………………………………………………………………………...… 13

3.1.1 Selection of Main Wing ……………………………………………………….. 13

3.1.2 Selection of Overall Aircraft Configuration …………………………………… 14

3.1.3 Winglet Design ………………………………………………………………… 15

3.2 Aircraft Stability and Control ………………….……………………………………… 16

3.2.1 Longitudinal Stability and Control …………………………………………….. 16

3.2.2 Lateral Stability and Control …………………………………………………… 17

3.2.3 Directional Stability and Control ………………………………………………. 17

3.2.4 Control Surface Sizing …………………………………………………………. 18

3.3 Aircraft Performance …………………………………………………………………… 19

3.4 Payload Analysis ..………………………………………………………………………. 21

3.4.1 Payload Prediction ……………………………………………………………... 22

3.4.2 Payload Drop …………………………………………………………………… 22

4.0 Component Structure and Manufacturing …………………………………………… 23

4.1 Main Wing …………………………………………………………………………….… 23

4.2 Empennage ……………………………………………………………………………… 24

4.3 Fuselage ………………………………………………………………………………….. 25

4.4 Landing Gear ……………………………………………………………………………. 26

4.5 Weight Management ……………………………………………………………………. 27

5.0 Conclusion ………………………………………………………………………………. 28

6.0 References ……………………………………………………………………………….. 29

Appendix – Technical Plans ……..…………………………………………………………. 30


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List of Figures and Tables_________________________________________________

Figure 2.3A – Mini Plane demonstrating hybrid wing construction …………………………………. 9

Figure 2.3B – Wing bending stress test ……………………………………………………………… 9

Figure 2.3C – Ideal layup configuration for fuselage and landing gear ……………………………... 10

Figure 2.3D – Static thrust test stand with O.S. 46 AXII ……………………………………………. 10

Figure 2.3E – Static thrust performance of the O.S. 46 AXII ………………………………………... 11

Figure 2.3F – Prototype Fuselage and tail boom …………………………………………………….. 11

Figure 2.3G – Experimental polar apparatus …………………………………………………………

Figure 3.1A – Non-dimensionalized geometry of the Selig 1223 airfoil …………………………….. 13

Figure 3.1B – 2D and 3D Polars and Lift Slope Curves of the AMAT 2013 Main Wing …………... 14

Figure 3.1C – Effect of Winglet Addition on 2013 AMAT design ………………………………….. 15

Figure 3.2A – Center of gravity and aerodynamic center for entire configuration ………………….. 16

Figure 3.2B – Aileron chord required to counteract a range of gust velocities ……………………… 18

Figure 3.3A – Comparison of available power to power required for equilibrium flight ……………. 20

Figure 3.4A – Effect of increasing density altitude on maximum payload cabability ……………….. 21

Figure 3.4B – Sandbag trajectory prediction for drop from cruise speed ……………………………. 22

Figure 4.1A – Main wing half section ……………………………………………………………….. 23

Figure 4.2A - Empennage without monokote ………………………………………………………... 24

Figure 4.3A – Fuselage and landing gear ……………………………………………………………. 25

Figure 4.4A – Machining of the airshock ……………………………………………………………. 26

Figure 4.5A - Glass/Epoxy Slurry, Aluminum Fasteners, 5-axis foam mill …………………………. 27

Table 2.1A - Summary of 2013 Aero Design Advanced Class Flight Score Criteria ……………….. 6

Table 2.1B – Design matrices from AMAT design process …………………………………………. 7

Table 2.2A – Lessons from Previous Designs ……………………………………………………….. 8

Table 2.3A – Rear Landing Gear Strut Coupon Mass Comparison ………………………………….. 10

Table 2.3B – Fuselage Wall Coupon Mass Comparison …………………………………………….. 10

Table 2.4A – Final Selection of Various Components ………………………………………………. 12

Table 3.1A – Zephyrus Lifting Surfaces …………………………………………………………….. 15

Table 3.1B – Final Configuration ……………………………………………………………………. 15

Table 3.3A – Zephyrus Fight Characteristics ………………………………………………………... 19

Table 4.2A – Weight (lbf.) Comparison of Zephyrus to 2012 AMAT Design ……………………… 23

Table 4.5A – Zephyrus Component Weight Allocations ……………………………………………. 27

Nomenclature___________________________________________________________ Symbols Symbols Cont. Symbols Cont.

Main wing chord c Coefficient of _ C Force F

Thickness t Equivalent flat plate area f Weight W

Main wing span b Density ρ Thrust T

Main wing area S Newton’s Constant gc Rotation per Minute RPM

Main wing aspect ratio AR Angle of attack α Subscripts

Center of Gravity CG Deflection angle δ Main wing m

Aerodynamic Center AC Velocity V Horizontal stabilizer h

Leading Edge LE Rate of climb RC Vertical stabilizer v

Trailing Edge TE Radius R Fuselage f

Length l Pitch p Aileron a

Longitudinal distance x Power P Take off TO

Lateral distance y Lift L Initial 0

Moment M Drag D Induced i


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1.0_Introduction_________________________________________________________

The Advanced Modeling Aeronautics Team is proud to present the 2013 AMAT aircraft to the SAE

