Wichita State University

Design Report AIAA/Cessna/Raytheon

Design Build Fly Competition April 2012

Flight Mechanics –Alfredo Gimenez

Team Lead - Tawny Blumenshine

Aerodynamics – Mitchell Nord

Propulsion – Wesley Lambert Structures – James Winkel

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AC Aerodynamic Center

AOA Angle of Attack in Degrees (also α)

AR Aspect Ratio

AVL Athena Vortex Lattice

CAD Computer Aided Drawing

CAM Competition Altimeter for Models

CBA Computerized Battery Analyzer

Acronyms, Abbreviations, and Symbols

Cd0 Coefficient of Drag at Zero Lift

CG Center of Gravity

CL Coefficient of Lift

Clo Airfoil Coefficient of Lift at Zero AOA

CNC Computer Numerical Controller

DBF Design Build Fly

DC Direct Current

e Oswald Efficiency

ESC Electronic Speed Controller

FEM Finite Element Analysis

FOM Figures of Merit

fps Feet per Second

FS Factor of Safety

ft Feet

G Acceleration due to Gravity

GPS Global Positioning System in Inches

Kv RPM per Volt

L Liters

lbs Pounds

m Meters

M1 Mission 1 Score

M2 Mission 2 Score

M3 Mission 3 Score

MAC Mean Aerodynamic Chord

mAh Milli-Amp Hours

min Minutes

MLG Main Landing Gear

MTOW Maximum Takeoff Weight

NICAD Nickel-Metal Cadmium

NIMH Nickel-Metal Hydride

OIA Other Important Aspects

oz Ounces

PDIP Preliminary Design Iteration Process

Prop Propeller

psi Pounds per Square Inch

R/C Radio Controlled

RAC Rated Aircraft Cost

rad/s Radians per Second

Re Reynold’s Number

RPM Revolutions per Minute

s Seconds

SM Static Margin

t/c Airfoil Thickness to Chord Ratio

T/O Takeoff

T/W Thrust to Weight Ratio

Tavg Average Overall Time to Climb to 100m

TTeam Team Time to Climb to 100m

™ Trademarked

W/S Wing Loading

WSU Wichita State University

WTT Wind Tunnel Test

α1 Steady State Angle of Attack

β Sideslip Angle

δa Aileron Deflection

δe Elevon Deflection

δe1 Steady State Elevon Deflection

% Percent

˚ Degrees

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Contents 1.0 Executive Summary ................................................................................................................................ 7 2.0 Management Summary .......................................................................................................................... 8 2.1 Team Organization ........................................................................................................................... 8 2.2 Design Schedule .............................................................................................................................. 9 3.0 Conceptual Design ................................................................................................................................. 9 3.1 Mission Requirements1 .................................................................................................................. 10 3.2 Competition Scoring Analysis ........................................................................................................ 10 3.3 Competitive Design Requirements ................................................................................................ 12 3.4 Conceptual Design Selection Process ........................................................................................... 12 3.5 Configuration Selection .................................................................................................................. 13 3.5.1 Propeller Location Selection................................................................................................. 14 3.5.2 Landing Gear Selection ........................................................................................................ 15 3.6 Selected Conceptual Design .......................................................................................................... 16 4.0 Preliminary Design................................................................................................................................ 16 4.1 Critical Design Parameters ............................................................................................................ 16 4.1.1 Aerodynamic Critical Parameters ......................................................................................... 16 4.1.2 Flight Mechanics Critical Parameters ................................................................................... 17 4.1.3 Propulsion Critical Parameters ............................................................................................. 17 4.1.4 Structural Critical Parameters .............................................................................................. 18 4.2 Mission Model ................................................................................................................................ 19 4.3 Optimization Tools and Methodology ............................................................................................. 19 4.4 Initial Sizing .................................................................................................................................... 20 4.4.1 W/S vs T/W .......................................................................................................................... 20 4.4.2 Preliminary Design Iteration Process (PDIP) ....................................................................... 20 4.4.3 Initial Weight Build Up .......................................................................................................... 21 4.5 Aerodynamics ................................................................................................................................ 22 4.5.1 Airfoil Selection ..................................................................................................................... 22 4.5.2 Aerodynamic Performance Predictions: Lift and Drag ......................................................... 23 4.6 Flight Mechanics ............................................................................................................................ 25 4.6.1 Control Surface Sizing, Placement and Authority ................................................................ 25 4.6.2 Longitudinal Stability ............................................................................................................ 26 4.6.3 Trim Drag.............................................................................................................................. 27 4.6.4 Lateral-Directional Stability................................................................................................... 27 4.6.5 Dynamic Stability .................................................................................................................. 28 4.6.6 Sustained Turn Dynamics .................................................................................................... 29 4.6.7 Servo Selection .................................................................................................................... 29 4.6.8 Stability Derivatives .............................................................................................................. 29 4.7 Propulsion ...................................................................................................................................... 29 4.7.1 Battery Selection .................................................................................................................. 29 4.7.2 Motor Selection .................................................................................................................... 30 4.7.3 ESC Selection ...................................................................................................................... 30 4.7.4 Wire Selection ...................................................................................................................... 30 4.7.5 Propeller Selection ............................................................................................................... 31 4.8 Structures ....................................................................................................................................... 32 4.8.1 Critical Loads ........................................................................................................................ 32 4.8.2 Materials ............................................................................................................................... 33 4.8.3 Spar Trade Study ................................................................................................................. 33 4.8.4 Spar Carry Through Structure .............................................................................................. 34 4.8.5 Motor Mount Optimization .................................................................................................... 34 4.9 Landing Gear ................................................................................................................................. 34 4.9.1 Dimensions ........................................................................................................................... 35 4.9.2 Attachment ........................................................................................................................... 35 4.9.3 Material ................................................................................................................................. 35 4.10 Preliminary Mission Performance Predictions .............................................................................. 35

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5.0 Detail Design ........................................................................................................................................ 36 5.1 Wing Iteration ................................................................................................................................. 36 5.2 Hoerner Tips .................................................................................................................................. 36 5.3 Fuselage Aerodynamics................................................................................................................. 37 5.3.1 Fuselage Nosecone ............................................................................................................. 37 5.3.2 Tail Fairing Assembly ........................................................................................................... 37 5.3.3 Wing Fillets ........................................................................................................................... 38 5.4 Water Tight Fuselage ..................................................................................................................... 38 5.5 Water Release Mechanism ............................................................................................................ 38 5.6 Flight Performance Summary ........................................................................................................ 39 5.7 Drawing Package ........................................................................................................................... 39 5.8 Aircraft Component Weight and CG Buildup ................................................................................. 44 6.0 Manufacturing Plan and Processes ...................................................................................................... 44 6.1 Manufacturing Methods.................................................................................................................. 44 6.2 Prototypes ...................................................................................................................................... 45 6.3 Mission Ready Model ..................................................................................................................... 45 6.3.1 Tooling .................................................................................................................................. 46 7.0 Testing Plan .......................................................................................................................................... 47 7.1 Half-Scale Wind Tunnel Testing .................................................................................................... 47 7.2 Propulsion System Testing ............................................................................................................ 48 7.2.1 Battery Testing ..................................................................................................................... 48 7.2.2 Total System Testing ............................................................................................................ 49 7.2.3 Results.................................................................................................................................. 50 7.3 Structural Testing ........................................................................................................................... 51 7.3.1 Final Spar Validation ............................................................................................................ 51 7.3.2 Wing Tip Test ....................................................................................................................... 52 7.3.3 Final Wing Structural Validation ........................................................................................... 52 7.4 Full Scale Wind Tunnel Test .......................................................................................................... 53 7.5 Ground Testing .............................................................................................................................. 54 7.6 Flight Testing .................................................................................................................................. 55 8.0 Performance Results ............................................................................................................................ 56 9.0 References ........................................................................................................................................... 58

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Figure 1 - Team Organizational Chart .......................................................................................................... Figures

8 Figure 2 - Design Schedule4 ......................................................................................................................... 9 Figure 3 - Total Score Contour Plot ............................................................................................................ 11 Figure 4 - Percent Change Method ............................................................................................................. 12 Figure 5 - Concept FOM Analysis ............................................................................................................... 14 Figure 6 - Propeller Location FOM Analysis ............................................................................................... 15 Figure 7 - Landing Gear Configuration FOM Analysis ................................................................................ 16 Figure 8 - Conceptual Design Aircraft ......................................................................................................... 16 Figure 9 - Wing Loading Chart .................................................................................................................... 19 Figure 10 - PDIP Flow Chart ....................................................................................................................... 20 Figure 11 - Power Required - Power Available ........................................................................................... 21 Figure 12 - Liebeck la2573a Reflexed Airfoil Performance Data18 ............................................................. 23 Figure 13 – Methods Used for 3-D Performance Predictions ..................................................................... 24 Figure 14 – Component Parasite Drag Buildup .......................................................................................... 24 Figure 15 – Total Aircraft Drag Prediction at Re = 250,000 ........................................................................ 25 Figure 16 - Athena Vortex Lattice Model .................................................................................................... 25 Figure 17 –M1 Moment Balance Illustration ............................................................................................... 26 Figure 18 - Longitudinal Trim Plots for M3 .................................................................................................. 27 Figure 19 - Trim Drag Optimization ............................................................................................................. 27 Figure 20 - Control Surface Crosswind Trim Authority for M3 T/O ............................................................. 28 Figure 21 - Root Locus for M1 Cruise19 ...................................................................................................... 28 Figure 22 - Wire Weight per Power Drop .................................................................................................... 31 Figure 23 - Propulsion Flow Chart .............................................................................................................. 31 Figure 24 - Power Required ........................................................................................................................ 32 Figure 25 - Load Path Diagram ................................................................................................................... 32 Figure 26 - Aircraft Component Weight Breakdown ................................................................................... 33 Figure 27 - Motor Mount FEM Fringe Plot .................................................................................................. 34 Figure 28 - Landing Gear Breakaway Plate Design ................................................................................... 35 Figure 29 - Nose Cone Side View ............................................................................................................... 37 Figure 30 - Water Visibility Test Drop System ............................................................................................ 39 Figure 31 - Ground Test Photo of Water Release Mechanism ................................................................... 39 Figure 32 - The Piranha Bending the MLG ................................................................................................. 46 Figure 33 - Main Landing Gear Construction .............................................................................................. 46 Figure 34 - Half-Scale Model Mounted in WSU’s 3ftx4ft Wind Tunnel ....................................................... 47 Figure 35 - Computerized Battery Analyzer System ................................................................................... 48 Figure 36 - Cold Soak Test Results ............................................................................................................ 48 Figure 37 - Propulsion Test Apparatus ....................................................................................................... 49 Figure 38 - Actual Recorded Static Power .................................................................................................. 50 Figure 39 - Actual Recorded Thrust Compared to Predictions ................................................................... 50 Figure 40 - Actual Recorded Cruise Power ................................................................................................ 51 Figure 41 - Spar Whiffletree Setup ............................................................................................................. 51 Figure 42 - Wing Tip Test ............................................................................................................................ 52 Figure 43 - Final Wing Whiffletree Test ...................................................................................................... 53 Figure 44 - Two Point Mounting System with Fairing for Dynamic Tare (Left) ........................................... 53 Figure 45 - Tufts Showing Separation Near Stall........................................................................................ 54 Figure 46 - First Ferry Flight January 13, 2012........................................................................................... 56 Figure 47 - Lift and Drag Comparison for Half-Scale and Full-Scale WTTs ............................................... 56 Figure 48 - Recorded GPS Speed (Wind 12 mph) ..................................................................................... 57

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Tables Table 1 - Total Score Sensitivity Study Results .......................................................................................... 11 Table 2 - Structural Weight Build Up ........................................................................................................... 22 Table 3 - Propulsion System Weight Build Up ............................................................................................ 22 Table 4 - Airfoil Screening Requirements ................................................................................................... 22 Table 5 - Final Five Airfoils .......................................................................................................................... 23 Table 6 - Demonstrated Elevon Trim Authority ........................................................................................... 26 Table 7 - Mission Specific Static Margins ................................................................................................... 26 Table 8 - Sustained-Turn Dynamics Characteristics .................................................................................. 29 Table 9 - Stability Derivatives for M3 T/O with a 21-knot Crosswind .......................................................... 29 Table 10 – Motor Comparison .................................................................................................................... 30 Table 11 – ESC Comparison ...................................................................................................................... 30 Table 12 - Spar Shape Trade Study ........................................................................................................... 34 Table 13 – Preliminary Mission Performance ............................................................................................. 36 Table 14 - Final Aircraft Dimensional Parameters ...................................................................................... 36 Table 15 - Flight Performance Parameters ................................................................................................. 39 Table 16 – Component CG and Weight Buildup ......................................................................................... 44 Table 17 – Mission Specific CG and Weight Estimates .............................................................................. 44 Table 18 - Half-Scale Wind Tunnel Test Matrix .......................................................................................... 47 Table 19 - Setup Information....................................................................................................................... 49 Table 20 - Propulsion Test Plan Matrix ....................................................................................................... 49 Table 21 - Full-Scale Wind Tunnel Test Matrix ........................................................................................... 54 Table 22 - Pre-Flight Checklist .................................................................................................................... 55 Table 23 - First Flight Test Matrix ............................................................................................................... 55 Table 24 - Flight Information ....................................................................................................................... 57

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1.0 Executive Summary

This document examines the design, testing, and manufacturing of Wichita State University’s (WSU)

Team Black to prepare for the 2011-2012 AIAA/Cessna/RMS Design/Build/Fly (DBF) competition. Team

Black’s primary objective is to design and build a winning aircraft by maximizing the total score, which is a

function of the report score and the flight score. The flight score is determined by the performance in

three missions: Ferry Flight, Passenger Flight, and Rate of Climb Flight. The Ferry Flight focuses on the

speed of the aircraft where the scoring equation is a function of the number of laps completed. The

Passenger Flight focuses on the payload to weight fraction of the aircraft where the scoring equation is a

function of the flight weight. Finally, the Rate of Climb Flight focuses on the time it takes to climb to 100

meters (m), where the scoring equation is a function of the team’s time to climb and the competitive

average time to climb1. A sensitivity analysis on the Flight Score shows that the Rated Aircraft Cost (RAC)

is the most critical design parameter.

