Agent Blue
2010-2011 Design/Build/Fly Competition Cessna Aircraft Company Raytheon Missile Systems AIAA Foundation
TABLE OF CONTENTS 1. 2.
3.
4.
5.
6.
7.
8.
9.
EXECUTIVE SUMMARY .............................................................................................................. 4 MANAGEMENT SUMMARY ........................................................................................................ 5 2.1 TEAM ORGANIZATION............................................................................................................... 5 2.2 SCHEDULE .............................................................................................................................. 5 CONCEPTUAL DESIGN .............................................................................................................. 6 3.1 MISSION REQUIREMENTS ......................................................................................................... 6 3.2 SITE ANALYSIS ........................................................................................................................ 8 3.3 DESIGN OBJECTIVES ............................................................................................................... 9 3.4 CONCEPT W EIGHTING, SELECTION PROCESS, AND RESULTS ................................................... 10 3.5 CONCEPTUAL SYSTEM SUMMARY ........................................................................................... 16 PRELIMINARY DESIGN ............................................................................................................. 17 4.1 DESIGN AND ANALYSIS METHODOLOGY .................................................................................. 17 4.2 MISSION MODEL .................................................................................................................... 19 4.3 OPTIMIZATION TOOLS AND METHODOLOGY ............................................................................. 19 4.4 INITIAL SIZING........................................................................................................................ 20 4.5 AERODYNAMICS..................................................................................................................... 22 4.6 FLIGHT MECHANICS ............................................................................................................... 23 4.7 PROPULSION SYSTEM ............................................................................................................ 25 4.8 AVIONICS .............................................................................................................................. 26 4.9 STRUCTURES ........................................................................................................................ 26 4.10 AIRCRAFT MISSION PERFORMANCE PREDICTIONS ................................................................... 31 DETAIL DESIGN ........................................................................................................................ 31 5.1 DIMENSIONAL PARAMETERS ................................................................................................... 31 5.2 ELECTRICAL COMPONENT SELECTION .................................................................................... 32 5.3 AIRCRAFT COMPONENT W EIGHT AND CG BUILDUP ................................................................. 32 5.4 COOLING SYSTEM DESIGN ..................................................................................................... 33 5.5 STRUCTURES ........................................................................................................................ 34 5.6 TOTAL SYSTEM W EIGHT ........................................................................................................ 35 5.7 FLIGHT PERFORMANCE SUMMARY .......................................................................................... 36 5.8 MISSION PERFORMANCE SUMMARY ........................................................................................ 37 5.9 DRAWING PACKAGE ............................................................................................................... 37 MANUFACTURING PLAN AND PROCESSES ......................................................................... 43 6.1 MANUFACTURING FIGURES OF MERIT ..................................................................................... 43 6.2 COST ANALYSIS .................................................................................................................... 45 6.3 MANUFACTURING SCHEDULE ................................................................................................. 45 6.4 PROTOTYPE .......................................................................................................................... 46 6.5 AIRCRAFT MANUFACTURING ................................................................................................... 47 TESTING PLAN .......................................................................................................................... 48 7.1 SUBSYSTEMS ........................................................................................................................ 48 7.2 SYSTEM ................................................................................................................................ 51 7.3 MASTER TEST SCHEDULE ...................................................................................................... 53 PERFORMANCE RESULTS ...................................................................................................... 54 8.1 SUBSYSTEMS ........................................................................................................................ 54 8.2 SYSTEM ................................................................................................................................ 58 REFERENCES ............................................................................................................................ 59
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Acronyms, Abbreviations, and Symbols A
Amps
L/D
Lift-to-Drag Ratio
AC
Aerodynamic Center
Lb
Pounds
AIAA
American Institute of Aeronautics and
mAh
Milliamp Hours
MDO
Multidisciplinary design optimization
Astronautics AOA
Angle of Attack (also α)
Min
Minutes
AR
Aspect Ratio
MPH
Miles Per Hour
AVL
Athena Vortex Lattice
NACA National Advisory Committee on
b
Span
Aeronautics
c
Chord
NEU
c/4
Quarter-Chord
NiCad Nickel Cadmium
CA
Cyanoacrylate
NiMh
Nickel Metal Hydride
CAD
Computer Aided Design
Oz
Ounces
CC
Castle Creations
PSI
Pounds per Square Inch
CDO
Parasite Drag Coefficient
RC
Radio Controlled
CG
Center of Gravity
Re
Reynolds Number
CHT
Horizontal Tail Sizing Coefficient
RMS
Raytheon Missile Systems
CL
Lift Coefficient
RPMs Revolutions per Minute
Cl
Roll Coefficient
SHT
Horizontal Stabilizer Area
Cm
Pitch Coefficient
SM
Static Margin
Cn
Yaw Coefficient
SVT
Vertical Stabilizer Area
CNC
Computer Numeric Control
Sw
Wing Area
CVT
Vertical Tail Sizing Coefficient
t 1/2
Time to Half (Also Time to Double)
DBF
Design/Build/Fly
T/W
Thrust to Weight Ratio
Deg
Degrees
UF
University of Florida
Eq
Equation
V
Voltage
ESC
Electronic Speed Controller
W
Weight
FOM
Figures of Merit
W.R.T With Respect To
Ft
Feet
W/Sw
Wing Loading
Ft/s
Feet per Second
WDM
Weighted Decision Matrix
Hz
Hertz
WF
Weighting Factor
IHT
Distance between 1/4 chord of main
β
Sideslip Angle
ζ
Damping
wing and horizontal stabilizer
Neutronics
In
Inches
η
Real Response
IVT
Distance between 1/4 chord of wing and
ω
Frequency
vertical stabilizer
Page 3 of 59
1. EXECUTIVE SUMMARY This report details the design, testing, and manufacturing efforts of University of Florida team Agent Blue in preparation for the 2010-2011 AIAA/Cessna/RMS Design/Build/Fly (DBF) competition. Agent Blue’s primary objective is to design a winning aircraft by obtaining a maximum total overall score, which is a function of written report score and total flight score. The total flight score is determined by the performance in three missions: Dash to critical target, Ammo Re-Supply, and Medical Supply Mission. The Dash to critical target focuses on flight time, requiring low drag and a high cruise speed. The Ammo Re-Supply demands optimization to find a balance between the payload weight and flight weight. The External Payload Flight requires high payload capabilities as the score is normalized to the maximum number of golf balls flown. Design requirements constrain aircraft size to fit inside a commercially produced carry-on bag measuring a total of 45 dimensional inches. A flight score study reveals system weight and the payload capacity as the most critical design parameters. Aircraft design concepts are developed to meet all mission requirements and achieve a maximum total flight score. Four aircraft configurations are evaluated for a conceptual design: Conventional, Flying Wing, Canard, and Biplane. Figure of Merit (FOM) analyses are utilized to select aircraft configuration and components, which yield a conventional configuration with a single tractor propeller, wheel/skid combination landing gear, V tail, and rectangular wings. Each component is selected to achieve adequate aircraft performance while minimizing system weight. The conventional body configuration with a single tractor propeller provides sufficient internal room for payload storage and superior stability at minimal weight. The wheel/skid landing gear minimizes weight and drag while preventing damage during landing. The V-tail provides the best stability for its size and weight. Rectangular wings improve manufacturability, are lightweight, and provide excellent stability. A multidisciplinary design optimization (MDO) architecture is developed to aide in the analysis of critical design parameters including airfoil selection, wing span, wing area, tail size, and propulsion system selection. An S2091-101-83 airfoil is selected because of its high lift-to-drag ratio and minimal 2
thickness. The wing is sized to a 50 in span and a 500 in area to meet hand launching requirements. 2
Longitudinal and Lateral-directional stability is provided by two 70 in V-tail stabilizers designed with a NACA 0009 airfoil. Propulsion analysis yields a NEU 1105/3Y motor with a 4.4:1 gear ratio, 11x6 Graupner propeller, and 13 Elite 1500 cells as the best combination for missions 2 and 3. A 10x6 Graupner propeller is used for mission 1. The aircraft disassembles to fit inside a 20 in × 10 in × 11 in suitcase. The predicted performance capabilities of Missions 1 / 2 / 3 are as follows: thrust-to-weight ratios are 0.71 / 0.46 / 0.46, stall speeds are 26.4 / 35.2 / 35.2 ft/s, and cruise speeds are 58.7 / 51.3 / 51.3 ft/s. The predicted unloaded aircraft weight is 2.61 lb capable of carrying a 2.25 lb steel bar and 22 golf balls. These predicted performance and scoring parameters result in a highly competitive mission oriented vehicle with an estimated competition score of 2.69.
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2. M ANAGEMENT SUMMARY 2.1
Team Organization The 2010-2011 2011 Agent Blue team is broken down into four technical groups: aerodynamics,
propulsion, structures, and manufacturing. Each group has a lead engineer responsible for organizing tasks and group meetings. The Documentation Lead and CAD Lead improve team collaboration and efficiency. The team organization is shown in Figure 2.1.
Chief Engineer James Lasater
Aerodynamics Lead
Propulsion Lead
Jason Rue
James Lasater
Documentation
Structures Lead
Manufacturing Lead
Kevin Rausch
Abe Balmori
Carlos Caballero
Ed Reed
Jason Cantrell CAD Laura Rose
Jason Cantrell
Kevin Rausch
Figure 2.1 – Team Organization Each group has unique responsibilities. •
Aerodynamics Group is responsible for aerodynamic force resolution, airfoil analysis, wing planform design, and flight mechanics.
•
Propulsion Group performs analysis and testing of all electronic systems in the aircraft, including propulsion and avionics systems.
•
Structures Group is responsible for designing the airframe and storage case, selecting construction materials, and testing structural compo components.
•
Manufacturing Group analyzes and executes methods of construction for each component of the aircraft and case.
The responsibilities of the remaining positions are detailed below. •
The Chief Engineer guides the team throughout the entire design and ma manufacturing nufacturing process while serving as the primary link between the students and the faculty adviser.
•
The Documentation Lead works with members from every technical group to document in a report the design, testing, and manufacturing of the aircraft and case case.
•
The CAD Lead instructs and manages members from each group in constructing a CAD model of the aircraft and storage case.
2.2
Schedule A detailed schedule of production is constructed for completion of the project in a timely manner. manner
Each group lead is responsible sible for overseeing and managing the tasks of their groups.. The Gantt chart in Page 5 of 59
Figure 2.2 displays the project’s planned schedule as well as the actual completion time for the corresponding tasks. Projected times for the remaining tasks are also displayed.
Figure 2.2 – Design Schedule 3. CONCEPTUAL DESIGN The conceptual design phase begins by determining mission requirements and design constraints. To achieve the best design, Figure of Merit(FOM) analyses are performed for the various aspects of the design process. Each of the considered designs meets the general contest requirements and complies with the mission requirements laid out by competition rules. 3.1
Mission Requirements The competition score is the product of the written report score (100 points) and the flight score,
which is the summation of the normalized scores for Mission 1 (1 point), Mission 2 (3 points), and Mission 3 (2 points). The missions are weighted such that the latter two missions are of greater importance to the flight score, and they must be completed in order with no scores being awarded for partially completed missions. These constraints ensure that all designs will be optimized for all missions. General contest specifications and mission specifications are listed below. •
The aircraft must be hand thrown.
•
The maximum battery weight is 3/4 lb. Only NiCad or NiMH batteries are allowed.
•
Current is limited by a 20 Amp slow-blow fuse.
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•
One commercial suitcase meeting airline carry-on bag rules is allowed. The aircraft with all assembly tools and support equipment must fit into this case.
•
Team selected and supplied steel bars of varied lengths and weights must be carried by the aircraft. The steel bars must be a minimum 3-in width x 4-in length
•
Team selected quantity of competition-supplied golf balls must be carried internal to the aircraft.
•
The aircraft must complete a successful landing on the runway for any mission to be scored.
3.1.1
Mission 1: Dash to Critical Target
Mission 1 requires the airplane to complete as many laps as possible in a 4 minute flight time. The
score is given by Eq. 3.1, where ܰ௦ is number of laps completed by an individual team and ܰ௦ ௫ is
the maximum number of completed laps obtained by any team. The flight time begins when the aircraft leaves the launcher’s hand and ends at the completion of the last lap when the plane passes over the finish line while in the air within the 4 minute flight time. ܯ1 = 3.1.2
ܰ௦
ܰ௦ ௫
Eq. 3.1
Mission 2: Ammo Re-Supply
Mission 2 requires the aircraft to complete three laps with a payload of team selected steel bar payload. The steel bars must have a minimum 3-in width x 4-in length. The mission score is given by Eq.
