Project Report, Stingray

California State University SacramentoCollege of Engineering and Computer Science

ME-191 Senior Project

Project Report, Stingray

Prepared by:

John Gyurics

Nicholas Rossi

Timothy Burkhard

1 st Edition • May 20 th , 2008

California State University SacramentoCollege of Engineering and Computer ScienceME-190 Senior ProjectProject Report, Stingray

Table of ContentsAbstract..........................................................................................................................................................................5Introduction....................................................................................................................................................................6Technical Discussion........................................................................................................................................................7

Design Criteria............................................................................................................................................................7Design Features.........................................................................................................................................................9Analysis....................................................................................................................................................................29Manufacturing Plan...................................................................................................................................................42Test Plan..................................................................................................................................................................69

References....................................................................................................................................................................71Appendix A, Competition Rules.......................................................................................................................................72Appendix B, Drawings....................................................................................................................................................73

John Gyurics • Nicholas Rossi • Timot hy Burkhard2


Table of Figures

Figure 1..............................................................................................................................................6Figure 2..............................................................................................................................................9Figure 3............................................................................................................................................11Figure 4............................................................................................................................................11Figure 5............................................................................................................................................12Figure 6............................................................................................................................................14Figure 7............................................................................................................................................17Figure 8............................................................................................................................................18Figure 9............................................................................................................................................18Figure 10..........................................................................................................................................19Figure 11..........................................................................................................................................20Figure 12..........................................................................................................................................21Figure 13..........................................................................................................................................22Figure 14..........................................................................................................................................23Figure 15..........................................................................................................................................24Figure 16..........................................................................................................................................25Figure 17..........................................................................................................................................26Figure 18..........................................................................................................................................27Figure 19..........................................................................................................................................28Figure 20..........................................................................................................................................29Figure 21..........................................................................................................................................29Figure 22..........................................................................................................................................29Figure 23..........................................................................................................................................30Figure 24..........................................................................................................................................30Figure 25..........................................................................................................................................31Figure 26..........................................................................................................................................31Figure 27..........................................................................................................................................32Figure 28..........................................................................................................................................32Figure 29..........................................................................................................................................36Figure 30..........................................................................................................................................37Figure 31..........................................................................................................................................38Figure 32..........................................................................................................................................39 Figure 33.........................................................................................................................................39Figure 34..........................................................................................................................................40Figure 35..........................................................................................................................................43Figure 36..........................................................................................................................................44Figure 37..........................................................................................................................................44Figure 38..........................................................................................................................................45Figure 39..........................................................................................................................................46Figure 40..........................................................................................................................................47



Figure 41..........................................................................................................................................48Figure 42..........................................................................................................................................49Figure 43..........................................................................................................................................50Figure 44..........................................................................................................................................52Figure 45..........................................................................................................................................53Figure 46..........................................................................................................................................57



AbstractJohn Gyurics, Nicholas Rossi, and Timothy Burkhard have decided to enter the 2008 AUVSI student UAV

(Unmanned Aerial Vehicle) competition in order to fulfill their ME-190/ME-191 senior project class

requirements. The team was required to design and build a small aircraft and autopilot system capable of

autonomously navigating several waypoints and retrieving data regarding targets at each waypoint.

The UAV team has designed and fabricated all major components of the airframe in order to be

competitive at the competition. The UAV team designed the aircraft using 3D CAD software. The UAV team

also built two airframes. The first airframe was built using conventional “balsa and tissue” construction in

order to further validate the design. The second airframe, which will be used in competition, was be

manufactured using “advanced composite materials”.

The autopilot system was purchased from Micropilot, a major manufacturer of UAV components. The

Micropilot system is capable of performing all the necessary functions to complete the AUVSI competition.

The UAV team spent the fall semester designing and analyzing the aircraft as well as creating a

manufacturing and testing plan. The winter session was spent building the first airframe. The beginning of

the spring semester was spent testing the first airframe and modifying the second airframe design. The

rest of the semester was spent building and testing the second airframe.

The results of the testing of our final product indicated that the Stingray met the aerodynamic objectives of

our project. The aircraft was stable and controllable across the range of airspeeds and performance

parameters set forth in ME-190.

Monetary support for the project was sought both on and off-campus. The cost of the project was

accurately determined and supported through various grants and scholarships.



IntroductionThe face of the United States' military is changing. Combat systems are improving the military's

communication and intelligence capabilities by becoming more “network-centric”. These systems include

various unmanned robotic devices capable of collecting information or even attacking targets remotely.

Unmanned systems improve the safety of our

troops and improve their capabilities. Future

military investments will include an increasing

budget for the development of unmanned

systems. In 2006, the Department of Defense

requested $1.7 billion for these systems. The

Army's Future Combat System (FCS) will also rely

heavily on UAV's and other unmanned vehicles

(Protecting America, 2006).

As a group of students in the field of mechanical engineering, the UAV team was presented with

the opportunity to be involved in this exciting new area of engineering. In order to prepare themselves for

employment in the future of aerospace research, foster aeronautical interest at California State University

Sacramento, and meet the objectives the ME-190/191 class series, John Gyurics, Nicholas Rossi and

Timothy Burkhard have decided to compete in the AUVSI UAV competition in Spring 2008. The main goal

of the AUVSI UAV competition is to create an aircraft capable of stable, autonomous flight and navigation,

autonomous take-off and landing, and actionable intelligence gathering. To complete the project, the

teams will need implement navigation equipment and design and fabricate an airframe.

Since the start of this project, the UAV team had the opportunity to apply much of the acquired knowledge

during their engineering career at CSUS, and implemented mechanical and dynamics concepts as well as

aerodynamic design.

Some students and faculty have contributed time and or monetary participation towards the Hornet

UAV project. Will Landreth and Pritpal Singh assisted the UAV team by applying concepts from ME-138

Concurrent Product and Process Design.


Figure 1: The Predator UAV has been used in multiple theaters around the world. In 2002, the DOD allocated $37 million for new Predators. (Beavin, 2002)


Technical Discussion

Design Criteria

Airframe Primary Requirements

The primary requirements of the airframe solution are as follows:

● Positive or neutral stability in all axes

● 1 hour of flight time

● Cargo bay for electronics

● Cutout for GPS antenna

● Ability to carry 3lbs of cargo

● Camera mount with clear visibility

Airframe Secondary Requirements

The secondary requirements of the airframe solution are as follows:

● Capable of stable slow flight (<25MPH with no danger of stalling)

● Fast cruise speed (>30MPH straight and level)

● Low manufaturing cost, DFM (<$15,000 at 100 units per year)

● Easy maintenance (All components serviceable)

● Portable (removable wings/tail surfaces)

● Durable (Recoverable after loss of control over flat asphalt surface with less

than $500 damage)

● Access to electronics, DFA

● Autonomous takeoff

● Autonomous landing



Autopilot Primary Requirements

The primary requirements of the autopilot/data retrieval system are as follows:

● Directs sustained, stable autonomous flight (>1 hour flight time, no loss of control)

● Accurate waypoint navigation (on course +-50 meters)

● On-the-fly waypoint changes

● On-the-fly altitude changes

● Identifies targets as shown in Table 1 of Appendix A

● Accepts course changes from the Ground Control Center (GCS)

● Uses safety precautions as stated in page 5 of Appendix A

Autopilot Secondary Requirements

The primary requirements of the autopilot/data retrieval system are as follows:

● Capable of autonomous takeoff and landing

● Capable of directing stable slow flight as well as cruise

● Capable of sending live video of sufficient quality to identify targets while the aircraft is still

in flight (aircraft at altitude 750' AGL, target <50 meters laterally from course, target as

defined in Appendix A)



Design FeaturesThe Stingray is a small, lightweight, strong, portable, and useful piece of equipment. It has a wingspan of

75 inches, is 65 inches long and weighs about 8.25 lbs. The Stingray can be assembled in a few minutes

with simple tools. The aircraft can be disassembled to fit inside of most vehicle's trunks, taking up about

9.7 cubic feet. Overall, the Hornet UAV meets all of the design requirements of the AUVSI competition.

Electronic Components

Autopilot System

In choosing the right auto-pilot for the job, the Hornet II UAV team researched for the following criteria:

● Meet the minimum requirements for the competition and project

○ capable of autonomous takeoff and flight

○ waypoint navigation (GPS coordinates)

○ maintain stability, specific altitudes, and directions

○ autonomous landing in same location as takeoff


Figure 2: Micropilot MP 2028G; the autopilot system that will be used in the Stingray.


● In case of loss of GPS signal, loss of RC signal, engine failure, loss of datalink, and or low battery

voltage, programmable error handling

● Weight: minimal weight for maximum efficiency / flight time

● Cost: meets the predetermined budget

● Size: compactness for fuselage size reduction and thus drag and weight reduction

The MP2028 is one of the smallest UAV autopilots in the world. Due to the compact nature of the MP2028

Autopilot, the Hornet II UAV team was able to minimize the necessary payload space by reducing the size

of the fuselage, and as a result benefited in drag and weight reduction. The MicroPilot – MP2028 weighing

only 28 grams including the GPS receiver, contributes towards wing loading reduction plus increased

efficiency and performance.