Vehicle concepts are created to meet mission requirements and obtain a maximum total score. A

screening analysis minimizes the number of concepts to six viable options: low wing conventional aircraft,

high wing conventional aircraft, the bumblebee (a low wing conventional aircraft where the chord length is

the same length as the fuselage with and without a horizontal tail), and a blended wing body (with and

without a tail). Figure of Merit (FOM) analyses are applied to select an aircraft configuration and layout,

which yields a blended wing body configuration with no tail. This concept includes a single tractor

propeller with a reverse tricycle landing gear configuration. Both of these components reduce the empty

weight while still fulfilling the mission requirements.

Multi-disciplinary optimization architecture is developed using each core engineering discipline which

performs an analysis on critical design parameters. A Liebeck la2573a airfoil is utilized due to the

thickness to chord ratio (t/c) and the amount of lift at zero angle of attack (Clo), allowing the aircraft to

cruise at a low angle of attack (AOA). The wing is sized at 739.2 in2, based on the lowest wing loading

(W/S) the aircraft experiences. A root chord of 16.5 in. is chosen based on the center section dimension

for the passenger payload and a 12° maximum fuselage deviation in order to reduce flow separation2. A

wing sweep of 15° is chosen due to previous wind tunnel tests conducted at WSU which shows that the

maximum coefficient of lift (CL), for a swept wing, is obtained at 15° leading edge sweep3. Propulsion

analysis yields a Rimfire .10 motor with a propeller of 10 inches in diameter. The predicted RAC weight is

1.6 pounds (lbs).

The predicted performance capabilities of Missions 1/2/3 are as follows:

• Mission 1 (M1): With a cruise speed of 70 feet per second (fps), the aircraft is capable of flying

six laps in the required four minutes of flight.

• Mission 2 (M2): The aircraft propulsion system and aerodynamic characteristics are chosen to

ensure successful takeoff (T/O), fly three laps, and land successfully.

• Mission 3 (M3): With a rate of climb speed of 10 fps, the aircraft will climb to 100 m in 100

seconds (s) and successfully land.

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2.0 Management Summary 2.1 Team Organization

A team hierarchy and project schedule is created to ensure an efficient design timeline. The WSU

Team Black consists of five senior Aerospace Engineering students with the help of underclassmen,

graduate students, and faculty. This structure organizes the five senior engineering students, the

advisors, and the underclassmen into essential disciplines as shown in Figure 1.

Figure 1 - Team Organizational Chart

Each discipline is accountable for important aircraft parameters and developing analysis tools to aid

in the overall design effort. Cross-functional roles are also given to each section lead in order to help

promote team communication and collaborative design processes. Each discipline and the corresponding

responsibilities are as follows:

• Project Manager: Takes into account other important aspects (OIA) of design that do not

directly correlate within the core disciplines. In charge of organizing design optimization.

Requests weekly updates from section leads on design progress. Mediates design discussion

while maintaining a level of healthy conflict to promote new ideas.

• Flight Mechanics: Evaluates aircraft logitudinal, lateral and directional, and dynamic stability in

conjunction with determining the longitudinal, lateral, and directional stability derivatives. Control

surface sizing, placement and procurement of needed hardware for the optimal design.

• Propulsion: Evaluates propulsion system performance by developing techniques to analyze the

necessary requirements for selecting the ideal motor, propeller, and batteries. Oversees the

acquiring of the transmitter, receiver, and any other needed electronic hardware to operate the

aircraft efficiently and safely.

Tawny BlumenshineProject Manager

Performance LeadStructures Staff

Alfredo GimenezFlight Mechanics Lead

Aerodynamics Staff

Wesley LambertPropulsion Lead

Flight Mechanics Staff

Mitchell NordAerodynamics LeadPerformance Staff

James WinkelStructures LeadPropulsion Staff

Dr. L S MillerDepartment Chair

Chief Advisor

Troy LakeDBF Manager

AdvisorsKevin Kelly

Dr. James Steck

AdvisorsJonathan Mowrey

Josh Nelson

AdvisorsJonathan Krenzel

Phil Meikel

AdvisorsDr. Suresh Keshavanarayana

Iwan Broodrӱk

Adriana BarraganAdam MaurathMichael Staab

Michael LambCameron Schwanke

James Tennant

Troy LakeAaron MaurerCristina Wilson

Joseph GraybillArnold Durel Deffo Nde

Miguel Correa

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• Aerodynamics: Responsible for the selection and designing of the wing, airfoil, and geometry.

Works in conjunction with the Flight Mechanics and Propulsion staff in order to optimize the

aircraft through means of lowering the drag and increasing the lift. Compromises with the

structures staff on an aircraft that can perform efficiently without exceeding the realm of what

can be structurally possible.

• Structures: Responsible for designing primary and secondary structure that efficiently handles

all loads. Seeks out new, light weight materials that can improve the final score. Calculates

stresses and deformations from loads that are received by the supporting staff. Relays back to

team with improvements that can increase performance.

2.2 Design Schedule

A fast paced schedule with set deadlines is developed in order to keep the team on pace to stay

competitive. Historical trends are examined regarding the areas where WSU DBF traditionally

exceeds the predicted time. Extra time is placed in these areas to ensure the same mistakes are not

repeated. Figure 2 is the project Gantt chart that tracks design, manufacturing, testing, and report

progress.

Figure 2 - Design Schedule4

3.0 Conceptual Design The 2011-2012 DBF Contest rules1 outline three flight missions. These mission constraints define

competitive design requirements. FOM analyses determine the optimum vehicle configuration.

Apr-12Mar-12Feb-12Jan-12Dec-11Nov-11Oct-11Sep-11Aug-11Jul-11Jun-11May-11

DesignTeam Dynamics

Read and Outline Raymer Design Textbook

Mission Requirements Review

Conceptual Design

Preliminary Design

Detailed Design

Design Optimization

9/2 2/28ReportEntry Form

Report Draft

Report Editing

Report Due

9/28 4/1TestingBattery/Fuse Testing

Propulsion System Testing

Materials Testing

Prototype WTT

Prototype Flight Test

Full-Scale Structural Test

Full-Scale WTT

Flight Tests

Inter-University Fly Off

DBF Competition

Design Schedule 2011 - 2012

PredictedActual

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3.1 Mission Requirements1

Contest specified mission and vehicle requirements are as follows:

• Maximum T/O distance of 100 feet (ft)

• Maximum battery weight of 1.5 lbs

• Each mission must be flown in order

• Allowed four flight attempts to complete three missions

• Eight 1x1x5 inch (in) aluminum bars, weighing a total of at least 3.75 lbs, simulate passengers

with the 5 in dimension being vertical to the aircraft during flight

• Two liters (L) of water must be carried to an altitude of 100 m then released by a servo operated

dump valve

• All payloads must be carried within the mold lines of the aircraft

• All payloads and aircraft assembly must be completed in under 5 minutes (min)

Mission 1: Ferry Flight

The aircraft is flown with no payload. Flight time begins when the throttle advances and runs for four

min. The aircraft will fly as many laps as possible in the flight time and must land on the runway to receive

a score.

𝑴𝑴𝑴𝑴 = 𝑴𝑴 + 𝑵𝑵𝑵𝑵𝑵𝑵𝑵𝑵𝑵𝑵𝑵𝑵 𝒐𝒐𝒇𝒇 𝑳𝑳𝑳𝑳𝑳𝑳𝑳𝑳

𝟔𝟔

Mission 2: Passenger Flight

The aircraft takes off and flies three laps while configured with 8-1x1x5 in aluminum bars. The

aircraft must then land on the runway to receive a score.

𝑴𝑴𝑴𝑴 = 𝑴𝑴.𝟓𝟓 + 𝟑𝟑.𝟕𝟕𝟓𝟓

𝑭𝑭𝑭𝑭𝑭𝑭𝑭𝑭𝑭𝑭𝑭𝑭 𝑾𝑾𝑵𝑵𝑭𝑭𝑭𝑭𝑭𝑭𝑭𝑭

Mission 3: Time to Climb Flight The aircraft will be loaded with a payload of 2 L of water. Flight time begins when the throttle

advances. The aircraft takes off and climbs to 100 m. Once the altitude is reached, a rule specified

altimeter will de-activate a servo that will open a dump valve, dropping the water out of the aircraft. The

judges will stop the flight time once the water is visible. The aircraft must then successfully land on the

runway to receive a score.

𝑴𝑴𝟑𝟑 = 𝑴𝑴 + �𝑻𝑻𝑳𝑳𝒂𝒂𝑭𝑭𝑻𝑻𝑭𝑭𝑵𝑵𝑳𝑳𝑵𝑵

3.2 Competition Scoring Analysis

The total competition score is equal to the product of the report score and the sum of the flight

scores divided by the square root of the RAC. RAC is defined as the maximum empty weight of all three

missions. Before each mission, the aircraft will be weighed and the maximum of the three weights will be

considered as the RAC.

𝐓𝐓𝐓𝐓𝐓𝐓𝐓𝐓𝐓𝐓 𝐒𝐒𝐒𝐒𝐓𝐓𝐒𝐒𝐒𝐒 = 𝑹𝑹𝑵𝑵𝑳𝑳𝒐𝒐𝑵𝑵𝑭𝑭 𝑺𝑺𝑺𝑺𝒐𝒐𝑵𝑵𝑵𝑵∗(𝑴𝑴𝑴𝑴+𝑴𝑴𝑴𝑴+𝑴𝑴𝟑𝟑)√𝑹𝑹𝑹𝑹𝑹𝑹

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Capitalizing on the report score is essential to increasing the total score and winning the competition.

Initial tech inspection and flight queue order is based on the ranking earned on the report. Scoring highly

on the report provides the opportunity to fly in calm weather. It is imperative to fly early in the morning or

later in the evening in the state of Kansas. This statement is based on historical trends5 for the month of

April. A survey of historical competition data and WSU DBF aircraft data determines nominal values for

this year’s scoring parameters. Sensitivity studies, shown in Table 1, performed with the 2k Factorial

Reduction Method, quantify each parameter’s effect on total score. Ultimately, scoring analysis results in

the determination of realistic and competitive design requirements and constraints.

Table 1 - Total Score Sensitivity Study Results

The 2k Factorial Reduction Method outputs a contour plot (Figure 3) which displays a range of

parameters to design to, while still receiving the same total score6.

Figure 3 - Total Score Contour Plot

Percent change analysis is performed to ensure that the results in Table 1 are accurate and the

trends are valid. The input parameters are both increased and decreased to understand how to achieve

the highest total score (Figure 4). Each trend line validates the respective percent effects determined via

the 2k Factorial Reduction Method. RAC and Number of Laps are the most influential factors in the total

Parameter Nominal Best Value Nominal Worst Value % ContributionRAC 1 lb 7 lb 95.44%Number of Laps 8 1 2.58%Report Score 98 90 0.73%Tteam 23 s 60 s 0.67%Tavg 40 s 75 s 0.29%Flight Weight 4.75 lb 10.75 lb 0.27%

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score other than the report, which already has a high priority. The report score range is able to be bound

smaller following historical WSU report score trends.