Aero Design West competition. AMAT participates in the Advanced Class, and its mission this year is

to design and construct an aircraft with the purpose of aerial delivery of humanitarian aid. For the

competition, an aid package is represented by a 3 lbf sandbag, because of its similarity in size, weight,

and physical properties to a package of food or supplies. The aircraft is also required to lift a 15 lbf static

weight representative of fuel reserve, since on actual humanitarian missions, the aircraft may be required

to travel long distances to reach those in need. Aircraft design is also dictated by other factors including

an empty weight restriction of 8 lbf and a Data Acquisition System (DAS) capability requirement of

real-time altitude measurements and First Person View (FPV) telemetry transmission. To meet this

challenge, AMAT has created Zephyrus, a radio-controlled, fixed wing aircraft designed by students at

the University of California, Davis. Zephyrus’s characteristic features are its composite material

construction, its angled lifting tail, and its highly sophisticated stability augmentation system. Each of

these traits work in combination with other design features to help the aircraft meet three main design

criteria: light structure, high-lift capability, and precise flight control. Maneuverability of the aircraft is a

new focus for SAE Aero Design, where the previous objective was pure heavy lifting, and was the

justification of AMAT’s previous double element wing design. This year’s competition provides the

interesting challenges of building a stable aircraft, optimizing its control, and strategizing drop

trajectories and pilot communication. Accurate delivery of the package is important because the aid

must arrive intact and undamaged as well as cause no collateral damage. The primary pilot is not

allowed to receive any FPV data from the aircraft they control and must be directed to the target by a

secondary pilot. This could be representative a situation where the aircraft is deployed in the field by the

primary pilot away from the base at which telemetry data can be received, forcing flight by verbal radio

instruction. AMAT has developed Zephyrus for ideal fulfillment of the design requirements and hopes

to be a top competitor at SAE Aero Design West.


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2.0 Design Process________________________________________________________

AMAT utilized past experience, research, and experimentation in its design process to progress from an

assessment of the design criteria to an aircraft design idealized for this challenge.

2.1 Design considerations

As might be done in the design of a full-size aircraft, AMAT divided the design process into two

categories: aircraft systems and electronic systems.

2.1.1 Aircraft Design

AMAT defines a successful airplane design as one that adheres to all stated design objectives, which are

in this case determined from the flight scoring below in Table 2.1A:

AMAT incoperated the three components of the flight score FS into three primary design criteria:

precise flight control (S1), minimized aircraft weight (S2), and lifting capability of maximum takeoff load

WTO = 26 lbf (S3). These criteria drove for weight minimization in all structures and materials and for a

focus on aircraft stability, both in design and control.

To make selections based on these criteria, AMAT elected to utilize decision matrices, which

allowed the team to organize and evaluate the many design options in a systematic manner. This method

for meeting the customer’s design specifications is a common practice in industry, and its application to

the prescribed rules of this competition was appropriate. A design matrix compares design options by

ranking them against one another in various categories, such as weight and strength. The importance of

each category relative to the design criteria is assigned a percentage, which is used to weigh design

option rankings. The points for each design option are then added and the most beneficial design

((

) ∑

)

( )

( )

Table 2.1A – Summary of 2013 Aero Design Advanced Class Flight Score Criteria


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determined. One example of importance percentage assignments was the team’s favor toward categories

that affected flight stability (such as wing surface precision) because of the definite zero FS received for

a payload drop not accurate to within 50 feet. Similarly, the team predominantly favored weight

reduction because an empty weight of greater than 12 pounds also receives a definite zero FS. Although

many such matrices have been completed, only a select number have been shown below.

AMAT’s rankings in Table 2.1B determined a preliminary selection of aircraft components specialized

to the design criteria. This selection was further refined in the team’s design process.

Main Wing Design Matrix

Double Element Single Element

Carbon Fiber

and Foam

Core

Monokote and

Balsa Frame Hybrid

Carbon Fiber

and Foam

Core

Monokote and

Balsa Frame Hybrid

Importance

(%)

Weight 2 4 3 1 6 5 40

Surface

Precision 4 1 3 6 2 5 30

Strength 5 3 1 6 4 2 15

Manufacturing 3 5 1 4 6 2 5

Survivability 3 2 1 6 5 4 5

Previous

Experience 6 1 3 5 2 4 5

Score 3.35 2.75 2.5 3.85 4.25 4.3 100

Landing Gear Design Matrix

Tricycle Tail Dragger Importance (%)

Weight 1 2 20

Stability 2 1 20

Need to Accelerate at Zero Incidence 2 1 20

Feasibility of Shock Absorbers 2 1 15

Survivability 1 2 15

Potential for Propeller Damage 2 1 10

Score 1.65 1.35 100

Fuselage Design Matrix

Flat Plate Enclosed Box Rounded Geometry Importance (%)

Weight 3 2 1 40

Strength 1 3 2 20

Compatibility with Enclosed Payload 1 3 2 12

Surface for Attachments 1 3 2 10

Safety for Internal Circuitry 1 3 2 10

Previous Experience 3 2 1 5

Aerodynamics 2 1 3 3

Score 1.93 2.49 1.58 100

Table 2.1B – Design matrices from AMAT design process


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2.1.2 Avionics Design

The 2013 Advanced Class competition calls for number of Data Acquisition System (DAS) functions.