3.2, where ݈݃݅ܨℎݐ௪௧ is the weight of the loaded aircraft recorded immediately after completion of a
successful flight, and ݈ܲܽ݀ܽݕ௪௧ is the weight of the removed payload recorded immediately after completion of a successful flight. ܯ2 = 3.1.3
3 ∗ ݈ܲܽ݀ܽݕ௪௧ ݈݃݅ܨℎݐ௪௧
Eq. 3.2
Mission 3: Medical Supply Mission
Mission 3 requires the airplane to fly three laps with a payload of a team selected quantity of golf
balls. The mission score is shown in Eq. 3.3, where ܰ௦ is the number of golf balls carried by the team
aircraft and ܰ௦ ௫ is maximum number of golf balls flown by any team. ܯ3 =
3.1.4
2 ∙ ܰ௦ ܰ௦ ௫
Eq. 3.3
Score Optimization Analysis
A score sensitivity study is performed to evaluate the effects of mission scoring parameters on the overall flight score. The equation for the overall competition score is shown below in Eq. 3.4. ∙ ݁ݎܿܵ ݐݎܴ݁ = ݁ݎܿܵ ݊݅ݐ݅ݐ݁݉ܥ
൬
݈ܰܽݏ
݈ܰܽݔܽ݉ ݏ
+
3∗ܲܽ݃݅݁ݓ݈݀ܽݕℎݐ ݈݃݅ܨℎ݃݅݁ݓݐℎݐ
ඥܴܥܣ௪௧
+
2∙ܾ݈݈ܰܽݏ ൰ ܾ݈݈ܰܽݔܽ݉ ݏ
Eq. 3.4
Best score estimates are needed for each score parameter to account for score normalization and score parameter estimates shown in Table 3.1.
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Table 3.1 – Score Parameter Estimates Parameter
Estimate
Aircraft Weight
2.25 lb
Mission 1: Completed Laps (N1,ref)
6
Mission 2: Payload Weight (w2,ref)
2.25 lb
Mission 3: Number of Golf Balls (Nmax,ref)
22
Data from the previous competitions indicate that the fastest anticipated lap time is approximately 35 seconds however new regulations are expected to increase this time to 40 seconds. Thus, a perfect mission 1 score can be obtained by completing 6 laps. The estimated maximum payload for Missions 2 and 3 is estimated in Motocalc. The outputted results are based off of wing loading typical in hand launched aircraft. Because these two missions are not timed, the only requirement for a perfect score is that the aircraft is able to complete all 3 laps with the highest payload. These best score estimates are incorporated into the score sensitivity analysis shown in Figure 3.1. 2.50
Estimated Score
2.00
Mission 1 Mission 2 Mission 3 Total Score
1.50 1.00 0.50 0.00 42.00
43.00
44.00
45.00
46.00
47.00
48.00
49.00
50.00
Wingspan (in.) Figure 3.1 – Comparative Score Sensitivity Analysis Figure 3.1 illustrates that in order to obtain the highest score the wingspan must be made as large as possible. While the mission 1 score is constant as the wingspan increases, the mission 2 and 3 scores increase significantly. Due to storage space limitations, a wingspan greater than 50 inches is considered unfeasible. Further optimization will only be conducted on a 50 in. wingspan airplane. 3.2
Site Analysis Temperature, barometric pressure, wind speed, and dew point data are collected over a ten year
period for the month of April to help predict the weather conditions for the competition. The density altitude and air density is calculated using the collected data as shown in Figure 3.2. The highest density altitude that the plane is predicted to fly at is 4,539 ft occurring at 4:00 PM. The collected data helps
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predict the actual performance of the aircraft under conditions present at the competition site as opposed
Altitude (feet)
to standard atmospheric conditions. 4800 4600 4400 4200 4000 3800 3600 3400 3200 8:00
9:00
10:00 11:00 12:00 13:00 14:00 15:00 16:00 17:00 18:00 Figure 3.2 – Density Altitude
The site analysis also includes a collection of wind speed data. Since it is difficult to predict wind direction, the maximum wind speed is assumed to be a crosswind. As shown in Figure 3.4, the wind speed steadily increases throughout the day, reaching its maximum speed of 13 mph at between 3:00 and 4:00 PM.
Wind Speed (mph)
14 12 10 8 6 4 2 0 8:00
9:00
10:00 11:00 12:00 13:00 14:00 15:00 16:00 17:00 Time of Day Figure 3.3 – Wind Speed
The best weather conditions occur at the beginning of the day. Since the flight line order is based on report scores, a high report score improves the chances of successful flight missions. 3.3
Design Objectives After concluding the score and weather analysis, the team concludes that the design solution must
have the following requirements for the different missions. •
Mission 1: Dash to Critical Target: Minimize aircraft weight to improve dash time.
•
Mission 2: Ammo Re-Supply: Optimize propulsion system, wing size, and airfoil to minimize overall aircraft weight.
•
Mission 3:Medicine Supply Mission: Optimize the payload bay size to reduce weight and drag. Page 9 of 59
3.4
Concept Weighting, Selection Process, and Results A successful aircraft begins with a conceptual design appropriate for the missions at hand. A Figure
of Merit (FOM) analysis is performed to discern among the design concepts. The FOMs used in evaluating concepts are listed below. •
Weight: The weight includes component and assembly tool weight.
•
Ease to Build: Aircraft components must be built and repaired in a short amount of time.
•
Storage Options: The plane must fit into a commercially produced suitcase
•
Stability: A measure of relative stability properties of a concept.
•
Payload Feasibility: The aircraft must accommodate all payloads.
•
Aerodynamic Efficiency: A measure of the lift and drag properties of a concept.
•
Launch Clearance: Hand launching must allow clearance for the safety of the thrower.
•
Ease of Landing: The pilot must land the plane without damaging the aircraft.
•
Impact Tolerance: The aircraft must not sustain damage during landing on the runway.
•
Assembly Time: The aircraft must be assembled within five minutes prior to each mission.
•
Reliability: The aircraft and components must function properly every time it is assembled.
The FOMs are assigned a Weighting Factor (WF) based on its importance such that the sum of the WFs is 1.00. Each design alternative is evaluated based on how well it performs for a particular FOM. The concept rating criteria are shown in Table 3.2. Table 3.2 – Rating Criteria Rating Value -1 0 1
Concept Characteristics Poor performance with respect to design criteria Moderate performance with respect to design criteria Exceptional performance with respect to design criteria
The rating values are multiplied by the WFs and summed for an overall score. The highest overall score is selected for the conceptual design. 3.4.1
Aircraft Configuration
Conceptual design begins with the determination of an aircraft configuration. The selected aircraft configuration must give the team the best chance of success. Aircraft configurations are considered by their ability to successfully complete all of the missions. Research is conducted on previous competition planes, and the field is narrowed down to the four alternatives shown below. •
Conventional: A conventional aircraft provides good stability in flight while allowing the payloads to be carried easily. The conventional configuration breaks down into several smaller pieces to fit in a small storage case.
•
Flying Wing: A flying wing configuration is lightweight and takes up minimal storage space as it does not have a tail, but it has poor stability characteristics. It may be difficult to fit all the golf
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balls inside the aircraft and great care must be taken to ensure that the aircraft is properly constructed. •
Canard: The canard is similar to the conventional configuration but the building complexities of a canard aircraft are greater than that of a conventional design. These complexities, as well as having little information available on designing canards, are obstacles the team must overcome.
•
Biplane: Biplanes provide additional lift for a small span and are stable, but the additional wings increase weight, build time, and limit the aircraft’s storage options.
The Weighted Decision Matrix (WDM) shown below in Table 3.3 gives a comparison between the aircraft configurations. This matrix identifies the advantages and disadvantages of each concept. Each option is scored based on the FOMs to emphasize the critical parameters. Table 3.3 – Aircraft Configuration Weighted Decision Matrix
Aircraft Configuration
FOM Weight Ease to Build Storage Options Stability
WF
Conventional
Flying Wing
Canard
Biplane
0.25 0.10 0.30 0.25
0 0 1 1
0 0 -1 -1
0 0 1 -1
-1 -1 -1 1
Payload Feasibility Total
0.10 1.00
1 0.65
0 -0.55
0 0.15
0 -0.40
A conventional aircraft configuration is selected by the team. The conventional configuration has good stability characteristics, is easy to manufacture, and can fit into a small storage case. Additionally, the weight of a conventional plane can be minimized. 3.4.2
Fuselage Configuration
Due to the score optimization analysis that identified mission 2 and 3's payload capacity as the critical score parameters, it is decided that the design of the fuselage is paramount to a high finish. The fuselage must have enough volume to carry an ample amount of golf balls and be wide enough to accommodate the steel bars width. The primary goal of the fuselage design is for the optimization of the golf ball configuration but, design of the fuselage was also driven by maximum length of the storage case. Assembly time must also be taken into account as the aircraft must be fully assembled within five minutes. The options for fuselage configurations are shown below. •
Single Column: A single wide column is the simplest configuration for the golf balls. However it is troublesome for the steel bar due to its minimal width. Carrying a large amount of golf balls would add considerable length to the aircraft and assembling the aircraft would take longer as golf balls would be loaded into the aircraft one at a time. Page 11 of 59
•
Double Column: A double column provides ample width for the steel bar payload and for the golf balls. The double column would hinder the number of golf balls carried because of storage limitations.
•
Quad Column: A quad column provides ample width for the golf ball the shortest fuselage length. The quad column does increase drag due to a larger cross section but will allow the most storage in the smallest area.
The WDM is shown in Table 3..4 and compares the landing gear configurations based on the FOMs. Table 3 3.4 – Fuselage Weighted Decision Matrix
Fuselage Configuration
FOM Assembly Time Storage Options Aero Efficiency Payload Feasibility Total
WF
Single
Double
Quad
0.20 40 0.40 0.15
-1 -1 1
0 0 0
0 1 -1 1
0.25
-1
1
1
1.00
-0.70
0.25
0.50
The Quad column configuration was selected due to the highest payload capacity and storage options. This configuration will allow both the steel bar and golf ball payloads to be easily stored within the th fuselage. 3.4.3
Wing Configuration
An aircraft’s wings are vital to the success of an airplane as they provide lift. The following wing alternatives are considered base on their ability to give the aircraft the best chance of a successful flight. •
Rectangular: Rectangular wings are easy to manufacture and provide good stability.
•
Elliptical: Elliptical wings have high aerodynamic efficiency but are difficult to construct. Elliptical wings are also prone to the entire wing stalling at once.
•
Swept: Swept wings provide ide lateral stability but see most of their benefit in transonic flight.
•
Tapered: Tapered wings have high aerodynamic efficiency and provide good stability. However, tapered wings do require more precision to manufacture.
The WDM is shown in Table 3..5 below and compares wing configurations based on the FOMs.
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Table 3.5 – Wing Weighted Decision Matrix
Wing Configuration FOM Weight
WF
Rectangular
Elliptical
Swept
Tapered
0.25
1
1
-1
1
Ease to Build Aero Efficiency Stability Total
0.25 0.20 0.30 1.00
1 0 1 0.80
-1 1 -1 -0.10
0 0 1 0.05
0 1 1 0.75
The Agent Blue team selected rectangular wings due to their ease of manufacturing and light weight construction. 3.4.4
Tail Configuration
Many tail configurations are considered because of the desire to minimize storage space. The tail needs to provide good stability characteristics while allowing for the smallest possible storage case. The tail alternatives are considered based on their potential to provide stability to the aircraft. The alternatives for the tail are listed below. •
Conventional: The conventional tail provides low weight and good stability.
•
V-tail: The V-tail is lightweight due to having only two control surfaces instead of three. It also has all surfaces above the tail boom providing better launch clearance.
•
U-tail: The U-tail has two vertical stabilizers that allow great yaw stability at a lower vertical height. The additional support needed for two vertical stabilizers adds extra weight however.
•
T-tail: The T-tail requires extra strengthening of the vertical stabilizers due to the horizontal stabilizer being at the top of the tail. It also has limited storage options.
The WDM shown in Table 3.6 shows the tail configuration decision. Table 3.6 – Tail Weighted Decision Matrix
Tail Configuration FOM
WF
Conventional
V-Tail
U-Tail
T-Tail
Weight
0.20
0
1
-1
-1
Storage Options
0.20
1
1
-1
0
Ease of Build
0.10
1
1
0
0
Launch Clearance
0.30
0
1
-1
1
Stability
0.20
1
0
1
1
Total
1.00
0.50
0.80
-0.50
0.30
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The V-tail is the best configuration for the aircraft. The team selects the V-tail for its excellent storage options, weight, and launch clearance. 3.4.5
Motor Configuration
Various motor configurations are compared to determine the best propulsion system for the aircraft. Motor configurations must be efficient and lightweight to obtain the highest possible score. Alternatives selected for motor configuration are shown below. •
Single Tractor: The single tractor configuration is lightweight, easy to manufacture, and provides adequate ground and payload clearance but disturbs airflow over the fuselage.
•
Double Tractor: The double tractor has excellent ground and payload clearance, but it disrupts airflow over the wings and requires two motor mounts to be built.
•
Single Pusher: The single pusher disturbs airflow over the tail and requires twin tail booms. This configuration is lightweight and easy to construct. The pusher option is a problem for hand launched vehicles as the propeller is likely to be in the way of thrower's body.