Figure 4 - Percent Change Method

3.3 Competitive Design Requirements

Competitive scoring analysis results determine a combination of requirements for a winning design:

• RAC less than 1.3 lbs

• Number of Laps greater than six

• Report Score greater than 93

• Tteam less than 180 s

• Flight Weight less than 5.05 lbs

3.4 Conceptual Design Selection Process Screening and Scoring of a variety of different concepts helps ensure that only the best concept is

chosen to further develop in preliminary design. If this process is not carried out in a well thought

out manner, the concept chosen can fall apart later during preliminary design, thus severely

hindering the chances of being competitive. The design schedule (Figure 2) allows for extra time to

determine the most viable concept. Well thought out criteria for selecting concepts are compromised

upon by all team leads to ensure that each has the capability to achieve the set goals. Evaluation of

the competitive scoring analysis (Figure 3 and Figure 4) results in the following FOM with the

respective weightings:

• Multi-purpose Structure (40%): This year’s rules call for an aircraft structure that can complete

a wide range of missions without the addition of extra weight. Emphasis is put into designing a

structure for multiple uses. “Multi-Purpose Structure” is treated as a primary screening

parameter of different concepts.

-35%

-30%

-25%

-20%

-15%

-10%

-5%

0%

0% 10% 20% 30% 40% 50% 60%

% C

hang

e in

Tota

l Sco

re

% Change in Scoring Parameter

RAC = 1

Number of Laps = 8

Tavg = 60

Tteam = 23

Flight Weight = 4.75

Report Score = 98

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• Payload Configuration (30%): The ability to hold the needed payload without adding large

volume while tying into the already designed structure is crucial. One of the design requirements

is to assemble and load the aircraft in under five min. Therefore, the payload must be simple to

load. Another design requirement is that the payload must be kept within the mold lines of the

aircraft. Therefore, the geometry of the payload cannot hinder the aircraft’s performance. During

M3, the payload is dropped. The ability to have the payload at the aircraft’s center of gravity

(CG) to eliminate CG shifting during flight is considered.

• Manufacturing and Assembly (10%): The capability to manufacture the design without major

variation enables calculations to match the actual aircraft performance. The aircraft will be easy

to load due to the five min time limit during pre-flight.

• Risks (10%): The number of uncertainties on the aircraft can cause unforeseen problems further

into the design. The conceptual design phase is the start of aircraft design. A determination that

a concept is not viable during preliminary design will set the design schedule back.

• Wind Effectiveness (10%): Kansas historically has high winds during the competition month5.

The aircraft must be able to handle high crosswinds during each mission as well as handle any

blanketing of control surfaces.

3.5 Configuration Selection Initial screening considers 11 team developed concepts. Each concept is screened with a ± rating for

each FOM outlined above. Subsequent down-selections and concept combinations result in six viable

configuration concepts. Scoring analysis is then performed on the remaining six, thus concluding on one

main concept to develop in preliminary design.

The conventional monoplane with a low wing is chosen as the baseline configuration. The FOM

analysis considers a five-point scale, where 1 = poor, 2 = below baseline, 3 = baseline, 4 = above

baseline, 5 = superior. Assigning point values to the concepts’ FOMs determines the best of the following

configurations:

• Conventional Low Wing (baseline): A monoplane with a fuselage to hold the payload and a

boom to attach the tail. The wing is mounted under the fuselage so that the landing gear can be

attached to the spar. The tail is composed of a low-mounted horizontal stabilizer and a single

vertical stabilizer. This concept follows a traditional aircraft configuration making it simple to

manufacture and low on risk.

• Conventional High Wing: A monoplane with a fuselage to hold the payload and a boom to

attach the tail. The wing is mounted above the fuselage allowing for more stability. The tail is

composed of a low-mounted horizontal stabilizer and a single vertical stabilizer. This concept is

relatively simple to manufacture and assemble.

• Blended Wing Body without End Plates: A tail-less configuration with a fuselage blended into

the wing. Elevons are used for lateral and longitudinal control. This concept is typically the

lightest configuration for a given payload requirement3.

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• Blended Wing Body with End Plates: A tail-less configuration with end plates on the wing tips.

End plates will help with adverse yaw during flight. However, it causes an addition of weight.

• Bumble Bee without a Horizontal Tail: A monoplane where the fuselage is the length of the

wing chord. The tail is composed of a single vertical stabilizer. This reduces the weight of having

a boom lead out to a conventional tail.

• Bumble Bee with a T-tail: A monoplane where the fuselage is the length of the wing chord. The

tail is composed of a high mounted horizontal stabilizer with a single vertical stabilizer. The

horizontal tail is used for longitudinal control.

Figure 5 presents the scoring analysis results for the six concepts listed above.

Figure 5 - Concept FOM Analysis

The blended wing body without end plates wins the initial FOM analysis due to its weight

advantages. However, end plates are still considered up to the preliminary design phase. If analysis

shows that end plates are a necessity, they will then be added into the concept and analyzed properly.

3.5.1 Propeller Location Selection

Once the blended wing body configuration is chosen, the propeller, number of motors and

placements are considered. An FOM analysis is performed to compare three configurations. A single

motor combined with a tractor propeller is chosen as the baseline configuration for the analysis. Point

values are assigned to each configuration’s FOMs to determine the most viable:

• Single Tractor: This is a traditional single motor/propeller configuration found on most aircraft

today. This system benefits from being lightweight and is less prone to propeller strikes. The low

risk factor is the main advantage. It is safe and can be depended on.

• Tractor and Pusher: A combination of a tractor and a pusher propeller mounted along the

centerline. Smaller propellers result in shorter, lighter landing gear. Having dual motors will

offset the benefits gained from having the extra thrust. This system can become very complex

and will add new issues if one motor fails during flight.

15

• Single Pusher: This is a setup found on many blended wings. A single motor/propeller

configuration with an aft placement. The motor will have to be pod mounted which incurs extra

structure. Propeller ground clearance presents an issue during takeoff as well.

Figure 6 presents the scoring analysis results for the three concepts listed above.

Figure 6 - Propeller Location FOM Analysis

The single tractor configuration wins the FOM analysis. This configuration will work well with the

chosen concept. The concept itself is currently deemed risky thus, adding more risk without much added

gain is disadvantageous.

3.5.2 Landing Gear Selection The aircraft must be able to take off in under 100 feet. Landing smoothly, without bouncing or

veering off the runway, is required to receive a flight score. Four concepts are selected as viable

solutions.

• Taildragger (baseline): Taildragger gear is lightweight and has a small frontal area due to its

small tail wheel2. If set up incorrectly, this design can have severe ground handling issues.

Extra precaution must be made to ensure that added risk does not exceed the benefits of the

weight savings.

• Tricycle: This landing gear concept allows for significant improvement of ground handling

before takeoff. However, tricycle gear has a prevalent nose gear compared to the small tail gear

of a taildragger concept. The nose gear protrusion increases the aircraft’s weight and drag2.

• Bicycle: Two centerline landing gear are accompanied by smaller outer gear to balance the

aircraft during taxiing. Weight and drag are similar to tricycle gear; however, unstable ground

handling becomes as issue just like the taildragger2.

Figure 7 presents the scoring analysis for the three concepts listed above.

16

Figure 7 - Landing Gear Configuration FOM Analysis

Despite having a disadvantage in ground handling, the taildragger gear wins the FOM analysis. This

design is selected for conceptual design due to the weight improvement and the elimination of extra

bulkheads that would be needed to attach the landing gear to the fuselage.

3.6 Selected Conceptual Design

The selected conceptual design (Figure 8) is a blended wing body with taildragger landing gear and

a single tractor propeller.

Figure 8 - Conceptual Design Aircraft

4.0 Preliminary Design Initial aircraft sizing begins at the preliminary design phase. This aids in the initial development of

aerodynamic, flight mechanic, propulsive, and structural characteristics of the aircraft. Trade studies are

performed in order to evaluate competitive design alternatives and find the optimal design point. Finally,

optimization in each core discipline is carried out based on trade study results. Final validation of the

analysis predictions are done with a wind tunnel test (WTT).

4.1 Critical Design Parameters

Competitive design requirements (Section 3.3) drive the analyses of the selected design. Constant

desire to decrease the overall aircraft weight imposes tight constraints on the allowable performance

characteristics. Each discipline lead then assesses the conceptual design to determine the design

parameters.

4.1.1 Aerodynamic Critical Parameters

• Wing Area: T/O distance is a major influence in the required wing sizing. The T/O distance

dictates that the aircraft will need a large wing area combined with a small propulsion system or

vice versa. A large wing area helps reduce the wing loading and will shorten the T/O distance for

17

the aircraft. However, a large wing area requires more structure thus increasing the aircraft

weight and drag.

• Wing Taper: A tapered wing allows for the majority of the wing mass to be located towards the

center of the aircraft. Mass at the wing root provides advantages in terms of flight

characteristics/stability. In regards to wing stability, a tapered wing is favored due to the natural

dihedral that is produced by the smaller t/c ratio in the span-wise direction. Taper sets the

standard for the absolute thickness at the wing root. Increasing the wing taper maximizes the

usable spar height at the wing root. This is beneficial due to the largest bending moments

occurring at the wing root. Taper reduces the size of the spar at the wing tip, thus decreasing the

aircraft weight. Low torsion characteristics are also a benefit where the torsion, generated by the

wing, is centered at the wing root allowing for more torsion to be absorbed at the wing root7.

• Wing Sweep: The aircraft aerodynamic center (AC) is aft of the CG. Wing sweep moves the

location of the AC in the aft direction. However, as the wing is swept in the aft direction the CG

moves aft as well. Depending on the CG location, the batteries may need to be placed further

forward to shift the CG forward, thus allowing for a larger static margin (SM). A proper amount of

sweep is needed to balance the location of the CG and AC.

• Wing Span: The wing span directly affects aerodynamic parameters such as aspect ratio (AR)

or the location of the mean aerodynamic chord (MAC). However, a larger wing span requires

more structure, thus increasing the structural weight. Once the wing area is derived, chord and

span combinations will be analyzed to ensure maximum performance.

4.1.2 Flight Mechanics Critical Parameters

• Static Margin: Greater than 5% SM8 is needed to minimize control surface deflections, avoid

dynamic instability, and be neutrally stable during the worst case scenario.

• Deflections: Necessary control surface deflections are set at ±20° based on the linearized

methods of Roskam9 in order for the flow to remain attached.

• Number of Control Surfaces: The number of necessary control surfaces is to be minimized in

order to avoid weight penalties.

• Stability: Longitudinal and lateral stability are to be achieved for flight performance as well as

better flight handling qualities.

• Cross Wind: The aircraft must have enough control authority to take off in a 21 knot cross-wind5

4.1.3 Propulsion Critical Parameters

• Batteries: Proper battery understanding is required to achieve the absolute lightest propulsion

weight. Rules regulate that only batteries with the nickel-metal hydride (NIMH) and nickel-metal

cadmium (NICAD) chemistries are allowed1. NIMH batteries have better energy density, allowing

for a greater endurance time, while NICAD batteries allow for a higher amperage draw. Both

types have a nominal cell voltage of 1.2 volts; however, a voltage of one is used during

conceptual analysis to account for any loss of voltage due to amperage draw. NIMH batteries

18

are favored over NICAD due to a lighter weight and lack of memory effects as compared to the

equivalent NICAD setup10. Initial sizing of the propulsion system uses M1 for the maximum

endurance required, and M3 is used for maximum power required.

• Motor: Consideration starts by determining the power required curve, matching a tier of motors

for that power, then establishing the lightest weight motor for that tier. The regulations require all

motors to be direct current (DC) brushed or brushless1. They may be direct drive or geared by

means of belt or gear reduction. Brushless motors are preferred due to less resistance and

internal friction in a lighter system. They can be further broken down into two separate

categories, inrunners and outrunners. Outrunners have higher torque and lower revolution per

minute (RPM) which leads to a direct matching of prop to motor shaft. When using an inrunner,

RPMs are too high to match directly with a prop, thus a gear reduction takes place in the form of

a gearbox. In a pure form, inrunners offer slightly better efficiency and component weight;

however the addition of a gearbox increases system complexity as well as the weight. WSU

history has shown that gearbox failure is a frequent occurrence. Due to the increased complexity

of inrunners, only outrunners are used for calculations during conceptual design.

• Propeller: Propeller diameter is calculated using an equation from Raymer2 which derives a

diameter of 14 in. However, the equation over predicts the diameter due to the simplicity of the

variables.

4.1.4 Structural Critical Parameters

• Loads Sizing: Finding a minimum load case that every aircraft in the competition will withstand

this year is crucial to sizing a competitive, lightweight structure. The wing-tip test ensures that

every aircraft has to support its maximum gross T/O weight using only simple supports located

at the wing tips1. This test loads up the aircraft structure to high stresses at the wing root which

represents what will occur during flight. Flight loads may not exceed the wing root stresses

caused during the wing-tip test of the aircraft.

• Spar Design: A two spar design is needed to resolve all the primary loads in the aircraft. The

front spar follows the lifting line of the wing to stop any major torsion from developing in the wing.

The two spar design carries all the lifting loads and provides an attachment point for all structural

elements to join into. A continuous rear spar will serve as an attachment point for the control

surfaces while providing less glue joints.