Some additional benefits to choosing the MP2028 are as follows:

● VRS Editor allows effortless changes to flight configuration.

● Datalog Viewer allows you to analyze flight performance and get your UAV in the air faster.

● Aircraft Editor allows you to model your airframe for more accurate simulations.

● MP Joystick functions provide long range manual control for versatile flight management.

● 1,000 programmable waypoints or commands

● Powerful command set allows tremendous flexibility when describing your mission

● Fully integrated - all sensors required for complete airframe stabilization are integrated into a single

circuit board

● Controls up to 24 servos or relays

● Flexibility in Autonomous takeoffs and landings

● Supports flaps, flaperons, elevons, v-tail, x-tail, split rudders, split ailerons and flap/aileron mixing

● Extensive user programmable feedback gains and flight parameters tailor the MP2028g to the

airframe and mission

● Extensive data log capability simplifies post flight diagnostics and analysis



● Low battery warning: both on the ground and in flight

Servos

A wide array of servos is available for specific applications or needs. In servo fitting, the UAV team

members had to once again base their decisions on the requirements, as well as the physical, mechanical,

and electrical demands of the aircraft. The following are some of the criteria used by the UAV team

during the servo selection process:

● Torque (calculated to be 25 oz-in for the aileron – biggest

control surface)

● Weight

● Dimension

● Cost

● Response time / Speed

● Quality / Type (digital / analog)

● Voltage Range

From the final four, Futaba S3150, Hitec RCD HS-125MG, Hitec

HS-645MG, and JR’s DS368BB, the team decided to select the

Hitec 125MG and the JR. Both of these servos come with metal

gears for reliability and longer life expectancy; ball bearings

instead of bushings are common in both of these servos for

performance, quality, and reliability. Both the DS368BB and the

125MG are Digital, high-torque micros servos that operate on both

4.8 and 6 volts. Preference of digital over analog servos was due

to the nature of digital servos being more likely to hold their

position, by making their specified torque available for correction

for the smallest change or deflection. The high torque micro


Figure 3: The JR DS368BB will be used as the aileron servos in the Stingray

Figure 4: The Hitec RCD HS-125MG will be used to control the Stingray's tail surfaces.


quality places these servos in the sought out category when it comes to model aircraft design; they only

weigh 0.83 and 0.84 oz respectively, and are narrow enough to fit into slim cavities under one-half inch.

Due to the compact size of these servos the team was able to design the aircraft keeping the servos,

horns, and linkages buried inside the wigs and ailerons for less drag and better aesthetics. Often, gear

reduction for higher torque results in loss of angular acceleration and angular velocity, however both the

125MG and the DS368BB are still some of the faster servos considering the competition; at 4.8 Volts, for

60 degrees of motion the Hitec is at 0.17seconds, while the JR is at 0.21; at 6 volts there is improvement

to 0.13 and 0.18 seconds respectively.

Motor

Motor selection was fairly straight forward. The team

determined that the less than one hour flight time specified

by the AUVSI competition would be the operational time

requirement for both the motor and the Battery setup. The

target weight of the composite aircraft including the payload

necessary to operate both radio-controlled and autonomously

was determined to be around 8 lbs. Accounting for the high

aspect ratio wing profile design, taking the requirements into

account, and adding a 37% factor of safety to allow for

unaccounted weight, the UAV members reduced the available

selection to the reliable, low maintenance Axi 4120 Motor.

The following information was provided by the manufacturer at: <http://www.modelmotors.cz/index.php?

page=61&product=4120&serie=20&line=GOLD>

Specification

No. of cells 5 - 6 Li-Poly

RPM/V 465 RMP/V

Max.Efficiency 87%


Figure 5: The AXI 4130/20 Gold outrunner brushless motor


Max. efficiency current 13 - 37 A (>82%)

No load current / 10V 1.5 A

Current capacity 52 A/60 s

Internal Resistance 82 mohm

Dimensions (diameter. x lenght) 49,8x55,5 mm

Shaft diameter 6 mm

Weight with cables 320 g

“This brushless motor with neodymium magnets and a rotating case is manufactured using the latest

technology from the finest quality materials. The hardened steel shaft supported by three ball bearings,

and overall robust yet lightweight construction, ensure a long service life. The optional radial mount set

includes: mounting flange, propeller adaptor, securing collar, and screws. The unique design of this motor

gives extremely high torque levels to turn large diameter and high pitch propellers with a high level of

efficiency. New AXI 4120/20 GOLD LINE has been developed especially for use with 5s and 6s lipo

batteries for 3D models up to 3500g and for sailplanes up to 5000g in weight.”



Batteries

Battery selection was dictated by

numerous requirements predetermined by

other already selected components such

as motor, servos, and other electronics

onboard. The team sought information

from professionals in the modeling world,

and narrowed the search to the final pick,

the EVO25 V-Power 5000 mAh by

FlightPower. A pair of the 5000 mAh

EVO25s will be adequate to enable the

aircraft for a minimum of one hour flight

time for the AUVSI competition. The

weight and location of the batteries will

be used to balance and adjust center of gravity and stability.

The following are some manufacturer specs for the batteries found at:

<http://www3.towerhobbies.com/cgi-bin/wti0001p?&I=LXRWA9&P=FR#tech>

FlightPower Evo 25

● 5000 3S 11.1V LiPo Battery Pack with Balance Connector.

● FEATURES: Optimum power and weight for helis and aerobatics

● 25C Continuous Discharge

● 50C Burst

● Charge and Discharge leads are compatible with FlightPower and Thunder Power

● INCLUDES: FlightPower Evo 25 5000mAh 3S 11.1V LiPo Pack w/Balance Connector

● REQUIRES: FlightPower V-Balance system (FPWM0120)


Figure 6: EVO25 V-Power 5000 mAh Battery by FlightPower.


● Charger - TRITON2 (GPMM3153) or TritonZX (GPMM3154)

● SPECS: Size: 6.5 x 1.8 x 1.1" (167 x 48 x 27mm)

● Weight: 14.oz (399g)

● Capacity: 5000mAh

● Continuous Discharge Current: 25C

● Rated Voltage: 11.1V

● Number of Cells in Series: 3

● Max. Charge Voltage: 4.25V per cell

● Min. Discharge Voltage: 3.0V per cell

● Maximum Recommended Charge Current: 1C



Camera

In the AUVSI competition, some of the requirements are to identify both pre-specified targets, as well as

unspecified targets in route. A micro video camera system will be imbedded into the Hornet UAV for the

identification purpose. After researching the available options, the UAV team decided to utilize the

PC207XP video camera. The following specification information is available at:

<http://www.supercircuits.com/index.asp?PageAction=VIEWPROD&ProdID=4205>

“The PC207XP is the world’s smallest video camera...an astounding 0.472” square by 0.669”. Latest

generation 1/4 color CMOS imager gives you a super sharp 380 lines of resolution and low light

performance of 3 lux...remarkable for a security camera this size. Precision pinhole lens gives you an 80

degree field of view. Power requirements are 12 volts DC at a miniscule 50 milliamps. Output is standard

NTSC video, compatible with all of our VCRs, transmitters and monitors. Advanced on-board signal

processing seamlessly controls exposure, shutter speed and more ensuring trouble free video in most any

lighting condition. Hurry, quantities may be limited.”

Key Features

• Incredible 0.472” square X 0.669”

• 380 Lines of resolution

• 3 Low lux rating

• Built-in 3.6MM pinhole lens

• 12 Volts DC

• Low 20mA power consumption

• 30 day MBG


Illustration 1: PC207XP Analog video camera.


Early Concepts

During the preliminary design phase of Stingray, the design team developed several designs for all of the

many different aspects of the aircraft and its accompanying components. Below are several concepts and

the reasoning behind scrapping them.

The Flying Wing

This concept mimics the modern day design of the flying delta wing. Its low profile, low drag aerodynamic

body provides an aesthetically pleasing airframe. Another main advantage that this airframe provides is

speed. The low aspect ratio of this aircraft allows for extreme high speed capabilities. Given all of this, the

design was ultimately scrapped due to several design requirements such as slow, stable flight, and the

ability to transmit clear video to the ground station.


Figure 7: Flying wing concept.


The Swept-Wing UAV

This conceptual wing design was a concept which provided at real idea of the dihedral angle incorporated

into an accurate air profile. This concept is aesthetically pleasing and is also functional. This wing would

have met the design requirements set forth by the Stingray design team. The design was scrapped but the

idea was retained. This concept would have much too difficult to effectively manufacture.

The Box Fuselage

This preliminary fuselage design was a concept that had several features that were enticing. First, the ease

of manufacturing. This semi-conventional fuselage design would have minimal swept surfaces, and remain

in the manufacturable realm. Also, this design had the feature of a drop-out electronics bay/wing design

that would allow for the minimizing of manufactured parts. Given all of these traits, the design was

scrapped due to the lack of aerodynamic characteristics on the OML. The portability and serviceability

aspects of this design, though, were retained and incorporated into the final design.


Figure 8: The swept-wing concept.

Figure 9: The box-type fuselage concept.