To meet these requirements, AMAT set a goal that the DAS would be able to display the altitude of the

plane in real time and log the altitude when the payload is dropped. Additional DAS goals were to

record GPS and air speed data as well as provide a real time stream from a camera to assist the dynamic

payload dropping. Flight score component S1 drove the team to also incorporate an aircraft control

system that would be able to assist flight with negative feedback from gyroscopes as well assist payload

dropping using telemetry data.

Previously, an Arudino Uno microcontroller was used to run a simple DAS. This year, to

achieve the more demanding DAS capabilities, the team upgraded to an Ardupilot 2.5, which has built in

gyroscopes, accelerometers, a digital compass, a GPS, and a high resolution altimeter. Considering the

precision nature of this competition, an off board GPS module, the uBlox LEA-6h, was chosen over the

onboard Mediatek GPS for higher accuracy. AMAT also incorporated a pitot tube to be able to use

airspeed data in drop calculations.

2.2 Previous Design

Though the current Advanced Class mission is considerably different than that of last year, AMAT was

able to use the knowledge it derived from its 2012 design to intelligently improve the 2013 Zephyrus.

Before other research, the team first considered which features of the previous design should be further

developed and which would not be effective in the new competition, as summarized in Table 2.2A:

Table 2.2A – Lessons from Previous Design

Effective Previous Features Features to Improve or Replace

Lifting tail Complex and heavy composite double element

Tricycle landing gear Flat plate, open fuselage

Winglets Heavy carbon fiber tail boom

Composite materials Heavy composite empennage


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Figure 2.3B – Wing bending stress test

Figure 2.3A – Mini Plane

demonstrating

hybrid wing construction

2.3 Experiments and Prototyping

AMAT devised experiments to test the feasibility of the designs determined in section 2.1. These tests

included the construction of a prototype mini plane, a wing stress test, carbon fiber coupon sampling,

and engine testing.

2.3.1 Prototyping and Material Testing

The team decided to test the hybrid wing design innovation

by constructing a miniature aircraft, called the “mini plane,”

so that qualitative analysis could be performed on its

practicality. The result was a flying aircraft that proved

manufacturing such a wing was possible and that the

connections between hybrid portions were acceptably strong.

AMAT also chose to stress test a full scale section of what would be the hybrid wing for strength. The

team reasoned that the hybrid monokote/composite skin would provide adequate torsional resistance, so

the wing would be most likely to fail first in bending.

AMAT represented wing loading as point forces applied

to the main spar, simulating the bending moment caused

by lift force. The result was total failure at 135 lbf of

applied point force, corresponding to a bending moment

of 63.28 ft*lbf, which is more than the spar will ever

experience under the design conditions. The team

considered this result a validation of the hybrid wing

design.

To be able to meet the weight requirements, AMAT also experimented with various composite

material arrangements of foam, wood, and carbon fiber to determine the lightest and strongest custom

materials to use on the aircraft. The fuselage was to be made out of a wood/composite sandwich for


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rigidity, and the rear landing gear was to be made out of a foam/composite sandwich to achieve the

strength of last year’s gear at a fraction of the weight. The team conducted layups as detailed in Tables

2.3A and 2.3B to determine the area density of each material, and qualitatively tested each for strength.

Table 2.3A – Rear Landing Gear Strut Coupon Mass Comparison

Layup Order m/A (g/in^2)

[0°/90°, 45°/-45°, 0°/90°, foam]s 1.2526

[0°/90°, 45°/-45°, 0°/90°, 45°/-45°, foam]s 1.6816

[0°/90°, 45°/-45°, 0°/90°, 45°/-45°, 0°/90°, foam]s 1.9664

Table 2.3B – Fuselage Wall Coupon Mass Comparison

Layup Order Wood (finish) m/A (g/in^2)

[0°/90°, 45°/-45°, wood, 0°/90°]s 1/8” Base (Mylar) 2.2933

[0°/90°, 45°/-45°, wood, 0°/90°]s 1/8” Balsa (Mylar) 1.5107

[0°/90°, 45°/-45°, wood, 0°/90°]s 1/8” Balsa (peel ply) 1.5249

[0°/90°, 45°/-45°, 0°/90°, wood]s 1/8” Balsa (peel ply) 0.932

[0°/90°, 45°/-45°, 0°/90°, wood]s 1/16” Balsa (Mylar) 1.304

Along with composite layer structure variation, the team also designed these coupons to determine the

effect of wood types and finishes of either smooth from impermeable Mylar or rough from permeable

peel ply. The tests showed the rigidity differences between wood types is negligible (unlike the weight

differences), with balsa being superior, and peel ply finishes allow the maximum amount of epoxy

removal from the layup, and thus result in lighter materials. In both cases, AMAT selected the material

with the least area density (Figure 2.3C) since it was decided that each would be sufficiently strong for

their application.