• Fuselage Configuration: The regulations require all payloads to be carried within the mold

lines of the aircraft1. However, the payloads are large, thus making a fuselage that can tie into

already existing structure a desirable option in order to keep the aircraft weight down. The

fuselage needs to be designed as a non-structural element to reduce weight where the loads are

resolved into the wing spar structure.

• Material Selection: Selecting the materials that carry significant loads without adding

significant amounts of weight is essential to reducing weight. The use of a composite main wing

19

structure comprised of foam and balsa is a viable option due to the lightweight properties of the

foam. However, incorporating additional materials, such as fishing line and Micro-Lite™ coating,

can be designed into the structure. Fishing line and Micro-Lite™ coating have lighter weight

properties than either foam or balsa.

4.2 Mission Model Each flight mission is modeled in four phases: T/O, climb, cruise, and turns (180° and 360°).

• Takeoff – In order to incorporate a margin of safety, the T/O distance is set to 90 ft. Inability to

takeoff within the regulation 100 ft distance will result in an incomplete score. Wing sizing and

propulsion requirements are set to meet the T/O requirements. A study between wing loading

(W/S) and thrust to weight ratio (T/W) is conducted to determine the wing area and thrust

needed for T/O.

Figure 9 - Wing Loading Chart

• Climb – A full throttle climb to 328 ft (100 m) is assumed. Time to climb is calculated based on

the M3 T/O weight and determining the excess power for climb angle. Excess power is derived

from the greatest difference between the available power and the power required inside the flight

envelope.

• Cruise – M1 velocity is optimized for a high scoring flight with minimal weight addition. Cruising

at 75 fps meets a competitive score of six laps.

• Turning – 180° and 360° turn rates are calculated using a 2.5G load factor as well as the throttle

settings used in cruise.

4.3 Optimization Tools and Methodology

Separate leads have designed codes to automatically update when design changes occur. These

codes enable a more efficient optimization process. During each optimization, the wing loading chart is

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8 4.0

T/W

W/S

Stall

Climb

Turn

Cruise

Takeoff

Design Area

20

updated, providing a new score. The new score is then compared with the most recent score to ensure

that improvement is made.

4.4 Initial Sizing 4.4.1 W/S vs. T/W Raymer’s method of T/W vs. W/S is utilized to find the minimum W/S needed for all flight regimes2.

Utilizing the wing loading chart (Figure 9), a minimum W/S of 1.1 is determined with a T/W of 0.2. The

W/S established a wing area of 758.88 in2 in order to complete all flights listed in the figure above.

4.4.2 Preliminary Design Iteration Process (PDIP) From a scoring standpoint, it is clear that a light weight aircraft that has the ability to fly all three

missions will be competitive in this year’s competition. The lightest possible aircraft that is able to carry

the 4.4 lbs of payload is the determining factor for this year’s competition. Team Black set up an iterative

process that takes into consideration all of the structural, propulsion, aerodynamic, and stability critical

parameters. PDIP is outlined below (Figure 10). An initial guess of one pound is considered based on a

preliminary scoring analysis. After several iterations, converging on this point is not considered possible

due to the design constraints. A second guess of 1.5 lbs is considered and then iterated to an empty

weight of 1.3 lbs. Team Black intends to start at this weight and continue to reduce through the course of

preliminary design.

Figure 10 - PDIP Flow Chart

In determining the initial size of the aircraft, iterations from the propulsion standpoint begin at a

maximum amount of battery. This amount is decreased to a minimum weight while still ensuring that the

thrust required is met for all three missions. This is done with outputs from the aerodynamic and

structural disciplines, which are then put into the next propulsion iteration until an absolute minimum is

achieved. To iterate in this fashion, assumptions are made for the propulsion system. First, a propeller

efficiency of 75%, a cell voltage of 1.0 volt, and a motor efficiency of 90% are all factored into the

theoretical maximum power achievable. Batteries are assumed to not hold the manufacturer’s advertised

energy density. Therefore, an efficiency of 85% is factored into the amp draw, except for the Kan 700

Done

Reconfigure dimensions

Determine Minimum Wing Loading Needed YES

No, Add Battery

Initial Weight Guess

Minimum Propulsion System Required

Minimum T/W needed to fly all 3 missions

Calculate wing dimensions needed

Cg < A.CStatic Margin > 6% YES

Calculate Minimum structure needed for maximum load case

Determine Total Structural Weight

Needed

New Guess

Weight Guess =

New Weight

Preliminary Design Iteration Processor (PDIP)

21

and Elite 1500 batteries, in which actual experimental data is used11. The second step requires plotting

the power available curve against an approximate power required curve derived from Anderson12. At the

velocity where the power available crosses the power required, a thrust can be estimated (Figure 11).

Assumptions made in the power required equation are, Cd0 of 0.04, Oswald efficiency factor (e) of .7, AR

of 5, and a wing area of 5 ft2. Thrust values computed with this curve match past historical data more

closely than by using a frontal area drag curve. Determining the final aircraft weight involves grouping the

propulsion systems from lightest to heaviest. For each group, an assembly of motors, batteries,

Electronic Speed Controller (ESC), receiver, fuse, and three micro servo weights are added together to

obtain a total electronic weight.

Figure 11 - Power Required - Power Available

The structural portion of PDIP requires inputs of the wing area and all wing dimensions. An

assumption is made that lift is continuous across the span. The wing tip test is the key aspect in the

amount of material volume that is needed to resolve the maximum bending moment and the shear

loads13. The bending and shear volumes are then added together to obtain a total spar volume. The rib

weight is the minimum number of ribs needed, multiplied by the volume of a rib at the MAC location. The

lightest weight covering material, Micro-Lite™, is factored into the weight for covering the total wing area.

Finally, the landing gear weight is input as seven ounces in order to ensure conservatism is built into the

model. The landing gear weight will decrease. However, the weight of the bulkheads and other items not

considered in this phase of the design will be encompassed by the conservatism.

4.4.3 Initial Weight Build Up After PDIP is completed, a structural and propulsion weight build up are tabulated (Table 2 and Table 3).

PDIP converges at an aircraft empty weight of 1.3 lbs with an uncertainty of 0.26 lbs. The uncertainty is

determined by analyzing a previous WSU DBF plane14 through PDIP. This shows that PDIP can possibly

under-predict the model by approximately 20%. Team Black determines that this tool provides a

reasonable estimate of where the aircraft weight should be.

0

20

40

60

80

100

120

140

160

180

200

0 10 20 30 40 50 60 70 80 90 100

Pow

er (

Wat

ts)

Velocity (ft/s)

Power Required

Power Available

22

Table 2 - Structural Weight Build Up

Table 3 - Propulsion System Weight Build Up

4.5 Aerodynamics 4.5.1 Airfoil Selection A unique trait of a flying wing concept is the ability to fly without the need of a tail. To accomplish this

task, there are several different approaches that can be taken. One approach involves using a cambered

airfoil with a large amount of wing sweep. Wing twist can also be applied to this approach thus enabling

the wing to stall at the root before the tip7. This will enable the pilot to reach stall while still having control

over the aircraft. A second approach involves using a reflexed cambered airfoil with a small amount of

sweep or taper7.

After the initial wing sizing parameters are identified and a value for each is obtained, airfoils are

chosen and screened to determine which will give the aircraft the best aerodynamic performance qualities

for the intended missions. Airfoil screening parameters are shown in Table 4.

Table 4 - Airfoil Screening Requirements

Half Span (in) 32.21 MTOW (lb) 5.80 Spar Weight (oz) 0.99Empty Weight (lb) 1.30 Max. Shear (lbs) 2.90 Rib Weight (oz) 0.44Payload Weight (lb) 4.50 Max. Moment (in-lb) 93.41 Micro-Lite™ Weight (oz) 0.70Root t/c (in) 1.00 Ashear Needed (in2) 0.24 Landing Gear Weight (oz) 7.00Wing Area (in2) 758.88 Abending Needed (in2) 0.11 Structural Weight (oz) 9.13

Heavy Weight (oz) Light Weight (oz)ESC 18 amp 0.60 ESC 12 Amp 0.5020 amp fuse 0.20 20 amp fuse 0.20Receiver 0.25 Receiver 0.25Receiver Battery 0.60 Receiver Battery 0.60Motor 1.45 Motor 0.70Wires 1.00 Wires 1.00Battery Elite 1500 6.64 Kan 700 7.07Prop 10x4 0.70 Prop 8X4 0.40Servos 3 Hitech HS-65-mg 1.30 Servos 3 Hitech HS-65-mg 1.30Total Electronic weight (oz) 12.74 Total Electronic weight (oz) 12.02lbs 0.80 lbs 0.75Avg Thrust (lb)** 1.60 Avg Thrust (lb)** 1.00Thrust per Prop Weight 2.01 Thrust per Prop Weight 1.33

Parameter Value Reasoning

Cl-max > 1.2 Must be above the indicated value to carry the maximum payload weight during the mission

Clo > 0.149 A cruise of 70 ft/s dictates that the lift coefficient be above the indicated value to cruise at 0° Angle of Attack

Cmo -0.12 < Cm < 0.12-A CG and an AC location are derived giving the aircraft a range for a pitching moment. -Reduced stabilization required in longitudinal trim

t/c ratio < 15% -Due to a large root chord, a smaller t/c ratio is desired. -Smaller t/c ratio contains less drag.

Cdo ~ 0.0 Smaller drag coefficients enable better aerodynamic performance

23

After an initial screening of 100 airfoils, 16 airfoils are chosen for further analysis. The airfoils that

are screened are either cambered, symmetric, or reflexed and designed by Selig15, Gottingen16, or

NACA17. These 16 airfoils are then narrowed down to the five airfoils shown in Table 5, of which two have

a “large” t/c ratio. The final airfoil is the Liebeck la2573a reflexed airfoil18.

Table 5 - Final Five Airfoils

The Liebeck la2573a reflexed airfoil is designed to fly at a low Reynold’s numbers (Re). This airfoil

gives the best aerodynamic performance when utilized at Re of 650,000 and below. However, the

disadvantage of this airfoil is that a laminar bubble is produced during flight. The laminar bubble creates

excess drag18.

Airfoil data, presented in Figure 12, indicates a transition at a Cl of roughly 0.2. This transition could

present some difficulty for designing the aircraft near the trim condition. However, upon multiple wind

tunnel tests, it is determined that the Re transition is not crossed.

Figure 12 - Liebeck la2573a Reflexed Airfoil Performance Data18

4.5.2 Aerodynamic Performance Predictions: Lift and Drag The transition from 2-D data to 3-D data is obtained using two different methods: the 88% method

from Raymer2 and the linear lift-curve slope method from Anderson12. The method from Raymer takes

each value of 2-D airfoil Cl data and multiplies the value by 88%. This method allows for the trends seen

in the 2-D airfoil data to be reproduced in the 3-D wing data. The method from Anderson uses the lift-

curve slope from the 2-D airfoil data, combined with e and AR, and incorporates it into the following

equation:

Name Cl-max Cl-cruise at α < 2° Cd at Stall Angle Cdo Cm at α < 2° t/cClark-Y(B) 1.3 0.4 0.0225 0.023 -0.08 11.70%Falcon 56 Mk II 1.25 0.25 0.03 0.0175 -0.03 13.68%SA7036 (B) 1.3 0.35 0.0435 0.017 -0.095 9.20%SD7037 E 1.28 0.25 0.04 0.016 -0.07 9.20%Liebeck la2573a 1.325 0.2 0.02 0.02 0.09 13.90%

24

𝑎𝑎 =𝑎𝑎𝑜𝑜

1 + 57.3𝑎𝑎𝑜𝑜𝜋𝜋𝜋𝜋𝜋𝜋𝜋𝜋

where “ao” is the lift-curve slope of the airfoil and “a” is the lift-curve slope of the wing12.

Figure 13 – Methods Used for 3-D Performance Predictions

Figure 13 represents the 2-D airfoil data compared to the two different methods of estimating 3-D

wing data. There is a clear difference between the two 3-D estimation methods. The 3-D estimation

method using Raymer2 follows the trends of the 2-D airfoil data better than the 3-D estimation method

from Anderson12.

Figure 14 represent the drag build up of the aircraft. The majority of the drag is produced at the wing.

Further optimization of the wing, fuselage, and landing gear are in the sections that follow. Figure 15

represents the 3-D drag estimates obtained from the Raymer and Anderson methods.

Figure 14 – Component Parasite Drag Buildup

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

-4 -2 0 2 4 6 8 10 12 14 16

CL

AoA (deg)

2D

3D Raymer

3D Anderson

Wing - 65%

Fuselage - 30%

Landing Gear -5%

Component CDo % of totalWing 0.0262 65%Fuselage 0.0121 30%Landing Gear 0.0019 5%Total 0.0402 100%

25

Figure 15 – Total Aircraft Drag Prediction at Re = 250,000

4.6 Flight Mechanics The chosen reflexed airfoil performance characteristics, along with Team Black’s design efforts to

keep the empty weight of the aircraft as low as possible, determines the sizing and placement of control

surfaces. Dynamic stability is assessed using Athena Vortex Lattice (AVL)19 (Figure 16). The aircraft is

controllable and compliant with stability requirements for all three missions.