Initial Tail Boom Attachment Design

From all of the concepts the Hornet UAV design team brainstormed and developed, the tail-boom

attachment method is the design which deviated the least from the concept to the final design. This

attachment feature was modeled around one of the conceptual fuselages which explains the feature at the

LH side of the figure above. The clocking feature, thought was a necessary feature and was retained. This

design is simple and manageable. The design team modified this design and implemented it into our final

fuselage design. It incorporated a purchase part, the AN nut, and only two machined fittings. This

accomplished the necessary DFM and DFA aspects that the design team was attempting to incorporate into

the Stingray aircraft.


Figure 10: An early tail-boom attachment design.


Horizontal Stabilizer Attachment

The interface between the horizontal stabilizer and the vertical stabilizer was a main area of concern during

the design phase of the Stingray aircraft. Given the requirements defined by the design team of portability,

removeablility and serviceability, several ideas were tossed around; some even ranging from attaching a

guy-wire from the vertical stabilizer/tail-boom interface out to the horizontal stabilizer. Understanding the

forces that this section of the aircraft will have to endure, a more robust interface was necessary. As

shown in Figure 11 the current, and most likely final, interface was determined to be just what the team

was after. This 6 screws interface will allow for any moment that acts on the tip of the horizontal stabilizer

to be reacted evenly and distributed throughout the composite laminate and hardware.


Figure 11: Horizontal stabilizer attachment detail.


Conceptual Top Level Body

This hand-sketched concept is the core behind the final top level design. It virtually mimics what the

Stingray final top level deliverable system will look like.


Figure 12: An early Stingray concept. This is the concept that we selected.


Tail Boom Attachment

FWD

From early on of the design phase of the Stingray, the design team agreed on a

hollow, cylindrical tail attachment method. This method is relatively unconventional

yet the design will allow the team to achieve several goals. First, the portability of

the aircraft is one of our major design requirements. This removable tail boom

allows for the aircraft to be properly disassembled and stored. The tail boom is also

a standard purchased part which eliminates the need for another manufactured

custom part thus decreasing manufacturing time and design. One hurdle that has to

be jumped is the fact that the boom is round which introduces the factor of

clocking. When the boom is attached to the fuselage, vertical stabilizer has to be

perfectly vertical with respect to the top plane of the aircraft. Thus the design team

had to incorporate a clocking feature into the connection interface. This will allow

for the boom to be properly clocked each time it is attached and removed.


Figure 13: Final tail-boom attachment design.


AFT:

The AFT connection presented several problems concurrently with the FWD

attachment. The design team needed a flat surface to mount the vertical stabilizer

to the tail boom. Evolving from this need, the AFT Tail Sleeve was designed. This is

a 6061-T6 Aluminum machined part which incorporates vertical tabs mounted

inside the vertical stabilizer. This lightweight, rigid component provides all of the

necessary functions and also incorporates visual aesthetics into the final design.


Figure 14: Final tailwheel and vertical stabilizer attachment design.


Wing Design

Wing Planform

In order to ensure the the Stingray would stall at a speed less than 25 mph, the aircraft has been designed

with a large wing area. The Hornet UAV team designed the aircraft with a wingspan of less than 7 feet so

that the aircraft could be broken into three 25 inch sections that could fit easily into the trunk of a vehicle.

As a result of the desired stall speed, the weight of the aircraft, and the size restrictions on the wings, the

wing design has been finalized as shown in the figure below. The wingspan is 75 inches, the average chord

length is 12 inches, and the taper ratio is .70.


Figure 15: Wing planform design.


Wing Profile

The wing design includes geometric washout and aerodynamic wash-in. Geometric washout combats tip-

stalling, thereby reducing the probability of crashing the aircraft at low airspeed and increasing stability

while landing. The aerodynamic wash-in counters the inefficiency created by geometric washout by

creating an elliptical lift distribution at the design airspeed. The result is a low-drag wing with good

characteristics in both slow and high speeds.


Figure 16: Wing profile design.


Horizontal Stabilizer Design

Horizontal Stabilizer Planform

The horizontal stabilizer was designed using based upon the size of the wing. The surface area of the

horizontal stabilizer is 12% of the total wing area. The aspect ratio of the horizontal stabilizer is 4.


Figure 17: Horizontal stabilizer planform.


Horizontal Stabilizer Profile

The horizontal stabilizer profile is designed based in order to create a balanced airframe at the design

airspeed with no trimming necessary. In the Stingray, the horizontal stabilizer has an angle of incidence of

-1° to provide the downward lift needed to keep the aircraft balanced at the design cruising speed. The

symmetrical profile shown was selected in because it was the thinnest profile that would completely

contain the tail's internal components.


Figure 18: Horizontal stabilizer profile design


Vertical Stabilizer Design

The vertical stabilizer was designed based upon the stability calculations in the “Analysis” section of this

report. The taper ratio is designed to support the horizontal stabilizer and provide ample strength at the

root of the stabilizer. The aspect ratio is 4.3.


Figure 19: Vertical stabilizer planform design.


Analysis

Aerodynamic Stability

Static stability is the tendency of an object to accelerate

toward or away from an equilibrium position when it is

disturbed. A positively stable object will accelerate toward its

equilibrium position when disturbed, while an unstable object

will accelerate away from its equilibrium position when

disturbed. A neutrally stable system will not tend to accelerate

toward or away from equilibrium. Static stability can be

illustrated by the “ball and bowl” analogy shown in figures 20,

21 and 22.

A ball inside a bowl is an example of positive static stability. A

ball on an upside-down bowl is an example of a statically

unstable system. A ball on a flat table is a system with neutral

static stability.

In aerospace terms, stability is understood with respect to

three axes of motion. These axes are illustrated in figure 5. An

infinite axis drawn from the nose to the tail of an aircraft is

referred to as the lateral axis and stability around this axis is

known as roll stability. An axis that runs from wingtip to

wingtip is called the longitudinal axis and stability around this axis is called pitch stability. An axis that runs

vertically through the fuselage of an aircraft is called the directional axis and stability about this axis is

called yaw stability. In order to fully comply with the requirements set forth above for aerodynamic

stability, the Stingray must be positively stable in all three axes.


Figure 20: Statically unstable.

Figure 21: Neutral static stability

Figure 22: Positive static stability.


Roll Stability

Stability about the lateral axis of an aircraft is governed by the mass

distribution of the aircraft and the aerodynamic properties of the

aircraft's shape. The location of an aircraft's center of mass affects

the its roll stability in the same way that a mass distribution affects

the stability of a pendulum. An aircraft that has a center of mass

that is higher than its center of lift is analogous to an upside-down

pendulum; it is statically unstable. Conversely, an aircraft whose

center of mass is below its center of lift is positively stable, as is a

pendulum whose center of mass is located below its pivot point. In the Stingray, the center of gravity of

the airframe is .625” below the center of lift, accomplished by locating heavy internal components such as

the batteries and autopilot low in the fuselage and by attaching the wing root slightly above the center of

the fuselage. This distance will significantly contribute to the Stingray's positive roll stability.

Roll stability is also affected by the shape of the aircraft's wing shape. When viewed from above, backsept

wings have a stabilizing affect. Dihedral is the angle of the wings with respect to a line drawn between the

wingtips. The Stingray's dihedral angle of 3° will also cause the aircraft to be stable in roll.

With the combination of dihedral and low center of mass, the Stingray with definitely be stable in roll. Most

aircraft are designed to

have nearly neutral roll

stability, often with

between 1° and 2°

dihedral. The Stingray's

larger dihedral angle will

cause the aircraft to be

more stable in roll than

most aircraft.


Figure 24: The Stingray's dihedral angle is 3°.

3°

Figure 23: Aerodynamic axes.


Pitch Stability

Pitch stability is determined by a value called the Longitudinal Stability Margin (LSM). The LSM is the

distance along the lateral axis from the mean aerodynamic center to the center of mass divided by the

mean chord length of the wing. An aircraft with an LSM of less than -0.05 will be unstable in pitch. Most

aircraft have a LSM of around zero, as does the Stingray. In pitch stability, the Stingray will be positively

stable.

Directional Stability

Directional stability is determined by a value called the Directional Stability Margin (DSM). The DSM is

calculated by taking the moment about the center of mass of the aircraft's area projected on the airframe's

right plane and dividing by the total projected surface area of the airframe. The resulting value refers to

the distance from the center of mass to the center of lateral area. The directional stability margin is often

expressed as a percentage of the vertical stabilizer's moment arm with respect to the center of mass.


Figure 26: Determining the center of lateral area.

Figure 25: LSM comparison.

Qualitative Stability LSMVery Stable 0.2Stable (Hornet UAV II) 0Less Stable -0.05Neutral -0.1


As shown in the preceeding figures, the Stingray's directional stability margin is 25.67%, providing good

directional stability.

Airfoil Selection

Airfoil Principals

An airfoil is a body that is designed to produce desired forces under a range of airflow conditions. Airfoils

are usually designed to produce lift at certain angles of attack while minimizing profile drag. Standard

airfoil designs are usually expressed as two dimensional “profiles”. The most important factors that affect

the behavior of an airfoil profile are thickness, thickness position, camber, and camber position. These

values are usually expressed as percentages of the wing's chord length. Airfoil profile behavior can be

approximated using computational fluid dynamics, but the most accurate method of predicting an airfoil

profile's behavior is by wind tunnel testing. Large volumes of airfoil wind tunnel testing data are now


Figure 28: Determining the DSM.