Figure 2.3C – Ideal layup configuration

for fuselage and landing gear

Figure 2.3D – Static thrust test stand

with O.S. 46 AXII


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2.3.2 Engine Performance and Propeller Sizing

This year, AMAT selected the O.S. 46 AXII engine as its power plant. To efficiently size and configure

its aircraft, AMAT needed an estimation of engine performance. The team derived the results in Figure

2.3E by running propellers of various diameter and pitch at the maximum safe RPM attainable by the

engine and recording static thrust with a fish scale attached to a sliding platform on which the engine

was mounted, seen in Figure 2.3D on the previous page. The resulting maximum thrust measurement

allowed AMAT to predict takeoff equilibrium conditions and determine that a gearbox would not be

necessary for sufficient thrust.

Figure 2.3E – Static thrust performance of the O.S. 46 AXII, yielding maximum of T0,max=8.5 lbf.

2.3.3 Model Testing

Through the use of a model fuselage, the team adequately sized the location of various components. This

produced a more efficient layout that minimizes unused space while reducing fuselage material.

Figure 2.3F – Prototype Fuselage and tail

boom.

Figure 2.3G – Experimental polar

apparatus


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The team also had great ambitions for obtaining an accurate experimental lift slope curve and

drag polar, but failed to obtain the proper sensors. The planned test involved attaching a model wing

segment to a car with the apparatus seen in Figure 2.3G so as to measure lift and drag at various speeds

and various angles of attack. The wing mounts would put pressure on sensors in both the vertical and

horizontal direction.

2.4 Final Selection

With the results from the design matrices validated through testing, the team finalized the selection of

the various aircraft components. The results are summarized in Table 2.4A:

Table 2.4A – Final Selection of Various Components

Single element, hybrid wing Box fuselage Angled truss tail boom Lifting tail

Tricycle landing gear Airshock front gear Composite rear gear Single engine, Un-geared

AMAT selected these features to combine into a design that is competitively optimized for the current

Advanced Class mission. The hybrid wing strikes an ideal balance between weight, strength, and airfoil

precision. The addition of winglets improves the efficiency of the untwisted, rectangular main wing,

bringing it very close to the ideal elliptic wing efficiency. The lifting tail configuration allows the

standard tailplane design to have the aerodynamic efficiency of a canard on takeoff, where the tail lifts

with the wing. The tail boom angle provides the added benefit of increased horizontal tail lifting

capability by elevating it above the downwash of the main wing ahead of it. The tricycle landing gear

and angled tail boom allow for a large degree of rotation on takeoff, which means the plane can

accelerate at low incidence to minimize induced drag. This efficiency in takeoff reduces the need for

additional power, so the gearbox design was rejected due to weight restrictions. The culmination of

these finalized selections results in an aircraft that is best suited for the 2013 design criteria.


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Figure 3.1A – Non-dimensionalized geometry of the Selig 1223 airfoil

3.0 Aerodynamic Analysis_________________________________________________

This section describes the analytical processes the team used to size the aircraft, ensure its stability, and

predict its performance in payload transport and delivery.

3.1 Aircraft Sizing

To size the aircraft, AMAT set a primary design criterion of achieving the mission-required lifting

capability with the lowest weight configuration possible.

3.1.1 Selection of Main Wing

AMAT opted for a single element wing equipped with the Selig 1223 airfoil. The team calculated the

viscous polar of this airfoil with XFOIL at Reynolds number 200,000. The maximum lift coefficient was

found to be CL,max=2.1 as seen in Figure 3.1A. The team began the sizing process by using a rapid

prototyping code to

dimension the main wing,

since it generates the

majority of the overall

configuration’s lift.

Optimization was achieved by iterating takeoff performance for successively increasing chord lengths c

with a given wingspan b until the desired takeoff condition was achieved. This process ensured

minimum wing volume and thus weight. Inputs included CL,max of the airfoil, static thrust, and thrust

slope of the engine. In the code, lift and induced drag were calculated, given zero-lift drag of the wing

and other aircraft components CD,0, and the best chord was selected when a 3° climb angle (to ensure the

aircraft can clear the runway) was achieved on takeoff. The team then created a 3D viscous polar for the

finite design wing using Prandtl lifting-line theory, which applies the 2D viscous polar along the span.

Figure 3.1B shows the effect of induced drag on the 3D polar, which is shifted further along the drag

axis than its 2D counterpart. Airfoil CL,max = 2.1 and aircraft CL,TO = 1.7 and CL,cruise = 0.62 can be seen

in Figure 3.1.B, and Figure 3.1A illustrates the more efficient L/D value in cruise compared to takeoff.


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3.1.2 Selection of Overall Aircraft Configuration

With the main wing geometry constrained, AMAT proceeded to configure the empennage using

equilibrium analysis of takeoff, the most demanding stage in the mission. The lifting tail design requires

that the center of gravity location xcg is behind the main wing center of pressure at takeoff, so that the

aircraft weight is balanced between the two lifting surfaces, as seen in Figure 3.2A. This is

accomplished by moving backward the aircraft aerodynamic center AC, on which CG location is

dependent for static stability. The team moved the AC aft by increasing the horizontal tails’s

aerodynamic moment contribution with a larger tail lifting area and a longer tail boom moment arm.

Empennage sizing was governed by minimizing weight increases so that the tail would be able to at least

lift its own weight in the final takeoff configuration. For stall stability, AMAT constrained the tail to

have an aspect ratio AR less than main wing, so that first stall would occur at the forward lifting surface

and create a nose-down pitching moment, ending the stall. As with rapid prototyping, designs were

iterated and trimmed for takeoff with desired main wing CL,TO and climb angle. AMAT validated each

iteration with an acceleration analysis of takeoff roll to determine that takeoff velocity could indeed be

reached. The results of sizing are given in Tables 3.1A and 3.1B on the following page, and it can be

seen that the tail indeed lifts its own weight.