•
Double Pusher: The double pusher configuration does not disturb airflow over the wings, providing a high aerodynamic efficiency, but it is heavy and difficult to construct. The double pusher, like the single is a hazard to the launcher.
The WDM used to determine the motor configuration is shown in Table 3.7. Table 3.7 – Motor Configuration Weighted Decision Matrix
Motor Configuration
0.30 0.15
Single Tractor 1 1
Double Tractor 0 -1
Single Pusher 1 1
Double Pusher -1 -1
0.30 0.10 0.15 1.00
1 0 1 0.90
1 -1 0 0.05
-1 0 0 0.15
-1 1 -1 -0.80
FOM
WF
Weight Storage Options Launch Clearance Aero Efficiency Ease to Build Total
A single tractor motor configuration is selected for the aircraft. It is lightweight, easy to store, and provides optimum launch clearance. 3.4.6
Landing Gear
Optimal configuration and placement of landing gear is necessary for proper ground handling. Other requirements involve the landing gear being lightweight and having sufficient strength for irregular landings. Landing gear alternatives are considered based on their ability to allow the aircraft land successfully on a paved runway. Clearance of the landing gear is not a major issue as the team has decided a folding propeller will be used allow for a larger propeller size.
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•
Tail Dragger with Wheels: The tail dragger with wheels has a high impact tolerance due to the cushioning of the wheels.
•
Tail il Dragger with Front Wheels/Back Skid: The wheels/skid option has a high impact tolerance while being lightweight and providing the adequate storage options.
•
Skids only: The skids only configuration provides a lightweight option. However, it has no impact tolerance and the probability of damage from an irregular landing is high.
The WDM is shown in Table 3..8 and compares the landing gear configurations based on the FOMs. Table 3.8 – Landing anding Gear Configuration Weighted Decision Matrix
Landing Gear
FOM
WF
Wheels Only
Weight
0.25
-1
Front Wheels/ Back Skid 0
Aero Efficiency Impact Tolerance Ease to Build Total
0.15 0.35 0.25 1.00
0 1 0 0.10
0 1 0 0.35
Skids Only 1 0 -1 1 0.15
The front wheels/ back skid option is selected for its storage options and impact tolerance. 3.4.7
Assembly Options
While aircraft assembly time is not a factor in the scoring, the aircraft must be assembled within five minutes prior to each flight light attempt. An aircraft that breaks down into several pieces fits into a smaller, lighter case although it takes longer to assemble. Assembly options are considered based on their ability to allow the aircraft to fit in a small storage case and are comp compared ared against their respective counterparts. For example, a fixed landing gear is compared solely against a removable landing gear, not against wing and tail assembly options. •
Fixed Landing Gear: A fixed landing gear is reliable and requires no tools but may ma take up excessive storage space.
•
Removable Landing Gear Gear: A removable landing gear saves space but may require tools and requires care to ensure that the landing gear is properly attached to the aircraft.
•
Folding Wings: The folding wings see a significant benefit in assembly time, but the folding and locking mechanisms increase weight.
•
Sliding Wings: Sliding wings take little time to assemble and are lightweight.
•
Fixed Tail: A fixed tail is reliable and lightweight but takes up a lot of storage space.
•
Extendable Tail: An extendable tail could contract into the fuselage or telescope. This configuration is heavy and may not be reliable, but has a fast assembly time. Page 15 of 59
Removable Tail: A single--boom boom configuration allows for the booms to be easily attached and
•
removed oved from the fuselage while maintaining a minimal weight. The removable tail may not be reliable however. Table 3.9 below displays the WDM for landing gear, w wing, ing, and tail assembly options. Table 3.9 – Assembly Options Weighted Decision Matrix Landing Gear Wings Tail Assembly Options FOM Weight Assembly Time Storage Options Ease to Build Reliability Total
WF 0.25 0.10 0.40 0.10 0.15 1.00
Fixed Removable Folding 1 0 1 1 0 1 0 1 0 0 0 -1 1 0 -1 0.50 0.40 0.1
Sliding 0 -1 1 0 1 0.45
Fixed 1 1 -1 1 1 0.20
Extendable 0 1 1 -1 -1 0.25
Removable 1 0 0 0 -1 0.10
The FOM analysis shows that an aircraft with removable lan landing ding gear, sliding wings, and a removable tail is ideal for aircraft assembly. 3.5
Conceptual System Summary Upon completing an analysis of mission requirements and scoring, the Agent Blue team develops a
conceptual design for the aircraft system. •
A conventional al aircraft design provides great stability characteristics.
•
Rectangular wings are lightweight and easily constructed.
•
A V-tail tail reduces storage space and weight while keeping all surfaces above the launcher.
•
A single tractor motor configuration is lightwei lightweight ght and provides ample launch clearance.
•
The wheel/skid landing gear configuration provides for exceptional impact tolerance.
•
Fixed landing gear, removable wings, and an extending tail allow for optimum storage options.
Preliminary estimates are made of th the e size, weight, and performance of the aircraft and storage case and are listed in Table 3.10.. This data provides a foundation on which to begin the optimization for the aircraft. Table 3.10 10 – Initial Size and Performance Estimates Size and Weight Performance Wingspan Cruise Velocity 42 – 50 in 45 – 55 5 ft/s Aircraft Length Max Thrust 34 – 44 in 2 – 5 lb Tail Span Fastest Lap Time 18 – 22 in 40 – 50 s Empty Weight Assembly Time 2.25 – 3 lb 2 – 3 min The conceptual CAD model in Figure 3.4 shows the complete conceptual design solution in the assembled and stowed configuration.
Page 16 of 59
Figure 3.4 – Assembled and Stowed Conceptual Aircraft Design 4. PRELIMINARY DESIGN The conceptual design configuration is divided into five groups for analysis and development, including: aerodynamics, flight mechanics, propulsion, avionics, and structures. A market survey surve of DBF aircraft from previous competitions is used to form initial sizing estimates for the aircraft. A mission model and trade studies are then performed to evaluate and select competitive design alternatives. Performance evaluations are conducted on the e selected design alternatives to ensure the successful completion of all three missions and the highest possible score. 4.1
Design and Analysis Methodology A list of critical design parameters is developed to guarantee a competitive design. Elevated
importance e is placed on reducing the aircraft weight as it is one of the largest contributors to the total flight score. Specific critical design parameters are considered in the appropriate analysis and development group. The methods by which trade studies are per performed formed are discussed below. 4.1.1 •
Aerodynamics Critical Parameters Wing Area: The fact that the plane will be launched by hand is a significant factor when considering the wing area. A large wing area allows for a greater payload capacity and lower wing loading, but is more troublesome to store and has more drag than a smaller wing area. A large wing area is desired as score sensitivity analysis shows that payload capacity is one of the most important factors in the overall flight score.
•
Wing Span: An aircraft’s wing span is critical in aerodynamic performance as it determines several parameters, including the aspect ratio of the wing. Although a larger wing span yields a better aspect ratio, the wing must be built much stronger, thereby increasing the weight, to accommodate for a larger bending moment. The optimal balance between aspect ratio and weight must be achieved to maximize the flight score.
•
Airfoil: Airfoil selection is critical as it affects the flight transition period (takeoff), cruise velocity, and stall all characteristics of the aircraft. An airfoil with a high lift lift-to-drag drag ratio, high stall angle, and a small airfoil thickness is desired.
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4.1.2 •
Flight Mechanics Critical Parameters Static Margin: The static margin provides a balance between pitch stability and control surfaces deflections for trim. Values for stable airplanes typically range from 5% to 40%. The static margin is desired to be in the range of 5% to 20% to ensure that the plane is stable and responsive to control surface deflections.
•
Tail Size: The tail is an important aspect in the static and dynamic stability of an aircraft. An appropriately designed tail provides longitudinal and lateral-directional stability while keeping weight and trim drag to a minimum.
•
Control Surface Location and Size: Tail control surface deflections are limited to ±20˚ to prevent flow separation that may take place at large deflection angles. Aileron deflections are limited to ±20˚. Ailerons are located at the outboard ends of the wings to maximize effectiveness.
4.1.3 •
Propulsion Critical Parameters Motor and Gearbox: Primary factors in motor selection are the total system weight and thrust requirements hand launching of the aircraft. Brushless electric motors are preferred because they are lighter and more efficient than brushed motors. The motor and gearbox are selected together as gear ratio availability for specific motors is limited. MotoCalc software is used to obtain estimated thrust curves for each motor and gearbox combination.
•
Batteries: Of the two types of batteries allowed for the competition, nickel metal hydride (NiMH) cells are preferred over nickel cadmium (NiCad) cells due to their higher energy density. Although multiple battery packs are allowed, using just one battery pack is desired due to cost restraints. The battery pack will be optimized for a high flight score by balancing weight and performance.
•
Propeller: The propeller is the component of the propulsion system that produces thrust. Important factors to consider when selecting a propeller are diameter, pitch, and material. Larger diameter propellers produce more thrust but also draw more current from the batteries. A high pitch is desired for maximizing cruise speed, but it produces lower thrust at low speeds which is a critical component for hand launched aircraft. Carbon-fiber folding propellers are considered as they are some of the more common and efficient propellers.
4.1.4 •
Avionics Critical Parameters Electronic Speed Controller: The ESC allows the pilot to control the throttle in flight. It is selected in conjunction with the propulsion system to ensure that it can carry the voltage and current loads of the propulsion system.
•
Servos: Servos must be lightweight and provide the torque required to deflect the control surfaces at cruise conditions. Servos are examined in groups of control surfaces according to required torque. Selecting servos in this manner allows for interchangeability and easy replacement in the event of servo damage from a crash.
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4.1.5 •
Structural Critical Parameters Spar Location: Two wing spars are desired for structural stability. The main spar is the aircraft’s primary load path and is located at 25% of the chord, near the aerodynamic center of the airfoil. The rear spar assists in preventing the wing from twisting in flight.
•
Spar Cross-section: The primary concern for the wing’s structure is the bending moment caused by lift. Circular carbon tubes are strong and readily available for purchase, but rectangular crosssections, when oriented correctly, are more efficient under bending loads. The complexities associated with rectangular cross section manufacturing and design led the team to choose carbon fiber tubes for wing spars.
•
Material Selection: Light materials with sufficient strength for each component of the aircraft are selected. It is essential that these materials can be obtained and manufactured.
4.2
Mission Model Each flight mission is modeled in four phases: Launch, climb, cruise, and turns (180° and 360°). •
Takeoff: The model assumes a full-throttle setting and a 0 ft/s wind speed.
•
Climb: A full throttle climb to 75 ft is assumed. Time-to-climb is calculated based on individual mission take-off weight.
•
Cruise: Level flight is modeled to find cruise speed for minimum thrust required. Mission 1 maintains a full-throttle position for cruise to minimize flight time. Missions 2 and 3 are untimed missions but will also maintain a full-throttle position to ensure sufficient battery capacity.
•
Turns: Turn rates are calculated with a 2G load factor and the throttle settings used during cruise. The turn rates for both the 180° and 360° turns are modeled using these parameters.
4.3
Optimization Tools and Methodology A multidisciplinary design optimization (MDO) system aids in the design process. Each group on the
team develops and maintains modules to control the design parameters. The design parameters are compiled into aircraft solutions and evaluated by the mission model program to determine the optimal design. The MDO architecture is displayed in Figure 4.1.
Figure 4.1 – Design Optimization Architecture
Page 19 of 59
•
Propulsion Module: Motor, gearbox, propeller, and battery pack combinations are analyzed using MotoCalc. Thrust and current curves are recorded for further evaluation in the mission missio model program.
•
Aerodynamics Module: The aircraft's lift, drag, and moment coefficients are re calculated from selected airfoil and wing planform specifications. The air density used in this module is based on the lowest anticipated air density calculated in the weather analysis.
•
Flight Mechanics Module: The aircraft’s tail and control surfaces are sized to ensure the aircraft is stable and controllable in anticipated flight conditions, including potential crosswinds.
•
Material Properties: The weight and stren strength gth of potential construction materials are measured and recorded.
•
Structures Module: Structural layouts are developed using appropriate materials to provide sufficient strength while minimizing weight.
•
Aircraft Solution: Candidate solutions obtained from each module are compiled and further evaluated in the mission model program.
•
Mission Model: Stall speeds, cruise velocities, battery capacities, and flight times are simulated for each aircraft solution. Predicted flight scores are recorded and compared.
4.4
Initial Sizing
4.4.1
Fuselage Sizing
The initial parameters of the fuselage are completely dependent upon the volume of the internal payload and avionics components required for competition. The internal payloads with the most volume are the golf balls. Having conceptually onceptually determined that the fuselage design would consist of a stacked 2x2x6 configuration, this volume determined the fuselage's geometric parameters. To reduce drag as much as possible, the cross sectional area of the fuselage is limited to approxima approximately tely 1.25 times the area of a 2x2 grid of golf balls plus an additional 0.75 inch of clearance for wing attachments in the vertical direction and 0.25 inches in the horizontal direction. The length of the fuselage is limited to also approximately 1.25 timess the length of 6 golf balls lined up in a row plus an additional 5.25 inches for avionics components and fairings. After calculations the fuselage was determined to have a cross section of 5x4.5 inches and a length of 18.00 inches. Figure 4.2 shows a general top-down down view of the fuselage.