Figure 16 - Athena Vortex Lattice Model

4.6.1 Control Surface Sizing, Placement and Authority Initially, two inboard elevons of 39% span and 25% chord are used to provide static longitudinal and

lateral control through channel mixing. Optimization leads to a wing iteration and empty weight re-

evaluation thus producing newly sized elevons of 19% span and 25% chord. The elevons are modeled as

flaps and analyzed using custom linearized tools by methods of Raymer2, Roskam9, and Etkin20. These

are confirmed using AVL and demonstrated through a full-scale 7ft x 10ft WTT. Though proving to be less

effective than predicted, demonstrated control authority is shown to handle the critical longitudinal trim

case, M3 T/O (heaviest payload), at -8.12° (Table 6) . All mission control surface deflections are within the

±20° allowed for flow attachment purposes9. The AOA required for trim flight for all three missions stay

below a stall angle of 15°.

26

`

Table 6 - Demonstrated Elevon Trim Authority 4.6.2 Longitudinal Stability

A minimum required SM of 5%7 for longitudinal stability is calculated accounting for inherent tail-

heavy pitching moment of the airfoil, CG offsets (Figure 17) from the thrust line, propeller wash, and

respective cruise and takeoff velocities for all three missions.

Figure 17 –M1 Moment Balance Illustration

Team Black’s design efforts ensure a final SM of 6.85% (Table 7) for the critical longitudinal trim

case by arranging internal aircraft components.

Table 7 - Mission Specific Static Margins

Trim plots are predicted using custom linearized tools, verif ied with AVL and

demonstrated with WTT results (Figure 18).

MissionCase Takeoff Cruise Takeoff Cruise Takeoff Cruiseδe (°) -1.28 3.48 -7.32 1.21 -8.12 -6.98α (°) 2.19 0.75 10.60 4.35 11.80 8.22

31 2

Mission 1 2 3AC (in) 6.05 6.05 6.05CG, X-direction (in) 4.89 4.79 5.21CG, Z-direction (in) 0.26 -0.44 -0.36Static Margin (%) 9.46 10.3 6.85

27

Figure 18 - Longitudinal Trim Plots for M3

4.6.3 Trim Drag A trim drag analysis using WTT results is performed to determine the increment in drag due to

trailing edge down critical drag case, M1 (α = 0.75°, δe = 3.48°). A total CD of 0.0019 (19 counts) is

added to the total aircraft drag buildup (402 counts). With the order of magnitude being minute, this is

neglected and thus, no further trim drag optimization is needed.

Figure 19 - Trim Drag Optimization

4.6.4 Lateral-Directional Stability

Based on the methods of Raymer, a vertical stabilizer is sized to compliment the inherent yaw

stability of the swept wing while yaw control is neglected2. Lateral-directional stability is assessed with

AVL for the critical lateral-directional trim case, M3 T/O. According to the methods of Roskam9, an 85th

percentile crosswind of 21 knots is introduced in the analysis to verify the aircraft’s lateral trim ability with

predominant Kansas crosswinds in the month of April5. The aircraft’s stability coefficients demonstrate the

ability to yaw into and roll away from the sideslip. A ±18.6° of aileron deflection Figure 20 is needed to

-0.1

-0.08

-0.06

-0.04

-0.02

0

0.02

0.04

0.06

0.08

0.1

0 2 4 6 8 10 12

CM,C

G

AoA (deg)

Predicted Cruise δe = -5.28Predicted Takeoff δe= -7.58Demonstrated Takeoff δe = -8.12Demonstrated Cruise δe = -6.98

28

compensate for the lack of rudder, canceling the yawing and rolling moments, while keeping the flow

attached9. Flight tests using X-Plane 921 accounts for prop wash effects. The aircraft is ultimately deemed

laterally and directionally stable, as well as controllable.

Figure 20 - Control Surface Crosswind Trim Authority for M3 T/O

4.6.5 Dynamic Stability

Mass and inertia properties are obtained using Catia22, and longitudinal and lateral-directional

dynamic mode approximations are made for the M1 cruise case (high speed, forward-most CG) based on

the methods of Roskam9, with AVL adding confirmation. It is concluded that the aircraft is dynamically

stable with relatively low undamped natural frequencies for the phugoid, dutch roll, and highly damped roll

modes, though a high undamped natural frequency of 5.49 rad/s for the short period mode is displayed

(Figure 21). This is attributed to the aircraft’s low moments of inertia. The vertical stabilizer aids in the

spiral mode stability.

Figure 21 - Root Locus for M1 Cruise19

0

5

10

15

20

0 2 4 6 8 10 12 14 16 18 20

Aile

ron

Def

elct

ion

()

Crosswind (Knots)

29

4.6.6 Sustained Turn Dynamics

Maximum allowable structural G-loads are provided by the structures lead. The turn radius, bank

angles, and turn time are calculated accordingly. The bank angles are iterated until the maximum

allowable and actual G-loads are approximately within 0.1 G’s of difference (Table 8) in order to make

mission turns at the fastest times.

Table 8 - Sustained-Turn Dynamics Characteristics

4.6.7 Servo Selection

Hinge moments are estimated using chase-around charts in Etkin20. An arbitrary factor of safety (FS)

of 2 is applied to the estimated hinge moments and subsequently, the lowest-weighing servo that

provides approximately 22 oz-in is selected. This avoids significant weight penalties in the scoring

equation and provides more than necessary torque.

4.6.8 Stability Derivatives

Table 9 summarizes the stability derivatives for the aircraft:

Table 9 - Stability Derivatives for M3 T/O with a 21-knot Crosswind

4.7 Propulsion Preliminary design of the propulsion system consists of trade studies that are expanded on results

from conceptual design. This contains a motor, battery, wire, ESC, and propeller selection. The selection

process for all propulsion components is tailored around minimizing weight.

4.7.1 Battery Selection Meeting the required power with the lightest weight battery pack, while still being able to meet the

M3 endurance, at full power, is the task for battery selection. During conceptual design it is concluded

that the Elite 1500s and 2000s are exceptionally light for the amount of energy they contain23. A trade

study is conducted to see if the Elite 1500s are capable of drawing enough amperage while maintaining a

high enough voltage to achieve the required power. This would allow for a replacement of the Elite 2000s,

which will lower the battery weight by two ounces. However, during testing, the Elite 2000s drastically

underperformed their predicted current draw. They did not achieve the predicted 24 amps while

maintaining the required voltage thus rendering the trade study inconclusive. Both batteries types are

selected for further testing.

Mission 1 2 3G-Load Allowed 7.5 2.5 2.35G-Load Actual 7.4 2.4 2.15Bank Angle (°) 78 65 63Turn Radius (ft) 2.6 61 28.1180° Turn Time (seconds) 0.15 3.2 2.1360° Turn Time (seconds) 0.23 6.7 4.2

CLα 3.325 CMα -0.326 Cyβ -0.042 Cyδa 0 Cyp -0.035 Cyr 0.021CLδe 0.011 CMδe -0.088 Clβ -0.112 Clδa 0.181 Clp -0.292 Clr 0.076CLq 5.379 CMq -2.029 Cnβ 0.048 Cnδa -0.002 Cnp 0.032 Cnr -0.015

30

4.7.2 Motor Selection

It is apparent in conceptual design that motor selection will revolve around a motor that can handle

constant amperage of 18-25 and a power of 200 watts. This is determined from the W/S, and is later

refined during iterations of conceptual design. Once a required power for the motor is known, a list of

motors, in that range, can then be compared. The first parameter compared is the weight due to the

contribution it plays in the scoring. The lightest motor for the given maximum amperage is assigned high

priority. Secondly, the motors’ RPM/volts (KV) are taken into consideration. Any KV rating greater than

1400 and fewer than 900 is undesirable due to unrealistic RPM performance for a 10 volt pack. After

tabulating many motors it is apparent that the Rimfire motors are constructed remarkably lighter than

other mainstream motors24, 25. A comparison of motors is shown in Table 10. The Rimfire .10 is chosen

due to its high max amperage per low weight and high reliability. Reliability is proven by its heavy use

within the WSU small aircraft prototype lab.

Table 10 – Motor Comparison

4.7.3 ESC Selection

A list of ESC’s that can handle 25 amps are compared and selected based on weight and

programmability24. The ability to program the cutoff voltage will ensure complete battery performance for

M3. The Phoenix 2526 is selected because it excelled in both areas. A comparison is shown in Table 11.

Table 11 – ESC Comparison

4.7.4 Wire Selection

Wire weight from the battery pack is a contributor to the total system weight. Historically, a heavy

gauge is required for the amps produced by the battery. A trade study is performed to determine the

lightest wire that can be used without a substantial power drop. Data consisting of resistance, diameter,

and weight of wire gages are tabulated ranging from 10 to 22 gauge wire27. Using Ohm’s law and the

equation of resistivity, power drop per wire weight can then be plotted (Figure 22). A gauge of 14 could

be used without substantial power drop while allowing for a weight reduction of 0.16 ounces per 1ft of

wire from the standard 12 gauge wire.

Motor Weight (oz) KV Max Current (amps)Rimfire 400 1.9 950 20Rimfire 10 2.5 1250 35Turnigy 2217 2.5 1050 18AXI 2808/20 2.6 1490 22

ESC Weight (oz) Amps ProgrammablePhoenix 0.6 25 YesFlight Power 0.78 25 NoGreat Planes 0.95 25 No

31

Figure 22 - Wire Weight per Power Drop

4.7.5 Propeller Selection

Propeller selection is determined by using a combined blade element and momentum program. A

java application called Java Prop28 is utilized to perform multi-analysis on a given propeller geometry and

RPM. Propeller geometry is determined by calculating the pitch and measuring the chord for every blade

section. The analysis follows the process outlined in Figure 23.

Figure 23 - Propulsion Flow Chart

The propeller pitch, diameter, and RPM are imported and analyzed through Java Prop28. If the

propeller stays within the motor-produced power of 185 watts, while keeping an RPM above 7000, then

data is plotted for consideration. Motor-produced watts and minimum propeller RPM are derived from the

KV of the Rimfire .10 motor. The battery pack selected is 10 Elite 2000 cells due to a previous analysis of

required power. After analysis, selected propellers are plotted against each other for efficiency and the

0

1

2

3

4

5

6

7

8

9

10

0 0.1 0.2 0.3 0.4 0.5 0.6

Pow

er D

rop

(wat

ts)

Wire Weight (oz)

32

power required (Figure 24). Verification for the power and torque data is shown by the use of equations

provided by Zamora for the corresponding coefficients29. Total propeller performance is verified by

comparing the results with actual experimental data provided by Merchant and Miller11. After verification

the propellers selected are the APC Electric 10X7 and 11X5.5.

Figure 24 - Power Required

4.8 Structures

Preliminary structural layout is determined utilizing the least amount of structure. Team Black

researched the structural members that were, historically, the heaviest parts of past DBF planes. The

emphasis is placed on optimization of these structural members. Further discussions of these members

are in the proceeding sections. Critical load paths are identified on the model and used for further

analysis (Figure 25). The ultimate goal is to keep the structure simple to manufacture while maintaining a

FS of approximately 1.2.

Figure 25 - Load Path Diagram

4.8.1 Critical Loads

Structural analysis is performed on the wing using methods presented by Allen and Haisler30. Limit

loads are found using the wing-tip test and then translating these loads into equivalent flight loads. A

modified version of Shrenk’s Approximation2 is combined with a constant rectangular distributed load to

obtain the lift curve that is used to size the wing structure. Ultimate flight loads are then scaled up from

the limit loads using a FS of 1.5, which is standard practice in the aerospace industry2. The maximum

0

20

40

60

80

100

120

140

160

180

200

0 10 20 30 40 50 60 70 80 90

Pow

er (

Wat

ts)

Velocity (fps)

Power Required

Battery Output

10X7

11X7 EF

10X6 EF

11X6 Experimental

33

bending stress, found at the root of the wing, is 1,767 pounds per square inch (psi) and the maximum

shear is 5.38 psi.

4.8.2 Materials

Identifying the best materials that can be used is another key step to ensure this aircraft is as light as

possible. Referring back to the scoring analysis, specific strength of a material, also known as strength

per unit density, is the best way to ensure that the weight of the aircraft is as low as possible. Ease of

manufacturing a specific material is another aspect that is considered for the selection. The aircraft will

be built to match the drawings and, therefore, the analysis. From these screening criteria, pink foam and

balsa wood are selected to be used for primary and secondary structure. Balsa works well in tension and

compression; however, it can become too strong for the stresses that are endured in the aircraft. This

leads to unrealistically thin parts that have buckling problems therefore pink foam is utilized. Pink foam

does not have the strength capabilities that balsa contains; however, it is approximately nine times lighter

than balsa. The structural properties of each material are passed down from past WSU DBF teams.