Figure 3.1B – 2D and 3D Polars and Lift Slope Curves of the AMAT 2013 Main Wing


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3.1.3 Winglet Design

To reduce the amount of induced drag at takeoff, AMAT incorporated winglets into the 2013 design.

With inputs of CL,TO, VTO, and the 2D viscous polar, the team used an optimization code to design

winglets that are 12% of half of the wingspan in height and mounted so as to have a toe-in angle of 6.5°.

The toe-in angle causes the symmetric profile winglet to develop a circulation distribution that forces a

constant downwash on the main wing that is less than that of the elliptic wing for the same global lift.

Figure 3.1C shows the loading redistribution effect of winglets on the 2013 AMAT wing.

Figure 3.1C – Effect of Winglet Addition on 2013 AMAT design

The ideal distribution of circulation is an elliptical profile, and is represented as a dashed line. To

simulate this distribution, winglets shift the AMAT wing from the red circulation curve to the blue,

maintaining the area under the curve by giving an effective elongation to the wingspan. The result is a

10% increase in efficiency at takeoff, which allows the plane to reach takeoff speed sooner.

Table 3.1A – Final Lifting Surfaces

Main

Wing

Horizontal

Tail

Vertical

Tail

Span [in] 92.9 47.75 23.25

Chord [in] 17.0 17.75 17.75

Table 3.1B – Final Configuration

CG Percentage of cm 51.55%

Aircraft AC Percentage of cm 88.65%

Main and Horizontal LE Separation 75.65 in

Lift Contribution of Tail at Takeoff

1.36F

1.36 lbf


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3.2 Aircraft Stability and Control

The controllability of the aircraft and the accuracy of the payload drop are extremely dependent its

stability, which is defined into two categories. Static stability refers to a natural tendency of the aircraft

to produce a restoring moment once displaced from equilibrium. Dynamic stability refers to the ability

of these restoring moments and forces to return the aircraft to equilibrium over some time.

3.2.1 Longitudinal Stability and Control

Longitudinal stability refers to an aircraft’s stability about the pitch axis. Pitching is rotation that causes

variance in the angle of the attack. The center of gravity CG is highly influential on the natural stability

of the aircraft, and proper placement of its location was required for Zephyrus to be stable. The team

used the equilibrium code to determine the optimum locations of center of gravity xcg and aerodynamic

center xac (shown in Figure 3.2A) for static stability.

To be longitudinally stable, the aircraft must respond to any disturbing pitching moment with an induced

pitching moment in the opposite direction. The slope of the pitching moment coefficient is the product

of the lift curve slope

of the aircraft, which is positive for positive chamber airfoils, and the static

margin

, as in Eqn 1.

Figure 3.2A – Center of gravity and aerodynamic center for the

entire configuration. Lift locations at takeoff.


17

(

) (1)

For longitudinal stability, SM should be positive, meaning the CG must be ahead of the AC. The SM of

Zephyrus was optimized to a value of 8% of the fuselage length as a reference area to give a desirable

amount of stability without severe impediment of maneuverability. Because payload delivery could

potentially alter xcg of the aircraft, AMAT chose to place the dynamic payload as close to the CG

location as possible to minimize its disturbance of stability.

3.2.2 Lateral Stability and Control

Lateral stability refers to the aircraft’s stability about the roll axis. Many parameters influence roll

stability including the height of the CG, dihedral angle, vertical tail size, rudder size, and ailerons.

It is unstable to locate the CG above the AC, as this will create a moment response in the direction of a

roll moment perturbation, increasing the perturbation. For this reason, AMAT chose to place the wings,

and therefore the AC, above the fuselage and all components. This will ensure the CG is far below the

AC, contributing a great deal of roll stability. Dihedral angle, or the angle at which the wings meet the

fuselage, can also increase the roll stability of an aircraft, but due to the complications in manufacturing

and the added weight of increased mounting mechanisms, the team decided to not utilize this form of

stabilization. The team compensated for any shortcomings in lateral static stability by incorporating

ailerons, which are roll-producing control surfaces that can be used to actively stabilize the aircraft with

commands from the pilot and control system.

3.2.3 Directional Stability

Directional stability refers to the aircraft’s stability in the yaw direction. The team decided to use a

single vertical tail and rudder for yaw stability. For any deflection from trim on the yaw axis, the vertical

tail will become angled relative to the wind, which will induce a lifting force that creates an opposing

moment to the yaw deflection. The winglets also provide some stability in this manner.


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3.2.4 Control Surface Sizing

For an aircraft to make a turn or climb and stray from it natural equilibrium, it must use control surfaces

to induce repositioning moments. Zephyrus uses a combination of ailerons, rudder, and elevator to

create these moments. To be effective, ailerons must be large enough to create a roll moment equal in

magnitude to the roll moment created from a gust of wind on the vertical tail. The force of the wind is

found from the time rate of change of the wind moment (Eqn 1) and is non-dimentionalized and set

equal to the coefficient of roll of the aileron (Eqn 2):.

