Team VOLocity - Trace: Tennessee Research and Creative Exchange

University of Tennessee, Knoxville
Trace: Tennessee Research and Creative
Exchange
University of Tennessee Honors Thesis Projects
University of Tennessee Honors Program
5-2010
Team VOLocity
Beth Clement
University of Tennessee - Knoxville, bclement@utk.edu
William Schuman
University of Tennessee - Knoxville, wschuman@utk.edu
Follow this and additional works at: http://trace.tennessee.edu/utk_chanhonoproj
Part of the Aerospace Engineering Commons
Recommended Citation
Clement, Beth and Schuman, William, "Team VOLocity" (2010). University of Tennessee Honors Thesis Projects.
http://trace.tennessee.edu/utk_chanhonoproj/1383
This Dissertation/Thesis is brought to you for free and open access by the University of Tennessee Honors Program at Trace: Tennessee Research and
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Team
VOLocity
Team
VOLocity
TABLE OF CONTENTS
1 EXECUTIVE SUMMARY .......................................................................................... 4
2 MANAGEMENT SUMMARY..................................................................................... 5
3 CONCEPTUAL DESIGN........................................................................................... 7
3.1 MISSION REQUIREMENTS ...............................................................................................................................7
3.2 INITIAL DESIGN POINT ....................................................................................................................................9
3.3 STRUCTURES ................................................................................................................................................10
Fuselage Design .........................................................................................................................................10
Construction Techniques.........................................................................................................................11
Payload Loading Configuration..............................................................................................................12
Landing Gear...............................................................................................................................................13
3.4 PROPULSION SYSTEM ..................................................................................................................................14
Motor Selection...........................................................................................................................................14
Battery Selection ........................................................................................................................................14
Propeller Selection ....................................................................................................................................14
Motor Configuration ..................................................................................................................................14
3.5 AERODYNAMICS ...........................................................................................................................................15
Wing Configuration....................................................................................................................................16
Wing Location .............................................................................................................................................17
3.6 STABILITY AND CONTROLS ..........................................................................................................................17
Empennage Configuration.......................................................................................................................17
4 PRELIMINARY DESIGN......................................................................................... 20
4.1 STRUCTURES ................................................................................................................................................20
First Iteration ...............................................................................................................................................20
Second Iteration .........................................................................................................................................21
Third Iteration..............................................................................................................................................21
Fuselage Construction .............................................................................................................................21
Wing and Tail Construction.....................................................................................................................22
Weight Estimate..........................................................................................................................................22
4.2 PROPULSION SYSTEM ..................................................................................................................................22
Motors and Electronic Speed Controllers ...........................................................................................23
Propellers .....................................................................................................................................................24
Batteries .......................................................................................................................................................26
4.3 AERODYNAMICS ...........................................................................................................................................26
Wing Airfoil Selection ...............................................................................................................................26
Wing Sweep Angle .....................................................................................................................................27
Aspect Ratio and Taper Ratio .................................................................................................................27
Detailed Drag Analysis .............................................................................................................................28
Wing Incidence ...........................................................................................................................................29
4.4 STABILITY AND CONTROLS ..........................................................................................................................30
Controls Design/Analysis Method .........................................................................................................30
Control Sizing Trades ...............................................................................................................................31
Controls and Stability Characteristics..................................................................................................31
4.5 MISSION MODEL AND UNCERTAINTIES ........................................................................................................31
4.6 INITIAL AIRCRAFT MISSION PERFORMANCE EVALUATION..........................................................................32
5 DETAILED DESIGN................................................................................................ 34
5.1 STRUCTURES ................................................................................................................................................34
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Final Fuselage Structure ..........................................................................................................................34
Cargo Section..............................................................................................................................................34
Bat Payload Mechanism ...........................................................................................................................35
Weight Breakdown.....................................................................................................................................35
Wing Structure ............................................................................................................................................36
Landing Gear...............................................................................................................................................38
5.2 PROPULSION SYSTEM ..................................................................................................................................39
5.3 CONTROLS DETAILED DESIGN.....................................................................................................................40
Control Surface Sizing..............................................................................................................................40
Stability Calculations ................................................................................................................................40
Servos ...........................................................................................................................................................41
5.4 WEIGHT AND BALANCE ................................................................................................................................41
5.5 DRAWING PACKAGE.....................................................................................................................................42
5.6 FLIGHT PERFORMANCE ................................................................................................................................42
5.7 MISSION PERFORMANCE..............................................................................................................................43
6 MANUFACTURING PLAN AND PROCESSES...................................................... 43
6.1 M ANUFACTURING PROCESSES ....................................................................................................................43
Fuselage, Wing, and Control Surfaces .................................................................................................43
Battery Compartment and Landing Gear .............................................................................................44
Cargo Compartment ..................................................................................................................................44
6.2 M ANUFACTURING TIMELINE .........................................................................................................................44
7 TESTING PLAN ...................................................................................................... 45
7.1 STATIC TESTING ...........................................................................................................................................45
7.2 FLIGHT TESTING ...........................................................................................................................................45
7.3 CHECKLISTS .................................................................................................................................................47
8 PRE-COMPETITION PERFORMANCE RESULTS ................................................ 48
8.1 ROLLING RESISTANCE TESTING ..................................................................................................................48
8.2 STRUCTURAL TESTING.................................................................................................................................48
8.3 SYSTEMS TESTING .......................................................................................................................................52
9 PRE-CRASH CONSTRUCTION .............................................................................. 52
10 CRASH ................................................................................................................. 54
11 TESTING................................................................................................................ 56
11.1 STATIC THRUST ..........................................................................................................................................56
11.2 ROLLING RESISTANCE ................................................................................................................................60
12 REVISED PERFORMANCE PREDICTIONS ......................................................... 60
13 WORKS CITED..................................................................................................... 64
APPENDIX A: DRAWING PACKAGE........................................................................ 65
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Team
VOLocity
1 Executive Summary
The aircraft designed by team Volocity for the 2010 American Institute for Aeronautics and
Astronautics Design Build Fly competition was designed to efficiently carry a payload of 10 softballs or
two baseball bats around the competition course. The aircraft is a single engine, high aspect ratio,
mono-wing design with a tubular fuselage, conventional tail, and tail dragger landing gear. The ten
softball payload is housed within the body of the fuselage in such a way that it is distributed evenly
about the aircraft’s empty center of gravity. The payload bay is accessed via clamshell doors designed
for quick loading capability. Additionally, two hard point mounting devices are built into the wings in
order to externally carry two baseball bats.
When initially designing the aircraft, the team analyzed the relative importance of the three
factors involved in flight scoring: weight, time to complete the course, and number of bats carried. The
priority of these three factors was established by determining the score’s sensitivity to each category.
Analysis of the scoring system showed that the time to complete the course was the most critical factor
followed by aircraft weight and the number of bats carried. With speed in mind, it was decided that two
bats would be a conservative mission three payload because the performance requirements for a ten
softball payload were closest to carrying two baseball bats. If more than two bats were carried, the
weight of the aircraft would require a more powerful propulsion system and more robust structure.
Since the weight of the aircraft was a higher priority parameter, the team decided to carry a bat payload
similar to the softball payload.
Maximizing speed meant increasing thrust and minimizing drag. Thus, the motor was selected
such that it could function at the maximum amperage available, 40 amps. It was also critical that this
motor be operating at close to its maximum capacity at this amperage in order to maximize propulsive
efficiency. The motor found to most closely meet these requirements was the AXI Gold 4130/20
Brushless Motor. Drag was reduced primarily through the reduction of the wing’s induced drag. This
was accomplished through the use of a high aspect ratio wing. Analysis showed that the best balance
of aerodynamic efficiency and structural simplicity was met by employing an aspect ratio of eight and a
taper ratio of one. Additionally an optimization scheme was employed to minimize the wing area to
minimize skin friction drag.
To guarantee that the aircraft would maintain a high level of thrust and complete all three
missions under powered flight, great consideration was given to battery selection. Analysis of the flight
time required a battery that would supply 40 amps for at least 95 seconds. This required batteries rated
to at least 1500 mAh. To allow for discrepancies between actual and ideal flight performance and to
enable the 40 amp current limit to be fully utilized, the team decided to employ a battery rated to at least
4500 mAh.
Takeoff performance was also a key design factor as a maximum takeoff distance of 100 feet
was specified by the competition requirements. A powerful motor and light airframe were critical for
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minimizing takeoff distance. Also, the aircraft was designed to roll with an angle of attack of 15 degrees
during takeoff in order to decrease takeoff distance. This configuration was achieved by employing a
tail dragger landing gear system.
Weight reduction was also a goal of the team, since it was the second leading factor in scoring
and had an effect on takeoff distance and maximum flight speed. Weight was reduced by employing
lightweight composite construction for the fuselage and a wooden rib and spar design for the wings and
control surfaces. A program was created to minimize the structural elements required by determining
the wing loading at 7g’s with a factor of safety of two.
Key performance and geometry data for the aircraft is shown in Tables 1.1 and 1.2.
Table 1.1 Key Performance Data
Parameter
Weight
Stall Velocity
Maximum Velocity
Takeoff Distance
Empty
9.5 lbf
32 ft/s
142 ft/s
28 ft
10 Softballs
13.7 lbf
38 ft/s
142 ft/s
65 ft
2 Bats
13.5 lbf
38 ft/s
134 ft/s
61 ft
Table 1.2 Airplane Geometry
Component
Wing Area
Wing Span
Wing Chord
Wing Aspect Ratio
Wing Taper Ratio
Fuselage Length
Dimension
2
6 ft
6.93 ft
0.866 ft
8
1
46 in
2 Management Summary
Team VOLocity 2010 was composed of eight seniors and twelve underclassmen, 19 of whom
were aerospace engineers with one mechanical engineer. The seniors were responsible for designing
the aircraft and subsystems as well as supervising all technical activities while underclassmen assisted
with construction and testing. Seniors were divided into four technical competencies: aerodynamics,
propulsion, structures, and controls. The team was lead by a team coordinator, whose primary role was
to facilitate communication between technical design groups to ensure efficient design and
development. The team coordinator was also responsible for leading team meetings, managing team
finances as well as purchases and acquisitions. The Team VOLocity 2010 organizational chart is
shown below in Figure 2.1.
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Figure 2.1 Team Organizational Chart
Seniors were divided into the four technical subsections in order to provide an efficient division
of labor for the design process based on individual competencies and preferences. The aerodynamics
group was responsible for airfoil selections, determining wing and tail geometry, and drag and lift
estimations. Their primary objective was to maximize the design’s aerodynamic efficiency by minimizing
drag and maximizing flight handling qualities. The propulsion team was responsible for engine, battery,
and propeller selection, and the performance modeling of these systems. Specific emphasis was
placed on selecting a motor and propeller combination that would most effectively balance take off and
cruise performance. The controls group determined control surface sizes and selected control servos.
The group was also responsible for estimating stability characteristics of the aircraft. The structures
group was responsible for determining construction methods and material selections for the aircraft as
well as providing weight and center of gravity estimations. Additionally, they were tasked with
determining how to secure and stow payloads. The structures group also determined the landing gear
geometry and its construction method. Underclassmen were each involved in assisting multiple
disciplines, and their assignments were based on a combination of team needs and personal
preferences.