Displayed below, Figure 26, is a breakdown of each aircraft component, and displays the percent of the

total weight of the aircraft it contains.

Figure 26 - Aircraft Component Weight Breakdown

4.8.3 Spar Trade Study The front spar is the primary load bearing structure in the aircraft. The airfoil chosen for the aircraft

is 13.9% thick which translates to 2.25 in. of useable spar height at the root. The maximum height of the

spar allows the bending rigidity of the spar to increase without the addition of large amounts of structural

weight. A simplified analysis using the Euler-Bernoulli beam theory31 and Castigliano’s second theorem30

calculates beam stresses and tip deflections which are utilized to evaluate several different spar shapes.

After the spar shapes are screened, different combinations of materials are applied to each of the spar

Motor 9.9%

ESC 2.4%

Batteries 38.1%

Receiver 1.4%Propeller 2.8%

Wires + Servos 17.9%

Spars 3.5%

Landing Gear 7.2%

Wings 12.4%

Fuselage 3.2%Vert. Tail 0.1% Misc. 1.0%

Propulsion

Structure

Control System

34

caps and shear webs to determine the lightest possible structure while still meeting the needed

requirements.

Table 12 - Spar Shape Trade Study

4.8.4 Spar Carry Through Structure

The carry through structure has to be compact in order for the payload to be properly spaced and

minimize weight penalties. The carry through structure is located specifically to tie into the spar and

transfer the wing loads through the fuselage. Two ¼ in. x ¼ in. balsa rods are placed on the inside of

each spar. These are the resolving points for the landing gear loads as well as the main handling point for

the fuselage box to attach. The use of two structural supports to resolve different loads is another

example of how structure in this aircraft must perform multiple purposes in order to score the highest in

competition1.

4.8.5 Motor Mount Optimization After preliminary sizing of the motor mount with hand calculations, a finite element model (FEM) is

constructed using the Abacus plug-in for Catia22. The results illustrate that the balsa rods can be made a

1/16 in smaller in diameter, reducing the structural weight of the mount.

Figure 27 - Motor Mount FEM Fringe Plot

4.9 Landing Gear The scoring analysis displays the RAC as a critical factor in this year’s competition1. A historical

analysis is performed on previous DBF winning aircraft landing gear weights, resulting in an average

Type Restriction I-Beam Box Beam Sandwich BeamBending FS > 1.00 2.01 1.73 1.53Shear FS > 1.00 1.44 1.08 1.08Worst Case Flight Deflection (in) < 5% * Span 1.26 1.30 1.47Wing Tip Test Deflection (in) < 5% * Span 1.49 1.68 2.00Total Spar Weight (oz) Lightest 1.40 1.58 1.86

35

weight of a ½ lb in landing gear1. The result of this analysis determines that a new method of designing

landing gear is needed in order to improve the RAC.

4.9.1 Dimensions

Taildragger landing gear has a tendency to ground-loop during taxiing if not set up properly. Thus,

the dimensions during design and manufacturing must be accurate. Raymer’s method determines landing

gear placement for a tail-down angle of 10°, allowing sufficient ground handling stability near stall AOAs2.

The CG of the aircraft is between 16° and 25° behind the vertical main landing gear (MLG) dimension to

keep the aircraft from ground looping or going nose over during T/O and landing2. The lateral gear

spacing is determined by creating a 25° angle off of the CG to prevent the aircraft from overturning2. A

computer aided design (CAD) part is created in order to shift the CG forward, and re-size the landing gear

efficiently during design changes.

4.9.2 Attachment A breakaway plate attaches the MLG to the fuselage, which transfers the load to the main spar

(Figure 28). Hard landings will break the gear away from the flight vehicle, but will leave the aircraft

structure intact.

Figure 28 - Landing Gear Breakaway Plate Design

4.9.3 Material Due to the RAC impact on the total score, light weight landing gear is needed. Steel wire landing

gear is designed and built using Logan’s FEM analysis, to deflect and absorb 24% of a 4G landing load

for Mission 232. The landing gear is not designed to land with the max payload of M3, given that it is

dropped during flight. This is a necessary risk in order to improve the RAC score.

4.10 Preliminary Mission Performance Predictions

Performance estimations are shown in Table 12 for each of the three mission profiles. The critical

case for the majority of parameters is M3, with a T/O weight of six lbs. These performance predictions are

validated through flight testing.

36

Table 13 – Preliminary Mission Performance

5.0 Detail Design

After preliminary design is frozen (with wing sizing parameters as listed in Table 14), optimization

begins. Solutions are developed for the payload release mechanism, the tail fairing assembly, and overall

weight reduction. Weight and CG build-up calculations are improved based on manufacturing experience,

while the RAC is constantly assessed to ensure competitive design requirements are met.

Table 14 - Final Aircraft Dimensional Parameters

5.1 Wing Iteration

Within the wing iteration, certain changes are made in the dimensions. The root chord is kept at the

current length of 16.5 inches. Making the root chord smaller is a possibility for Team Black; however,

doing so will force the fuselage to shift down exposing more of the front face. This will inevitably add more

drag to the aircraft. The major changes to the wing dimensions are within the span and the taper ratio.

The wing area is decreased from 766.08 in2 to 739.20 in2 and the taper ratio is changed from 0.4 to 0.6. A

combination of these two changes allows Team Black to derive a wing span of 55.125 in. The

disadvantage to the change in taper ratio is a resulting increase in induced drag. This increase is within

the tolerance set earlier in preliminary design.

5.2 Hoerner Tips

A main component of an aircraft’s total drag is known as induced drag or drag-due-lift. This type of

drag is produced when high pressure on the bottom of the wing travels towards the wingtip where it

combines with low pressure on the top of the wing to create a vortex. According to Hoerner33, as the tip

CLMaxOswald Efficiency FactorCD0(L/D)MaxTakeoff Weight (lbs) 1.60 5.35 6.06W/S (lb/ft2) 0.31 1.04 1.21T/W 1.25 0.37 0.32Takeoff Distance (ft) 5 75 90Cruise Speed (ft/s) 75 73 69Stall Speed (ft/s) 32 32 32Total Flight Time (sec) 240 70 75Capacity Required (mAh)

0.04

1450

9.60

Performance Parameter Mission 1 (Ferry Flight)

Mission 2 (Passengers)

Mission 3 (Water Climb)

1.100.85

37

vortex forms on the lateral edge of the wing, the effective aspect ratio actually becomes smaller than the

geometric aspect ratio. To counteract this problem, Team Black is looking into employing Hoerner tips on

the aircraft. Hoerner tips will essentially help reduce the size of this vortex. However, due to the design of

the wing structure, a Hoerner tip will not allow for any reduction of weight or increase of AR. Constructing

the Hoerner tip will extend the manufacturing process as well.

5.3 Fuselage Aerodynamics 5.3.1 Fuselage Nosecone

The fuselage originally had a flat plate on the front face. A flat plate in flow, according to Hoerner33,

has a parasitic drag coefficient of 2.05. This number is based upon the aircraft’s reference area. There is

also a decrease in lift over the midsection of the wing due to the “knife-edge corners” of this flat face. This

“knife-edge” does not allow the flow to stay attached, thus causing turbulence along the bottom side of

the fuselage and along the root of the wing. A nose cone, in the shape of a two-dimensional half-circle,

has the potential to reduce the parasitic drag coefficient by approximately 50%. This nosecone will not

only reduce the parasitic drag but reduce flow separation as well, thus promoting lift along the bottom of

the fuselage and root of the wing. However, while adding the nose cone bolsters aerodynamic

performance, the motor mount must be shifted to accommodate the new hardware.

Figure 29 – Nose Cone Side View

5.3.2 Tail Fairing Assembly

After the preliminary design WTT, a tail fairing is needed to reduce base drag33. The tail fairing is

added, but with minimal structure to avoid weight penalties that are not included in the initial weight build-

up. The fairing is also needed as a platform to attach the steerable rear landing gear, the vertical tail, and

the water drop mechanism. Emphasis is placed on designing a structure that can hold the items listed

above and support any loads that are produced. The fairing must follow the 12° upsweep as stated in

Raymer2; however, significant loads will not be resolved through it. Foam serves as the perfect material

for this application with the addition of balsa cross members that will handle the impact loads of the rear

landing gear as well as hold the tail wheel servo in place. Two stiffeners are added in order to prevent

the foam from flexing from side to side during a turn. The entire assembly is covered in Micro-Lite™ in

order to provide the lightweight skin for aerodynamics.

38

5.3.3 Wing Fillets

Another contributor to the total aerodynamic performance is interference drag. This is the drag that is

produced between the wing and the fuselage33. However, upon further inspection of Hoerner, the

decrease in interference drag is not large enough to counteract the additional weight that these fillets will

add to the aircraft. This year’s scoring is based highly on aircraft weight, thus removing weight at the cost

of having a small amount of extra drag serves as a benefit.

5.4 Water Tight Fuselage

The mission requires the water to be carried within the mold lines of the aircraft as well as the use of

a non- pressurized vessel. In order to carry the payload during M3, the fuselage will be constructed to

hold water. Four full-scale concepts are built and tested for water tightness and weight. The concepts

include laminated paper that is hot glued at the seams, a combination of foam and epoxy, balsa sheets

with Micro-Lite™, and a balsa truss structure with Micro-Lite™. The laminated paper does not effectively

store water without leaking. However, the rest of the concepts hold water successfully. After weighing

each water tight fuselage, it is determined that the foam and epoxy design is the lightest. During the

prototype building process, lightening holes are added to the sides of the foam fuselage in order to lose

weight. The sides of the fuselage are then covered in Micro-Lite™ in order to keep the water tightness

and allow for a smoother surface.

5.5 Water Release Mechanism

A water tight door design is needed to carry water to 100 m and then be released for time to stop. A

trade study is conducted between the door size and the weight gained in order to seal a larger area. The

size of the door directly affects the M3 flight score as well as the RAC. During conceptual and preliminary

design, the fuselage contains a sealed door that swings open thus dropping all of the water

instantaneously. A rubber o-ring provides the seal around the door. A prototype test of this concept

determined that the seal is ineffective, due to the porous and flexible qualities of the foam. The o-ring is

unable to make a proper seal.

A simple cork attached to a servo is designed to solve this issue. This allows the manufacturing of

the fuselage to be more efficient. The servo will attach to the pre-existing structure of the tail fairing. This

also allows for the landing gear to be shortened from seven inches to two inches, thus saving

approximately an ounce of weight. The time to drop the water increases from an instant drop to five

seconds to completely release all of the water. It is determined that this will not cause any instability

during flight.

A concern of water visibility arose due to the smaller release area. A group of underclassmen

designed a basic water-drop system and released water from a personal radio controlled (R/C) aircraft

(Figure 30). This test is designed to determine how small of a release area can be utilized to effectively

see the water when released at 100 m. The result of this test verifies that the water is capable of being

seen when released from a four square inch hole at the 100 m mark. This test proves that a smaller

release area can be used as effectively as a larger area.

39

Figure 30 - Water Visibility Test Drop System

The cork fits in a 1.77 in2 area hole and is easily extruded by the servo once the CAM releases it.

Ground testing is used to ensure water tightness around the cork and release effectively multiple times

(Figure 31).

Figure 31 - Ground Test Photo of Water Release Mechanism

5.6 Flight Performance Summary Table 15 details flight and mission performance parameters. Applicable final predictions are

compared to actual test results in Section 8.

Table 15 - Flight Performance Parameters

5.7 Drawing Package

The drawing package includes: 3-view (dimensioned), Structural Arrangement, Aircraft Systems Layout,

and Mission Configuration.