Item Area (in^2) Moment Arm (in)Wing 1 21.4 -4.95 -105.93Wing 2 21.4 -4.95 -105.93Fuselage 86.4 3.59 310.18Left Landing Gear 15.6 5.38 83.93Right Landing Gear 15.6 5.38 83.93Tail Boom 24.9 -30.45 -758.21Tail Gear 5.2 -50.29 -261.51Vertical Stabilizer 44.1 -50.01 -2205.44Horizontal Stabilizer 1.4 -50.57 -70.8

DSM% 25.67

Moment of Area (in^3)

Figure 27: Directional stability comparison (Lennon, pg. 46).

Super directional stability 22Good directional stability 25Neutral directional stability 28Mild directional instability 30Very directionally unstable 33 and up


publicly available through the NASA website. Because of the wide availability of NACA (now NASA) airfoil

wind tunnel testing data, the Stingray team decided to use a NACA airfoil profile on the Stingray.

Airfoil wind tunnel data is expressed in terms of the following terms:

● Coefficient of Lift (CL)

The coefficient of lift is a dimensionless parameter that relates the lifting force that

an airfoil creates at 25% of the chord length from the leading edge of an airfoil to

the velocity and density of air passing over it.

CL = 2 FL / ρ x V2 x S

Where ρ is the density of air, V is the free airstream velocity, and S is the total

wing area.

● Coefficient of Drag (CD)

The coefficient of drag is a dimensionless parameter that relates the drag force that

an airfoil creates at 25% of the chord length from the leading edge of an airfoil to

the velocity and density of air passing over it.

CD = 2 FD / ρ x V2 x S


wing area.

● Coefficient of Pitching Moment (Cm)

The coefficient of pitching moment is a dimensionless parameter that relates the

moment that an airfoil creates at 25% of the chord length from the leading edge of

an airfoil to the velocity and density of air passing over it.

CM = 2 M / ρ x V2 x S


wing area.

● Reynold's Number (Re)



The Reynold's number is a dimensionless parameter that takes into account the size

of the airfoil being considered and the density and velocity of the air passing over

it.

Re = V x L / ν

Where V is the free airstream velocity, L is the chord length, and v is the kinematic

viscosity of air.

● Angle of Attack (AoA or α)

The Angle of Attack is the angle between the airfoil's chord line and the airflow

incident on that airfoil.

Requirements

Airfoil selection is perhaps the most important aspect of aircraft design. An aircraft's airfoil determines that

aircraft's behavior over the range of conditions that it must operate in. Airfoil selection is determined by

the particular requirements of the aircraft being designed. An aerobatic airplane, for instance, requires the

use of a very different type of airfoil than a glider. In the case of the Stingray, the airfoil profile was

selected based upon the requirements set forth earlier in this report. The requirements that affect airfoil

selection are as follows:

● Can carry 3 lbs of cargo

The estimated weight of the Stingray is 8.8 lbs including the 3 lb cargo. The heavier

the aircraft is, the more lift the wings must create to keep it aloft. Therefore, the

cargo weight requirement directly affects the choice of airfoil profile.

● Capable of stable slow flight (<25MPH with no danger of stalling)

The selected airfoil profile must be able to maintain a lift force equal to the total

weight of the aircraft at slow speeds. Stall behavior should also be docile.

● Fast cruise speed (>30MPH straight and level)

The selected airfoil profile must be able to maintain a lift force equal to the total

weight of the aircraft at higher speeds with minimal drag.



Method

The first step to selecting an airfoil is to express the most extreme flight phases required in terms of the

dimensionless parameters governing airfoil behavior shown above. The next step is to use iteration to find

an airfoil that meets the requirements of each flight phase with minimal drag. A trade off is usually

required between slower stall speed versus low-drag at higher speeds. In the case of the Stingray, slow

stall speed has been favored over high-speed performance in airfoil selection.

The wind tunnel test data shown below is for the NACA 3413, which will be used in the Stingray. This

airfoil profile was selected by reviewing dozens of similar profiles in an interative process and narrowing

the search down to 6 profiles. The NACA 3413 was then selected as a trade off between low and high-

speed performance.

Landing Phase:

Velocity: 8.94 m/s

CL = 2 FL / ρ x V2 x S = 2 x (35.6 N) / (1.23 kg/m3) x (8.94 m/s)2 x (0.58 m2)

= 1.26

Re = V x L / ν = ( 8.94 m/s) x (0.31 m) / (1.51 x 10-5 m2/s)

= 183,500



The NACA 3413 has a low drag when creating a coefficient of lift between 0 and 1.5. At the Reynold's

number that the above figure shows, the Stingray will be creating a coefficient of lift of 1.26. This

corresponds to a coefficient of drag of .025.


Figure 29: Coefficient of Lift versus Coefficient of Drag for NACA 3413 from wind tunnel test data.

-0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.60

0.02

0.04

0.06

0.08

0.1

0.12

NACA 3413Re = 183,500

Coefficient of Lift

Coe

ffici

ent o

f Dra

g


The stall characteristics of the NACA 3413 at Re = 183,500 are smooth and there is no historesis. The

profile stalls at a coefficient of lift of 1.5. At 20 mph, the Stingray will be stall safe with a safety factor of

1.2.


Figure 30: Coefficient of Lift versus Angle of Attack for NACA 3413 from wind tunnel test data.

-6 -4 -2 0 2 4 6 8 10 12 14-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

NACA 3413Re = 183,500

AoA

Coe

ffici

ent o

f Lift


High Speed, Straight and Level:

Velocity: 22.35 m/s

CL = 2 FL / ρ x V2 x S = 2 x (35.6 N) / (1.23 kg/m3) x (22.35 m/s)2 x (0.58 m2)

= 0.20

Re = V x L / ν = ( 22.35 m/s) x (0.31 m) / (1.51 x 10-5 m2/s)

= 458,800

NACA 3413 has a low drag when creating a coefficient of lift between 0 and 1 with a Reynold's number of

458,800. At the Reynold's number that the above figure shows, the Stingray will be creating a coefficient


Figure 31: Coefficient of Lift versus Coefficient of Drag for NACA 3413 from wind tunnel test data.

-0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.80

0.01

0.01

0.02

0.02

0.03

0.03

0.04

0.04

0.05

NACA 3413Re = 458,800

Coefficient of Lift

Coe

ffici

ent o

f Dra

g


of lift of .02. This corresponds to a coefficient of drag of .013.

With a coefficient of lift of 0.02, the AoA of NACA 3413

is -3°. At the design max airspeed, the wing chord must

have a downward angle of incidence. As a result, the

Stingray's wing is rigged at a 3° downward angle with

respect to the fuselage to reduce drag at this airspeed.


Figure 33: Wing chord angle of incidence

Figure 32: Coefficient of Lift versus Angle of Attack for NACA 3413 from wind tunnel test data.

-6 -4 -2 0 2 4 6 8 10 12 14-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

NACA 3413Re = 458,800

AoA

Coe

ffici

ent o

f Lift


Wing Planform

Wing loading is the ratio of an aircraft's total weight to it's wing area.

Wing Loading = Weight / Wing Area

Wing loading can be estimated by examining aircraft with performance characteristics similar to the

desired values. Figure 22 shows that most model aircraft have a wing loading of about 22 oz/ft2. Because

the Stingray will have a lower stall speed than many of these models, the aircraft will have a wing loading

of 20 oz/ft2. Once wing loading has been determined, total wing area can be easily calculated. The Stingray

has a total wing area of 950 in2.

Wing Area = Weight / Wing Loading

Wing Area = (132 oz) / (20 oz/ft2) = 6.5 ft2 = 950 in2

The aspect ratio of the aircraft can now be determined based upon the wingspan limitation. Since the

Stingray's maximum wingspan is 75 in, the necessary average chord length must be 12” in order to

maintain the correct wing area. Therefore, the aspect ratio is 6.25.

Aspect Ratio = Wing Span / Chord Length

Aspect Ratio = (75 in) / (12 in) = 6.25


Figure 34: Wing loading and aspect ratio comparison.

Model Type Aspect RatioHigh-Speed, highly maneuverable 24 5Hornet UAV II 20 6Moderate-speed sport 19 7Low-speed trainer 14 9Slope gliders 13 9Soaring gliders 10 12

Approximate Wing Loading


The Stingray's taper ratio was determined by the stable slow-speed requirement. Tapered wings reduce

drag and more closely approximate the ideal elliptical wing planform than a rectangular wing. However,

too much taper can cause tip stalling. A 65% taper ratio is often used, but the Stingray uses a 70% taper

ratio in order to be sure that the wings will not tend to tip stall in slow flight, thereby facilitating stable

slow flight.

Airfoil wind tunnel data represents the behavior of airfoils with infinite aspect ratio. Induced drag from the

airfoil can be calculated if we know the coefficient of lift and aspect ratio of the airfoil.