( ) (2)

∫

(

)

(3)

To solve for the ratio of aileron chord to main wing chord τ, all other values were determined using

equilibrium analysis or through assumption. The results are shown in Figure 3.2B.

The team assumed a max perpendicular gust velocity equal to the forward velocity of the aircraft,

approximately 32 ft/s. This resulted in τ = 0.162 and the chord of the aileron equal to 5.0 in, given a

factor of safety of 1.8.

Figure 3.2B – Aileron chord required to counteract

a range of gust velocities.

0 5 10 15 20 25 30 35 40 45 500

1

2

3

4

5

6

7Chord of Aileron Required for Various Perpendicular Gust Velocities

Velocity of Perpendicular Gust, ft/s

Aile

ron C

hord

, in


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3.3 Aircraft Performance

AMAT used equilibrium analysis as in part 3.1.2 to calculate aircraft performance at important flight

conditions like takeoff and cruise, as seen in Table 3.3A.

Table 3.3A – Zephyrus Fight Characteristics

Velocity

[ft/s]

Thrust

[lbf]

Lift

[lbf]

Viscous

Drag [lbf]

Induced

Drag [lbf]

Takeoff 34.4 5.41 25.1 0.6 2.5 8.41

Cruise 60.7 2.81 26.5 1.8 1.01 9.5

The high lift state of Zephyrus’s takeoff causes it to experience 150% more induced drag force than in

cruise. This is due the high CL required to create the necessary lift force at low speed. In cruise, the

thrust is exactly equal to the drag, and aircraft is traveling almost twice as fast as in takeoff. The

required value of CL decreases to a point of larger L/D on the 3D viscous polar in Figure 3.1B.

AMAT determined the limits of Zephyrus’s aerodynamic performance by comparing predicted

engine power performance with calculated overall drag behavior of the aircraft. First, the team

calculated the drag contributions of the lifting surfaces as a sum of each surface’s zero-lift drag CD0 and

the induced drag CD,i. CD0 is a combination of drag from friction, profile, and interference:

(

(

)

)

(4)

The team calculated induced drag CD,i from its dependency on the total lift coefficient CL, which was

weighed proportionally for each lifting surface. The vertical stabilizer had no lift and therefore no

induced drag in the trim flight:

;

;

(5), (6), & (7)

The effect of the winglets on induced drag occurs in Eqn (5), where the increase in efficiency e causes a

decrease in CD,i. The total drag of each wing was calculated as in Eqn (6).

The fuselage and tail boom were considered to have negligible lift, so only CD0 was calculated

for each. The greater length of these bodies allowed the formation of longer boundary layers that

became turbulent, necessitating the use of the following relationship:


20

; ;

(

( ⁄ )

)

(8)

The team treated the remaining external components (landing gear and engine) as bluff bodies, so that

drag depended on frontal area and tabulated drag coefficients. Equivalent flat plate areas f were

calculated to represent each component’s drag in a form that could be directly added, and then a range of

total drag forces were found by multiplying this sum by various dynamic pressures. In flight, thrust

opposes drag, and performance is a direct function of their difference. The team assumed propeller

power output P to be constant at low speeds and dependent on an experimentally measured static thrust

T0 =8.5 lbf found for a propeller of radius R=0.5 ft and pitch p=4 in. The team also determined the

required power Preq to sustain equilibrium by drag balance:

⁄

√ ( )

(9)

The team calculated thrust T to decrease linearly from its static maximum according to the following:

( )

⁄

⁄ ⁄ ⁄ (10)

Figure 3.3A – Comparison of available power to power required for equilibrium flight.


21

The power performance demonstration in Figure 3.3A shows the feasible range of flight conditions for

Zephyrus. The top of the range is limited to the intersection of the two curves at Vmax, where the entire

power of the propulsion system is needed to maintain cruise. Figure 3.3A also allows the calculation of

rate of climb RC, with the maximum value corresponding to the greatest excess power available ΔPmax:

( )

(11)

3.4 Payload Analysis

Zephyrus is tasked with two main payload-related requirements: general lifting and delivery accuracy.

3.4.1 Payload Prediction

The maximum lifting performance requirement for Zephyrus is the combination of weight of the

aircraft, payload for delivery, and payload representative of fuel. Since fuel capacity directly affects the

operational range of the aircraft, AMAT performed a payload prediction analysis to demonstrate mission

capability in a variety of conditions.

UC Davis Advanced Modeling Aeronautics Team

Team 215

Figure 3.4A – Effect of increasing density altitude on maximum

payload cabability.


22

Figure 3.4B – Sandbag trajectory prediction

for drop from cruise speed.

Using equilibrium analysis with the trim setting for highest takeoff lift available, the team found

maximum payload values at various densities, which are plotted in Figure 3.4A. A linear fit was applied

to the plot to determine a maximum payload prediction equation for the 2013 AMAT aircraft. This

relationship predicts that a gain in 4000 ft of density altitude from sea level at standard conditions will

reduce Zephyrus’s payload by less than 7.5%, showing that the aircraft will be mission-capable in a

variety of environments and conditions that could require humanitarian aid.