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Development and construction of the aircraft was initiated during August, 2009, as the seniors
began independently developing possible design concepts. The individual design period was
completed in November, 2009. The team then began reconciling individual concepts into an initial
conceptual design, which concluded in December, 2009. Preliminary design immediately began and
was finished by mid-February. The design freeze for the detailed design was set for February 19, 2010.
Figure 2.2 details the sequence and progress of each design and construction phase.
08/01/09
09/20/09
11/09/09
12/29/09
02/17/10
04/08/10
Review of rules and requirements
Actual Review of rules and requirements
Formal Formation of Design Build Fly Team
Actual Formation of Design Build Fly Team
Rough Draft of C onceptual Design
Actual Rough Draft of C onceptual Design
Obtain Funiding
Actual Obtain Funding
Individual Design Process
Actual Individual Design Process
Formal C onceptual Design Phase
Actual Formal C onceptual Design Phase
Formation of sub- teams
Actual Formation of sub- teams
Preliminary Group Designs
Actual Preliminary Group Design
Detailed Design Phase
Actual Detailed Design Phase Begin
Report Writing
Actual Report Writing
Design Finalized
Actual Design Finalized
Report Submitted
C onstruction
Flight Testing
C ompetition
Planned
Actual
Figure 2.2 Project Schedule
3 Conceptual Design
3.1 Mission Requirements
The mission requirements set forth by the 2010 Design Build Fly competition necessitated an
airplane capable of carrying a payload of ten 11 and 12 inch circumference softballs or one to five “bats”
over a rectangular course as shown in Figure 3.1 in as little time as possible.
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Figure 3.1 Flight Course Layout (Ref. 2)
The aircraft was restricted to a maximum gross weight of 55 pounds. An electric motor had to
be used with a maximum allowable current of 40 amps per motor, and propulsive batteries were limited
to a weight of four pounds or less. The batteries were also limited to either nicked cadmium or nickel
metal hydride types. Additionally, the aircraft was required to take off in a distance of no greater than
100 feet. The final, and perhaps most limiting restriction, was that the entire aircraft had to be able to fit
within a box with external dimensions of two feet by two feet by four feet.
The design of the aircraft was primarily driven by attempting to maximize speed and payload
capacity while minimizing weight, as these factors were the scoring criteria for the competition. As
shown in Figure 3.2 the flight time is the most critical parameter for maximizing overall score, as it
displays the highest sensitivity.
Figure 3.2 Score Sensitivity for Each Scoring Parameter
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Team
In order to successfully carry ten softballs, it was necessary for the aircraft to be capable of
lifting an internal payload of at least 4.5 pounds. Also, it was conservatively estimated that the aircraft
would need a range of 24,000 feet in order to complete three laps around the competition course at full
power.
In addition to meeting the competition requirements, the aircraft also needed to be capable of
maintaining safe flight in the seasonal weather conditions for Wichita, KS. The aircraft was designed to
yield satisfactory performance at density altitudes up to 3,500 feet and peak winds up to 25 miles per
hour and up to 15 mile per hour crosswinds. These values represent the most extreme flight conditions
expected. The average meteorological conditions are shown below in Table 3.1.
Table 3.1 Average Meteorological Conditions for Wichita, KS
Weather Properties in Wichita, KS
16-Apr
17-Apr
18-Apr
Average Wind Speed (mph)
15.0
14.3
15.5
Maximum Wind Speed (mph)
31
25
33
Maximum Wind Gust Speed (mph)
38
33
43
Average Temp (˚F)
63.8
61.6
59.8
29.922
29.892
29.862
Average Density (slugs/ft )
0.002246
0.002255
0.002264
Average Density Altitude (ft)
2000
1800
1600
Pressure at Sea Level (in-Hg)
3
The aircraft also needed to be built with ease of construction in mind in order to allow more time
to be devoted towards testing and refining the final design rather than building the aircraft. A simple
design was preferred so that any damage incurred during flight could be readily repaired. In addition to
ease of construction, the plane needed to be capable of being quickly assembled and loaded. This
meant that the design team was required to incorporate loading mechanisms and structural breakpoints
into the aircraft. This was especially critical for the wings, which would not be able to fit into the box as
a single unit attached to the fuselage.
3.2 Initial Design Point
In order to establish an initial design point, eight individual members performed designs and
these were combined to form an “average” aircraft. This aircraft reflected the averaged values of each
individual’s weight estimate and surface sizing, and is shown below in Table 3.2. This initial design point
was used as the basis for each group to begin analysis.
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Table 3.2 Average Aircraft Geometry
Parameter
Empty Weight (lb)
Wingspan (ft)
Wing Chord (ft)
Aspect Ratio
Taper Ratio
Value
12
8
1
8
1
3.3 Structures
Fuselage Design
For the fuselage initial design, three main configurations were considered. The tube with wings,
twin boom, and flying wing designs were evaluated on the criteria of ease of construction, ease of
loading, weight, and structural integrity. Construction simplicity was a very important factor due to the
team’s inexperience in model aircraft fabrication. Analysis of mission requirements showed that quick
loading capability was critical due to its influence on scoring. Thus, an easily accessible payload bay
was required. Minimizing weight was critical to maximizing both the flight speed and weight score.
Therefore, the lightest possible construction methods were preferred. When analyzing the structural
integrity of the fuselage configuration, the structural break points necessary to allow the aircraft to fit
within the box were key factors in the final selection. These break points would most likely be the
weakest feature of fuselage assembly. An evaluation of each configuration is shown below in Table 3.3.
•
Twin Boom Design: The internal payload could be stored in two booms, and could hold an
engine in each nose. The advantage of this design was that two smaller fuselages could
reduce the total drag due to the smaller form factor of the thinner fuselages. It also had several
disadvantages. First, the landing gear would be either complex or limiting in performance. For
a twin boom design, it would be necessary to use bicycle type landing gear on each one of the
fuselages, making ground handling difficult. Second, construction of all the connection points
for a twin boom aircraft creates additional points of failure. Additionally, the twin engine design
introduces increased complexity to the propulsion system.
•
Flying Wing Design: While this fuselage offered the potential for the greatest aerodynamic
efficiency, there were several disadvantages to this design. Storing the payload, batteries,
servos, and structure inside of the wing’s cross section would require a greater wing thickness
than the airfoil geometry would permit without an excessively long chord length. This would
offset the aerodynamic advantage that a flying wing would hold over a normally configured
aircraft.
•
Tube with Wings Design: This configuration called for an approximately cylindrical fuselage,
tapering into a conic tail section. The placement of the wing in this design could be low, mid, or
high wing. There were several advantages to the tube with wings approach. The first was the
ease of construction. A cylindrical fuselage is fairly simple to construct, compared to a flying
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wing or a dual fuselage design. The only connection to the fuselage that had to be taken into
account was the intersection of the wing with the main body of the aircraft. The second was the
ease of internal payload loading. This design was ultimately chosen because of the ease of
construction and payload loading advantages.
Table 3.3 Fuselage Configuration Figure of Merit
Figure of Merit
Weight
Ease of Loading
Structural
Integrity
Ease of
Construction
Total
Weight Factor
25
30
25
Tube With Wings
23
27
20
Twin Body
17
20
18
Flying Wing
20
18
20
20
17
15
8
100
87
70
66
Construction Techniques
Several different construction methods were considered for the aircraft. It was desired that the
aircraft be structurally robust, and easy to build and repair in the field. Emphasis was placed on using
methods that were familiar to the build team members, thus ensuring the highest possible quality of
construction. A construction figure of merit is shown below in Table 3.4.
Composite Fuselage and Wing: This method would provide a sturdy and lightweight fuselage
and wing. The fuselage structure would be especially robust due to the added structural
stability of the strong composite outer skin. It would, however, be fairly difficult to construct and
repair in the field, should a major accident occur due to the time required for epoxy to set.
Foam Fuselage and Wing: This method had merit in that it would be very lightweight and fairly
simple to construct. The principle detriment to foam construction was the durability of the
material. Foam tends to be a brittle material and would likely suffer heavy damage in the event
of a crash. Also, the foam construction would necessitate that each section of the aircraft be of
uni-body design. Thus, replacing a damaged component would require a major rebuilding
operation.
Composite Fuselage with Rib and Spar Surfaces: A hybrid design featuring a composite
fuselage shell and rib and spar wing and control surfaces was also proposed. This design
combined structural integrity of a solid composite fuselage with the familiar construction
techniques of the rib and spar wings. Also, this design provides for easier repair to damaged
surfaces than a purely composite aircraft.
Rib and Spar Construction: Another construction method considered was a traditional rib and
spar design, using a composite carbon fiber laminated wooden main spar for both the wing and
the fuselage. This method was most familiar with the team, and would be easier to repair at the
competition than a pure composite design. This also seemed to be the most straightforward of
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Team
the three methods to construct when accounting for the team’s building experience. The group
ultimately chose to go with this design because of the familiarity with the construction technique
and the ease of repair in the field.
Table 3.4 Fuselage Construction Methods Figure of Merit
Structural
Integrity
Ease of
Construction
Ease of Repair
Familiarity
Total
Weight
Factor
Composite
Fuselage and
Wing
Foam Fuselage
and Wing
Composite
Fuselage with
Rib and Spar
Surfaces
Rib and Spar
Construction
20
15
8
18
18
30
8
16
20
25
30
20
100
16
8
47
10
10
44
20
15
73
24
17
84
Payload Loading Configuration
Payload loading was determined to be a critical design factor, as it had a direct impact on the
time score for Mission 2. Thus, the payload system needed to be both structurally feasible and easily
operated. Principle design requirements were that the system had to be structurally robust, easily
loaded, and secured from opening during flight. Design concepts are discussed below.
Clamshell Nosecone or Tailcone: This design had the advantage of not weakening the
fuselage around the wing where the lift from the wing would be transmitted to the fuselage in a
composite design. Two disadvantages to this loading design were the loading time and the
securing method. Loading the fuselage through the nosecone does not allow the team access
to the interior of the fuselage to secure the cargo. In addition, it only allows loading of one or
two softballs at a time, unlike the top fuselage clamshell that could allow loading of all of the
softballs at the same time.
Clamshell Door Design: The second option considered was to have a clamshell hatch on the
top of the fuselage, regardless of the final fuselage configuration chosen. This had the
advantage of being the fastest way to load because the softballs could be dropped directly into
position at the same time. It had the disadvantage of providing a weak zone in the fuselage for
composite shell construction techniques, which would be located near where the wing attached
to the main fuselage. This disadvantage would not be present in the rib and spar fuselage
construction technique because the lift would not be transmitted through the skin of the aircraft.
However, due to the importance of loading to the mission score, it was decided that the
clamshell door was the best design. Also, since the rib and spar design was chosen, the
structural detriments of the clamshell doors were greatly reduced.
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For all of the designs it was decided that two hard points under each wing would serve to hold
the baseball bats during that scoring flight.