A A

B B

4

4

3

3

2

2

1

1

DRAWN BY

J. WinkelCHECKED BY

T. BlumenshineAPPROVED

WSU

DATE

DATE

DATE

2/6/2012

2/7/2012

2/8/2012

WICHITA STATE UNIVERSITYShockin Stingray

TITLE

Aircraft Overview

SIZE DWG NO

2012-001REV

XSCALE 1:6B

Report Page SHEET 1 OF 4

DIMENSIONS ARE IN INCHES

40 of 59

6.5

6.1

Front view

56.1

3.1

13.2

26.3

3.9

26.29.9

15

16.5

Top view

4.5

1.0

2.0

.8

18.7

12 3.5

Right view

10.0

25% Chord Elevons

Engineering Bill of MaterialsItem Part Number Qty Material1 Half-Wing Cutout 2 Foam2 Main Spar Cap 2 Balsa 3 Rear Spar Cap 2 Balsa4 Foam Nose Cone 1 Foam5 Motor Mount 1 Balsa/Spruce6 Main Landing Gear

Support Rod2 Balsa

7 Wing Box CarryThrough

2 Balsa

8 Main Rib 2 Balsa9 Rear Landing Gear

Support Rod2 Balsa

10 Tail FairingStiffener

2 Balsa

11 Rear Landing GearAlignment Panel

1 Balsa

12 Tail Fairing 1 Foam13 Fuselage 1 Foam

3

1

2

8

10

13

9

11

7

4

5

6

12

A A

B B

4

4

3

3

2

2

1

1

DRAWN BY

J. WinkelCHECKED BY

T. BlumenshineAPPROVED

WSU

DATE

DATE

DATE

2/6/2012

2/7/2012

2/8/2012

WICHITA STATE UNIVERSITYShockin Stingray

TITLE

SIZE DWG NO

2012-002REV

XSCALE XXB

Report Page 41 of 59 SHEET 2 OF 4

Structural Arrangement

Aircraft Components List

Item Part Type QTY

1 Propeller APC 10 x 7 E 1

2 Elevon Servo Hitech HS-65MG 2

3 Tail Wheel Servo Futaba S3114 1

4 Motor Rimfire 0.10 1250 kV 1

5 Battery Pack 5-Cells Elite 1500 2

6 Reciever Battery 4-Cells KAN 180 1

7 Receiver Spectrum 6 Channel 1

8 Speed Controller Phoenix 25Amp 1

9 Fuse ATO 20 Amp Slow-Blow 1

10 Release Servo Hitech HS-65MG 1

11 Altitude Sensor CAM Altitude Sensor 1

12 Front Wheel 2" Dubro Super Lite 2

13 Rear Wheel 3/4" Dubro Super Lite 1

A A

B B

4

4

3

3

2

2

1

1

DRAWN BY

J. WinkelCHECKED BY

T. BlumenshineAPPROVED

WSU

DATE

DATE

DATE

2/6/2012

2/7/2012

2/8/2012

WICHITA STATE UNIVERSITYShockin Stingray

TITLE

Aircraft System Layout

SIZE DWG NO

2012-003REV

XSCALE 1:7B

Report Page SHEET 3 OF 442 of 59

Front view

3

4

10

Isometric view12

2

A

Detail AScale: 1:2

1

5

6

7

8

9

11

Landing GearProjection View

13

8 - 1" x 1" x 5" Aluminum Bars

1/4" Foam Cutouts

Mission 2 Payload ConfigurationAluminum Bars

Mission 3 Payload Configuration2 Liters of Water

1" Vertical Spacing1/2" Horizontal Spacing

Release Servo

1" Cork

1/16" Steel Wire

Water Release SystemDetail View

NOTE: All views have hidden components for clarity

2 L Water Tank

A A

B B

4

4

3

3

2

2

1

1

DRAWN BY

J. WinkelCHECKED BY

T. BlumenshineAPPROVED

WSU

DATE

DATE

DATE

2/6/2012

2/7/2012

2/8/2012

WICHITA STATE UNIVERSITYShock'n Stingray

TITLE

Mission Configuration

SIZE DWG NO

2012-004REV

XSCALE XXB

Report Page SHEET 4 OF 443 of 59

44

5.8 Aircraft Component Weight and CG Buildup

Accurate weight and CG locations are presented for each mission (Table 16). Consistent flight

handling characteristics result from minimal change in SM between missions. Batteries and payloads are

relocated between missions to keep the same SM. Particular attention is paid to M3, where constant CG

location is desired after the water is released.

Table 16 – Component CG and Weight Buildup

Table 17 – Mission Specific CG and Weight Estimates

6.0 Manufacturing Plan and Processes

Practice refines the manufacturing approach throughout the course of the design process. The

manufacturing team includes the five senior engineers and underclassmen.

6.1 Manufacturing Methods

The RAC is a key factor for the type of material used to construct the aircraft. Trade studies show

that a foam leading edge, foam spar webs, balsa spar caps, and a foam fuselage are lighter than the

traditional balsa and MonoKote™ model. The majority of the aircraft is constructed from foam. This allows

the design to be cut into four main parts using the hot wire computer numerical controller (CNC) foam

45

cutter. The four parts include two half-wings, the front and sides of the fuselage, and the bottom of the

fuselage with the tail fairing. This reduces glue joints and adhesive weight.

6.2 Prototypes

Five prototypes are constructed prior to the initial mission ready model. A half-scale, solid foam

prototype is fabricated using the CNC foam cutter. Initial efforts demonstrate that the taper results in extra

burning on the wing tip causing the tip chord and thickness to be inaccurate. A code is created to take

into account this inaccuracy to produce proper sized components.

The second prototype is a full scale, solid foam wing aircraft for flight testing. This construction

validated the code for the wing cut outs. Due to the structural test not being completed at the time of this

construction, extra structure is added to the joint between the wings. The fuselage is constructed similar

to the final mission ready model to accommodate for water drop mechanism testing. This fuselage

construction also validates the two part cut out technique.

The third and fourth prototypes are full scale, structural models of the wings. The first wing is used in

the wing tip deflection test and the second wing is used in the whiffletree test. The fuselage is not

constructed for these models because the fuselage is designed not to contain any structural elements.

The final prototype is a full scale, solid foam wing, constructed for a full-scale WTT. In order to

reduce the interference from the mounting system, the two-point mount is fixed to the aircraft inside the

fuselage. To structurally support the mount, the fuselage and tail fairing are created from a 4 in x 6 in

solid wood post in place of the foam model. With this final prototype construction, the mission ready

model, including all the tooling, is finalized and tested.

6.3 Mission Ready Model

Each wing is cut out of two sheets of solid foam that are glued together with 3M77 spray adhesive to

attain the max thickness of the airfoil. The areas between the foam ribs are cut out and the leading edge

is hollowed out of each wing. Each wing has a groove inscribed where the spar caps are placed allowing

for each wing to line up along the spar. The sides of the fuselage and front foam bulkhead are created

using the foam cutter followed by the bottom of the fuselage with the tail fairing. Spar caps as well as two

balsa ribs are cut out with the laser cutter. CAD drawings are created for the laser cutter to read and cut

out 1/8 in thick sheets. The spar caps are laser cut out given that the fuselage is straight and the wings

are swept. This prevents the aircraft from having a glue joint at the location of the maximum bending

moment.

The MLG is constructed once, and is used on all prototypes due to the breakaway plate design.

Steel wire is first cut to the correct length for each separate MLG. A Piranha hydraulic press brake is used

to bend and flatten the wire achieving the correct angles (Figure 32). In order to attain the correct

dimensions, a full-scale CAD drawing is printed and the fabricated landing gear utilizes the sheet as

reference tooling.

46

Figure 32 - The Piranha Bending the MLG

Two 1/8 in balsa sheets are fabricated with the laser cutter in order to construct the breakaway plate.

Each sheet is then routed where the landing gear will be placed. The landing gear rods can then be

sandwiched between the plates to prevent twisting. Epoxy is used to adhere both sheets of balsa together

with the landing gear wire placed in between (Figure 33).

Figure 33 - Main Landing Gear Construction

6.3.1 Tooling

Early build experiences reveal that the classic “ruler and square” layout methods are inaccurate and

lengthy. A rapid, precise, and repeatable manufacturing method is sought. Tooling is a viable and simple

solution to this manufacturing problem. When the wings are built from the solid pieces of foam, they have

a starting point of 1.75 in from the bottom and two inches into the block enabling the remaining foam,

without the wing, to be used as the tooling device. This tooling does not only line up the wings, but also

provides support on the thin trailing edge when spar caps are glued. This prevents the foam from being

crushed. A third foam block is cut to the width of the fuselage, enabling the block to sit between the two

wing tools. This sets the correct spacing between the wings.

During detailed design, lightening holes are analyzed in order to lose weight. In order to accurately

cut out the lightening holes with a router or razor blade, tooling is required. The tooling is laser cut out of

1/8 in plywood to the proper dimensions. The tooling is then placed on the aircraft to be utilized as a

stencil during the cutting process. This same principle is then used when cutting out the passenger

constraints.

47

7.0 Testing Plan

Preliminary design testing focuses on validation of the airfoil, AC location, and propulsion system

predictions. Detail design testing begins with a full scale, flight ready prototype to test ground and flight

handling qualities.

7.1 Half-Scale Wind Tunnel Testing

Two half-scale wind tunnel tests are conducted in WSU’s 3ftx4ft Low Speed Wind Tunnel.

Figure 34 - Half-Scale Model mounted in WSU’s 3ftx4ft Wind Tunnel

These tests are conducted to validate aerodynamic and stability calculations for the Liebeck

la2573a. The test plan matrix is shown in Table 18.

Table 18 - Half-Scale Wind Tunnel Test Matrix

Run # TareModel

Configuration upsweep = 20 deg

q = takeoff Alpha Sweep = -4° to 12°

Δα = 2°, then 12° to 20° by Δα = 1°

q = cruise Alpha Sweep = -4° to 12°

Δα = 2°, then 12° to 20° by Δα = 1°

q = 0 Alpha Sweep = -4° to 20°

Δα = 2°1 Dynamic No Model x x2 Static No deflections x3 No deflections x x4 Repeat Run 3 x x5 Static Deflection = -5° x6 Deflection = -5° x x7 Repeat Run 6 x x8 Static Deflection = +5° x9 Deflection = +5° x x10 Repeat Run 9 x x11 Static Deflection = -10° x12 Deflection = -10° x x13 Repeat Run 12 x x14 Static Deflection = +10° x15 Deflection = +10° x x16 Repeat Run 15 x x

Run # TareModel

Configuration upsweep = 12

q = 1.717 Alpha Sweep = -4° to 12°

Δα = 2°, then 12° to 20° by Δα = 1°

q = 5.6 Alpha Sweep = -4° to 20°

Δα = 2°

q = 0 Alpha Sweep = -4° to 20°

Δα = 2°1 Static No deflections x2 No deflections x x3 Repeat Run 2 x x4 Static Deflection = +15° x5 Deflection = +15° x6 Repeat Run 5 x7 Static Deflection = +20° x8 Deflection = +20° x9 Repeat Run 8 x

Primary Test Matrix---1/2 scale model

Install foam fuse extension pieceSecondary Test Matrix----1/2 Scale Model

48

7.2 Propulsion System Testing

An all inclusive propulsion system test is necessary to validate predictions from the preliminary

selection as well as address various uncertainties. Motor and battery performance, in addition to

reliability and endurance of the total system, are evaluated at this point of the design phase.

7.2.1 Battery Testing

Batteries are thoroughly assessed though a computerized battery analyzer (CBA) (Figure 35). This

gives voltage values per given discharged amperage. Results are plotted against time. Reliability and

performance degradation are assessed by analyzing cold soaked and elevated temperature batteries

along with deep cycled batteries. Reliability proves essential because lower weight cells tend to be

sensitive to environmental factors, such as excessive temperatures. Figure 36 shows an average of 25%

decrease in performance of cold batteries.

Figure 35 - Computerized Battery Analyzer System

Figure 36 - Cold Soak Test Results

8

8.5

9

9.5

10

10.5

11

11.5

12

0 50 100 150 200

Volts

Time (S)

Cold Soak Pack

Room Temperature Pack

49

7.2.2 Total System Testing

The entire propulsion system is tested in the Wichita State 3ftX4 ft windtunnel. This test allows for

the verification of the propeller and battery selection. Data is gathered by the internal windtunnel balance

as well as a battery data logger. Setup information is shown in Table 19. The basis of the test is to

perform two runs with the power supply at the predicted battery wattage to verify the 10X7 and 11X5.5

propellers. After these runs, the better performing propeller is selected for further evaluation with the

batteries. The batteries are selected based on their weight and predicted performance. Table 20 shows

the test matrix.

Table 19 - Setup Information

Table 20 - Propulsion Test Plan Matrix

Figure 37 - Propulsion Test Apparatus

Motor Rimfire 10 1250KVESC Phoenix 45Batteries (2 5-Cell Packs in Series) Elite 1500 4/5 A or Elite 2000 AAData Logger Eagle Tree V3Prop APC EF 10X7 , 11X5.5

Run # Propeller Battery Type Motor Run Time (min)1 10X7 Power Supply 10V 20amp Exploratory Run, q sweep 22 10X7 Power Supply 10V 20amp q Sweep 23 11X6 Power Supply 10V 20amp q Sweep 24 10X7 10 Cells Elite 2000s Static, Cruise q Endurance 45 10X7 10 Cells Elite 1500s Static, Cruise q Endurance 4

50

7.2.3 Results

Data from this test is conclusive. The 11X5.5 propeller proves inadaquate for the wattage required

for the motor to operate efficiently. Furthermore, the Elite 1500s pack performed as predicted, matching

the output power of the Elite 2000s while being two ounces lighter for the entire pack. The Elite 2000s

pack, however, underperformed the predicted output. Data is compared to Java prop28 for the given

batteries predicted power and similarly matched predictions for the Elite 1500s output. Results are shown

in Figure 38, Figure 39, and Figure 40.