3.4.2 Payload Drop

The succeess of AMAT’s mission also depends on the aircraft’s ability to deliver its humanitarian aid

package accurately, so that its contents arrive intact

and no damage is caused by its arrival, which is the

basis for the accuracy score S1. AMAT used a

viscous free-fall code to simulate the package’s

trajectory and predict the optimal conditions for its

release, as shown in Figure 3.4B. These

calculations were incorperated into the DAS so that

it will be able to supervise the payload drop in real

time. When the payload release is armed, the

Ardupilot will calculate the expected impact locations and advise the secondary pilot on when to release

the payload. The secondary pilot will use this data along with video stream to guide the primary pilot

and determine when to drop the payload.


23

4.0 Component Structure and Manufacturing_________________________________

To meet the design intent, AMAT employed a number of structural and manufacturing innovations in

the construction of the aircraft.

4.1 Main Wing

Zephyrus’s main wing is a balsa rib and composite hybrid wing that combines the lightness of traditional

balsa construction with the strength and precision of composite materials. The team laser cut all balsa

airfoil components to ensure exact shape. These were attached to the primary wing spar, which was

made of sandwiched base wood and bi-ply carbon fiber with layers oriented at 0°/90° and 45°/-45° to

provide axial and shear strength under bending. To reduce weight, the team cut holes from the spar

center, where the material has less bending resistance. AMAT produced a sharp trailing edge TE from

hotwire-cut foam attached to the wing with an underneath composite panel to provide torsional

resistance. Only a single side is composite because the team determined that a fully enclosed cross-

section of twice the weight would be a strength overdesign. The ailerons are fully enclosed in composite

to guarantee rigidity so that shape deformation does not occur when they are actuated. Tape secures the

ailerons to the wing and allows rotation.

Figure 4.1A – Main wing half section

view


24

The team milled winglets from plates of composite/foam sandwich material for simplicity and

lightness. The flat plate satisfied the symmetric airfoil design condition. The TE’s were sharpened to

produce a Kutta condition, where the streamlines leave tangent to the TE, reducing form drag. The

leading edges LE were made of rounded foam to reduce the front curvature at the airfoil, improving

boundary layer attachment. Bumpers were extended downward from the winglets for wingtip strike

protection, and these areas were strengthened with an additional layer of carbon fiber. Drag from these

small extensions is negligible.

4.2 Empennage

AMAT constructed the tail lifting surfaces with the same streamlined flat plate shape as the winglets and

a similar hybrid structure design as the main wing. The TE design of the main wing was used on the

horizontal and vertical stabilizers to give rigid shape to the empennage control surfaces. This year’s

hybrid design and reduced thickness of the stabilizers compared to AMAT’s previous empennage has

resulted in a significantly lighter tail, as summarized in Table 4.2A.

The tail boom structure has also changed significantly from the heavy carbon/epoxy tube in AMAT’s

previous design. The team’s current light-weight solution is a structure of three small carbon/epoxy rods

joined at intervals by wood spacers, which, in combination, provide a cross-section that is highly

Figure 4.2A - Empennage without Monokote


25

resistive in bending. Torsion is controlled by a Monokote covering. This design also simplified tail

mounting, with low-profile composite brackets in the tail and direct insertion into the fuselage. This

allowed for easy angling of the tail boom and reduced weight as seen in Table 4.2A.

Table 4.2A – Weight (lbf.) Comparison of Zephyrus to 2012 AMAT Design

Main Wing Horizontal Tail Vertical Tail Tail Boom

AMAT 2012 3.8 1.4 0.56 1.05

Zephyrus 2.5 1.0 0.50 0.50

Difference -1.3 -0.4 -0.06 -0.55

Table 4.2A shows weight reduction in all listed components of Zephyrus despite each having larger

dimensions than its predecessor.

4.3 Fuselage

AMAT’s fuselage is a box shape made of balsa/composite sandwich material detailed in Figure 2.3C

that encloses the aircraft’s operational components to reduce drag. The team chose a box cross-section

so that minimal material would be required to provide adequate bending resistance. Further weight

reduction was accomplished by removing excess center material.

The fuselage provides mounting for the landing gear, engine, and tail boom and contains the avionics,

fuel tank, and payloads. The static payload is mounted in slots in the fuselage walls to allow for CG

Figure 4.3A – Fuselage and landing gear


26

adjustment but does not contribute to the airframe’s structure, which will be structurally sound upon

static payload removal. The dynamic payload bay is a suspension member that supports the payload in

tension and is located at the CG for reasons discussed in section 3.2.1. The dynamic payload is held by a

pin that is removed by a servo at the time of drop.

4.4 Landing Gear

Zephyrus is equipped with a tricycle landing gear, which allows it to accelerate to takeoff speed with

minimum induced drag and then rotate into the required takeoff incidence. The rear landing gear are

tapered foam/composite struts that descend at 30o from vertical to provide the best compromise between

ground clearance, gear load distribution, and stable stance (layup in Figure 2.3C). The foam was cut

with a 5-axis CNC mill allowing for complex geometry and rapid repeatability. AMAT placed the rear

gear near and aft of the CG to balance the aircraft weight on the landing gear and to minimize the

moment required from the tail for rotation. For increased control of landing rollout, the team designed

minimalist brakes for the rear gear consisting of composite pads that are pushed against the rear wheels

with servos.