Landing Gear
Three landing gear configurations were considered: tricycle, tail dragger, and quadricycle. For a
conventional tube-with-wings fuselage, both tricycle and tail dragger gear were considered. The
quadricycle design was considered exclusively for a twin boom concept due to the awkward position of
the booms with respect to the aircraft’s center of gravity. These designs were analyzed based on their
structural robustness, weight, ease of construction, handling qualities, and drag as shown in Table 3.5.
Quadricycle: This design was by far the most complex, least efficient, and worst handling of
any concept. However, it was the most practical solution for the twin boom arrangement.
Tricycle: The tricycle landing gear configuration is typically preferred amongst pilots because of
its inherent stability and controllability. However, the lengthy nose strut adds additional drag
and is typically the weakest structural element of the aircraft.
Tail Dragger: The design favored by the design team was the tail dragger configuration. A tail
dragger arrangement is inherently unstable, but this problem is typically not a limiting factor for
skilled pilots. The team identified two primary advantages to the conventional configuration.
First, the nose gear can be eliminated and replaced by a much shorter tail wheel assembly.
This provides a reduction in drag and a more robust structure. Secondly, the tail dragger
landing gear guarantees that the aircraft wing will immediately be at a high angle of attack when
takeoff is initiated. This allows the aircraft to minimize its rolling resistance and shorten its
takeoff distance, which is a key design parameter.
Table 3.5 Landing Gear Configuration Selection Figure of Merit (Ref. 6)
Figure of Merit
Weight
Structural
Integrity
Ground
Handling
Drag
Construction
Total
Weight
Factor
20
25
Tricycle
Tail Dragger
Quadricycle
7
13
17
20
14
13
25
20
11
22
20
10
100
7
5
52
18
8
74
12
8
69
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3.4 Propulsion System
The propulsion system selection was a rigorous process which required several motor, battery,
and propeller configurations. The mission scoring focused heavily on the weight of the aircraft. Thus,
the group focused on creating the lightest possible propulsion system. Also the competition limitation of
four pounds for the batteries influenced the motor options available. Early weight predictions estimated
the plane would range from 11 to 18 pounds, meaning up to 36% of the weight would be the propulsion
system. The aim for the propulsion team was to find a system that was lightweight, powerful enough to
meet the takeoff requirements, and approach its operating limits at 40 amps of current so as to assure
no excess capability would be wasted.
Motor Selection
The competition allowed for both brushless and brushed motors. Brushless motors were
favored because of their ability to achieve higher powers with less internal friction and heat generation.
Initial designs for the aircraft used the AXI 5330 F3A because it was already available in house and was
rated for aircraft of similar weight to the team’s average aircraft.
Battery Selection
The battery weight and motor run time were the main factors in the conceptual battery design.
The batteries needed to maintain high performance for the duration of the three missions. To minimize
weight, amp hour ratings were minimized while attempting to maintain the optimum motor operating
voltage throughout the flight. The batter pack configuration initially chosen was a single barrel pack
because of its simplicity. Initial estimates of the required voltage were established based on
manufacturer data for the AXI 5330 F3A. This was found to be approximately 30 volts in order to attain
a maximum current of 40 amps.
Propeller Selection
Propeller efficiencies dictated the overall performance of the aircraft and current draw from the
batteries. It was important to evaluate separate propellers for the different missions since speed and
climb performance are changed with pitch and diameter alterations. A propeller with superior low speed
performance, and therefore a low pitch angle, would be necessary for the heaviest payload missions
because of the decreased takeoff and climb performance of the aircraft. Conversely, high speed
performance would be a priority for empty missions where takeoff and climb performance is not a
concern, necessitating a higher blade pitch.
Motor Configuration
The number and placement of the motors was determined during conceptual design. Center of
gravity, weight, and construction ease were the main issues considered. A comparison of each motor
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Team
configuration evaluated is shown in Table 3.6. The advantage and disadvantages associated with each
system are listed below.
•
Pusher: This design has the advantage of increased aerodynamic efficiency because the
aircraft does not lie in the propeller’s wake thus decreasing skin friction drag. However, the
aerodynamics of the plane would produce non-uniform airflow near the propeller leading to
inefficiencies. Potential ground clearance issues and limitations on control surface location
were also detractors of this design.
•
Twin: The twin engine design offered several benefits. First, the propeller wash could be
relocated from the surface of the fuselage to the surface of the wing. This would reduce the
skin friction of the aircraft and increase the lift distribution along the wing section immediately aft
of the propellers. However, there were some distinct disadvantages to this configuration.
Principally the added complexity both in terms of construction and operation detracted from this
design. Preliminary research of available motors suggested negligible weight savings over a
single engine design.
•
Tractor: The primary advantage of the tractor configuration is that the propeller does not
adversely affect takeoff rotation. Also, the airflow entering the propeller is undisturbed,
improving the propeller efficiency. However, the propeller wash across the fuselage would
increase skin friction drag. Ultimately, this design was chosen because of its simplicity of
construction, maximization of takeoff performance and increased propeller efficiency.
Table 3.6 Motor Configuration Figure of Merit (Ref. 3)
Figure of
Merit
Weight
CG Location
Aerodynamic
Interference
Construction
Familiarity
Total
Weight
Factor
30
25
Tractor
Pusher
Twin
25
24
20
10
30
12
20
7
14
17
15
10
100
14
10
80
5
4
53
12
1
72
3.5 Aerodynamics
Throughout the conceptual design stage the aerodynamics team concentrated on the wing
configuration and placement. Taken into consideration were stability concerns, lift, construction
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complexity, and the drag associated with the design. Weight and aerodynamic efficiency were of
primary importance because of their direct relationship to the final mission scores.
Wing Configuration
Three different wing configurations were initially considered by the design team. These were
straight mono-wing, delta wing, and straight biplane designs. These concepts were evaluated based on
their aerodynamic efficiency, structural weight, and ease of construction. Each of these factors had an
impact on the final flight score due to their direct impact on mission flight time and total weight. Below
are the pros and cons of each configuration and the figure of merit in Table 3.7:
•
Straight mono-wing: This design offered the advantage of being both straight forward and
proven. The design would be simple to construct and is aerodynamically efficient at expected
flight speeds.
•
Delta wing: This wing would be similar in weight to the mono-wing design, but would be more
difficult to construct. A delta wing is usually used for transonic and supersonic flight regimes, but
it does have beneficial low speed flight characteristics such as an increased critical angle of
attack. However, the delta wing is generally considered a poor choice for aircraft operating in
the low subsonic flight regime. This, along with the added difficulty of construction led to the
delta wing being discounted.
•
Flying Wing: The flying wing design offered the best aerodynamic efficiency of any concept
proposed. Due to the absence of a tubular fuselage and horizontal stabilizer, the skin friction
drag of this design was by far the lowest. Also, this design was not plagued with the trim drag
of a conventional wing and tail arrangement. However, this design also had distinct
disadvantages. First, the stability of a flying wing is marginal when compared to wing and spar
of twin boom designs. Moreover, the construction of the wing would be difficult due to the thick
chord required to store the softball payloads. Thus, the wing would have to be longer than the
box’s maximum length. This would require at least one major break point in the structure,
reducing its robustness. Finally, the team was unfamiliar with the construction techniques
required for building a remote controlled flying wing aircraft. These factors lead to the
disqualification of this concept.
•
Bi-plane: This design offered several advantages in both limiting the total wing span of the
aircraft and adding to its structural integrity. It would be more structurally sound due to the
shortened wing span, but also was more difficult to construct. Additionally, the design would
likely be heavier than a mono-wing aircraft. Aerodynamically, the biplane design was also at a
disadvantage due to the effects of interference between the two wings. Great care would have
to be taken to ensure enough vertical separation was left between the wings so as not to
dramatically reduce their efficiency. The greatest advantage of the biplane with regards to the
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mission requirements would be the potential to create a robust aircraft capable of fitting within
the box without any structural break points. Despite this, the added structural weight and
complexity, as well as the decreased efficiency of the wing, led to this concept being rejected.
Table 3.7 Wing Configuration Figure of Merit
Figure of Merit
Weight
Drag
Characteristics
Lift
Characteristics
Construction
Familiarity
Total
Weight
Factor
20
Mono Wing
Bi-plane
Delta Wing
Flying Wing
16
11
18
17
30
23
18
12
25
30
23
14
10
25
10
10
100
10
10
82
6
7
56
4
5
49
2
2
71
Wing Location
Three scenarios were considered for wing placement on the fuselage; they were high, mid, and
low wing designs. The high wing design was the most ideal from a stability and controls perspective
due to the additional roll stability. However, a high wing would impede the loading of the softballs from
the top of the fuselage due to additional structural requirements. This design was rejected since the
team decided top loading would be the best scenario. The low wing design was the least desirable from
a stability perspective, requiring a proper amount of dihedral to correct rolling moments. This concept
did have benefits, such as allowing the undercarriage to be directly mounted to the wing spar, but it was
deemed to have too many detractors to be optimum. The ideal design, which balanced stability and
structural concerns as well as loading, was the mid wing arrangement. This configuration would afford
more roll stability than the low wing design, but would still permit the payload to be loaded from the top
of the aircraft.
3.6 Stability and Controls
Empennage Configuration
Several tail designs were considered for the aircraft. These included conventional tail, T-tail, Vtail, dual vertical stabilizer, and canard configurations. These tails were considered for several different
fuselage configurations, particularly twin-boom and tube-with-wings designs. These configurations were
evaluated based on ease of construction, weight, aerodynamic efficiency, and control authority. The
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advantages and disadvantages of each configuration along with the figure of merit are listed below and
shown in Table 3.8.
•
Dual Vertical Stabilizer: This design was especially applicable for the twin boom fuselage
layout, but it offered little practical benefit for a conventional fuselage layout due to the
increased complexity of having two rudders.
•
V-tail: A V-tail design was considered for both tube-with-wings and twin-boom aircraft. The Vtail was ultimately rejected because aerodynamic data suggested that a V-tail would ultimately
need to have almost the same surface area as a conventional tail in order to maintain the same
level of control authority; thus defeating the purpose of eliminating the vertical stabilizer. It may
have been possible to design an inverted V-tail for the twin boom aircraft that would have been
superior to a dual rudder design, but because the twin boom design was eliminated from
consideration for structural and aerodynamic reasons this option was never explored.
•
T-tail: This design was applicable to the tube-with-wings fuselage. It was thought that the
efficiency of the tail could be improved by removing the horizontal stab from the propeller wash
and reducing the effect of the wing’s downwash. The T-tail design would also allow for a
shorter vertical stabilizer, thus offering aerodynamic efficiency gains. Despite these
advantages, the configuration was also plagued with complications. The T-tail design would
require a more substantial structure in the vertical stabilizer and would also create additional
difficulty for mounting control servos compared to a conventional tail. Also, T-tails are prone to
entering deep stalls which could prove unrecoverable without careful design attention.
•
Canard: This design was considered for both the tube with wings and twin boom concepts.