Coefficient of Induced Drag = Cdi = CL / π x Aspect Ratio

= .02 / 3.1416 x 6.25 = .001



Manufacturing Method

Prototype Manufacturing

By definition, “a prototype is a full-scale working model used to test a design concept by making

actual observations and necessary adjustments”. After all the research and conceptual design of the model

was completed, the decision of weather to build a prototype or to just go forward with the final product

had to be made. In spite the fact that prototype building would likely increase the cost and also exhaust

much of the already limited time, the importance of a successful outcome made choosing the building of

the prototype option the safer route. Some other factors that contributed to our prototype decision were

the following:

-Allow for last minute changes to our original design before much time, material, and money is spent;

Making carbon fiber tooling can be extremely expensive and time consuming, and though it is something

that needs to be done, avoiding the having to make it a second or third time is the goal.

-Material for a prototype build is usually cheaper and more readily available than carbon fiber.

-Keeping from having to use computer numerical control machines during the prototype can also allow

builders who don’t necessarily have access to such a tool to participate.

Why Balsa?

After the decision to build the prototype was made, choosing the right material for this task was just as

important. Material options for this portion of the project were: -Balsa with Monocote-, -High Density

Foam-, -Fiber Glass-, and any combination of some of these.

The material of choice was Balsa wood with Monokote coating. Availability, cost effectiveness, adequate

strength to weight ratio, and easy to work with were some of the factors contributing towards making the

choice. Being able to build a model that would accurately represent the characteristics of the final product

is all that was needed, and the balsa model promised all that with the exception of physical strength;

however, if built correctly, there would be more than enough strength to validate the design.

The high density foam would be a real appealing approach to a prototype solution if CNC machining the

mold would be accessible; the effort of not wanting to do mold machining more than once was one of the



factors contributing to our prototype first choice, it made no sense to use foam as the build material.

Using fiberglass as the material of choice would also require making of molds, thus the balsa method came

out on top.

After the choice was made, designing the balsa prototype took center stage. Same wing profile, same

scale, similar weight, and like dimensions, didn’t mean same design whatsoever. Some of the following

issues had to be address before the balsa plane would be built: wing bending strength, wing torsional

rigidity, fuselage bending and torsion, portability, component locations, feature resolution, and build-ability.

Maintaining the correct wing profile and scale was as important as building the prototype. To achieve the

desired profile, the wing was constructed by placing many parallel ribs that were the shape, size, and

orientation of the cross section of the wing at their location. An I-beam was constructed by gluing 1/16

inch balsa sheeting sections between two 1/4 by 1/8 inch spruce spars; the height of the balsa sheeting

controlled the I-beam height and also maintained the desired spacing. The sheeting section widths acted

as rib spacers, and their locations were chosen based on the pre-determined strength necessity as well as

desired wing profile resolution. Both the top and the bottom spruce spars of the I-beam then were mated

with unidirectional carbon fiber strips, to further strengthen the beam for added safety factor. At the

locations of the aileron hinges pockets were built up with 1/8 inch thick balsa sheeting, and later foam

filled for increased bonding surface area. The leading edges of the wings were sheeted with 1/16 inch

balsa sheeting; the sheeting ran from the leading edges back to the top and bottom of the I-beams to

form the D-tube wings. D-tube design is very effective in both bending and torsion.



Figure 35: Right hand wing in process.

A round aluminum spar was sized to protrude trough the center of the fuselage and into the wings; the

protrusions into the wings were 8.5 inches deep, penetrating the first 6 ribs of each of the wings. The

maximum allowable spar into the wing protrusion was limited by the dihedral angle. The wings were

designed with 3 degrees of dihedral, to match that of the final product. It was crucial to achieve a large

enough moment arm to steer away from the aluminum spar breakout.

Buildups with 1/8 in liteply were introduced for aileron housing, and also reinforcement for the wing to

wing strap mating.

The ailerons were carved out of solid balsa wood, and then further lightened by removal of unnecessary

material. The material removal was performed with a 2 inch hole saw.



Figure 36: Balsa aileron and aileron hinge pockets.

The ailerons ended up being much stronger than necessary, but the extra strength came at the cost of

extra weight.

The fuselage was first framed with 1/4 by 1/8 inch spruce spars, then completely sheeted with 1/16 inch

balsa sheeting.1/8 liteply bulkheads were installed to control the cross sectional shape and scale of the

fuse in the desired locations. The carbon tail-boom protruded to about half of the fuselage into a double

bulkhead support for a larger moment arm.

Figure 37: Carbon boom to fuselage interface.



Buildup for extra support was introduce in places such as landing gear mounting platform and firewall for

the necessary strength.

Figure 38: Carbon boom to fuselage interface.

The entire fuselage, hatch, and wings were covered with Monokote wrap to achieve the desired low-drag

body that would closely represent the final model.

Installation of the motor was done by mounting an aluminum motor bracket to the front of the firewall,

which was holding the back side of the Axi 4130/20 outrunner moto.



Figure 39: Electric motor location image.

Carbon Fiber Composite Lamination

The manufacturing plan for the Stingray final project has several components. Considering the wide array

of parts that this aircraft will employ, several manufacturing processes will have to be utilized concurrently

with on another to achieve the final deliverable product. Processes ranging from annealing, quenching, and

tempering to vacuum bagging, to manual machining, to CNC machining will all have to be utilized. This

plan will consist of the strategy to hurdle all of problems that will arise with the manufacturing of this

complete aircraft.

For the manufacturing of the airframe for The Stingray, we have determined that the material best suited

to fit all of our design requirements (light, strong, geometrically complex) is pre-impregnated woven

carbon fiber material. Carbon fiber is a material that was developed in the 1950’s but is just is the past

decade really experiencing the extent of its possibilities in manufacturing. Carbon fiber is being used all

over industry. Ranging from racing bicycle frames to supersonic aircraft to a basic accent panel in an

automobile, carbon fiber is a material that provides a wide range of manufacturing options that are not

feasible with conventional metal. The ability to efficiently manufacture complex geometric parts while

maintaining a lightweight and structurally rigid frame presents an unlimited amount of options for this

material.



Carbon fiber starts as a dry cloth but this will not assist in manufacturing because there is no

mechanism to permanently form the shapes of the required design. The mechanism that performs this task

is epoxy resin which hardens when cured. To date, it is possible to manufacture carbon fiber parts using

two different pre-cure configurations: Pre-Preg or a standard ‘Wet’ Lay-up. Pre-Preg (or pre-impregnated)

carbon fiber is a carbon cloth that comes from the manufacture of the raw material with the resin pre-

impregnated into the cloth. This configuration has advantages and disadvantages. The main advantage to

this method is that the resin content is controlled. The manufacturer of the material specs-out the

necessary amount of resin needed and applies it in a uniform coat throughout the woven cloth. This makes

controlling resin content much more manageable. You are able to determine just how much resin you want

to remove during vacuum bagging and how much you want to retain with in the cloth. This method is

especially important in weight sensitive applications. The ability to remove all the unneeded resin while still

retaining a rigid part is very appealing to certain applications. One downside to the pre-preg configuration

is that the material as to be stored in a freezer to ensure that the resin doesn’t begin to prematurely cure

before lamination. Another, less relevant, quality of the Pre-Preg is that the actual handling of the material

is much easier and cleaner due to the fact that you don’t have to mix and apply resin. Mentioning this

comes to the second configuration, ‘Wet’ Lay-Up. This configuration also has its advantages and

disadvantages. In this process, dry cloth is the baseline which is formed in to the desired shape. Followed

by this, the resin/hardener is applied to each layer of the dry woven cloth. After this application, the excess

resin is ‘squeegeed’ from the mat. This configuration is desirable if you want a quick, and cheap way of

laminating a composite part. The one major disadvantage of the process versus a Pre-Preg lay up is weight

and resin content. When the resin is applied and removed in the ‘Wet’ Lay-Up process, the operator has

very little control over resin content and thus the resulting weight if the part.

For our application in particular, we have decided that a Pre-Preg material is the only way that we will be

able to accurately predict the weight of the aircraft in its post-cure state. Taking into account that weight is

an aircraft’s worst enemy, this Pre-Preg material is choice for the aircraft. The content of the resin is

predetermined and with minimal preproduction parts needing to be made for testing, the correct part will

be able to me produced quicker and with less waste.




Figure 40: Pre-preg composite layup.



Figure 41: Carbon fiber composite layup.

Figure 42: Carbon fiber pre-preg uni-directional material.


Figure 43: Finishing the cured bottom horizontal stabilizer.