For the front landing gear, the team manufactured a custom air shock with spring and damper

characteristics so that landing impact is softened and the resulting oscillations are impeded. The front

gear is made primarily of aluminum, making it lighter than any metal shock available for purchase, and

it can be rotated with a servo for steering on the tarmac. AMAT purchased and manufactured foam

wheels for all gear to minimize weight.

Figure 4.4A – Machining

of the airshock


27

4.5 Weight Management

Fight score component S2 had a major influence on the structural design of Zephyrus. To manage final

weight, AMAT set allocations (shown in Table 4.5A) for structural systems, so that each could be

minimized individually, rather than trying to meet the empty weight requirement all at once.

Table 4.5A – Zephyrus Component Weight Allocations

Weight Allocation (lbf) % of Total

Main Wing 2.5 31.25%

Vertical Stabilizer 0.5 6.25%

Horizontal Stabilizer 1.0 12.5%

Fuselage 0.5 6.25%

Tail Boom 0.5 6.25%

Engine 1.1 13.75%

Landing Gear 0.5 6.25%

Avionics & Misc. 1.4 17.5%

To meet these strict weight restrictions, AMAT devised a number of weight-reducing innovations in

addition to the structural designs already mentioned. For instance, AMAT manufactured fasteners out of

aluminum, which is a third of the weight of steel. The team further reduced fastener weight by drilling

out center material. Composite/foam layups were lightened by applying a slurry of epoxy and glass

microballoons to foam surfaces. The addition of glass to epoxy effectively reduces its density without

severely reducing its strength. Foam pores absorb a significant amount of epoxy, so filling them with

slurry reduces the total amount of epoxy and therefore weight in the layup.

Figure 4.5A - From left to right: Glass/Epoxy Slurry, Aluminum Fasteners, 5-axis foam mill

used to machine lightweight landing gear and wheels.


28

5.0 Conclusion___________________________________________________________

AMAT’s aircraft design is optimized for the 2013 SAE Aero Design Advanced Class mission. The team

followed a thorough design procedure to select aircraft features that, in combination, provide ideal

aircraft functionality under the design constraints. AMAT then configured the aircraft design using

codes for rapid prototyping, equilibrium, and acceleration analysis. This configuration was analyzed for

stability and control as well as performance to confirm that it was suitable for the challenge. With this

determined, manufacturing began with an emphasis on minimum weight construction. For competitive

flight capability, the team designed a control system to simplify the pilot’s approach to the target and

supervise the conditions at which the payload is dropped. AMAT designed Zephyrus with the intention

of creating an ideal combination of aerodynamics, structure, and electronics, and the team is excited to

have the opportunity to see how its product performs at competition.


29

6.0 References___________________________________________________________

1. Bauchau, O.A. and Craig, J.I., Structural Analysis with Applications to Aerospace Structures, New York:

Springer Science+Business, 2009.

2. Chattot, J. J., Glider and Airplane Design for Students, Int. J. Aerodynamics, Vol. 1, No. 2, 2010

3. Etkin, Bernard, and Lloyd Duff Reid. Dynamics of Flight, Stability and Control. New York: John Wiley &

Sons, Inc., 1996.

4. Mattingly, Jack D, Elements of Propulsion: Gas Turbines And Rockets. AIAA. 2006

5. Moran, Jack, An Introduction to Theoretical and Computational Aerodynamics, New York: Dover 2003.

6. Nelson, Robert C. Flight Stability and Automatic Control. New York: McGraw-Hill, Inc., 1989.

7. Roskam, Jan. Airplane Design Part1: Sizing of Airplane Kansas: Design, Analysis and Research

Corporation, 2005.

8. Roskam, Jan. Airplane Aerodynamics and Performance Kansas: Design, Analysis and Research

Corporation, 2005.

9. Shigley, Joseph E. et al. Mechanical Engineering Design: Seventh Edition, McGraw Hill, Inc. 2003

15.00° 12.00

2.50

5.00

17.75

15.75

2.01

101.38

92.91

75.65 17.00

47.69

13.10

4.37

2.01

6.50°

23.66 9.28

Empty CGLoaded CG

23.25

7.15

49.58

Appendix A - Technical Plans

Wingspan 92.91 In

Approximate Empty Weight 7.9lbf

Engine OS 46 AXII

Cargo Bay Volume 66.0 cu. in

Ballast for Empty CG Balance 15.0 lbf. placed 1.71. in behind Main Wing Leading Edge

D

C

B

AA

B

C

D

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8 7 6 5 4 3 2 1

DIMENSIONS ARE IN INCHESTOLERANCES:FRACTIONAL 1/32TWO PLACE DECIMAL .02THREE PLACE DECIMAL .005

MATERIAL

FINISH

DRAWN

CHECKED

ENG APPR.

MFG APPR.

Q.A.

COMMENTS:

DATENAME University of California, DavisAMAT #215

TITLE:

SIZE

BDWG. NO. REV

SCALE: 1:22

UNLESS OTHERWISE SPECIFIED:

Edelman 3/3/2013

AMAT Zephyrus

SHEET 1 OF 1DO NOT SCALE DRAWING

1SolidWorks Student Edition. For Academic Use Only.

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