The canard would be employed had a pusher style propeller been employed rather than a
tractor propeller. The canard was not considered for tractor style propulsion aircraft because
the canard would be directly placed in the propeller’s slipstream, increasing its parasite drag. In
addition to slipstream difficulties, the canard would also create additional downwash on the
wing, increasing drag regardless of the propeller’s location. A principle benefit of the canard
configuration is that it can be designed such that the aircraft will never be able to stall its wing.
However, this added safety was not significant enough to overrule the design’s detractors in the
eyes of the team.
•
Conventional Tail: The conventional tail was considered for the tube with wings fuselage
layout. This design featured a horizontal tail extending from the fuselage midline and a vertical
tail positioned above the fuselage. The primary advantage of this design was the reduced
structural weight when compared with the T-tail and dual rudder concepts. Also, the
construction of this design is simpler to execute than other options considered. Detractors to
this design included the additional skin friction associated with placing the control surfaces in
the propeller wash. This design was ultimately chosen due to its reduced structural weight and
simplicity of construction.
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Table 3.8 Empennage Configuration Figure of Merit (Ref. 6)
Figure of Merit
Weight
Factor
T-tail
V-tail
Conventional
Dual
Canard
Weight
25
7
22
18
12
18
Control Authority
30
20
16
25
25
14
Efficiency
20
17
15
15
7
10
Construction
25
9
9
22
16
17
Total
100
53
62
80
60
59
3.7 Storage Box
The design of the box was arrived at using same approach as for the airplane. The contest rules
stated that the box could not use tape, magnets, or Velcro in any way. The box could not sustain
damage during the missions or the flight attempt would be forfeited and it had to fit within the
dimensions specified by the competition (4x2x2 feet.) The box is a factor for the team’s score, so the
overall objective was to build a box as light as possible. It also had to be strong enough to withstand
damage during the missions as well as to carry the plane and everything the team would need.
Three designs were proposed. The first design was made completely of balsa with the base
being reinforced with a fiberglass laminate. Two hinges and a latch would form the functional lid. All
pieces would be tacked together with small nails. The weight of this design was estimated to be eight
pounds.
The second design used a balsa lid and bottom with the sides made of balsa strips evenly
space. The balsa strips would be attached with glue. The sides would then be wrapped with a durable
plastic in order to complete the box walls. The lid would be sheet of foam that would be cut to rest snug
inside the box with small, short strips of balsa for it to rest on. Estimated weight was 6 lbs.
The third design was a box made out of strong, durable foam. The lid would rest on top of the
box with scrap pieces of foam attached to its underneath which slide inside the box assuring that the lid
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was attached. All the sides would be glued together. Estimated weight was four pounds due to the box’s
reinforcements to ensure strength. The third design was chosen due to its simplicity to construct and
light weight.
4 Preliminary Design
After deciding that the aircraft would be a tube with wings style, single engine, mid-wing, monowing aircraft it was necessary to further investigate the performance and behavior of the aircraft in
greater detail. During the preliminary design, work was done to investigate the structural, propulsive,
aerodynamic and stability characteristics of the airplane. Several trade studies were performed to
ensure that the airplane performance would be optimized within the flight objectives.
4.1 Structures
For the preliminary design the team decided to use a cylindrical fuselage with a mid wing. For
this design, it was decided that the fuselage would have a main spar running down the centerline,
composed of balsa wood with a carbon fiber weave bonded to both sides. This would provide a strong
connection point for the wing, since all of the aerodynamic forces acting on the wing have to be
transferred to the fuselage through the wing’s connection. Also, it was decided that the wing would be
split into two sections, each connecting to one another through the main spar of the fuselage.
First Iteration
The fuselage was to have a rounded engine cowl, with a 20 inch long cargo section for the
softball mission, and a 20 inch long tapered tail section. The diameter of the cylindrical cargo section
was to be 8.5 inches. This dimension was chosen so that two softballs could be loaded side by side,
with the main structural spar separating the softballs from each other. The top of the cargo section was
designed to open via a clamshell hatch, allowing for speedy loading and unloading of cargo. This was
chosen over the nose loading due to speed of loading and the ability of the team to access the inside of
the fuselage. The ribs for the fuselage would serve the dual purpose of holding the fuselage shape and
separating the softballs from each other in a two by five grid. Initially the wing was to connect to the
fuselage via the main spar at the center of the cargo section, 14 inches from the nose of the plane. The
main spar of the wing was to be placed at its quarter chord, with the secondary spar located at the three
quarter chord which also attached to the fuselage spar. In addition, ribs were placed in the wing eight
inches apart to help support and shape the wing. It was initially decided that for the third mission the
cargo would be carried with a single bat under the center of the fuselage with any other bats attached to
hard points on the wing. The skin of the fuselage and wing was to be constructed out of Monokote, and
then heat shrunk to fit the structure. This method was chosen due to its ease of construction, repair,
and the team’s familiarity with the construction technique.
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Second Iteration
From the first iteration there were numerous items that needed improvement. The cargo
compartment design and wing connections with the fuselage needed to be reevaluated. Reducing the
rib weight was addressed along with the rib spacing for the fuselage, wing, and tail. For the first design
iteration the ribs used in the fuselage, wing, and control surfaces were presumed to be solid. For the
second iteration these ribs were hollowed out wherever structurally possible. This was done because
the main loads of the aircraft were carried through the spars, while the ribs served the purpose of
holding the shape of the fuselage and wing. This provided for a reduction in weight while still having a
fairly robust structure. In addition, the shape of the fuselage was changed from a cylinder to an ovoid
cross section. This ovoid section was eight inches wide, four inches tall and 20 inches long, with a
tapering nose and tail section. There were several reasons for changing the fuselage shape. First, the
cylindrical fuselage had a lot of empty internal space that was going unused. By reducing this internal
space into an ovoid, it had the effect of reducing the frontal area of the airplane. This in turn had the
effect of reducing the profile drag of the aircraft. Additionally, the structure needed for the ovoid
fuselage would decrease its weight.
Material selection for the construction of the airplane was also reexamined at this time. The first
iteration assumed the entire structure of the aircraft would be constructed from plain balsa wood. In this
second iteration the main spars for the fuselage and wing were constructed out of balsa wood
reinforced with carbon fiber. The weight and center of gravity for these components were recalculated
by hand. The new loaded weight was calculated to be 11.96 lbs and the center of gravity was located
ten inches from the nose. The empty weight was found to be 7.59 pounds and its center of gravity was
located 8.89 inches from the nose. Compared to previous competition entries, the design weight and
structure was more reasonable for accomplishing the team’s goals. The team felt that this was a better
estimate, but still had room for improvement.
Third Iteration
For the third iteration, a Matlab code was devised to give a weight estimate based upon the
airfoil selection, wing dimension, and desired rib perimeter thickness. The wing weight estimation
scheme was applied to both the wing and tail surfaces. The rib spacing for the wing for this iteration
was improved by using a rough estimate of the lift distribution and the weight to determine where the
highest loads would occur on the wing and then concentrate the ribs at those points.
Fuselage Construction
It was decided that the fuselage would be constructed using the rib and spar technique. The
spar of the fuselage would be carbon fiber laminated balsa wood. The balsa would be cut by hand to
the proper dimensions and then coated with carbon fiber. The composite lay-up would be achieved
using a vacuum bagging system in order to save weight. The fuselage ribs would be cut using a
computer numerical controlled (CNC) laser cutter.
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Wing and Tail Construction
The main spar of the wing was the most critical element of the aircraft’s structure after the main
fuselage spar. Preliminary analysis of the wing spar suggested that it could be constructed using either
carbon fiber laminated balsa or a carbon fiber tube. Both the carbon fiber tube and wooden spar would
likely need to be manufactured by hand due to the limited selection of off the shelf products matching
aircraft requirements. Again, the ribs of the wing and tail could be produced using a CNC laser cutter to
reduce construction time.
Weight Estimate
To more accurately analyze the aircraft’s performance it was necessary to evaluate the weight
characteristics of the plane. In Figure 4.1 one can see the preliminary breakdown of the weight
including all major structural components. The preliminary structural analysis estimated the total aircraft
weight would be 7.78 pounds empty with a center of mass 15 inches from the nose.
Component
WING RIBS(8)
FUSELAGE RIBS(8)
FIREWALL
MOTOR
BATTERIES
MISCELLANEOUS
Weight(lb)
0.8
0.48
0.75
1.5
4
0.25
Figure 4.1 Structural Component Weight Breakdown
4.2 Propulsion System
During the preliminary design phase the propulsion system was selected with an iterative
process. First the motor was selected followed by the batteries. The propellers were recommended by
the manufacturers and evaluated with a curve fit equation developed using propeller efficiency graphs
found in Anderson. A Matlab code was created to generate high order polynomial curve fit equations,
some to the twentieth order. This enabled the propulsion team to better predict the performance of the
propeller during flight by generating closed form equations describing propeller efficiency. From those
equations, thrust outputs were developed with various propeller pitches and diameters.
The flow chart in Figure 4.2 shows the selection process that was used to determine the final
propulsion system. The motor was chosen first followed by the propellers and batteries. Using the
Matlab codes produced, an iterative scheme was devised for the analysis and selection of the
propulsion system. This process was similar to the one used during the conceptual design. During each
iteration, initial competition requirements were checked to verify that all limits and design restrictions
were not surpassed. Closest attention was paid to the takeoff ground roll due its dependence on the
maximum thrust available. Also, the current draw by the motor from the batteries was a primary limiting
factor in the system design.
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Figure 4.2 Propulsion Design Process Flowchart
Motors and Electronic Speed Controllers
The motor was selected using a Microsoft Excel program developed early in the design
process. This program calculated the power each motor could achieve with the use of manufacturer
recommended propeller and battery combinations. The various motors that were reviewed and relevant
data are detailed below in Table 4.1.
Table 4.1 Motor Performance Characteristics
Motor
Weight (lbs)
AXI 4130/20
AXI 5320/28
AXI 5330/F3A
AXI 4130/16
0.902
1.091
1.437
0.902
KV Rating
(RPM/V)
305
249
235
385
Power Output
(W)
739.7
706.1
764.9
722.9
Power/Weight
(W/lb)
52.8
50.4
54.6
51.6
The motor selection focused on reducing the weight of the system as much as possible without
losing large amounts of shaft power. The AXI 5330/F3A was the first to be evaluated and became our
most powerful but heaviest motor option. However, given the current and battery weight restrictions,
this motor would be operated below its maximum capabilities. The AXI 4130/20 gave similar power
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predictions but was a half pound lighter than the AXI 5330/F3A. The maximum velocity of the empty
aircraft with the AXI 4130/20 was predicted to be 146 feet per second during preliminary design. This
motor was large enough to power the plane including a two pound margin over the initial design point
weight of 12 pounds. Smaller motors were also considered but they did not provide a sufficient margin
for error in the performance or weight calculations. In addition, these motors did not allow for the full
current limit of 40 amps to be utilized, thereby limiting power available.
The speed controller was found by comparing the voltage required by the motor and the
controller’s maximum allowable voltage and current to the limitations set forth by the competition rules. It
is an integral part of controling the propulsion system and enables the plane’s acceleration to be
controled much like that of a throttle of a gas powered engine. The speed controller does this by
varying the current and voltage that is drawn by the motor. The controller chosen was a Phoenix HV-45
manufactured by Castle Creations which was specifically designed for a brushless motor. The controller
takes the direct current coming from the batteries and converts it to a phased alternating current. In this
case it is a three phased alternating current. This also works to decelerate the plane by changing the
polarity of the magnets of the brushless motor, which is a useful characteristic during the landing phase.