Curing Cycle

Once that the composite laminate material has been selected, the curing cycle is one of the next plans that

was to be determined. The cure cycle is the amount of heat that is applied to the composite laminate in

order to ensure that the part will harden properly. This cycle included 3 major sections: Ramp-Up, Plateau,

and Ramp-Down. The Ramp-Up portion of the cycle is the rate at which temperature inside the

oven/autoclave must rise to ensure that the entire laminate and tool heats up uniformly. This portion of

the cure is also dependent on the laminate material, resin and tooling material. The composite laminate

material we are using has a cure temp range from 240ºF - 270ºF. This range was predetermined by the

manufacture of the material. During ramp-up, the only factor that will have to be monitored will be to

verify that the rate doesn’t exceed 1ºF / minute. This is the ramp-up rate that the manufacturer of the

tooling material we have chosen has recommended. This is less than the ramp-up rate of the composite

laminate material which is 5ºF / minute so this will not affect the laminate. The next section of the cure

cycle is the plateau. This is the portion of the cure will remain constant for 1 hour at 252ºF. This time will

allow for the resin to completely cure

uniformly and conform exactly to the

desired shape in the mold or plug. This

portion of the cycle is the mechanism for

the actual melting of the resin within the

woven fabric fibers. The final portion of the

cycle is the Ramp-Down portion. This

section of the cycle allows the part to

properly, uniformly cool at a safe rate where

the integrity of the visible and molecular

structures remain in the as-molded condition. This rate is also a number that is dependent on several

factors. The manufacturer of the pre-preg laminate that will be used has recommended a cool rate of

6ºF/min max. The tooling material has a Ramp-Down rate of 4ºF/ per minute. Understanding these two

values, the ramp-down portion of the cure cycle will be maintained at 4ºF/minute. If the part is cooled too

rapidly, warping and cracking of this brittle material can occur so the ramp-down portion of the cycle is

extremely critical.


Figure 44: Cure cycle: Temperature versus time.


One more critical aspect of curing is a material property known as the CTE (Coefficient of Thermal

Expansion). This is a property that is imbedded within the material that actually is the ammunition behind

the cure cycle and ramp rates. The ideal situation is to match the CTE of the tooling material to the CTE of

the composite laminate. This will ensure that the warping of the composite laminate is minimized because

the two materials will expand and compress at the same rate when under uniform heat.

Tooling Machining and Care

A technician is only as good as the tools he/she has at their disposal. This is the reason behind having

strict tool control and care. It is essential to the outcome of a manufacturing process that the tool being

use is going to produce exactly what was theoretically designed or envisioned. Tools, also, can be one of

the most expensive aspects of an operation which is why extreme care must be taken during machining

and storing of the manufactured tools.

Another aspect of the tooling design is deciding what type of composite laminate tooling is going to

be necessary to produce the parts for the aircraft. Given the timeframe that the Stingray design team has

been allotted, we have two options for tooling: molds or plugs. A mold, otherwise classified as a female

tool, is a negative cavity cut down into the raw, stock tooling material. This style of composite lamination

tooling allows the OML (Outer Material Line) to be controlled because the first ply of laminate will

coincident with the surface. Considering the aerodynamic design of the UAV, the OML is what needs to be

controlled to ensure that the air profiles and low drag bodies remain just that. This allows the engineers to

adjust thicknesses of the laminate without altering the OML. All of the tooling for the Stingray will be

female tooling. The other option for composite laminate tooling is a plug, otherwise classified as a male

tool. This style of tooling allows the IML (Inside Material Line) to be controlled. The tools is cut in a

configuration that provides a positive boss to be laid-up around and additional plies to be added to the


Figure 45: Ball-nose end mill.


outer, exposed surface. This is less desirable for the Stingray application due to the fact that the OML is

being controlled and the IML will be used for composite modification by method of ply build up.

Due to the complex shapes that the Hornet UAV II: Stingray employs, all of the molds will had to

be manufactured by CNC milling. Concurrently, the CNC programs had to be ran using an array of ball

nosed end mills. The round end of the mill allowed the complex, ruled surface to be void of any sharp

edges. The CAD files that have been generated for the individual mold parts were imported in the CAM

(Computer Aided Manufacturing) program known as MasterCAM. This is a program that will take all of the

profiles from the CAD files and create a matching NC (Numerical Code) which was then uploaded into the

CNC mill. This program should have provided all of the information necessary to machine the parts and

virtually eliminated the possibility of the human machining error. As stated, ‘virtually’ is a nice word to

claim. We had several different instances where a factor came into play where there ended up being some

sort of error during the machining process. On the manufacturing of the nose cone tool, for instance, the

machinist chose the incorrect tool and proceeded to machine the entire, overly complex tool with a flat end

mill when we had designed the tool for a ball end mill to be used to produce the joining ellipses. This

resulted in roughly 10 hours of wet sanding and polishing the tooled surface to achieve the mirror finish we

were looking for on the tool. This, along with a few other of the molds, had minor errors during the

machining processes but nothing that prevented us from having to re-run any of the tools.

The tooling material that was initially chosen for the composite parts of the Hornet UAV II: Stingray

was initially a high density, high temperate, machineable tooling polyurethane foam. This foam material is


Figure 46: Final left-lower wing mold design.


specifically designed for cost effective way of composite lamination. This foam is a substitute for more

traditional composite tooling like aluminum. Unfortunately, the sponsorship through Coastal Enterprises fell

through so we had to adapt to the different possibilities of tooling material out in the industry. After a little

research, we were able to find that drops of 6061-T6 aluminum were not that much more than that of the

predicted cost of the foam if we were to purchase it on the open market. This realization actually helped us

greatly because we were actually able to manufacture tooling the ‘correct’ way. These machined tools

would be able to produce thousands of parts and the only wear would be operator induced. The aluminum

tooling was one of the best decisions we made during this project.

Not only does this allow use to worry less about the tools being damaged from operator neglect or

warp during cure, we were able to produce an amazing part out of this tooling. The surface finish was

phenomenal and the strength was out of this world. The rigidity of an autoclaved, prepreg tooled carbon

part yields the most beneficial aspect of composite lamination. One limiting factor that we had to deal with

was the shear size of the NC machines that we had access to. The NC mill in the ECS Tech Lab has an

envelope size of roughly 20”x18”x10”. This machine was going to be able to accommodate out vertical

stabilizer tools, rudder tools and nosecone/nosecone plug tools. The NC machine in the Physics Tech Shop

was able to accommodate the top and bottom horizontal stabilizer molds.

The major issue that we ran into was the fuselage and LH/RH wing molds. Not only were these

molds extremely complicated is geometry, the parts were very large; too large for either of the machines

to accommodate. This quickly made us have to switch directions into figuring out to QUICKLY and

accurately manufacture the molds necessary. After several thoughts, we came to the conclusion that

manufacturing the molds out of maple hardwood was the cheapest way to go. This process consisted of

taking the model and splitting it into cross sections ( as shown in Figure XXXXX), plotting them out to

scale, cutting the out manually on a band saw and aligning them on rods which accommodated a common

area on the parts. These cross sections were glued together using standard wood glue. The next step was

the tedious part, shaping. The areas of concentration were the interface areas including the nosecone,

wings and tail boom attachments. These areas were tightly controlled using the model to validate against.

The main downfall using this process is that we were forced to use a ‘wet’ lay up for several different

reasons. The ‘wet’ lay-up parts use a room temperature curing resin system that does not require ramp

rate or elevated cure temperatures to achieve the cured state the end user is looking for. The other issue



was warping due to heat and pressure of an autoclave. The wood might have been able to withstand the

heat and pressure but we did not want to risk the tool/laminate warping, catching fire, or breaking during

the cure cycle. These were the main issues with the fuselage and wing tools. So taking all of this into

account, we decided to try 2 plies of the 12k dry twill material for the first of the wing lamination cures.

What we did was created a template of what the piece of carbon was supposed to be according to the tool

size. We then proceeded to cut out the material per the template and taped the edges of the dry mat

because of how easily the material frayed. The tape was removed during lamination. Next what we did was

lay the 2 pieces out on a table covered in wax paper so that the laminate wouldn’t stick to the table or get

contaminated with FOD (Foreign Object Debris). Lastly, we proceeded to mix the epoxy resin system with

a ratio of 5 parts resin and 1 part catalyst. This was per the manufacturer’s instruction of the epoxy resin

system. After to talking several composite company’s about the composite lamination using the wet lay-up

technique, we decided to go with a 60% mat weight and 40% resin weight ratio. This also was going to be

incorporated into a vacuum bagging envelope which helped draw the majority of the resin that was evenly

applied out into the breather cloth and leaving the correct amount of resin into the system. This allowed us

to achieve a light and strong part all at the same time.

Another down fall with using male tools and this wet lamination is that we wrapped the outside of

the tools with the laminate and proceeded to vacuum bag. This presented a few problems because of how

the parts needed top be de-molded in a certain way due to the complex geometry we laid LH and RH side

of the fuselage separately, de-molded and then bonded the two half together with an internal superseam.

This process is not preferable for several reasons. First off, the external surface finish ends up looking less

than desirable because this surface is in contact with the bagging material and not with the smooth surface

of the tool. This is merely aesthetics, though, and were be massaged to have a more desirable surface

finish. The wings were fabricated in a similar fashion and ended up with the same issues but they were

about to be worked through.



Figure 47: Wrapping the wing tool with two plies of carbon laminate.



Figure 48: The left and right-hand outboard wings curing under vacuum and a heating blanket.



Figure 49: Cured left-hand outboard wing before demolding.



Figure 50: The fuselage plug.



Figure 51: Applying mold release to the wing plugs.