Figure 38 - Actual Recorded Static Power

Figure 39 - Actual Recorded Thrust Compared to Predictions

51

Figure 40 - Actual Recorded Cruise Power

7.3 Structural Testing

Structural testing is needed to validate the theoretical analysis and locate areas where additional

structure can be removed to lighten the aircraft. Three structural test articles are built in order to test and

simulate different load cases.

7.3.1 Final Spar Validation

The front spar in the aircraft is the primary support for all flight loads. A whiffletree test is setup in

order to apply a close approximation of the calculated lift distribution loads to ensure that the theoretical

analysis is comparable to actual results. Deflections are measured with a laser extensometer to ensure

high accuracy when measuring the small deflection changes. The predicted results are 25% higher than

the actual results. Furthermore, the test proves that analyzing the foam as an isentropic material is an

over-simplification.

Figure 41 – Spar Whiffletree Setup

52

7.3.2 Wing Tip Test

A wing tip test is performed to validate the primary structural configuration when loaded in the critical

load case. The model is a full scale model without Mircro-Lite™. The wing tips are supported by ropes

hanging down and sand is applied to the center payload section in order to simulate the correct weight

placement. Six pounds is added to simulate the maximum T/O weight of the aircraft (Figure 42). The

configuration of the aircraft causes the CG not to fall in line with the wing tip. Therefore, a third support

must be added underneath the wing apex in order to balance the model. This invalidates the boundary

conditions of a true wing tip test. A more theoretically correct way of testing the aircraft for stress and

deflection validation is sought.

Figure 42 - Wing Tip Test

7.3.3 Final Wing Structural Validation

In this test a full scale half wing is constructed. This test utilizes the same whiffletree test setup that

the spar validation was performed with (Figure 43). Two laser extensometers are setup on both the front

and rear spar in order to validate the torsion resistance that is predicted for the wing. The wing held an

eight percent greater load than had been predicted and the maximum tip deflection is 16% greater than

predicted. A correction factor that was obtained from the spar testing still does not cover the deflection

difference. Looking further into the results, the divergence from the theoretical deflection does not occur

until the loads approach failure loads. This can be expected for a non-elastic material such as foam. The

wing is predicted to rotate seven degrees under the maximum lift distribution while the actual rotation is

two degrees. From this test, several design changes are made to improve the RAC score. The first

improvement involves resizing the front spar from a thickness of 0.375 in to 0.300 in. The rear spar is

reduced in size from 0.375 in to 0.250 in due to the torsion resistance being higher than predicted. This

allows the total aircraft weight to drop from 1.66 lbs to 1.60 lbs.

53

Figure 43 - Final Wing Whiffletree Test

7.4 Full Scale Wind Tunnel Test

This full-scale prototype is constructed using the solid foam build-up outlined in Section 6.2. All of

the electronics are installed so that deflections can be made using the trim settings on the transmitter. A

spinner is attached to the motor to simulate propulsion geometry during runs without an active propulsion

system. Testing is conducted in the NIAR 7ftx10 ft Walter H. Beech Memorial Wind Tunnel (Figure 44).

Figure 44 - Two Point Mounting System with Fairing for Dynamic Tare (Left)

Model Attached to Mounting System with Spinner (Right) Tufts are used to enable flow visualization. The tufts enable Team Black to determine which angle

the flow begins to separate along the wing and the fuselage (Figure 45).

54

Figure 45 - Tufts Showing Separation Near Stall

Table 21 - Full-Scale Wind Tunnel Test Matrix

Aerodynamic data is obtained and compared to predictions (Figure 47 in Section 8).

7.5 Ground Testing Proper operation of electrical subsystems and the propulsion system is verified upon completion of

the prototype aircraft. The water drop mechanism is tested multiple times to ensure there is no fatigue.

Run # Tare q (Psi) α (°) δe (°) Ψ (°) Notes1 Dynamic Takeoff -4,18,2 0 0 -----2 Dynamic Cruise -4,16,2 0 0 -----3 Static 0 -4,18,2 0 0 -----4 ----- Takeoff -4,18,2 0 0 Lift, Drag, and Moment Polars5 ----- Repeat Run 4 -4,18,2 0 0 "6 ----- Cruise -4,16,2 0 0 To determine if Re transition point is crossed7 ----- Takeoff -4,18,2 -10/-10 0 Control surface effectiveness8 ----- Takeoff -4,18,2 +10/+10 0 "9 ----- Takeoff -4,18,2 0 10 Sideslip

10 ----- Takeoff -4,18,2 0 20 Sideslip11 Static 0 -4,18,2 0 0 Propeller run12 ----- Takeoff -4,18,2 0 0 Propeller run13 ----- Cruise -4,16,2 -10/-10 0 Use following runs if Re transition is crossed14 ----- Cruise -4,16,2 +10/+10 0 "15 ----- Cruise -4,16,2 0 10 "16 ----- Cruise -4,16,2 0 20 "

17 ----- Takeoff -4,18,2 0 0 Flow visualization18 ----- Cruise 0 ? 0 Trim in real time19 static 0 -4,18,2 0 0 Remove verticle tail20 ----- Takeoff -4,18,2 0 0 "21 ----- Takeoff -4,18,2 0 10 "22 ----- Takeoff -4,18,2 0 20 "

Primary Test Matrix

Secondary Test Matrix

55

Ground handling is a concern in the design of the aircraft’s landing gear; therefore, ground handling

testing is needed before flights are conducted. This test validates that the aircraft does not ground loop

and the steerable tail wheel helps in the case of a crosswind.

7.6 Flight Testing A foam prototype model is built to validate the concept and verify aerodynamic, propulsion, and

stability and control parameters (Figure 46). The pre-flight checklist (Table 22) and flight test plan (Table

23) are designed for a DBF competition simulation.

Table 22 - Pre-Flight Checklist

Table 23 - First Flight Test Matrix

Multiple flight tests are scheduled to familiarize the pilot with the aircraft’s handling and performance.

Team Black Preflight ChecklistNight Before

1.) Set up reciever batteries on the charger

2.) Set up propulsion batteries on the charger

3.) Set up transmitter batteries on the charger

3.) Get supply box ready

Pre-Flight

1.) Attach receiver batteries to aircraft

2.) Attach propulsion batteries to aircraft

3.) CG aircraft

4.) Plug in receiver batteries

5.) Plug in propulsion batteries

6.) Plug in fuse

7.) Range check

8.) Elevon deflection check

9.) Door servo operation check

10.) Throttle up check

11.) Fail safe check

56

Figure 46 - First Ferry Flight - January 13, 2012

8.0 Performance Results

The half-scale and full-scale WTT results are compared in Figure 47. The mounting system in the

7x10 ft wind tunnel is much larger than the mount for the 3x4 ft wind tunnel. This reduces the amount of

lift that is produced by the fuselage and creates turbulence over the fuselage thus creating more drag.

Another reason for an increase in drag during the 7x10 test can be attributed to the landing gear being

added to the full-scale model and also a propeller or spinner added to the motor. Trim plots comparing

predicted and actual results can be found in Section 4.6.2 on Figure 18.

Figure 47 - Lift and Drag Comparison for Half-Scale and Full-Scale WTTs

To aquire data, a telemetry sensor is installed in the model which has the ability to attain global

positioning system (GPS) speed, altitude, and battery information. From this data, an average cruise and

T/O velocity are calculated. (Figure 48). Table 24 displays important information about the conditions and

setup of each of the three flights.

57

Figure 48 - Recorded GPS Speed (Wind 12 mph)

Table 24 - Flight Information

0

10

20

30

40

50

60

70

80

90

0 50 100 150 200 250

Mph

Time (S)

Flight # Battery Endurance-Max Power Total Time (min) CG (in) Temp (F) Wind (mph) Aircraft Weight (lbs) Type1 Elite 1500 3.9 4.75 5.375 30 S 12 2.22 Elite 2000 4.8 5.5 5.25 32 S 12 2.23 Elite 1500 3.9 4.25 5.125 34 S 12 2.2

Stabil ity and Propulsion

Analysis

58

9.0 References 1“2011/2012 Rules and Vehicle Design,” AIAA Student Design/Build/Fly Competition

[http://www.aiaadbf.org. Accessed 8/19/11]. 2Raymer, D.P., Aircraft Design: A Conceptual Approach, 4th ed., AIAA Education Series, AIAA, Reston, VA, 2006. 3WSU 2009 SAE Heavy Lift Team Buffalo Works, Final Report, Wichita State University, Wichita, KS,

2009. 4WSU 2009 AIAA DBF Team sUAVe, Scheduling Macro, Software, Brian Hinson, Wichita State

University, Wichita, KS, 2009. 5The Weather Underground, Inc., “Almanac for Wichita, KS,” Weather Underground, 2011,

http://www.wunderground.com/history/airport/KICT [Accessed 8/02/11]

6Montgomery, Douglas C., Design and Analysis of Experiments, 7th ed., John Wiley & Sons, Inc., Hoboken, NJ, 2009.

7Nickel, K. and Wohlfahrt, M., Tailless Aircraft in Theory & Practice, Butterworth-Heinemann, Oxford, Great Britain, 1994

8Hepperle, M., “Basic Design of Flying Wing Models,” Dr. Martin Hepperle Engineering, Germany,

January, 2002, [http://www.mh-aerotools.de/airfoils/flywing1.htm. Accessed 10/3/11]. 9Roskam, J., Airplanes Flight Dynamics and Automatic Flight Controls, Lawrence, KS, 2007. 10Buchmann, I., “Batteries in a Portable World,” Cadex Electronics, [http://www.buchmann.ca/. Accessed

9/9/11]. 11Merchant, M., and Miller, L.S., “Propeller Performance Measurement for Low Reynolds Number UAV

Applications”. AIAA paper 2006-1127, 2006 12Anderson, J.D., Introduction to Flight, 6th ed., McGraw Hill Anderson Series, Boston, 2005. 13Sarafin, T.P. (ed.), Spacecraft Structures and Mechanisms, Hawthorne, CA, 2007. 14WSU 2011 AIAA DBF Team MiniWheat, “Team MiniWheat Design Report,” AIAA DBF Report Archive,

Wichita State University, Wichita, KS, 2011. 15Selig, M. (2008, 02 19). UIUC Airfoil Data Site. Retrieved 10 01, 2008, from [www.ae.uiuc.edu/m-

selig/ads.html]. 16Miley, S.J. A Catalog of Low Reynolds Number Airfoil Data for Wind Turbine Applications. U.S. Dept. of

Energy, 1982. 17Stivers, L & Rice, F., Aerodynamic Characteristics of Four NACA Airfoil Sections Designed for

Helicopter Rotor Blades. NACA-RB-L5K02. Langley, Langley Field, VA: 1946. 18Liebeck, R.H., “Laminar Separation Bubbles and Airfoil Design at Low Reynolds Numbers,” AIAA Paper

92-2735, Jun. 1992. 19Drela, M., & Youngren, H. (2008, 08 04). AVL. Retrieved 09 14, 2011, from

[http://web.mit.edu/drela/Public/web/avl/].

59

20Etkin, B., and Reid, L.D. Dynamics of Flight: Stability and Control. 3rd ed. New York. John Wiley & Sons, Inc. 1996. p 19 -59.

21X-Plane, Software Package, Vers. 9.7, Laminar Research, Austin Meyer, United States, 2011. 22CATIA, Software Package, Vers. 5.18, Dassault Systemes, Vélizy-Villacoublay, France, 2008 23“Loose Cells.” CheapBatteryPacks. 10 September 2011 http://www.cheapbatterypacks.com19“ 24“Brushless Motors”, Towerhobbies. 10 September 2011 <http://www.towerhobbies.com 25“Brushless Motors.” 20Hobby King. 10 September 2011

http://www.hobbyking.com/hobbyking/store/index.rc 26“Phoenix 25.” Castle Creations. 26 October 2011 http://www.castlecreations.com/products/phoenix-

25.html 27“AWG.” American Wire Gauge. 28 November 2011 http://www.hardwarebook.info/AWG 28Hepperle, Martin . "Java Prop." 14 November 2011 http://www.mh-aerotools.de/airfoils/javaprop.htm. 29Zamora, A., et al, DBF Propulsion Manual, Wichita State University, Wichita, KS, 2007. 30Allen, D. & Haisler, W. Intro. to Aerospace Structural Analysis. College Station, TX: Wiley, 1985 31Hibbeler, R.C., Mechanics of Materials, 7th ed., Pearson Education Inc., New Jersey, 2008. 32 Logan, D.L., Finite Element Method, 5th ed., Cengage Learning Inc., Stamford, CT, 2012. 33Hoerner, S.F. Fluid Dynamic Drag. 1st ed. Bakersfield, CA: Hoerner, 1965.