Figure 4.2 – Fuselage Top Down View Page 20 of 59
4.4.2
Wing Sizing
When optimizing for wing size, there are many aspects to consider. Increasing wing area increases lift, drag, stall speed, and wing loading. A larger wing area improves payload capacity at the cost of greater drag and stall speed. A large aspect ratio for the wing provides efficient flight, where as a lower aspect ratio allows for aircraft maneuverability. Finding a balance between the two is essential. Another issue considered for wing sizing is large wingspans may be difficult to fit into the storage case. Due to the aircraft requiring hand launch capabilities wing loading was a major factor in initial sizing conditions. Research of hand launched aircraft showed that the maximum safe wing loading would be approximately 2
25 oz/ft . Once the wing loading was found, a wing area can be calculated using an estimated initial
weight, ܹ௧ , and a wing loading,
ௐ
ௌೢ
. Following the calculation of the wing area while knowing the
maximum span is 50 in, aspect ratio and chord can be calculated using Eq. 4.2 and Eq. 4.3 respectively. ܵௐ = ൬
ܹ ିଵ ൰ ∗ ܹ௧ ܵ௪
= ܴܣ ܿ= 4.4.3
Tail Sizing
Eq. 4.1
ܾଶ ܵ௪
Eq. 4.2
ܾ ܵ௪
Eq. 4.3
Upon determining the surface area of the wing, the tail size was then estimated using the traditional tail volume coefficient equations. This method uses predetermined sizing coefficients for the horizontal and vertical tails presented in Corke (2003). The initial horizontal and vertical tail areas were determined using Eq. 4.4 and Eq. 4.5 respectively. Since the aircraft design calls for a V-tail aircraft these equations are simply added together to find the total volume. The tail dihedral angle was found using Eq. 4.6. ்ܵ = ்ܿ ∙ ܵு் = ܿு் ∙
ܾ௪ ܵ௪ ்݈
ܿҧ௪ ܵ௪ ݈ு்
݈ܶܽ݅݅ܦℎ݁݀ି݊ܽݐ = ݈݁݃݊ܣ݈ܽݎଵ ൬
Eq. 4.4 Eq. 4.5 ்ܵ ൰ ܵு்
ଵൗ ଶ
൩
Eq. 4.6
The tail was placed directly behind and in the same plane as the wing according to the recommendations of Corke (2003). This allows the tail to not be positioned in the wing wake if a stall where to occur, allowing the aircraft to recover from stall. Although in this position the wing downwash affects the tail, this is countered with a minor adjustment of the trim condition.
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4.5
Aerodynamics
After initial aircraft sizing, analysis is performed to select the appropriate aerodynamic parameters. 4.5.1
Airfoil Selection
An airfoil is selected to fulfill the requirements stated in Table 4.1. The “optimal” airfoil was chosen to not only reach these values but have the highest lift and lowest sustained drag coefficient from a range of angles of attack (-5° to 15°). Table 4.1 – Airfoil Selection Requirements Parameter
ࢻ
,ࢇ࢞
Thickness Stall
Requirement ~ -0.05 ~ 0.0 > 1.2 ≤ 12% > 6°
Rationale Reduce control surface workload for longitudinal trim Minimize trim drag Decrease control force requirements Increase control surface effectiveness Launch attainable without high lift devices Minimize thrust required for takeoff The plane is kept small, so a thin airfoil is warranted High stall so that maneuvers will not result in stalling
After an initial survey of airfoils, three airfoils are selected for further analysis and reviewS2091-10183, SD7032-099-88, and Eppler 234 airfoils. Figure 4.3 shows the lift and moment polars for these airfoils while Figure 4.4shows the lift versus drag polars. As can be seen below, the S2091-101-83 meets the lift coefficient criterion while having a moment coefficient closest to the desired value. The S2091-101-83also
1.5 1.3 1.1 0.9 0.7 0.5 0.3 0.1 -0.1
CM
CL
has a lower drag coefficient than the SD7032-099-88 while having a higher lift coefficient.
S2091-101-83 SD7032-099-88 EPPLER 334 -4 -2 0 2 4 6 8 10 12 14 Angle of Attack (deg)
0.02 0 -0.02 -0.04 -0.06 -0.08 -0.1 -0.12
S2091-101-83 SD7032-099-88 EPPLER 334
-4
-2
0
2 4 6 8 10 Angle of Attack (deg)
12
14
Figure 4.3 – Lift and Drag Polars for Final Three Airfoils
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0.035 S2091-101-83 SD7032-099-88 EPPLER 334
0.03
CD
0.025 0.02
0.015 0.01 0.005 0 -0.4
-0.2 0.2
0
0.2
0.4
CL
0.6
0.8
1
1.2
1.4
Figure 4.4 – Lift and Drag Polars for Final Three Airfoils 4.5.2
Aerodynamic Performance Predictions
Preliminary aerodynamic amic performance predictions are based on airfoil data and drag buildup. Mission 2 and 3 are determined to be the critical mission as the aircraft carries the heaviest payload. The drag buildup is equal for all missions as the external configuration of the aircraft will not change. The drag buildup is shown in Figure 4.5.. In an effort to reduce parasitic drag on the aircraft, the fuselage is blended into the wing section at the corners. This also improves the lift distribution over the wingspan. Landing Gear
Drag Buildup Component
CDO
Percent
Wing
0.0205
40%
Fuselage
0.0137
27%
V-Tail
0.0108
21%
Landing Gear
0.0062
12%
Total
0.0512
100%
Wing
V-Tail
Fuselage
Figure 4.5 – Aircraft Drag Buildup 4.6
Flight Mechanics Agent Blue’s stability is analyzed using a combination of Athena Vortex Lattice (AVL) and MATLAB
programming. AVL is a vortex lattice program that determines stability and control derivatives which are analyzed to ensure that Agent Blue is stable a and nd controllable in the anticipated flight conditions. 4.6.1
Stability Characteristics
Initial analysis is performed to ensure a statically stable aircraft and an acceptable static margin. By creating a diagram for each of the critical stability derivatives, the static stability of the aircraft is established. Agent Blue's static stability is analyzed using a NACA 0009 airfoil for the empennage. Each ruddervator measures a span of 10 in and chord of 7 in. The ruddervators are placed such that its
Page 23 of 59
quarter-chord is 25 in behind the wing’s quarter-chord at with each ruddervator placed at a 35° angle from
0.02 0.015 0.01 0.005 0 -0.005 -0.01 -0.015 -0.02
0.05
Roll Moment Yaw Moment
Pitch Moment
0 Coefficient Value
Coefficient Value
horizontal. Figure 4.6 shows that the aircraft is statically stable.
-0.05 -0.1 -0.15 -0.2 -0.25 -0.3
-15
-5 5 Sideslip Angle, β (deg)
15
-5
0 5 10 Angle of Attack, α (deg)
15
Figure 4.6 – Static Stability Charts Using this data, the neutral point and static margin of the aircraft are determined. The neutral point for this aircraft is at 0.436 of the chord length, giving the aircraft a static margin of approximately 18.6%. The maximum crosswind that Agent Blue can sustain for landing is 22.5 ft/s, which exceeds the predicted maximum crosswind of 19.1 ft/s for the competition. Once the static stability is established, an analysis on dynamic stability is also performed. Stability derivatives are calculated for both longitudinal and lateral motion and are used to predict the five aircraft modes. Figure 4.7 displays the real response (η) and the damped frequency/imaginary response (ωD). 15 Aircraft Mode Eigenvalues Mode
10
η
ωN
ωD
ζ
t1/2
Short Period
-9.18
14.13
10.73
0.65
0.08
Phugoid
-0.06
0.55
0.55
0.11
11.17
Roll
-20.40
–
–
–
0.03
0 iωD -5
Dutch Roll
-1.03
6.53
6.45
0.16
0.67
-10
Spiral
0.01
–
–
–
48.52
-15
5
Short Period Phugoid Dutch Roll Spiral
-10
-8
-6
η
-4
-2
0
Figure 4.7 – Aircraft Mode Eigenvalues From this data, only the spiral mode is not in the left half plane. This is a typical flight dynamic quality, and the pilot should be able to correct for the spiral instability as it has a doubling time of 46.52 seconds. All other aircraft modes are stable and have reasonable damping ratios for this type of aircraft. 4.6.2
Control Surface Sizing
The control surfaces are sized using a market survey previous DBF aircrafts. The measurements are designed to keep the plane steady under potential unbalanced loads while staying small to minimize
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weight and reduce manufacturing difficulty. A ruddervator sizing of 35% of the chord with ±20˚ deflection is used for the entire length of the empennage. At 25% of the chord, each aileron runs 60% of the halfspan of the wings. 4.7
Propulsion System Propulsion system design focuses on optimization of battery, motor, gearbox, and propeller
combinations. The final design is selected based on both experimental and theoretical data. Missions 2 and 3 are the critical mission for the propulsion system design as is has the maximum payload weight. Table 4.2 displays the final propulsion system options and total predicted flight score which is the sum of the three missions divided by the square root of the total aircraft weight with the propulsion system. Table 4.2 – Motor and Propeller Selection Propulsion System Mission Motor (Gear Ratio) Batteries 2 Prop NEU 1105/3Y (4.4:1) 13 Elite 1500 11x6 MEGA ACn16/15/3 13Elite 1500 11x7 (4.4:1) NEU 1105/3.5Y (4.4:1) 10 Elite 2000 13x8 Hacker A2010 Elite 2000 13x7 6XL(4.4:1)
4.7.1
Performance Parameters Weight (lb)
Flight Scores
1.021
2.90
1.068
2.80
1.103
2.73
1.082
2.77
Motor and Gearbox
The Castle Creations NEU 1105/1.5Y with Maxon 4.4:1 gear box is the optimal motor and gearbox configuration. Analysis on a static test stand and MotoCalc calculations prove that the selected motor is capable of producing the thrust required to for hand launch. This motor has the highest predicted flight score based on weight and flight times for Mission 1. 4.7.2
Propeller
MotoCalc and static thrust tests are utilized to select a propeller that is most suitable to complete all three missions. A maximum diameter of 15 inches is set to comply with wing and ground clearance. Propeller diameter is varied from 8 to 15 inches and pitch is varied from 4 to 10 inches per revolution. The predicted flight score is recorded for each propeller. The propellers producing the highest flight score for Missions 2 and 3 for each motor is recorded in Table 4.2. The optimal propeller selection has a diameter of 11 inches and a pitch of 6 inches per revolution for Missions 2 and 3. An optimal propeller selection for Mission 1 has a diameter of 10 inches and a pitch of 6 inches. 4.7.3
Battery
Battery capacity and number of cells is optimized using thrust and current curves obtained from MotoCalc. Battery packs are limited to 13 cells in accordance with ESC availability and weight. Batteries included in the analysis are the Elite 1500, 1700, 2000, and cells. A battery pack consisting of 13 Elite 1500 cells is the lightest pack capable of completing all 3 missions, weighing 0.655 lb. Page 25 of 59
4.8
Avionics Preliminary avionics system design focuses on the electronic speed controller (ESC), and servo
analysis and selection. 4.8.1
Electronic Speed Controller
The ESC is critical for the control of an aircraft as it allows for variable power to be supplied by the batteries, controlling the thrust generated by the motor. The current and voltage supplied by the battery pack must not exceed the limits for each ESC as the ESC could be damaged during flight. A programmable ESC is desired because it allows the system to run at peak efficiency at all times. Table 4.3 displays the critical parameters for each ESC that meets the design constraints. The Castle Creations Phoenix 25 is selected as it has a light weight of 0.6 ounces. Table 4.3 - Electronic Speed Controller Selection Speed Controller Weight (oz) Max Voltage (V) Max Current (A) CC Phoenix Ice Lite 50 0.8 25 50 CC Phoenix 25 0.6 19.2 25 Great Planes Silver Series 25 0.92 20 25 Flight Power 25 0.78 20 25 4.8.2
Servos
Torque, weight, and cost are the critical parameters in servo selection. XFOIL is used to determine the necessary torque required to provide adequate deflection of control surfaces. Control surfaces requiring similar torque all use the same servos for interchangeability. Table 4.4 displays the servos considered for each set of control surfaces. Servo Hitec HS-55 Hitec HS-65 Hobbico CS-12 Spektrum 7.5G
Table 4.4 – Servo Selection Torque (oz-in) Speed (sec/60°) Ailerons and Ruddervators 18 0.14 31 0.11 36 0.11 19 0.11
Weight (oz) 0.28 0.39 0.7 0.26
The Hitec HS-65 servos are selected as the aileron and ruddervator servos since they are lightweight and have high torque value which is required for adequate control. These servos also have no cost or delivery time as they are inherited from previous UF DBF teams. 4.9
Structures The preliminary structural layout is designed to withstand expected flight loads and maintain the
desired aircraft shape. Weight is minimized by optimizing the structural critical design parameters. 4.9.1
Layout Studies
The layout studies yield a basic two spar configuration. The main spar is located at the 25% chord and provides most of the structural support. The aft spar is located at the 75% chord and is included to provide moment stability and aileron redundancy. Several critical load paths are identified, and the
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internal structure is modified to withstand these loads. The major structural components are produced in order to transmit the wing, g, motor, and landing loads to the main bulkheads. Figure 4.8 identifies the primary spars and bulkheads required for a sufficient aircraft structure.