The Phoenix HV-45 had a maximum allowable current draw of 45 amps which was more than that
permitted by the competition and met the voltage requirements for all motors considered. The
maximum input voltage certified by the manufacturer was 50 volts.
Propellers
During the conceptual design phase, the propellers considered were those recommended by
the manufacturers and propeller efficiencies were developed using propeller efficiency graphs from
Anderson. A curve-fit was produced and used to evaluate the propeller efficiency as it varied with
velocity. All of the equations had high order polynomials, some up to the twentieth order. Below in
Figure 4.3 are the curve-fit plots for blade angles of 12, 17 and 20 degrees, which corresponds to 16x8,
14x10, and 13x11 propellers, respectively. The equations generated from these graphs were used in
the thrust predictions to find the optimal propeller for our design.
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Figure 4.3 Propeller Efficiencies for Various Blade Angles
Although the curve-fit was not as accurate for the 14x10, the peak errors were on the order of
five percent. This level of accuracy was thought to be sufficient given the uncertainties in other
performance parameters. Thrust estimates showed that this propeller gave the highest peak values.
The thrust predictions, displayed in Figure 4.4 below, reveal the clear choice for the 14x10 propeller.
The 13x11 was predicted to have a peak thrust of 5.3 pounds while the 14x10 had a peak thrust of 8.7
pounds. The peak thrust for both propellers occurred near 40 ft/s velocity. These codes enabled the
propulsion team to work backwards by first finding a propeller for the selected motor, followed by its
battery pack.
Figure 4.4 Thrust Predictions for 20˚ (Left) and 17˚ (Right) Blade Angle Propellers
The thrust for the 14x10 propeller was superior to all other propellers tested. It gave the best
performance data and provided a margin of error incase drag or weight calculations were under
predicted. This propeller’s performance was verified using manufacturer provided data. Our choice was
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consistent with the manufacturer’s suggestions for the AXI 4130/20 motor. Also, the motor and
propeller combination selected was widely recommended for model aircraft of similar size and weight to
the team’s design.
Batteries
To begin the battery selection, a decision was made between choosing Nickel-Metal Hydride
(NiMH) or Nickel-Cadmium (NiCAD) batteries, which were the two types allowed for the competition.
Competition requirements specified that the weight of the battery packs be less than four pounds and
the motors and batteries to be limited to a 40 amp current draw via a 40 amp fuse. NiMH batteries were
chosen over NiCAD batteries because of their higher energy densities and the fact that they could
discharge at a higher rate. These reasons made NiMH batteries the best choice for the aircraft.
The next step was to choose the specific batteries for the selected motor and propeller
configuration. Using manufacturer specifications for the AXI 4130/20 and its recommended propellers,
it was determined that 31.2 volts was the optimal voltage required to achieve mission goals. This
voltage value was within 1% of the manufacturer’s recommendations. To achieve this voltage, twentysix 1.2 volt batteries were needed to power the aircraft. The battery pack was designed to be organized
into three rows to be placed on the bottom of the fuselage. It was also determined with a run time and
power consumption examination that a minimum capacity of 1500 mAh would be necessary to run each
phase of the course. However, it was found that a battery certified to at least 4500 mAh would be
necessary to achieve a 40 amp discharge rate.
4.3 Aerodynamics
Wing Airfoil Selection
Wing airfoil selection was one of the most important decisions in the design process.
Considering the mission requirements, the team decided that the ideal airfoil would need several key
attributes. First, the team desired an aircraft that could easily takeoff in the required distance. Thus, the
ideal airfoil would then have a high maximum coefficient of lift and low drag coefficient. Also, it was
desired for the airfoil to have smooth stall characteristics so the airplane would not abruptly stall,
potentially creating an unrecoverable loss of control. Finally, an aircraft capable of high cruise speeds
was desired. Again, this required an airfoil with a minimum drag coefficient. A comparison of the NACA
4415, NACA 4412, Clark Y, and Clark YM 15 airfoils considered for the aircraft wing can be seen in
Table 4.2.
Table 4.2 Airfoil Aerodynamic Characteristics
NACA 4415
NACA 4412
Clark Y
Clark YM 15
(CL)max
1.56
1.63
1.45
1.6
26
(L/D)max
51.25
55.71
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The NACA 4415, NACA 4412, Clark Y, and Clark YM 15 airfoils were all considered. After
analysis of the design requirements, the team initially selected the NACA 4412 greatly due to its high
maximum lift coefficient and favorable stall characteristics. However, structural concerns required a
thicker airfoil. The NACA 4415 would allow for a rib-and-spar system to be constructed more easily
within the wing. This would also yield a stronger wing, as a thicker main spar could be employed. The
NACA 4415 had similar lift and drag characteristics to the NACA 4412 and would represent an
appreciable sacrifice in performance from the first design choice. For these reasons, the NACA 4415
airfoil was chosen.
Wing Sweep Angle
The next factor analyzed was the wing sweep. Adding sweep angle adds an extra layer of
construction difficulty to the wing and unnecessarily reduces the effectiveness of the wing for minimal
drag reduction at slow speeds. For this reason an unswept wing was chosen.
Aspect Ratio and Taper Ratio
The influence of wing aspect ratio on induced drag and lift were also evaluated. A high aspect
ratio wing produces less induced drag and also increases the lift generated at a specific angle of attack.
While high aspect ratio wings are most efficient, they can become vulnerable to structural problems.
Calculations performed by the Aerodynamics group show, as seen below in Figure 4.5, that the optimal
aspect ratio for the aircraft, balancing both aerodynamic and structural interests, would be
approximately eight. As one can also see in Figure 4.5, there is not a very large benefit to introducing
taper to the wing, especially in light of the increased construction difficulty. The group decided that an
aspect ratio of eight and taper ratio of one was the best solution for the mission requirements.
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Figure 4.5 Induced Drag Coefficient vs. Taper Ratio
Detailed Drag Analysis
After setting the aspect ratio to 8 and constant chord wing, it was necessary to further
investigate the drag associated with the vehicle as a whole. To develop a more detailed study of the
aircraft drag, a new computer code was created. The updated code included empirical equations to
evaluate the drag of the various aircraft components including the fuselage, bats, landing gear, and
wing. The lift induced drag for the aircraft lifting surfaces was calculated using Prandtl's Lifting Line
Theory; while the profile drag of the wing and empennage were evaluated using a high-order curve fit of
historical NACA data. The code calculated the total drag coefficients, both zero lift and induced drag, as
a function of only wing planform area since the aspect ratios and taper ratios were set. This was useful
in evaluating the aircraft takeoff performance and stall speed while minimizing the total wing planform
area.
By looking at Figure 4.6 one can see the drag breakdown at cruise for the aircraft both loaded
with and without bats. The zero lift drag coefficient of the plane during mission three was found to be
0.0329, while the zero lift drag coefficient without the bats on the first and second missions was found to
be 0.0247. Figure 4.6 shows the significance of the added drag associated with the bats during mission
three at cruise.
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Figure 4.6 Drag Breakdown on Missions One and Two (LEFT) and Mission Three (RIGHT) at a Cruise
Angle of Attack of 6º
The drag polar of both missions, which can be seen below in Figure 4.7, was not only needed to
predict the necessary output of the propulsion system, but also to map the aircraft’s overall mission
performance. The program used to output the drag polar information was an integral part in the final
performance optimization code which was used to develop the airplane’s final aircraft dimensions and
layout.
Figure 4.7 Drag Polar of Bat-Loaded and Bat-Unloaded Missions
Wing Incidence
To analyze the incidence angle of the wing on the fuselage, a Matlab code was created using
the drag polar program described earlier. Vortex lattice method was used to find the wing lift curve
slope of 4.563 1/radians. This was used to display the drag polars of the wing and fuselage as a
function of angle of attack rather than lift coefficient. As seen below in Figure 4.8, the drag experienced
by the wing increases nonlinearly as the angle of attack is increased, while the drag associated with the
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fuselage stays approximately constant. The fuselage minimum drag condition is located at
approximately the same angle of attack as the cruise angle of attack which was determined to be
approximately -2 degrees. Thus, the wing could be mounted to the fuselage with zero angle of
incidence without incurring a significant drag penalty during cruise.
Figure 4.8 Wing and Fuselage Drag Polar Plot
4.4 Stability and Controls
Controls Design/Analysis Method
Control surfaces were designed by referencing previous aircraft designs and empirical data
from past design experiences. The sizing of the horizontal and vertical stabilizers was of primary
interest. These surfaces are critical to the aircraft because they directly control its pitch and yaw
stability and maneuverability. The sizes of these surfaces were established by matching tail volume
coefficients with historical data found in Raymer. Because our aircraft most closely resembled a single
engine, general aviation aircraft, it was decided to follow the sizing criteria established for this class of
airplanes. These sizing coefficients were used along with wing span, wing area, mean aerodynamic
chord, and distance to the tail aerodynamic center from the aircraft center of gravity to establish the area
of the stabilizers. Ailerons were also sized based on historical data.
Stability derivatives for the aircraft in the roll and yaw axes were also estimated using empirical
methods derived from historical data. Again, these methods were detailed in Raymer. In the yaw axis,
the wing, vertical tail, and fuselage were the primary factors in stability calculations. Equations supplied
by Raymer suggested that the average aircraft design would have a CNβ value of approximately 0.07
1/rad. This is about the same value as a Cessna 182, suggesting that the aircraft will be sufficiently
stable in yaw based on the sizing guidelines given in Raymer. The wing and vertical tail were
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considered to be the principle contributors to roll stability and these were also calculated based on
Raymer’s empirical formulas. Again, the average aircraft proved to be stable, with a Clβ value of -0.12
1/rad.
The neutral point of the aircraft was also calculated to ensure a proper stability margin for the
aircraft. The average aircraft showed a stability margin of approximately 18%. This value is on the
order of a general aviation trainer, suggesting ample stability in pitch.
Control Sizing Trades
The primary sizing trade for the aircraft’s controls system was stability versus weight and drag.
Larger control surfaces would make the aircraft more stable and lend more control authority; however,
they would also add weight and increase drag. It was decided that, ideally, the aircraft should have
similar stability characteristics as a general aviation trainer aircraft. This was important because of the
difficulties involved in piloting a marginally stable remote controlled aircraft.
Controls and Stability Characteristics
Preliminary controls analysis was performed based on the average aircraft. These calculations
showed that the aircraft would have a static margin of approximately 18% in pitch. It is recommended
that an aircraft have at least 5% margins and 15-25% margins are not uncommon in stable aircraft such
as general aviation trainers. Additionally, the average aircraft was predicted to have a lateral stability
derivative of 0.07 1/rad, which is again comparable to general aviation aircraft. The aircraft was
designed to have similar stability characteristics as single engine general aviation aircraft because they
share an analogous fuselage configuration and also lack sophisticated flight control systems.