Vacuum Bagging

The mechanism for actually forming the shapes within the machined tool is a process known as Vacuum

Bagging or De-Bulking. This is a process where a multi-layer ‘sandwich’ of several different application

specific materials are drawn by vacuum down onto the desired part shape(s). The main factor in the de-

bulking process that needs to be monitored is vacuum pressure and the two main contributing factors to

this monitoring is the bagging material and the tooling material. The tooling material that has been chosen

for this manufacturing process doesn’t allow for more than 35 psi to be drawn onto the tool otherwise,

deformation will most likely occur. The other factor is the vacuum bagging material itself. This material can

allow for up to 50 psi drawn before the bag will tear or puncture. Knowing these two factors, we will de-

bulk at 30 psi.

Within the de-bulking process, there are 5 main materials that each contribute to achieving the

desired laminate. First is called Peel Ply. This is an abrasive cloth that will be the first layer of the

‘sandwich’ of de-bulking materials. This layer will be coincident with the last layer of pre-preg applied to

the laminate. This material is used for producing an etched surface in post-cure configuration. This is ideal

for any surface that will be bonded post-cure. The second layer is the Bleeder Film. This material is a

perforated, thin plastic material that will lie on top of the Peel Ply. This Bleeder Film allows for resin to be

removed from the composite laminate under vacuum pressure. The sizes of the perforations in the cloth

control how much resin will be removed from the laminate. Thirdly is the Breather Cloth. The breather

cloth is a thick, polyester, woven fabric used to absorb and retain the excess resin that is drawn from the

composite laminate while under vacuum. Again, the engineer has the option to choose thickness/density of

this material depending on the application and amount of resin desired to remove. The second to last layer

of the ‘sandwich’ is the Bagging Film. This material is similar in chemical composition to the Bleeder Film

mentioned earlier but void of the perforations. This is the layer that actually provided the applied force

onto all of the aforementioned layers and compresses the laminate into the desired shape. The last

component of the vacuum bag is the chromate tape used to seal the bag. This tape is a thick, formable

tape which seals the bag to ensure a proper vacuum. A valve is attached to this layer which is in turn

attached to a vacuum pump. Vacuum pressure is monitored through the attachment of a standard Burdon

gauge. See Figure for a cross section of the vacuum bagging ‘sandwich’ on top of a Male tool.

There are several types of vacuum bagging techniques but the two that will be employed in this



project is the Standard configuration and the Envelope Bag configuration. The configuration shown in

Figure XXX is a standard configuration where the chromate tape is attached to the tool itself. The bagging

material is adhered to the chromate tape and drawn in one direction. The Envelope bag configuration is

similar but this is where the entire tool is placed inside a bag and vacuum is drawn. The bag is made by

sealing a sheet of bagging film to another sheet of bagging film and sealed with the chromate tape to

ensure vacuum. Both of these methods will yield accurate parts if used concurrently with the proper tool

configuration.

Within the de-bulking process, there are 5 main materials that each contribute to achieving the

desired laminate. First is called Peel Ply. This is an abrasive cloth that will be the first layer of the

‘sandwich’ of de-bulking materials. This layer will be coincident with the last layer of pre-preg applied to

the laminate. This material is used for producing an etched surface in post-cure configuration. This is ideal

for any surface that will be bonded post-cure. The second layer is the Bleeder Film. This material is a

perforated, thin plastic material that will lie on top of the Peel Ply. This Bleeder Film allows for resin to be

removed from the composite laminate under vacuum pressure. The sizes of the perforations in the cloth

control how much resin will be removed from the laminate. Thirdly is the Breather Cloth. The breather

cloth is a thick, polyester, woven fabric used to absorb and retain the excess resin that is drawn from the


Figure 52: Vacuum-bagging assembly cross-section.


composite laminate while under vacuum. Again, the engineer has the option to choose thickness/density of

this material depending on the application and amount of resin desired to remove. The second to last layer

of the ‘sandwich’ is the Bagging Film. This material is similar in chemical composition to the Bleeder Film

mentioned earlier but void of the perforations. This is the layer that actually provided the applied force

onto all of the aforementioned layers and compresses the laminate into the desired shape. The last

component of the vacuum bag is the chromate tape used to seal the bag. This tape is a thick, formable

tape which seals the bag to ensure a proper vacuum. A valve is attached to this layer which is in turn

attached to a vacuum pump. Vacuum pressure is monitored through the attachment of a standard Burdon

gauge. See Figure for a cross section of the vacuum bagging ‘sandwich’ on top of a Male tool.

There are several types of vacuum bagging techniques but the two that will be employed in this

project is the Standard configuration and the Envelope Bag configuration. The configuration shown in the

figure above is a standard configuration where the chromate tape is attached to the tool itself. The

bagging material is adhered to the chromate tape and drawn in one direction. The Envelope bag

configuration is similar but this is where the entire tool is placed inside a bag and vacuum is drawn. The


Figure 53: The vacuum-bagging process applied to a composite part.


bag is made by sealing a sheet of bagging film to another sheet of bagging film and sealed with the

chromate tape to ensure vacuum. Both of these methods will yield accurate parts if used concurrently with

the proper tool configuration.

Figure 54: Example of the nosecone layup.



Figure 55: Example of staggered peel ply sandwich lamination.



Figure 56: Vacuum bagging of the bottom vertical stabilizer.



Figure 57: The autoclave used to cure the Stingray's parts.



Ply Schedule

The pre-preg composite laminate is built in layers or plies. Each of these plies are coincident with

each other which is one of the ways of increasing the strength of the part that is being laminate. The ply

schedule is the determined orientation of the laminate plies with respect to a fixed axis as well as with

respect to each other. As the plies are laid up, the ability to stagger the plies orientations can allow for

minimum thickness (minimal weight), structural rigidity, and strength in the direction of the applied loads

to the surface. Carbon fiber is best in tension and worst in compression so to maximize the amount of the

part that will remain intension is critical.

Knowing that the pre-preg carbon fiber that will be used on the composite skins for the Hornet UAV

II: Stingray is 90° plain weave, we oriented the first ply parallel with the determined 0° axis. The next ply

was clocked 45° clockwise to the original 0° axis. This pattern of 45° staggering allowed the part to remain

strong in virtually all directions of loading.

The Hornet UAV II: Stingray is a lofting, slow flying aircraft which was manufactured from ultra

high strength composite material. This being said, the ply thickness will range from 3 to 6 plies thick

throughout the airframe. This need for this range of thicknesses is understanding and identifying the areas

of high stress. The areas of build up include but are not limited to countersunk holes, interfaces, and

corners. The areas of build-up were concentrated areas where specific sizes of laminate will be cut and fit

to that local area of high stress.


Figure 58: Ply direction is expressed in terms of a common

axis.


Non-Composite Manufacturing

Along side all of the composite laminate manufacturing, the Stingray team will also have to

incorporate traditional manufacturing techniques and methods as well. Considering the fact that the

majority of the parts in the Stingray are composite, the traditional manufacturing methods will be kept to a

minimum.

One of the most complex, non composite manufactured parts on the Stingray are the pivot pins

that will act as both anchors and pivots for the elevator and rudder control surfaces. To make these parts

as simple as possible, standard, off-the-shelf parts are going to be utilized as the baseline configuration.

These off-the-shelf standard 82° countersunk screws are going to have the tips of the threads machined

off while leaving ~3/4 of the original thread shank intact. The problem that arises with this manufacturing

process is that the screws are already hardened and cannot be accurately machined in the baseline

configuration. To accomplish this machining, the screws will annealed, quenched and tempered. This

process will allow us to alter the molecular structure of the metal before machining occurs and return the

structure to its baseline configuration after the machining is complete.

The remaining, complex non-composite manufactured part on

the aircraft is the AFT sleeve which will provides the flat interface for

the vertical stabilizer to attach to the tail boom. This is a part that

begins as a stock 6061-T6 piece of aluminum and milled and turned to

achieve the desired configuration. The specific alloy was chosen for

several reasons. 6061-T6 aluminum is an alloy made with magnesium

and silicon as the alloying elements. This will allow the alloy to be easily

machined and welded; both processes will be involved in the

manufacturing of this particular part. This alloy provides the strength

and lightweight properties that are essential to the design of the

Stingray aircraft. One downside to this alloy is that it is corrosive. To counteract this, the part in its post-

machined configuration will be anodized. Anodizing is an electrolytic passivation process used to increase

the thickness and density of the natural oxide layer on the surface of metal parts. This will improve the

corrosion resistance of the raw metallic part. The part will be anodized to the military specification MIL-


Figure 59: Stingray Pivot Pin.


A-8625, Class II, Type II, Black. This specification is a standard anodizing called out for military

applications. The specification reads as follows: MIL-A-8625 is the specification number, Class II is the

thickness of the anodizing, Type II is the hardness of the anodizing, and Black is the resulting color of the

plating. The Class and Type classifications have been chosen in the middle range of the possible selections

to not overkill the price and lead-time of the plating operation.

Figure 60: Tailwheel mount detail.



Standard Parts

To minimize the amount of custom manufactured, the Stingray will utilize as many off-the-shelf parts as

possible. The utilization of standard parts allows the project to not only employ tested and proven parts, it

also expedites acquiring the parts as they tend to be in high supply. This emphasizes a concurrent

manufacturing technique known as DFM (Design For Manufacturing). DFM is a technique where the actual

manufacturing of there system is taken into consideration during the initial design phase. This in turn

allows for the parts to be quickly and accurately manufactured and assembled with minimal problems on

the production floor.