Figure 4.8 – Structural Layouts 4.9.2
Materials Selection
A balsa frame is selected for the inner structure of the aircraft wing and tail due to its ease of manufacturing and repair. A carbon fiber tube is purchased for the wing spars due to its light weight and strong resistance to bending moments. Material properties of both the balsa and carbon fiber tubing that are used for the aircraft are shown below in Table 4.5 . These properties are used for load analysis of the aircraft. Table 4.5 – Wing and Tail Material Selection Material
4.9.3
Carbon Fiber Tubing 3
Balsa 3
Density
0.058 (lb/in )
0.00798 (lb/in )
Elasticity Modulus (longitudinal)
20.0 Msi
991.8 Msi
Shear Modulus
1.03 Msi
24.36 Msi
Yield Strength (longitudinal)
210 ksi
2987 psi
Shear Strength
13.5 ksi
653 ksi
Critical Load Analysis
Critical load analysis is conducted to determine the critical flight loads experienced by Agent Blue. The analysis is conducted for the critical conditions experienced during Mission 3 which are an estimated flight weight of 4.75 lb with h a maximum wing moment of 59.375 in-lb. An initial model for a wing structure Page 27 of 59
cross section is presented in Figure 4.9.The simplified model assumes an average camber across the chord of the airfoil. This creates a constant radius for stringer placement from the neutral axis of the cross section. Additionally, stringer distribution is assumed symmetric across the chord line, regardless of quantity, such that the neutral axis does not change. The material used for stringers is square stock balsa wood with sides of 0.125 in with the properties mentioned in the Materials Selection section.
Figure 4.9 – Simplified Wing Structure Model
For each stringer and spar combination, the mass of the model must be calculated according to Eq. 4.7. As shown in this equation, mass increases linearly with the number of stringers in the wing and quadratically with the increase in spar diameter. ݉ = ݉ ∗ ݕݐݍ௦௧ + ݉௦ = ߩ ∗ ݕݐݍ௦௧ ܾℎ݈ +
ߨߩ௦ ݈ ሺܱ ܦଶ − ܦܫଶ ሻ 4
Eq. 4.7
As the mass equation is a function of the spar and stringer, a surface plot can be created using these elements as input parameters. The range of spar diameters for consideration is 0.125 in to 0.750 in, and the quantity of stringers ranges from 0 to 14.The result of the surface plot is shown in Figure 4.10.
Page 28 of 59
0.2
Mass
0.15
0.1
0.05
0 1 0.5 OD
0
0
10
5
15
Nstr
Figure 4.10– Mass Plot of Spar Size and Stringer Quantity From the plot, it can be seen that the combination of small diameter spar and low stringer quantity yields the lowest mass wing structure. During optimization of the design, the stresses incurred for each given configuration must be considered in order to establish a range of valid combinations on the plot. The calculation of maximum stress requires a maximum in both bending moment and sample radius. The maximum bending moment in the beam occurs at the center of the wingspan, implying that the maximum stresses occur at the upper and lower surfaces of the beam at the center of the plane. Once the optimization was completed with a factor of safety of 1.2 added it was found that the final wing had a spar with an outer diameter of 0.5 in and 10 stringers spaced equally along the wing. 4.9.4
Landing Gear
The location of the landing gear is limited by the fuselage and wing location.. Accounting for the weight of the loaded aircraft and potential asymmetric loads from an abnormal landing, a main gear width of 5 in is determined to be sufficient for landing. A straight steel wire landing gear ran through the fuselage is selected for the main landing gear because the gear is easily mounted without adding unnecessary weight to the aircraft. A replaceable layer of Kevlar is added to the rear of the fuselage as a skid. The layout of the main landing gear is shown in Figure 4.11.
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Figure 4.11 – Landing Gear Layout 4.9.5
Motor Loads
Missions 2 and 3 require the greatest thrust out of all mission and is considered the critical case for the motor mount bulkhead. ead. The maximum motor lloads are 2.5 lb forward from thrust, 0.625 lb vertically from a 2.5 g turn, and 5.53 in-lb torque. The motor bulkhead is carbon fiber and the motor is supported by a balsa structure. The he motor shaft protrudes through the carbon fiber motor mount and is secured with three machine screws. The motor is shown mounted on the aircraft in Figure 4.12.
Figure 4.12 – Motor Mount Solution
Page 30 of 59
4.10 Aircraft Mission Performance Predictions Performance estimations for each of the three mission profiles are shown in Table 4.6 below. These performance predictions are validated through flight testing. The critical performance predictions are for Mission 3, which requires the heaviest payload and a total takeoff weight of approximately 4.75 lb. As 2
seen below this weight gives a wing loading of approximately 22 oz/ft which is below the maximum the 2
team found acceptable of 25 oz/ft . The reasoning behind this is as wing loading increases so does stall speed and through research it was found the maximum speed an individual could launch an aircraft 2
without external assistance is roughly 25 mph. The team found that at a wing loading of 25 oz/ft the stall speed was above this mark thus reducing the maximum allowable loading. The score estimates for each mission are based on the predicted best score parameters found in the Score Optimization Analysis section in the Conceptual Design. The estimated total flight score is 4.42 out of a possible 6. Table 4.6 – Preliminary Mission Performance Predictions
(L/D)Max Takeoff Weight (lb) 2 W/S (oz/ft )
2.50 11.6
Mission 2 (Steel Bar) 1.19 0.0512 9.87 4.75 22.0
T/W
0.76
0.48
0.48
Cruise Speed (ft/s) Stall Speed (ft/s) Steel Bar Weight(lb) Number of Golf Balls
61.60 24.93 – – 1.00
54.26 33.7 2.25 – 1.42
54.27 33.7 – 22 2.00
Performance Parameter ࡸࡹࢇ࢞ ࡰ
Estimated Score
Mission 1 (Empty)
Mission 3 (Golf Balls)
4.75 22.0
5. DETAIL DESIGN The detailed design shows all the dimensional parameters as well as the full system level integration and optimization of the final design aircraft. The electrical components are selected based on optimized system performance. Solutions for propulsion cooling system, payload restraints, motor, assembly, and landing gear are developed based on design requirements and improvements. The total system weight is minimized with structural designs. Flight and mission performance summaries are revised to reflect all improvements made during the detail design process. 5.1
Dimensional Parameters System level integration and optimization begins once the preliminary design is complete. Table 5.1
provides primary dimensions and performance parameters for the system.
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Table 5.1 – Final Aircraft Dimensional Parameters Fuselage Length(in) Width (in)
Ruddervators 18 4.25
Height(in) 5.00 Landing Gear Width(in) 6.50 Aircraft Weight and Balance Airframe (lb) Propulsion System (lb) Control Systems (lb)
1.412 0.963 0.241
Max Payload (lb) Max Empty Weight (lb) Max Gross Weight (lb) CG Location (in)
2.250 2.616 4.844 9.5
5.2
Airfoil Span (in)
NACA 0009 10
Chord (in) 2 Area (in ) Aspect Ratio 2 Ruddervator Area (in ) Wing Airfoil
7 70 1.43 24.5 S 2091-101-83
Span (in) 2 Area (in ) Aspect Ratio Chord (in) 2 Aileron Area (in )
50 500 6.0 10.00 85
Electrical Component Selection Lightweight electrical systems are selected to reduce the total system weight. A total of four servos
are necessary for the flight control mechanisms. Table 5.2 contains a complete list of Agent Blue’s electrical components. The location of each component is detailed in the Drawing Package. Table 5.2 – Electrical System Components Component Battery Pack Propeller Mission 1 Propeller Mission 2 & 3 Motor (Gear Ratio)
5.3
Selection 13 Elite 1500 Graupner 10x6 Graupner 11x6 NEU 1105/3Y (4.4:1)
Speed Controller Transmitter Transmitter Battery Receiver Receiver Battery Aileron Servos
CC Phoenix 25 Futaba 6EXP 8 KAN 400 CC Berg 7P 5 KAN 400 Qty 2: HS-65
Ruddervator Servos
Qty 2: HS-65
Aircraft Component Weight and CG Buildup The weight and balance are essential components of the aircraft's flight. The airplane weight is taken
into consideration as the overall team score relies upon the weight. Each component of the aircraft is weighed and the distance from the nose of the airplane to the center of gravity of each component is determined. The sum of the distances multiplied by the component’s weight is calculated and then divided by the overall weight yielding the CG of the aircraft.
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Table 5.3 shows the payload weights and other essential data. The aircraft is designed so that the CG is at the wing’s quarter-chord for all missions. Table 5.3 – Component Weight and CG Buildup Component
Weight (oz) Location of CG w.r.t Nose (in) Avionics
Moment (oz in)
Aileron Servos (x2)
0.78
9.5
7.41
Ruddervator Servos (x2) Avionics Battery Receiver Speed Controller
0.78 1.41 0.28 0.60
35.5 1.75 1.75 1.15
27.69 2.47 0.49 0.69
9
22.05
Landing Gear Main Landing Gear (w/Wheels) Tail Landing Gear
2.45
0.40 36.5 Fuselage, Empennage, and Wings
Wing Fuselage Empennage Boom
8.50 7.65 2.15 1.65
Motor (w/prop) Batteries
5.25 10.48
14.60
10 9 36.25 23.5
85.00 68.85 77.94 38.78
0.35 5.25
1.84 55.02
Steel Bar 22 Golf Balls
36.00 9.50 35.64 9.50 Mission Configuration
313.5 307.8
Mission 1 (Empty) Mission 2 (Steel Bar) Mission 3 (20 Golf Balls)
42.38 78.38 78.02
9.50 9.50 9.50
Propulsion
Payloads
5.4
CG Location (no payload) CG Location (with payload) CG Location (with payload)
Cooling System Design A fuselage without ventilation can trap in heat, causing overheating of the motor, battery pack, and
ESC during flight. Static thrust testing as well as MotoCalc data reveals that the batteries, when kept out of the wake of the propeller, reaching a temperature of 200˚F. A cooling system is designed to keep the batteries from overheating and allow the propulsion system to operate at maximum efficiency. Air ventilation routes are designed inside the fuselage structure to provide cooling to the motor, ESC, and batteries as shown in Figure 5.1. Air vent openings are situated just behind the motor mount and a turbo spinner with a hollow center is used instead of a traditional spinner for additional airflow to all electrical components. Air flows in these openings, through the fuselage, and out openings in the back of the aircraft cooling all electrical components in flight.
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Figure 5.1 – System Cooling Design 5.5
Structures Weight reduction is the primary objective of the structural design. Structural optimization methods
are used to identify low stress areas of the aircraft structure where holes can be cut to lower weight. Ribs are distributed so as to ensure the proper shape of the aircraft is maintained while minimizing weight. 5.5.1
Wing Assembly Solution
The conceptual wing assembly mechanism incorporates two carbon fiber rods attached to the wings. These rods pass are mounted in the wings and pass into larger carbon fiber rods for a friction fitting. In addition to the friction fitting the wings are held in place by a wire clip placed between the forward and aft booms. Figure 5.2 shows the wing assembly solution.
Figure 5.2 – Wing Assembly Solution 5.5.2
Tail Assembly Solution
The tail assembly mechanism involves a telescoping carbon fiber rod serving as the tail boom and the fuselage. The rod slides into a channel along the fuselage and is held in place by a quick-release button. Kevlar string is wrapped around the opening in the carbon fiber rod to reinforce and prevent cracking. The empennage must clip on in addition to the boom sliding out to fit in the storage container. This is achieved by placing two release buckles on the top portion of the tail boom. These buckles allow
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the tail to release for storage torage and then clipped on during assembly for flight. The tail is kept in place during flight by a pin placed through the boom and ba base of the empennage. 5.5.3
Steel Bar Restraint Solution The fuselage of the aircraft is intended to carry two different payloads. The first payload is the
steel bar and in order to accommodate this payload safely a Velcro restraint system was developed. A smalll raised surface was installed in the fuselage using epoxy and then slits placed in it. Velcro straps are passed through these slits in both the forward/aft and left/right directions. These straps are wrapped around the steel bar to secure it during flight flight. Figure 5.3 shows the steel bar restraint solution mockup.