The performance of the control surfaces was estimated using a variety of techniques.
Specifically, Prandtl’s lifting line theory and vortex lattice matrix modeling were used to estimate the lift
curve slopes of each surface. This data was then used to calculate stability and control parameters.
4.5 Mission Model and Uncertainties
The aircraft was designed to carry a 4.5 pound payload over a distance of 24,000 feet. These
parameters correspond to carrying ten softballs of the largest size available in the competition over
three laps. The aircraft was also designed to carry a minimum of two bats through the course.
The greatest uncertainty in modeling the mission was the drag prediction for the aircraft in both
its empty and loaded configurations. It was believed that a reasonable drag prediction can be obtained
for the aircraft in its empty configuration. However, predicting the drag for the externally loaded bats
was much more uncertain. The bats will experience both form drag and skin friction drag, which were
difficult to quantify without wind tunnel testing an exact replica. Additionally, they will create interference
drag with the wing. These additional drag terms will reduce the performance of the aircraft and are
difficult to quantify without full scale wind tunnel or flight testing.
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Also, the performance of the propulsion system will have a significant impact on mission
success. The greatest uncertainty was the ability of the propulsion system to perform at the highest
power output throughout the duration of the flight. As the batteries become depleted, their voltage and
current outputs will be reduced, thus reducing aircraft performance. This will alter the flight times of the
actual aircraft from those predicted by the group’s analysis.
The final uncertainty facing the design team is the validity of the weight estimation scheme. The
uncertainty in the weight prediction stemmed from difficulties in estimating hidden weight factors such as
small connectors and glue. While not as heavy as the other structural components, these items could
make a significant addition on the order of 10% or more based on past construction projects. The total
weight of the aircraft contributes to the team’s final score through both the weight and time categories.
4.6 Initial Aircraft Mission Performance Evaluation
An initial analysis of mission performance was performed by using the predicted performance
values for takeoff, cruise, and maneuvering phases of flight. A 2g turn, corresponding to 60 degrees of
bank, was assumed for the 180 and 360 degree turns. The values for flight time presented in Tables
4.3, 4.4 and 4.5 were presumed to be a best case scenario, as the evaluation did not take the
accelerated phases of flight into account.
Table 4.3 Mission 1 Flight Performance
Mission 1- Empty Weight Flight
Mission Phase
Takeoff
Climb
Cruise
Lap First 180º turn
1
Cruise
360 degree turn
Second 180º turn
Cruise
Cruise
First 180º turn
Lap Cruise
2
360 degree turn
Second 180º turn
Cruise
Total
Velocity (ft/s)
0-40
95
142
55
142
55
55
142
142
55
142
55
55
142
32
Time (s)
1.7
4.1
0.6
3.1
7
9.7
3.1
3.5
3.5
3.1
7
9.7
3.1
3.5
62.9
Distance (ft)
28
388.3
83.7
170.4
1000
535.3
170.4
500
500
170.4
1000
535.3
170.4
500
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Table 4.4 Mission 2 Flight Performance
Mission 2- 10 Softballs
Lap 1
Laps 2 & 3
Mission Phase
Takeoff
Climb
First 180º turn
Cruise
360 degree turn
Second 180º turn
Cruise
Cruise
First 180º turn
Cruise
360 degree turn
Second 180º turn
Cruise
Total
Velocity (ft/s)
0-40
95
66
142
66
66
142
142
66
142
66
66
142
Time (s)
3.12
5.26
3.72
7.04
7.44
3.72
3.53
3.52
3.72
7.04
7.44
3.72
3.52
91.75
Distance (ft)
65
500
245.4
1000
490.74
3.72
500
500
245.4
1000
490.74
245.4
500
8767
Table 4.5 Mission 3 Flight Performance
Mission 3- 2 Bats
Lap 1
Laps 2 & 3
Mission Phase
Takeoff
Climb
First 180º turn
Cruise
360 degree turn
Second 180º turn
Cruise
Cruise
First 180º turn
Cruise
360 degree turn
Second 180º turn
Cruise
Total
Velocity (ft/s)
0-40
84
66
134
66
66
134
134
66
134
66
66
134
33
Time (s)
2.99
5.95
3.72
7.46
7.43
3.72
3.73
3.73
3.72
7.46
7.43
3.72
3.73
94.58
Distance (ft)
61
500
245.4
1000
490.74
245.4
500
500
245.4
1000
490.74
245.4
500
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5 Detailed Design
5.1 Structures
Final Fuselage Structure
The final fuselage design incorporated ideas from all of the iterations. The fuselage cargo
compartment was an ovoid cross section with tapered nose and tail sections. The main spar of the
fuselage was composed of 0.0938 inch balsa reinforced with carbon fiber. This main spar spans a
distance of 36.38 inches from the front firewall where the motor was attached to the tail. A firewall
composed of 0.0938 inch reinforced balsa was mounted to the front of the main spar. This firewall was
where the motor attached to the structure of the fuselage. The next main structural component of the
fuselage was the ribs. The fuselage had eight ribs, each with a slit in the middle where they attached to
the main spar. These ribs were constructed out of plain 0.0938 balsa, with a wall thickness of 0.25 inch.
The spacing between the fuselage ribs was based upon two things. First, the ribs were used as
separators in the cargo section to hold the softballs, so they had to be at least one large softball
diameter away from each other. Second, since a circular section would be used for the wing’s main
spar, there had to be sufficient spacing between the ribs to allow for the wing spar to tie into the
fuselage spar. The spacing for the ribs, measured from the firewall, is shown in Table 5.1. The width of
the fuselage was 8 inches, and the height of the fuselage was 4 inches. The nose held the ovoid shape
of the fuselage, with a radius of curvature of 4 inches. The cargo section of the fuselage starts at the
firewall and goes back 22.88 inches to the beginning of the tapering of the tail section. The top of the
cargo section was a clamshell hatch. The skin of the fuselage was compose of Monokote, heat shrunk
to the fuselage.
Table 5.1 Fuselage Rib Spacing
Rib
1
2
3
4
5
6
7
Distance From Firewall (in)
4.00
8.13
12.25
13.63
17.75
18.63
22.88
Cargo Section
The cargo occupied the majority of the fuselage. The cargo bay is shown in Figure 5.1. It
started at the firewall and ran 22.8 inches through the fuselage to the beginning of the tail section. It
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was comprised of several structural components. First, the cargo compartment used the structural ribs
of the fuselage and the fuselage main spar to separate the softball cargo into ten individual cells. This
prevented the softball payload from shifting during flight and creating problems with the aircraft’s center
of gravity. The bottom of these cells each had two 0.125 by 0.0938 inch balsa wood supports to hold
the softballs off the Monokote skin of the fuselage. These supports were attached to the ribs of the
fuselage. The top of the cargo section opened using hinges attached to the first and seventh rib in the
fuselage. These tied into two half ribs that formed the front and rear of the hatch. Three balsa wood
supports ran from the front to the rear of the hatch, which helped to hold the Monokote skin in the shape
of the fuselage. The hinges attached at the starboard side of the plane, while a simple push latch
attached on the port side to secure the hatch during flight.
Figure 5.1 Cross Section of Cargo Bay
Bat Payload Mechanism
The mechanism used to secure the bat to the aircraft was required constrain the bat in six
degrees of freedom. Additionally, a 0.1875 inch diameter pin had to register within a hole drilled in each
bat to ensure security. The mechanism designed by the group was essentially a sleeve with an
insertable pin. The sleeve secured the bat in the yaw and pitch axes. The pin secured the bat along
the longitudinal axis.
Weight Breakdown
Figure 5.2 shows the breakdown of the weight of the aircraft by each structural component. This
shows that almost 60% of the weight of the loaded aircraft is comprised of the batteries and the
payload. The glue and hardware portion included all of the screws, hinges, latches, and other fasteners
used in the construction of the plane.
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COMPONENT
WING
FUSELAGE
MOTOR
ASSEMBLY
LANDING GEAR
HORIZONTAL
STABILIZER
VERTICAL
STABILIZER
BATTERIES AND
CONTROLLER
PAYLOAD
MONOKOTE
GLUE AND
HARDWARE
TOTAL
WEIGHT (lb)
0.50
0.57
1.26
0.37
0.12
0.67
4.48
4.38
0.1
1.29
13.7
Figure 5.2 Final Aircraft Weight Breakdown
Wing Structure
The wing structure underwent several changes before arriving at the final design. First, the rib
spacing was re-evaluated. The third iteration of the wing structure had taken the lift distribution into
account, but did not consider its own weight or the forces from the external cargo. The rib spacing
method used for the final design accounted for the lift distribution, the moment about the wing spar due
to lift, the point loads of the external cargo, and the weight of the structure itself. Table 5.2 shows the
rib spacing, with the distance measured from the centerline of the fuselage.
Table 5.2 Wing Rib Spacing
Rib
1
2
3
4
5
6
7
8
9
Distance From Wing Root (in.)
4.00
8.00
12.25
16.50
21.25
26.00
31.25
37.25
41.60
The first part of the main spar of the wing was a 0.88 inch diameter carbon fiber tube. It was
attached to the fuselage main spar 13 inches from the firewall. This tube did not run the length of the
whole wing. The carbon fiber tube was attached to a carbon fiber reinforced balsa wood spar 13 inches
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from the fuselage main spar that ran for the rest of the wingspan. The rear spar of the wing was
constructed in the same manner and located at the three quarter chord of the wing. Figure 5.3 shows
the structure of the fuselage with the wing attached.
Figure 5.3 Aircraft Structure
To determine the load factor, the wing spars were assumed to be isotropic materials. This
analysis was performed using a higher compressive stress than would have been experience by the
airplane during flight. From previous testing it was known that the composite components would fail first
when undergoing compressive loading. Thus, analyzing the structure in this way gave a conservative
estimate of the stresses on the aircraft.
The maximum shear stress on the carbon fiber reinforced balsa occurred 12.8 inches from the
wing root, and was 277 psi. The samples of balsa composite tested failed at a compressive stress of
3900 psi. From assuming that the lift distribution curve remained nearly constant in shape a maximum
load factor on the composite spar was found to be 14. A factor of safety of 2 was used for the structure,
so the planned limit load factor was 7. A stress analysis was then done on the carbon fiber tube portion
of the main spar. The maximum compressive stress in this section, using the assumptions from before,
was 7300 psi occurring 4 inches from the wing root. Figure 5.4 shows the moment as a function of
distance from the wing root. The compressive failure strength of the carbon fiber tube was assumed to
be 17,400 psi. This gave a maximum load factor on that section of 23. Based on the predicted CLmax
and a limit load factor of seven, the corner velocity was found to be 92 ft/s.
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Figure 5.4 Wing Bending Moment
Landing Gear
A tail dragger configuration was selected for the final landing gear design for two primary
reasons. First, a tail dragger design increased the angle of attack of the wing during the aircraft’s
takeoff roll, which was found to significantly reduce takeoff distance. Second, the tail dragger
configuration eliminated the need for a long nose gear. This had two beneficial results. First, the added
drag of a long nose gear strut was eliminated. Second, the tail wheel design would be more structurally
sound due to the shorter strut. The dimensions of the landing main landing gear were established using
proven empirical methods to ensure that the aircraft center of gravity would not lie too far forward or aft
so as to inhibit ground handling and takeoff capabilities. The dimensions of the landing gear are shown
in Table 5.3.