Expanding on this principle, the aircraft will even minimize the number of standard parts that will

be used in the assembly. For example, all of the screws that will attach all of the composite components

together will all be 4-40 x 5/16” Torx screws. This will allow for the less likelihood of error during assembly

as well as the easeability of acquiring only one type of screw.

Mechanical components aside, all of the electronic components will be off-the-shelf, readably

available parts. Both of the batteries, the speed controller, the receiver, servos, and motor are all standard

parts. This has an extreme advantage due to the fact that this is a mechanical project and if an issue arises

with an electrical component, time shouldn’t have to be wasted trouble shooting and electrical failure.

These are issues can be taken up with the distributor of the standard parts and progress will proceed from

that point.

BOM

The Stingray has a vast amount of components to the top level, deliverable assembly. To keep track of all

of the parts, sub-assemblies and top level assemblies, a structured, cascading BOM (Bill Of Materials) has

been developed. The Stingray team has designed this accounting system of all of the parts and quantities

used on the aircraft. This system was more favorable versus a standard parts list due to the fact that the

aircraft has been truly designed for manufacturing.

A part numbering system has also been developed. The part numbering convention that has been

designed is to assist in the accounting of subassemblies and so that parts are not duplicated in said

subassembly. This systems allows for and accurate accounting of all parts that their associated next level



assembly.

Lastly, a tool numbering list has been created as well. This tool numbering convention is concurrent

with the system that was designed with the part numbering convention. Again, this is a system that will

allow the team to keep an accurate accounting of the tools and which parts that those specific tools will

yield.

Finishing

After all of the composite and non-composite manufactured parts are molded or machined, assembling all

of the components will be necessary. The assembly that will take place will reflect the structure of the BOM

(Bill of Materials).

The vast majority of the composite skin parts will be manufactured in top/bottom or left/right

configurations. This entails that the top/bottom or left/right skins will have to be bonded together after

cure. This is standard for small scale composite parts because laminating small scale parts in a continuous

laminate can present several issues; cost being the main contributing factor. The skins will be bonded

together using a 3M® epoxy that is specifically used for the bonding of carbon composite laminates.

It is the nature of composite manufacturing that hand working and manual processes are

incorporated into the manufacturing plan. The composite parts that will be manufactured and molded will

have flash that extends beyond the desired material limit. This is to ensure that that the edges are fully

built and cover the designed profile. The flash is then hand worked (sanded, ground) off and the desired

profile is achieved. This is only possible, though is the proper tolerance study is completed will allow for

small variances in the profile as the parts are hand worked and human error does become a contributing

factor.

Once all of the top/bottom ad left/right carbon skins are properly bonded together, the aesthetic

finishing work can begin. The bond seams will want to be covered up considering that the epoxy we will be

using is bright blue and the carbon laminate is metallic black. This is aesthetically unpleasing. The gap will,

though, be converted into a useful entity as it will be painted using a reflective, white paint which will allow

for the aircraft to be located in the sky during flight.



Manufacture Hurdles

Aside from all of the advantages the carbon composite manufacturing offers to the Stingray, there

remain challenges that need to be address and overcome before flight.

One of these hurdles that must be scaled is the fact that carbon fiber is a metallic material. Regardless of

the premonition that this wouldn’t present a problem, the issue of signal loss and morphing came

prevalent. The global design of this aircraft is to incorporate low-drag bodies throughout. This includes

fuselage, wings, vertical stabilizer and horizontal stabilizer. The plan was to keep the antenna inboard,

eliminating any external body to affect the aerodynamics of this aircraft. Accepting the fact that the

Stingray must be made of carbon fiber, this determination was made that external antennas were going to

be absolutely necessary to ensure no signal loss will occur during flight. The aircraft will have external

antennas mounted as flush as possible to reduce air drag. This aspect of the initial conceptual design was

overlooked and now that there is a better understanding of the non-mechanical properties of the carbon

composite laminate, this issue will not arise in the future.

Another manufacturing hurdle that arose was the issue

of convective cooling of all of the electronic

components inside the fuselage. Since the Stingray will

be operated under the power of a solely electronic

propulsion system, the electronics components need to

be properly cold to remain within the operating

temperature range determined by the manufacturer of

the components. To handle this, possible catastrophic

electronics failure, the design team has developed an

air inlet concentric with the opening in the nose come

for the propeller shaft. The air drawn into the nose

cone will then flow through the FWD motor bulkhead

which has been designed with several perforations.

This will in turn allow the air drawn into the fuselage to

convectively cool the electronics components during

flight operations. See the figure above for a section view of the nose cone and convective bulkhead. The


Figure 61: Nosecone assembly.


nose cone is shown transparent for clarity.

Test PlanThe testing plan that this aircraft will be subjected to several tests ranging from physical strength testing,

thermal electronics validation, theoretical FEA testing in Solidworks, and much more. The testing that this

aircraft will endure will validate that the airframe will be durable for the conditions that we have designed it

to withstand. Considering the fact that there are virtually infinite amount of situations that could present

problems, the Stingray design team will limit the testing to major areas of possible failure within the

airframe. The fact is that the design team is under the time constraint to deliver the completed system by

May 15th, 2008 and the amount of validation testing that could possibly be completed on the airframe

could prevent the system from being delivered. Below will be the main system test that will be ran outside

of the prototype airframe validation.

Thermal Electronics Validation

For the electronics portion of the systems testing, the Stingray design team validated the operation of the

electronics system after operating temperature were reached. All of the electronics components that the

Stingray utilizes are standard, off-the-shelf parts which have a predetermined temperature operating range

established by the manufacturer of the components. Understanding this, the design team used the

component with the lowest operating temperature as the maximum temperature to be achieved with the

inside of the fuselage. First will be a pre-flight test. The aircraft will be constrained to the ground and the

system will be tested by revving the motor in a variable operating range. Thermocouples will be placed

inside the fuselage which will validate that the temperature will not rise above the determined operating

temperature that the design team will set. Once this pre-flight test is validated, the design team will then

proceed to a fully operational flight test. The aircraft will be taken off, flown for ~20 minutes at variable

speeds, and brought back down to the ground for landing. Immediately after landing, the design team will

insert thermocouples into the fuselage and validation of the operating temperature will be tested and

recorded. Once this can be confirmed, the thermal electronics testing will be complete. If the temperature

range is not reached, additional aspects of the airframe which allow for convective cooling will have to be

implemented.



Airspeed Testing

The team had to determine a method to validate that the Stingray is controllable at the design airspeeds.

In order to accomplish this, the team flew the aircraft at various throttle settings at low altitude. The team

used a RADAR gun to verify the airspeed. In order to eliminate the effects of wind on the test, the aircraft

was clocked both upwind and downwind and the average speed was determined. The results of the speed

test follow:

Flight Condition: Airspeed:

Stall 18

Slow Flight 15

Cruise, ½ Throttle 55

Straight and Level, Full Throttle 65

Airframe Meets All Primary Requirements

● Positive or neutral stability in all axes. After releasing the controls, the

aircraft returns immediately to straight-and-level flight.

● The aircraft is capable of at least one hour flight time. After a 50 minute

flight, the batteries were still at 15 volts, well above the 14.8 nominal

voltage.

● The aircraft has a cavernous electronics bay that will fully accommodate the

autopilot, video transmitter, and camera.

● Cutout for GPS antenna

● Able to carry 3lbs of cargo

● Camera mount with clear visibility

Airframe Meets Most of the Secondary Requirements

● Capable of stable slow flight (<25MPH with no danger of stalling). Aircraft's

stall speed was gentle and occurred at less than 18 MPH.

● Fast cruise speed (>30MPH straight and level). Aircraft's cruise speed at



half throttle was demonstrated to be 55 MPH.

● Easy maintenance (All components serviceable)

● Portable (removable wings/tail surfaces). The aircraft fits in a standard

automobile trunk.

● Durable (Recoverable after loss of control over flat asphalt surface with less

than $500 damage). This requirement, thankfully, was not verified in a

destructive test. However, the team is confident that the aircraft would not

sustain significant damage under these conditions.

● Access to electronics, DFA



ReferencesBeavin, Micheal; Looney, Paul. (2002) Legislative Update. Retrieved October 1, 2007, from AIAA Aerospace

America Online Web site: http://www.aiaa.org/aerospace/Article.cfm?

issuetocid=272& ArchiveIssueID =32

Protecting America. (2006) Retrieved October 1, 2007, from Office of Management and Budget:

Department of Defense Web site: http://www.whitehouse.gov/omb/budget/fy2006/defense.html


http://www.aiaa.org/aerospace/Article.cfm?issuetocid=272&

http://www.aiaa.org/aerospace/Article.cfm?issuetocid=272&

http://www.whitehouse.gov/omb/budget/fy2006/defense.html

http://www.aiaa.org/aerospace/Article.cfm?issuetocid=272&ArchiveIssueID


Appendix A, Competition Rules



Appendix B, Drawings


Project Report, Stingray

Documents

Transcript of Project Report, Stingray