Figure 5.3 – Brick Restraint Solution 5.5.4
Golf Ball Restraint Solution According to Mission 3 parameters, the payload compa compartment rtment must be capable of carrying a team
selected number of golf balls. Therefore it was found that egg cartons where ideal to secure the golf balls as represented in Figure 5.4.
Figure 5.4 – Golf Ball Restraint Solution Egg cartons are chosen for the payload system due to their high availability, simplicity, and light weight. The foam provides a stable platform for the golf balls and allows for a balanced layout. 5.6
Total System Weight Scoring analysis identifies the total system weight as one of the two largest contributing factors to the
overall flight score. The weight of each aircraft component is recorded in Figure 5.5.. The system weight analysis shows that the majority of the weight comes from the fuselage, batteries, and wing. wing Battery
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weight is at its minimum and could not be reduced further, therefore, the team focuses on minimizing the weight of the airframe. Component Propeller Receiver Avionics Batteries Prop. System
Weight (oz)
Percent
1.1 0.28
2.60% 0.66%
1.41
3.33%
4.75
11.21%
Landing Gear Wing Servos Boom Tail Fuselage
2.85 8.5 1.56 1.65 2.15 7.65
6.72% 20.06% 3.68% 3.89% 5.07% 18.05%
Batteries Total
10.48 42.38
24.73% 100.00%
Propeller Receiver
Batteries
Avionics Batteries
Fuselage
Prop. System
Tail Wing
Boom Servos
Landing Gear
Figure 5.5 – Total System Weight Breakdown 5.7
Flight Performance Summary Table 5.4 details flight and mission performance parameters. Applicable final predictions are
compared to actual test results at the completion of flight testing. Table 5.4 – Flight Performance Parameters Performance (L/D)Max T/W Takeoff Weight (lb) 2 W/S (lb/ft ) Max Speed (ft/s) Stall Speed (ft/s) Max Turn Rate (deg/s) Max Rate of Climb (ft/min) Predicted Flight Score Predicted Final Score
Mission 1 ((Empty)
Mission 2 (Steel Bar)
Mission 3 (Golf Balls) Balls
9.85 0.71 2.65 0.7625 60.14 26.4 176
0.46 4.89 1.41 55.70 35.2 84.9
0.46 4.88 1.41 55.70 35.2 84.9
885 1.0
378 1.38 2.69
378 2.0
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5.8
Mission Performance Summary Table 5.5 displays the estimated battery capacity remaining, current draw, and time allotted for each
of the three missions. Table 5.5 – Mission Performance Predictions Flight Leg Hand Launch Climb Cruise to First Turn Cruise (6 laps) 180° Turn (x12) 360° Turn (x6) Error And Decent Total Available Energy Excess Hand Launch Climb Cruise to First Turn Cruise (3 laps) 180° Turn (x6) 360° Turn (x3) Error and Descent Total Available Energy Excess Hand Launch Climb Cruise to First Turn Cruise (3 laps) 180° Turn (x6) 360° Turn (x3) Descent Total Available Energy Excess 5.9
Velocity (ft/s) Time (sec) Mission 1 0-44.0 51.4 58.7 58.7 58.7 51.4 – – – –
1.0 5.5 3.6 195.9 12.3 10.8 7.0 236.1 – – Mission 2
Distance (ft)
Capacity (mAh)
5 281.4 119.8 11500 720 553.8
4.5 19.7 11 598.6 34.2 30.0
– 13273.8 – –
40.0 738 1300 562
0-40.0
1.0
5
5.4
46.0 51.3 51.3 51.3 46.0 –
10.0 0.7 107.2 12.9 11.0 7.0
460 35 5500 654 503.1 –
42.2 2.6 396.0 47.66 46.44 40.0
149.8 – – Mission 3
7157.1 – –
574.90 1300 725.1
– – – 0-40.0 46.0
1.0 10.0
5 460
5.4 42.2
51.3 51.3 51.3 46.0 – –
0.7 107.2 12.9 11.0 7.0 149.8
35 5500 654 503.1 – 7157.1
2.6 396.0 47.66 46.44 40.0 574.90
– –
– –
– –
1300 725.1
Drawing Package The drawing package includes dimensioned aircraft 3-view, structural arrangement, systems layout,
payload accommodation, and stored configuration drawings. Page 37 of 59
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6. M ANUFACTURING PLAN AND PROCESSES The manufacturing of Agent Blue adheres to a detailed plan to ensure the aircraft is constructed in a timely manner. A schedule is followed to provide the necessary time for both the construction and testing of major components. Several manufacturing techniques and materials are analyzed prior to construction. 6.1
Manufacturing Figures of Merit A conventional airplane can be broken into two main categories for manufacturing: fuselage and
lifting surfaces. Manufacturing processes are selected based on team experience, manufacturing time, and financial constraints. Each method is evaluated using FOMs which are weighted according to their importance. •
Weight: Weight varies depending on the manufacturing material and process. A low weight corresponds to a short flight transition period and a high flight score.
•
Strength: The plane must have the structural strength required to tolerate internal payloads, wing deflection, and have a high impact resistance. Ideal manufacturing techniques and materials result in using less or lighter weight materials, both reducing the overall weight.
•
Cost: Low cost methods and materials are necessary to remain on budget.
•
Ease to build: Aircraft components must be built and rebuilt quickly to allow time for tests and modifications. Simplicity of construction and technical skill advance this goal.
•
Maintenance: Low maintenance construction methods are desired to lessen repair time.
Several manufacturing techniques are described and analyzed below. 6.1.1
Fuselage
Agent Blue’s fuselage is a simple rectangular structure characterized by an internal compartment for payloads and housing for all critical electrical components. Weight reduction is a primary objective of the design since the fuselage is a major element of the aircraft. Methods deemed feasible for fuselage construction are listed below. •
Balsa Frame: A balsa frame is constructed for the fuselage and covered in MicroLite. Balsa beams and walls are joined together with adhesives such as CA and epoxy. The balsa frame is precision cut by a CNC machine to preserve the proper curvature of the fuselage.
•
Foam and Fiberglass: Foam sections are cut out using a hotwire cutter. Fiberglass and epoxy are applied to the foam and allowed to cure. A vacuum bag is placed around the fuselage during the curing process to ensure a smooth finish. This method provides high strength and weight.
•
Composite: A foam fuselage is first constructed as a male mold. Layers of pre-preg composite material such as carbon fiber or Kevlar are layered on top of Teflon cloth over the male mold. The mold is placed in a vacuum bag, and the composite is left to cure in an oven. Once hardened, the composite is removed from the mold, and the Teflon cloth is peeled off. Bulkheads are added where necessary to improve rigidity. Table 6.1 displays WDM for fuselage construction.
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Table 6.1 – Fuselage Construction Method Weighted Decision Matrix Fuselage Construction FOM
W.F.
Balsa Frame
Foam and Fiberglass
Composite
Weight
0.30
1
-1
1
Strength
0.40
0
1
1
Maintenance
0.15
0
-1
0
Ease to Build
0.15
0
0
-1
Total
1.00
0.30
-0.05
0.55
The composite fuselage provides an ideal balance between the FOMs. It is the lightest alternative and provides the highest strength to weight in case of adverse landings. A balsa fuselage is low cost and requires low maintenance, but does not provide the same strength as the composite structure. 6.1.2
Lifting Surfaces
Similar to the fuselage, the weight of the wing and tail structure is the critical factor of the FOM. A high strength to weight ratio is desired for the vehicle’s lifting surfaces. The methods considered for manufacturing Blue Agent’s lifting surfaces are detailed below. •
Balsa Frame: Ribs from the root to the tip of the wing and tail are cut using a CNC machine. The ribs are connected by the main and aft wing spars using CA. Stringers are added to maintain the proper airfoil shapes of the lifting surfaces. Once finished, the balsa frame is covered with MicroLite. The precision of the wing ribs are not a concern as the balsa is cut using a CNC machine.
•
Foam: Blocks of foam are shaped into the wing and tail with use of a hotwire cutter and sanding equipment. Foam lifting surfaces require the trailing edges to be replaced by balsa as they are too thin to stand alone. Great care must be taken to properly align airfoil templates to account for the airfoil shape.
•
Foam and Fiberglass: Shaped foam lifting surfaces are sheeted with fiberglass using epoxy and allowed to cure. This method does not require a replacement of the trailing edge but does have the same issue as foam wings in aligning airfoil templates.
•
Foam and Balsa: Foam lifting surfaces have thin balsa sheets carefully shaped and adhered to their surfaces to provide increased strength. This method has a high maintenance cost as both the balsa sheeting and foam core may be damaged in a crash.
Table 6.2 displays the WDM for lifting surface construction methods.
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Table 6.2 – Lifting Surface Construction Method Weighted Decision Matrix Lifting Surface Construction FOM
W.F.
Balsa Frame
Balsa and Foam
Foam
Foam and Fiberglass
Weight
0.30
1
0
0
0
Strength
0.25
1
1
0
1
Cost
0.10
1
0
1
-1
Maintenance
0.15
0
-1
1
1
Ease to Build
0.20
1
-1
0
0
Total
1.00
0.85
-0.10
0.25
0.30
The balsa frame is determined to provide the best weight, strength, cost, ease to build, and maintenance characteristics out of the manu manufacturing facturing methods evaluated. MicroLite skin is used to minimize weight and maintain the proper airfoil shape over the entire lifting surface. The foam and fiberglass construction method has high strength and requires low maintenance but is heavier than desired and has a high cost. 6.2
Cost Analysis Agent Blue is funded by University of Florida’s AIAA chapter and corporate sponsors. A breakdown
of total expenses to date adds to a budget of $944.87. It should be noted that items inherited from previous University ty of Florida DBF teams, such as servos and wheels, are not included in the costs. The propulsion system has the highest cost as it is optimized for the current and no materials can be obtained from previous UF DBF teams.
Category
Cost
Percent
Propulsion
$373.57
39.6%
Avionics
$104.99
11.1%
Construction
$200.60
21.2%
Tools
$265.71
28.1%
Total
$944.87
100%
Tools Propulsion
Construction Avionics Figure 6.1 – Cost Breakdown
6.3
Manufacturing Schedule The manufacturing schedule allot allots s sufficient time to construct and test individual components of the
aircraft, allowing the team to work efficiently. The Gantt chart in Figure 6.2 tracks the progress of both the prototype and final aircraft. Agentt Blue’s electronics are installed during the aircraft assembly.
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Figure 6.2 â&#x20AC;&#x201C; Manufacturing Schedule 6.4
Prototype A prototype aircraft is constructed to verify the design of the aircraft through flight testing. The
prototype construction process is selected to provide a fast, low-cost assembly of the aircraft. The materials and techniques used in the prototype construction will reflect those used in the final plane. This is to ensure that the construction of the final plane can be accomplished without complications. The fuselage of the prototype plane is constructed by laying up pre-preg carbon fiber over a cylindrical foam mold. The fuselage was attached to the wing by means of release buckles. The carbon fiber tail boom has telescoping pin releases that allow the tail location to be adjusted during flight testing. This feature also allows for easy storage. A balsa structure is used to construct the lifting surfaces on the prototype. A laser cutter is used to cut the rib structure along the span of the wing. The trailing edge of the wing is excluded from the bulkheads in order to implement ruddervators. Carbon fiber spars are inserted into the wings and the tail to help rigidity. All major elements are assembled and electronics installed to complete the prototype aircraft. Figure 6.3 shows parts of the manufacturing process for the prototype aircraft.
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Figure 6.3 â&#x20AC;&#x201C; Prototype Construction of Wing, Fuselage, and Empennage 6.5
Aircraft Manufacturing Design aspects are confirmed when the prototype aircraft is built and tested. The final aircraft is
constructed using the methods established by the WDMs. The manufacturing processes used to build the final competition are similar to the prototype to ensure a reliable aircraft structure is produced. 6.5.1
Tooling
A laser cutter is used to cut out the balsa frame components of the aircraft. The components contain holes and notches to help align the pieces for assembly. An iron and a heat gun are required to bond MicroLite to the balsa frame and shrink it to form. A composites lay-up lab provides all of the necessary equipment to construct the carbon fiber fuselage. Other tools are used to refine components through physical means (i.e. sander, drill, saw blade). 6.5.2
Airframe
The airframe consists of two major parts: fuselage and lifting surfaces. Balsa bulkheads are required in high stress areas of the inboard wing section to increase strength. The primary bulkheads are located where the fuselage release buckles attach to the wing. CA adhesive attaches the balsa components together for both the tail and wings. Epoxy is used to secure the main wing spar to the wing ribs. 6.5.3
Wing Joint
The wing joint consists of several different components: carbon fiber spars, wire clip , and the electronic contacts. The wing joint is fabricated as outlined in the Detailed Design. The two spar extensions slide through the inboard wing section while the wire clip springs into its position locking the wing in place. The electrical contacts are connected by lifting the magnetic cover on the inboard section of the wing and pressing the respective contacts together.