Table 5.3 Landing Gear Dimensions
Dimension
Value
Distance from A/C nose to Main Gear
13.3 in.
Main Gear Height
8.6 in.
Main Gear Width
10.3 in.
Distance from Tail Wheel to A/C C.G.
24.7 in.
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5.2 Propulsion System
More rigorous analysis of the aircraft design proved that the AXI 4130/20 was the best
combination of capability and light weight design. This motor’s limitations were similar to those imposed
on the propulsion system by the competition rules. Thus, the motor did not have any excess capability
and therefore additional weight, which was not utilized during the missions. Batteries were selected
based on the preliminary mission analysis’ times to complete the course and current draw restrictions.
This was used as the criteria for picking the milliamp-hour rating and cell size of the batteries. The
propulsion parameters for the aircraft are shown in Table 5.4.
Table 5.4 Propulsion System Details
Motor: AXI 4130/20
Weight (lb)
KV Rating (RPM/V)
Total Power In (W)
Shaft Power (W)
Efficiency (%)
Length (in)
Diameter (in)
Batteries: Nickel-Metal Hydride
Weight of Each Cell (oz)
Length (in)
Diameter (in)
Voltage (V/battery)
Capacity (mAh)
Number of Cells
Total Weight (lbs)
Total Voltage (V)
Propeller
Manufacturer
Diameter (in)
Pitch (in)
o
Blade Angle ( )
0.902
305
1224
1024
84
3.33
1.16
2.08
1.69
0.91
1.2
4500
26
3.38
31.2
APC
14
10
17
A wiring diagram for the aircraft’s electrical system, including propulsion components, is shown
below in Figure 5.5.
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Figure 5.5 System Diagram
5.3 Controls Detailed Design
Control Surface Sizing
Control surface sizing for the horizontal and vertical stabilizers was initially performed using
established empirical relations for general aviation aircraft. Once this initial sizing was completed, a
more detailed analysis was performed to verify pitch and yaw stability involving empirical methods and
vortex lattice matrix modeling. The rudder and ailerons were also sized using empirical relationships
proven through experience with general aviation aircraft. Control surface sizes are shown in Table 5.5.
Table 5.5 Control Surface Sizes
Surface
Aileron
Horizontal Stabilizer
Vertical Stabilizer
Rudder
Span
16.6 in
2.16 ft
1.19 ft
14.3 in
Chord
2.6 in
0.48 ft
0.6 ft
2.9 in
Area
2
43.2 in
1.04 ft
0.71 ft
2
41.5 in
Aspect Ratio
N/A
4.5
2
N/A
Taper Ratio
1
1
1
1
Stability Calculations
The neutral point of the aircraft was calculated using empirical relations to find the downwash
angle on the horizontal surface and a vortex lattice matrix code to calculate the lift curve slopes of the
horizontal stabilizer and wing. The neutral point calculations showed a static margin of greater than 5%
in all payload configurations. The yaw stability derivative was also calculated with proven empirical
methods and vortex lattice matrix codes. Stability calculations are displayed in Table 5.6.
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Table 5.6 Stability Parameters
Parameter
Static Margin
Cnβ (1/rad)
Clβ (1/rad)
Empty
6%
0.096
-0.13
10 Softballs
17%
0.095
-0.12
2 Bats
17%
0.095
-0.12
Servos
The servos selected for use on the aircraft were four Low Profile HiTec HS-77BB servos.
These servos are capable of producing 61.1 ounce inches of torque at 4.8 volts, which is estimated to
be sufficient for this aircraft.
5.4 Weight and Balance
The weight and balance and static stability charts for each payload configuration are shown in
Tables 5.7 and 5.8 respectively. The aircraft was found to have a stability margin of 15% or greater in
each softball payload configuration. The empty aircraft had a static margin of 5.6%, which was the least
stable configuration. The bat configuration was expected to have similar center of gravity and stability
characteristics to the empty design, because the center of gravity of the bats would coincide with that of
the empty aircraft.
Table 5.7 Aircraft Center of Gravity
Aircraft Center of Gravity (Inches from Front of Aircraft)
Small balls
#
0
0
18.
3
1
2
3
4
1
Big Balls
2
3
4
5
6
7
8
17.
5
17.
4
17.
3
17.
5
17.
4
17.
3
17.
2
17.
5
17.
4
17.
3
17.
2
17.
1
17.
5
17.
4
17.
3
17.
2
17.
1
17.
5
17.
4
17.
3
17.
2
17.
2
41
5
17.
5
17.
4
17.
3
17.
2
17.
1
6
17.
5
17.
4
17.
3
17.
2
17.
2
7
17.
5
17.
3
17.
3
17.
2
8
17.
3
17.
3
17.
2
9
17.
3
17.
2
10
17.
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10
17.
2
17.
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17.
1
Table 5.8 Aircraft Static Stability Margins
Static Margin (Percent)
Small balls
#
0
0
5.6
1
2
3
4
6
7
8
9
10
13
13
15
15
16
13
14
15
15
16
13
14
15
15
16
13
14
15
16
16
13
14
15
16
16
13
14
15
16
16
16
1
2
Big Balls
3
4
5
6
13
14
15
16
7
14
15
16
16
8
15
16
16
9
16
16
10
16
5
5.5 Drawing Package
A complete drawing package is attached as an appendix to the report.
5.6 Flight Performance
The flight performance of the aircraft was predicted by estimating the power available and
power required for the aircraft while in flight. Power available was calculated using motor data supplied
by the manufacturer combined with an in house curve fit of propeller efficiency. Power required was
estimated using the drag polar provided by the aerodynamics group. Results of these calculations are
shown in Table 5.9.
Table 5.9 Flight Performance Data
Parameter
Weight (lbf)
Empty
9.5
42
10 Softballs
13.7
2 Bats
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Stall Velocity (ft/s)
Maximum Velocity (ft/s)
Maximum Rate of Climb (ft/min)
Velocity for Maximum Rate of Climb(ft/s)
Maximum Lift to Drag Ratio
Takeoff Distance (ft)
2
Wing Loading (lbf/ ft )
Power Loading (Shaft W/ lbf)
Maneuvering Speed (ft/s)
32
142
2610
95
13.2
28
0.63
107.8
75
38
142
1750
95
13.2
65
1.44
74.7
90
38
134
1630
84
11.3
61
1.42
75.9
89
5.7 Mission Performance
The mission performance of the aircraft was analyzed using a spreadsheet program that
incorporated the aircraft’s calculated performance parameters. The program used the aircraft’s
maximum climb rate, cruise speed, and maneuvering speed as well as its thrust and drag curves. The
outputs of the spreadsheet were the distance traveled by the aircraft during each mission as well as the
time required to complete the course. The results of the calculations are presented in Table 5.10.
Table 5.10 Expected Mission Performance
Mission
1
2
3
Time (s)
72
118
120
Distance (ft)
6600
11500
11400
6 Manufacturing Plan and Processes
6.1 Manufacturing Processes
Fuselage, Wing, and Control Surfaces
For the final design, there were several different methods used to assemble the plane. The
blueprints for the final design were drawn in Autodesk Inventor, with the details of the structural
components fully laid out. A computer numerically controlled cutter (CNC) was used to manufacture the
ribs for the fuselage, wing, and tail. The designs for these components were uploaded from Autodesk to
the CNC machine, and it cut the ribs to specification. The ribs were constructed out of 0.938 inch thick
balsa wood.
The fuselage main spar was hand made of 0.1875 inch thick balsa wood reinforced with carbon
fiber. The main and secondary spars for the horizontal tail were also hand made from carbon fiber
reinforced balsa. The main spars for the wing were constructed using two interlocking carbon fiber
tubes and an additional balsa and carbon fiber tip section. The two tubes were constructed by layering
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carbon fiber over a cylindrical mold and then removing it from the mold. Vacuum bagging was used in
order to minimize the component’s weight. The tip section of the spars was made of carbon fiber
reinforced balsa, which was connected to the tube part of the spars through custom designed junction.
This reinforced balsa was constructed using a sheet of 0.938 inch balsa wood and then applying a coat
of resin and a layer of carbon fiber to each side. The ribs for the wing and fuselage were attached to
the spars using a high strength modeling glue, for ease of construction. The skin of the wing, fuselage,
and tail were made out of Monokote, shrunk with a heat gun to fit the structure of the plane. The last
quarter of the wing chord-wise was made of lightweight foam, from the secondary spar to the trailing
edge. This enabled easy construction of the aileron surfaces.
Battery Compartment and Landing Gear
The battery compartment beneath the fuselage was made completely from carbon fiber, and
attached to the aircraft via four screws. These screws attached to the fuselage’s structural ribs through
reinforced blocks of balsa wood. The compartment’s outer shell was constructed by layering the carbon
fiber over a mold then vacuum bagging. The landing gear was constructed using carbon fiber based
upon a mold of an aluminum landing gear bent to the desired dimensions. This was done once again
by making a mold and then layering carbon fiber and resin over it until reaching the desired thickness
and then vacuum bagging. The aluminum landing gear set was kept as a back up, in case of failure on
the part of the main gear due to crash.
Cargo Compartment
The cargo doors of the aircraft were constructed of two half ribs. These half rib sections
attached to the full ribs at the front and rear of the fuselage’s cargo area via hinges. The top of the
hatch was a Monokote skin, heat shrunk to the half ribs. The hinges were bought from a hobby store.
The latch was a simple press and release latch. The same hinges that were used for the cargo hatch
were used for all of the control surfaces on the aircraft. The base of the cargo compartment was two
0.125 inch by 0.0938 inch strips of balsa wood on each side of the main spar, running in between the
ribs on the bottom of the fuselage. The purpose of these strips was to hold up the weight of the
softballs without letting them rip through the Monokote skin of the fuselage. These were constructed by
hand as well, and glued into place using the same method by which the ribs were attached to the spars.
6.2 Manufacturing Timeline
The manufacturing plan for the aircraft is shown in Figure 6.1. Fabrication of the aircraft is
st
st
expected to initiate on the 1 of March and continue until April 1 . Details of the construction process
are presented in the figure.
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2/28/2010 3/5/2010 3/10/2010 3/15/2010 3/20/2010 3/25/2010 3/30/2010
Landing Gear Fabrication
Fabricate Wing Ribs and Spars
Assemble Wings
Monokote Wings
Fabricate Empennage Ribs and Spars
Assemble Empennage
Monokote Empennage
Fabricate Fuselage Ribs and Spars
Assemble Fuselage
Monokote Fuselage
Wiring
Complete Propulsion System Installation
Planned
Actual
Figure 6.1 Construction Timeline
7 TESTING PLAN
7.1 Static Testing
Static testing of the aircraft and its subsystems will be performed in order to verify flight
worthiness of the completed vehicle. These static tests will be completed prior to first flight of the
aircraft in order to increase the chances of mission success and provide time for modifications if
necessary. Individual tests are described below.