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6.5.4
Control Surfaces
Control surfaces are constructed with a balsa framework covered by MicroLite much like the wing design. The outer ribs of each control surface are manufactured out of 1/8 in balsa, while inner ribs are 1/16 in balsa. CA hinges connect the control surfaces to the airframe. 6.5.5
Landing Gear
The main landing gear is developed to have high flexibility and light weight to compliment a hand launch. The landing gear is comprised of a small diameter steel wire conformed to its desired shape. Small lightweight foam wheels are fitted to the wire to accommodate landing on a hard surface. 6.5.6
Fuselage Restraints
The fuselage is restrained with release buckles for rapid assembly. Due to size constraints, the fuselage cannot be permanently attached to the main wing section. Simple attachment mechanism such as release buckles allow a single person to prepare the aircraft for launch within the time limit. 7. TESTING PLAN Testing of the system and subsystems is performed to ensure functionality and reliability. This includes both strength and performance testing to verify the design of the entire aircraft system. Individual components of the prototype plane are analyzed prior to the construction of the final aircraft. 7.1
Subsystems Testing the aircraftâ&#x20AC;&#x2122;s subsystems is important to obtain the best performance results, analysis, and
validation for the design system. Each individual component test is used to improve the final aircraft. 7.1.1
Static Thrust
The propulsion system is first tested on a static thrust test stand that is assembled by mounting the motor on a Tahmazo thrust stand on a secure mounting bracket as shown in Figure 7.1.The entire propulsion system is mounted on this assembly, and the motor is tested at 100% throttle. Battery and motor temperatures are monitored for safety using a digital infrared thermometer held to both components. Thrust is recorded from the scale at 15 s intervals while voltage and current are measured using a Castle Creations Phoenix Ice Lite 50 ESC and a Castle Link.
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Figure 7.1 – Static Test Stand Configuration The test card in Table 7.1 is followed to ensure team member safety and accurate testing. Table 7.1 – Static Thrust Test Card X
Procedure
7.1.2
X
Procedure
1. Secure motor to test stand 2. Wire motor, ESC, and receiver 3. Verify all connections are secure
5. Set throttle to 0% and connect batteries 6. Increase throttle to 100% 7. Record data until the battery dies
4. Turn on the transmitter and receiver
8. Set throttle to 0% and unplug batteries
Dynamic Thrust Testing
The propulsion system was also tested to see how well it would perform in dynamic thrust tests. The motor was mounted in a wind tunnel and the motor ran at various throttle positions for this test. The stand used was the same test stand used for static testing. The motor is run at speeds from stall to cruise and data is collected on current draw and battery endurance for both prop selections. The mounting setup is shown below in Figure 7.2.
Figure 7.2 – Dynamic Thrust Test Stand The test card in Table 7.2 details the procedure of dynamic thrust testing.
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Table 7.2 â&#x20AC;&#x201C; Dynamic Thrust Test Card X
Procedure
X
Procedure
1. Secure motor to test stand 2. Wire motor, ESC, and receiver
6. Set wind tunnel speed. 7. Increase throttle to 100%
3. Verify all connections are secure 4. Turn on the transmitter and receiver 5. Set throttle to 0% and connect batteries
8. Record data 9. Increase wind tunnel speed 10. Repeat procedures 8 and 9 to cruise velocity
7.1.3
Materials
The material selection for the wing spar and tail boom is critical for the structural integrity of the aircraft. The selected materials are required to have high strength-to-weight ratios. Small beams are constructed from several materials for stiffness testing. The strength-to-weight ratios of these materials are calculated, and the best material is then selected for the aircraft. An Instron machine is used to subject the beam to three point bending. The beam is centered when place on the machine, and a force is placed at the center of the beam. The applied force is gradually increased until the beam fails. The stiffness of each beam is measured using the measured deflection and force. The strength-to-weight ratio of each material is calculated from the stiffness and weight of the test sample. Table 7.3 displays the test card used for material testing. Table 7.3 â&#x20AC;&#x201C; Materials Test Card X
7.1.4
Procedure
X
Procedure
1. Weigh the test sample
5. Place sample on Instron machine
2. Measure dimensions of sample 3. Mark the center of sample 4. Zero the Instron machine
6. Center sample on Instron machine 7. Run Instron machine 8. Save data
Vibration Test
After observation of aircraft at past competitions reaching wing and tail excitation frequencies during flight the team decided to find the excitation frequencies of the competition aircraft wing to better understand these phenomena. This will help the team avoid failure of the aircraft structure at competition. Ground vibration testing was performed on a balsa-carbon fiber wing representative of the final wing design. The wing was mounted on a soft support setup consisting of bungee cables to hold the weight of the wing, while a 2lb Baldor shaker was used to apply a point excitation to the structure at the root quarter chord as shown in Figure 7.3. A Polytech Laser Doppler Vibrometer (LDV) system was used as a noncontact method to measure the structural response of the wing. The system shines a laser onto the structure as it is vibrating under some excitation and measures the scattering of the light reflected back by the structure. The phase shift of the reflected light wave caused by the Doppler effect is proportional to the object's velocity and the beam wavelength and therefore may be used to generate the frequency response of the structure.
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Figure 7.3 –Vibration Testing Configuration Table 7.4 –Vibration Test Card X
Procedure
X
1. Suspend outboard portion of wing with bungee 2. Secure inboard portion of wing to the shaker plate
Procedure 4. Turn on the Baldor shaker 5. Record vibration response for range of frequencies
3. Align vibrometer with wing section 7.1.5
Payload Restraints
All payloads must be secured such that they do not move during flight. Tests are executed to ensure the payload restraint systems perform as desired. The golf ball restraints are loaded and the fuselage undergoes a shake test to determine if the restraints are adequate. The steel bar is then loaded and another shake test applied. The test card for payload restraint testing is shown in Table 7.5. Table 7.5 – Payload Restraint Test Card X
Procedure 1. Assemble the aircraft 2. Verify all connections are secure 3. Secure 22 golf balls inside fuselage 4. Close hatch and verify it is locked
7.2
X
Procedure 5. Apply a shake test and observe payload 6. Unload payload 7. Secure steel bar inside fuselage 8. Repeat procedures 4 through 6
System The entire system is tested once the subsystems tests are completed. The system testing verifies or
refutes predictions made during Agent Blue’s preliminary design process. 7.2.1
Flight Readiness
Flight readiness testing simulates the tech inspection portion of the competition. This testing guarantees the safety of all individuals for the remaining tests and prepares the Agent Blue team for a Page 51 of 59
smooth tech inspection at the competition, which allows the team to have the most possible flight attempts. Table 7.6 shows the test card for the flight readiness testing. Table 7.6 – Flight Readiness Test Card X
Procedure
X
Procedure
1. Assemble aircraft 2. Verify all components are secure 3. Verify propeller attachment integrity
7. Load payload combinations and mark CG 8. Wing tip test for heaviest load 9. Radio range check, motor off and motor on
4. Visual inspection of all electronic wiring 5. Verify controls move in the proper sense 6. Check integrity of the payload system
10. Radio fail-safe check 11. Taxi on pavement 12. Taxi on grass
Another important aspect of flight readiness is the aircraft assembly. The aircraft must be removed from the storage case and flight ready within five minutes. The test card shown in Table 7.7 is used to simulate competition conditions. The preparations guarantee that all aircraft components are secured in the case and ready for assembly when the team enters the staging area. Table 7.7 – Assembly Test Card X
Procedure 1. Verify all flight hardware is in the case 2. Start stopwatch and assemble aircraft 3. Stop stopwatch and record time
7.2.2
Flight Testing
Flight testing offers a final verification of an aircraft’s capabilities. The test card in Table 7.8 is used to check all of the aircraft components prior to flight and ensure the safety of both the aircraft and students observing the flight. Table 7.8 – Flight Testing Test Card X
Procedure
X
Procedure
1. Check the CG
11. Check all servo connections
2. Check the lateral balance 3. Check alignment of all surfaces 4. Verify full charge of propulsion battery pack 5. Verify full charge of receiver battery pack 6. Check attachment of control surfaces 7. Check screws on servo horns
12. Check for slop in the control surfaces 13. Check for inadvertent movement of servos 14. Listen for abnormal servo sounds 15. Verify proper movement of each control 16. Look for binding pushrods 17. Check rotation of propellers
8. Check main landing gear 9. Check skid 10. Check structural integrity of the aircraft
18. Range check (antenna down, 100 ft) 19. Check spotter and pilot readiness 20. Carry out flight profile
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Flight testing for Agent Blue begins with a maiden flight in which the pilot becomes familiar with and trims the aircraft. Once the airplane is trimmed, mission testing can commence. A flight schedule including the dates and goals of each test flight is outlined in Table 7.9. Table 7.9 – Flight Testing Schedule Flight
Date
1
2/12/2011
Empty aircraft, 4 180° turns and land
Hand Launch Feasibility Trim Airplane
2
2/16/2011
Fly Mission 1 course profile (empty)
Hand Launch Feasibility Record Lap Times
3
2/20/2011
Fly Mission 2 course profile (steel bar)
Hand Launch Feasibility Sufficient Battery Capacity
4
2/25/2011
Fly Mission 3 course profile (golf balls)
Hand Launch Feasibility Sufficient Battery Capacity
Mission Prototype Aircraft
Goals
Final Aircraft
7.3
5
3/15/2011
Empty aircraft, 4 180˚ turns and land
Hand Launch Feasibility Trim Aircraft
6
3/21/2011
Fly Mission 1 course profile (empty)
Time Aircraft Assembly Record Lap Times
7
3/27/2011
Fly Mission 2 course profile (steel bar)
Hand Launch Feasibility Sufficient Battery Capacity
8
3/30/2011
Fly Mission 3 course profile (golf ball)
Hand Launch Feasibility Sufficient Battery Capacity
Master Test Schedule Detail design testing begins at the subsystems level to ensure all components of the aircraft are
functioning properly. Upon completion of subsystems testing, the team can proceed to an entire system check and flight testing. The master test schedule displayed in Figure 7.4 makes effective use of the time remaining. Flight readiness testing and flight tests are scheduled for both the prototype and final aircrafts.
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Figure 7.4– Master Test Schedule 8. PERFORMANCE RESULTS Testing is a critical step in the design process. Performance results are compared to predictions formed during the preliminary design process. Design improvements are made based on the data collected from testing. 8.1
Subsystems The testing of subsystems is a critical period for the aircraft. Each subsystem must perform properly
to ensure suitable aircraft performance as a whole. 8.1.1
Static Thrust Testing
The designed propulsion system is tested for static thrust to aid in the approximation of flight endurance and determine theoretical data error. Figure 8.1 displays the predicted and measured static thrust for a Graupner 10×6 folding propeller and a Graupner 11×6 folding propeller. MotoCalc predicts
Thrust (lb)
static thrust and current as constant values over time and are represented as such in the figures. 4 3.5 3 2.5 2 1.5 1 0.5 0 0
11x6 Tested
11x6 Calculated
10x6 Tested
10x6 Calculated
50
100
150
200
250
300
Time (s) Figure 8.1 – Static Thrust Test
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The data shows that the tested static thrusts are much higher than the Motocalc estimations. This is good for the team as it allows the aircraft to perform better than previously anticipated. Table 8.1 shows the error in Motocalcs thrust, voltage, amps, and propeller RPM calculations versus the average of tested values. Table 8.1 â&#x20AC;&#x201C; Static Thrust Test Parameter
Tested Data
Calculated Data
Percent Error
Graupner 11x6 Thrust (lb)
3.01
2.25
25.3%
Voltage (V)
11.91
10.41
12.6%
Amperage (A)
17.7
19.3
9.04%
RPMs
7839
7656
2.33%
Graupner 10x6 Thrust (lb)
2.60
1.87
28.1%
Voltage (V)
13.11
11.3
13.8%
Amperage (A)
15.24
16.2
6.30%
RPMs
8759
8443
3.94%
As can be seen above the thrust generated was at least 25% above Motocalc's prediction. This along with all other data shows that Motocalc under predicted the motors performance at static. The Figure 8.2 displays the actual current draw and voltage draw by each propeller. Each test was completed
20
20 18 16 14 12 10 8 6 4 2 0
18 16 14 12
11x6 Amps 10x6 Amps 11x6 Voltage 10x6 Voltage 0
50
Voltage (V)
Amps (A)
with ample battery capacity to spare.
10 8 100
150
200
250
Time (s) Figure 8.2 â&#x20AC;&#x201C; Static Thrust Test Current and Voltage Draw
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8.1.2
Dynamic Thrust Testing
The propulsion system is tested in the wind tunnel to better understand the systems performance under flight conditions. The system was tested under full throttle conditions at various wind speeds to verify Motocalc's predictions at higher velocities. Each propeller was tested from 18-39 mph which is from
20 18 16 14 12 10 8 6 4 2 0
10000 9800 9600 11x6 Amps 10x6 Amps 11x6 RPMs 10x6 RPMs
9400 9200 9000 8800
Propeller RPMs
Amps (A)
the stall speed to cruise of Mission 1. Figure 8.3 shows current and motor RPMs for the data collected.
8600 8400 17.5
22.5
27.5 32.5 Air Speed (mph)
37.5
Figure 8.3 â&#x20AC;&#x201C;Dynamic Thrust Test Current Draw and RPMs After collecting current data anticipated battery endurance was measured by calculating the mAh/min used at a set airspeed then the total mAh with a factor of safety available by that number. It was anticipated that at full throttle the aircraft could fly 5.15 minutes at Mission 2 or 3 cruise speeds and 6.7 minutes at Mission 1 cruise speeds. Motocalc's predicted a Mission 1 full throttle time of 6.44 minutes at cruise and a full throttle time of 5.7 minutes for Missions 2 and 3. Giving an error of 4% for Mission 1 and 10% for Missions 2 and 3. 8.1.3
Materials
Material testing encompasses both the wing spar and tail boom materials. A small section of each material considered is placed on the Instron machine and tested until breaking. The test is designed to validate that each material is capable of holding under a 2.5G load using Mission 3 with a total weight of 12.5 lb. The results are displayed in Table 8.2. Table 8.2 â&#x20AC;&#x201C; Materials Selection Testing Results Material
Weight (lb)
Stiffness (lb/in)
Strength-to-Weight Ratio
Carbon Fiber/Poplar Carbon Fiber Rod Carbon Fiber/Balsa
0.0508 0.0278 0.0190
3415.0 3240.0 2180.9
67275 116547 115078
Carbon Fiber/Plywood
0.0419
3597.0
85792
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The best material for the tail boom and wing spar is determined to be the carbon fiber rod as the team predicted. The carbon fiber/balsa strength-to-weight ratio is close to that of carbon fiber rods and is actually a lighter material and would be adequate for use. However due to the manufacturability of the carbon fiber/balsa structure it was determined that carbon fiber rods where the safest material for use. 8.1.4
Vibration Testing
The wing proved difficult to test due to order of magnitude differences in the stiffness of the components of the wing. The carbon fiber spars were extremely stiff and experienced resonant behavior at frequencies between 1000-5000Hz, while the balsa wood assembly had multiple, larger magnitude resonances well below 750Hz. Therefore to ensure that both types of resonances could be measured, the final test was separated into two stages where a signal generator applied a sweep signal to a shaker from 100-2000Hz over 400ms and 1000-6000Hz over 1sec. The final results are shown in Figure 8.4. The first bending mode of the wood is clearly seen at 157 Hz, while substructure modes are present at 295Hz and 740Hz. The first bending mode of the carbon fiber spars is shown at 3896Hz as measured during the second test.
Figure 8.4 â&#x20AC;&#x201C; Composite Frequency Response and Mode Shapes Typically it is desired to prevent large magnitude, low (<100Hz) frequency resonances in the wing which could be excited during flight by a pilot moving a control surface or a wind gust at the resonant frequency of the wing. The final test reveals that the stiffness of the cylindrical carbon fiber spars gives the wing sufficient bending and torsional stiffness to prevent issues from the structural dynamics. The single spar forces the first bending mode of the wing over 100Hz, sufficiently preventing flutter from bending, while the 2-spar arrangement ensures significant torsional stiffness as long as the connection between the spars remains rigid. The only low frequencies of concern were generated by wing skin panel vibration where large deflections may cause aerodynamic inefficiency, but are unlikely to cause the plane to crash.
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8.1.5
Payload Restraints
The golf ball restraints are tested by loading 22 golf balls into the plane. A shake test is performed, observing golf ball movement. Flight requirements are met as all golf balls remain in place throughout the shake test. The steel bar restraints are then tested by securing the steel bar payload into the restraints and applying forces and moments in all 6 degrees of freedom. 8.2
System The aircraft must be consistently assembled under five minutes and have good flight qualities.
Extensive testing is performed to ensure acceptable assembly times. Flight testing is also performed to practice flying individual missions as well as to verify that the aircraft is stable and capable of completing all missions. 8.2.1
Flight Readiness
Pre-flight readiness requires the aircraft to pass tech inspection and be assembled in less than five minutes. The aircraft assembly time averages 3:15 minutes through 25 assembly tests. The aircraft passes the wing tip test for all possible payload combinations with nominal deflection. Finally, the aircraft is tested for hand-launch ability as well as the ability to land well on both grass for test flights and asphalt for competition. 8.2.2
Flight Testing
During the prototype aircraft’s maiden flight, a hand launch without payloads was performed. Agent Blue successfully completed several laps around the flying field. Several rolls and loops were performed to demonstrate maneuverability. During stall, the aircraft would nose up and then subsequently dive. Stall recovery was quick because the aircraft was unloaded. Finally, the pilot remarked that the ailerons at the wingtips lost effectiveness when under full downwards deflection. The pilot advised reducing the maximum downward aileron deflection by 5˚. The flight was considered a success and Figure 8.5 shows the Agent Blue just before launch on its maiden flight.
Figure 8.5 – Agent Blue's Maiden Flight
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9. REFERENCES Aerospace Design. (2010). Retrieved from Aerospaceweb.org: http://www.aerospaceweb.org/ AIAA. (2010, October 31). AIAA Design/Build/Fly Competition- 2010-2011 Rules. Retrieved from www.aiaadbf.org Airfoil Investigation Database. (2009). Retrieved from World of Krauss: http://worldofkrauss.com/ Anderson, J. D. (2007). Fundamentals of Aerodynamics (4th ed.). New York, NY: McGraw-Hill Companies, Inc. Anderson, J. D. (2008). Introduction to Flight (6th ed.). New York, NY: McGraw-Hill Companies, Inc. Corke, T. C. (2003). Design of Aircraft. Upper Saddle River, NJ: Pearson Education, Inc. Drela, M., & Youngren, H. (2006, April 29). AVL 3.26 User Primer. Retrieved from AVL: http://web.mit.edu/drela/Public/web/avl/ Drela, M., & Youngren, H. (2001, November 30). XFOIL 6.9 User Primer. Retrieved from XFOIL: http://web.mit.edu/drela/Public/web/xfoil/xfoil_doc.txt Ifju, P. (2010, October). Professor. (J. Cantrell, Interviewer) Laurence K. Loftin, J. (1985). Quest for Performance: The Evolution of Modern Aircraft. Washington, D.C.: NASA Scientific and Technical Information Branch. Lennon, A. (1996). R/C Model Aircraft Design. Ridgefield, CT: Air Age Media, Inc. Nelson, R. C. (1997). Flight Stability and Automatic Control. McGraw-Hill. Phillips, W. F. (2004). Mechanics of Flight. Hoboken, New Jersey: John Wiely & Sons, Inc. R. T. Compton, J. (1992, September). Fourteen Steps to a Clearly Written Technical Paper. Retrieved January 24, 2010, from http://www.philadelphia.edu.jo/pdf/r1.pdf Raymer, D. P. (2006). Aircraft Design: A Conceptual Approach. (4th Edition ed.). Reston, VA: AIAA. Reyes, C. (2009). RCadvisor's Model Airplane Design Made Easy. Albequerque, NM: RCadvisor.com. Tucson
International
Airport,
AZ.
(2010).
Retrieved
from
Weather
Underground:
www.wunderground.com Yechout, T. R., Morris, S. L., Bossert, D. E., & Hallgren, W. F. (2003). Introduction to Aircraft Flight Mechanics: Performance, Static Stability, Dynamic Stability, and Classical Feedback Control. Reston, Virginia: American Institute of Aeronautics and Astronautics, Inc.
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8
6
7
5
4
2
3
50.00 16.77
1
40.00 20.00
D
10.25
D
7.00
7.00
11째
4.50
17.00 C
C
2.50
25.00
10.00
B
B
10.00
10.00
University of Florida A
Design/Build/Fly 2010-2011
NOTE: DIMENSIONS IN INCHES PROPRIETARY AND CONFIDENTIAL THE INFORMATION CONTAINED IN THIS DRAWING IS THE SOLE PROPERTY OF THE UNIVERSITY OF FLORIDA DBF GROUP. ANY REPRODUCTION IN PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF THE UNIVERSITY OF FLORIDA DBF GROUP IS PROHIBITED.
6.50
SolidWorks Student License Academic Use Only 8 7
6
5
4
3
Project Title
J. Lasater
Agent Blue
Dr. R. Lind
Aircraft 3-View
Chief Engineer
Advisor
Drawn By
K. Rausch 2
SCALE: 1:8
2/24/2011
SHEET 1 OF 5 1
A
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8 15
16
2 14
4
8 B
7 3
1 10
ITEM NO. 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
PART
MATERIAL Fuselage Carbon Fiber / Balsa Inboard Wing Balsa / MicroLite Outboard Wing Balsa / MicroLite Aileron Balsa / MicroLite Boom A Carbon Fiber Boom B Carbon Fiber Fuselage Hatch Carbon Fiber Access Panel Carbon Fiber Main Gear Axle Steel Skid Plate Kevlar Tail Connector Carbon Fiber Ruddervator Balsa / MicroLite Rudder Balsa / MicroLite Inboard Main Spar Carbon Fiber Inboard Alignment Spar Carbon Fiber Outboard Main Spar Carbon Fiber Outboard Alignment Spar Carbon Fiber Wheel Foam
QTY 1 1 2 2 1 1 1 1 1 1 1 2 2 1 2 2 2 2
B
University of Florida A
Design/Build/Fly 2010-2011 9
Project Title PROPRIETARY AND CONFIDENTIAL THE INFORMATION CONTAINED IN THIS DRAWING IS THE SOLE PROPERTY OF THE UNIVERSITY OF FLORIDA DBF GROUP. ANY REPRODUCTION IN PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF THE UNIVERSITY OF FLORIDA DBF GROUP IS PROHIBITED.
18
SolidWorks Student License Academic Use Only 8 7
6
5
4
3
J. Lasater
Agent Blue
Dr. R. Lind
Structural Arrangement
K. Rausch
SCALE: 1:6
Chief Engineer
Advisor
Drawn By
2
2/24/2011
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8
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B
ITEM NO. 1 2 3 4 5 6 7 8 9
9
9
PART DESCRIPTION QTY Motor / Gear Box NEU1105-3Y (4.4:1 Gear Ratio) 1 Propeller Graupner 10x6 1 Battery Pack Elite 1500 (13x) 1 ESC Castle Phoenix 25 1 Receiver Battery Pack KAN 400 (5x) 1 Receiver Castle Berg-7P 1 Fuse Maxi Slow-Blow 20 Amp 1 Aileron Servo HiTec HS-65HB 2 Ruddervator Servo HiTec HS-65HB 2
University of Florida A
Design/Build/Fly 2010-2011 Project Title PROPRIETARY AND CONFIDENTIAL
SolidWorks Student License Academic Use Only 8 7
THE INFORMATION CONTAINED IN THIS DRAWING IS THE SOLE PROPERTY OF THE UNIVERSITY OF FLORIDA DBF GROUP. ANY REPRODUCTION IN PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF THE UNIVERSITY OF FLORIDA DBF GROUP IS PROHIBITED.
6
5
4
3
J. Lasater
Agent Blue
Dr. R. Lind
System Arrangement
Chief Engineer
Advisor
Drawn By
K. Rausch 2
SCALE: 1:5
2/24/2011
SHEET 3 OF 5 1
A
8
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1
D
D
C
C
22 GOLF BALLS STEEL BRICK
B
B
3.81
.45
6.00 12.19
4.19
3.00
University of Florida
A
Design/Build/Fly 2010-2011
NOTE: DIMENSIONS IN INCHES
Project Title PROPRIETARY AND CONFIDENTIAL
SolidWorks Student License Academic Use Only 8 7
THE INFORMATION CONTAINED IN THIS DRAWING IS THE SOLE PROPERTY OF THE UNIVERSITY OF FLORIDA DBF GROUP. ANY REPRODUCTION IN PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF THE UNIVERSITY OF FLORIDA DBF GROUP IS PROHIBITED.
6
5
4
3
J. Lasater
Agent Blue
Dr. R. Lind
Payload Accommodation
K. Rausch
SCALE: 1:2
Chief Engineer
Advisor
Drawn By
2
2/24/2011
SHEET 4 OF 5 1
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8
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D
D
20.00
C
C
10.00 B
B
11.00
University of Florida A
Design/Build/Fly 2010-2011
NOTE: DIMENSIONS IN INCHES
Project Title PROPRIETARY AND CONFIDENTIAL
SolidWorks Student License Academic Use Only 8 7
THE INFORMATION CONTAINED IN THIS DRAWING IS THE SOLE PROPERTY OF THE UNIVERSITY OF FLORIDA DBF GROUP. ANY REPRODUCTION IN PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF THE UNIVERSITY OF FLORIDA DBF GROUP IS PROHIBITED.
6
5
4
3
J. Lasater
Agent Blue
Dr. R. Lind
Stowed Configuration
Chief Engineer
Advisor
Drawn By
K. Rausch 2
SCALE: 1:4
2/24/2011
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A