Rolling Resistance Coefficient: The friction coefficient of the aircraft’s undercarriage will be
determined in order to establish a higher fidelity takeoff distance prediction for the final aircraft.
Static Thrust: Static thrust tests will be performed in order to establish the performance of the
propulsion system. This will be used to verify computational predictions made previously.
Structural Integrity: Critical structural elements such as the wing spar and fuselage will be pull
tested to the calculated structural limits of the aircraft in order to verify its robustness. These
tests will be critical for establishing the actual flight envelope of the aircraft.
Center of Gravity: The aircraft will be hung from each wingtip in order verify the location of the
empty center of gravity. This test will also verify that the aircraft will pass the competition center
of gravity test.
Flight Controls Check: The aircraft’s flight control system will be checked for proper operation.
This includes both flight control and engine control systems.
Aircraft Assembly Check: Verify that the aircraft can be constructed out of the box within the
five minute contest assembly limit.
7.2 Flight Testing
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Once the structural integrity and baseline performance parameters of the finished aircraft have
been established, the team will proceed with flight testing of the final design. The flight test program will
proceed through a series of envelope expansions described below. A chronological test plan is
presented in Figure 7.1.
Low Speed Taxi Test: The aircraft will be taxied at low speeds in order to verify the proper
operation of flight controls and to familiarize the test pilot with the aircraft’s systems.
High Speed Taxi Test: A high speed taxi test up to 80% of takeoff velocity will be conducted in
order to observe aircraft ground handling qualities at higher speeds and increase pilot
familiarity.
Center of Gravity Envelope Expansion: Upon the successful completion of taxi testing, a
flight envelope expansion phase will begin by flying the aircraft in its most stable configuration.
As the pilot gains familiarity with the aircraft, the center of gravity will be located at increasingly
aft fuselage stations until the aircraft’s aft center of gravity limit is reached. This will allow the
pilot to methodically increase familiarity with the aircraft’s handling tendencies in all flight
configurations. The center of gravity will initially be shifted by moving a light weight to various
fuselage stations. This will allow for the center of gravity to be shifted while also keeping the
plane light. Once the stability characteristics have been established for different center of
gravity locations, heavier payloads will be tested.
Endurance Testing: Once the aircraft has been deemed airworthy in all center of gravity and
weight configurations, the endurance of the vehicle will be established. This test phase will
require the airplane to cruise at full throttle with competition payloads until the batteries are
depleted. This test will ensure that the aircraft is capable of completing the competition flights.
Advanced Maneuvers: The aircraft will also be tested to the structural limits established during
static testing by engaging in steep turns. These turns will establish the pilot’s entry parameters
for the crosswind, base, and 360 degree turn legs of the flight. Also, the team will attempt to
perfect the short field takeoff technique for the aircraft so that a larger bat payload may be
carried and successful takeoff per competition minimums on all flights will be ensured.
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3/25/2010
3/30/2010
4/4/2010
4/9/2010
4/14/2010
Rolling Resistance Coefficent
Static Thrust
Structural Integrity
Center of Gravity
Flight Controls Check
Aircraft Assembly Check
Low Speed Taxi Test
High Speed Taxi Test
Center of Gravity Envelope Expansion
Endurance
Advanced Maneuvers
Planned
Figure 7.1 Testing Milestone Chart
7.3 Checklists
The day of flight checklist is shown in Table 7.1. This checklist was developed in order to
provide the maximum chance for successful flight during each mission.
Table 7.1 Aircraft Checklist
Check Lists
PREFLIGHT INSPECTION
Verify payload door SECURED
Verify external payload SECURED
Inspect propeller structural integrity
Verify propeller SECURED
Inspect aircraft wiring for damage
Verify all wiring connections
Verify battery cover SECURED
Verify aircraft power is ON
Verify engine cowl SECURED
Zero all servos
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Flight controls FREE and CORRECT
Radio range check
Verify aircraft area clear of FOD
ENGINE START
Clear propeller area
Perform engine run-up
8 Pre-Competition Performance Results
8.1 Rolling Resistance Testing
The rolling resistance coefficient of the aircraft’s landing gear was a critical component of
estimating takeoff distance. Estimated values for the rolling resistance were presented in Anderson for
both asphalt and concrete surfaces. However, these values were determined for aircraft tires
significantly larger than those used on the team’s aircraft. The team decided to test the rolling
resistance of another aircraft of similar size to the one being constructed for the 2010 competition in
order to have a higher fidelity estimate. The coefficient was determined by pushing the aircraft with a
small scale at a constant velocity. Using this testing method, the aircraft tires to be employed on the
current aircraft were found to have a rolling resistance coefficient of approximately 0.05 on a hard wood
floor. Due to the climatologic difficulties posed by a particularly wet winter, the concrete and asphalt
surface tests were unable to be completed.
8.2 Structural Testing
During the airplane design process, the team decided that it would be advantageous to
research the structural properties of several different composite material combinations built with similar
construction techniques as those available for use on the actual aircraft. It was decided that foam and
balsa wood pieces laminated with carbon fiber or fiberglass would be the most worthwhile materials to
examine. This was because the use of a hollow, composite covered skin was being considered as an
alternative to traditional rib and spar construction for the fuselage and wing structures. If this
construction technique was to be used, it was important to establish the strength of the finished
laminates prior to proceeding with the design. Additionally, it was important to research the material
properties of these composite lay-ups to see if there were any other applications the team may not have
considered previously.
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Balsa wood of 0.0938 inch thickness and Cellfoam 88 polystyrene foam of 0.125 inch thickness,
commonly used in model airplane construction, were selected as the substrate materials. Bi-directional
plain weave 5.8
oz
yd 2
carbon fiber and 6 oz yd 2 bi-directional weave fiberglass cloth were selected as the
overlay materials. By overlaying both the balsa wood and foam with carbon fiber and fiberglass four
samples were constructed; a carbon fiber-balsa sample, a carbon fiber-foam sample, a fiberglass-balsa
sample, and a fiberglass-foam sample. The substrate materials were laminated on both top and bottom
with the fiber cloths to insure that the samples would behave in a similar manner independent of the
direction of loading normal to the samples.
The construction method used to create the laminates was a squeegee technique which
removed excess resin through pressure applied by a tongue depressor. West System 105 epoxy resin
mixed with West System 205 fast hardener was applied to the surface of the substrate, either foam or
balsa wood, and the tongue depressor was then used to spread an even coating of resin. Once the
sample had been fully coated in resin, the depressor was then used to remove excess resin through the
application of pressure to the fiber blankets. The lay ups were completed on top of a smooth plastic
surface covered with a visqueen poly sheet. Despite using this resin removal method, some excess
resin collected on the backside of the samples during the curing process. Samples tended to have a
side with a greater amount of excess resin than the other, which was referred to as the dirty side.
It was decided that the samples would be tested as a beam simply supported at two points and
loaded at the midpoint between the supports. The setup used to test the samples consisted of wooden
supports, a dial gauge, and a digital scale arranged as shown in Figure 8.1.
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Figure 8.1 Sample Test Setup
Each sample was placed so that it spanned the two wooden supports spaced 5 inches apart. A
thin wooden strip spanning the width of the sample was placed in the center of the test section so as to
apply a more even force distribution across the sample. By pressing down on the wooden strip and
monitoring the readout on the digital scale, specified loads were applied to the sample. The dial gauge
displayed the displacement that, combined with the measured force, was used to find the elastic
modulus of the sample
The elastic modulus was calculated for the deflection associated with each applied load. These
values were then averaged to determine the average modulus of elasticity for each side, dirty (excess
resin side in compression) and clean (excess resin side in tension). The average modulus of elasticity
each sample type was then found by averaging the values found for the clean and dirty sides. The
results of these calculations are shown in Table 8.1. Additionally, to gage the weight contribution of
each laminate, the weight per unit surface area of each sample was calculated so as to negate the
differences in thickness between the carbon fiber and the fiberglass. These results are also recorded in
Table 8.1.
Material
Combination
Balsa/Fiberglass
Balsa/Carbon
Table 8.1 Results for Elastic Modulus Tests
Weight per Unit
Elastic Modulus Elastic Modulus
2
Surface Area (lb/in )
(Mpsi) clean
(Mpsi) dirty
0.0022
1.30
1.40
0.0016
2.18
50
2.31
Elastic
Modulus (Mpsi)
1.35
2.25
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Fiber
Foam/Fiberglass
0.0011
0.28
0.26
0.27
Foam/Carbon
Fiber
0.0012
0.34
0.35
0.34
The carbon fiber/balsa samples had the greatest modulus of elasticity of all the samples tested
and had a lower weight per unit surface area than balsa coated with fiberglass. It was also found that
by having the dirty side in compression, the modulus of elasticity was increased by approximately 6%
versus having it in tension. However, in actual construction of the plane, the excess resin would be
removed in order to save weight. Averaging the modulus of elasticity for the dirty and clean sides, an
average modulus of elasticity of the carbon fiber/balsa sample was found to be approximately 2.25
Mpsi.
The fiberglass/balsa sample had the second highest modulus of elasticity of the samples tested.
Using the same methodology as was applied to the previous sample, the average modulus of elasticity
was found to be 1.35 Mpsi. The dirty side in compression was found to have an elastic modulus
approximately 8% larger than in tension. As previously noted, the fiberglass/balsa samples had a
greater weight per unit surface area than the carbon fiber/balsa samples. This is thought to be because
it was more difficult to remove excess resin from the fiberglass than the carbon fiber using the squeegee
technique. The fiberglass cloth was more susceptible to being pulled or tattered during the lay up
process.
The foam substrate samples performed markedly different from the balsa substrate samples.
The foam samples were much less rigid and failed at much lower load than their balsa counterparts.
The carbon fiber/foam failed at approximately 12 lbs while the carbon fiber/balsa samples were
repeatedly tested to 40 lbs or greater without failure. The carbon fiber/foam samples possessed an
average modulus of elasticity of 0.34 Mpsi.
Testing of the fiberglass/foam samples revealed that it possessed and average modulus of
elasticity of approximately 0.27Mpsi, making it the least rigid of all the samples tested. Additionally, this
sample type routinely failed at approximately 9lb of loading.
Based on the data shown above, the carbon fiber/balsa samples exhibited the best compromise
between strength and weight of the material combinations tested. The carbon fiber/ balsa combination
weighed less and was stronger than its fiberglass counterpart. Additionally, it was only about 30%
heavier than the foam samples with an elastic modulus that was nearly an order of magnitude greater.
Thus, it was concluded that the major structural components requiring strength and low weight such as
the fuselage spar would be made out of balsa coated with carbon fiber. Additionally, it was decided that
due to the strength exhibited by the carbon fiber, it could be used to wrap the leading edge of the wing if
needed.
Both foam samples proved to be weaker than their balsa counterparts both in terms of modulus
of elasticity and ultimate tensile strength. Further examination of the mode of failure of the foam
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samples revealed that the foam substrate failed in compression first causing the overlaying material to
become delaminated from the substrate. This caused a drastic loss in strength. Despite these
shortcomings, it was still decided that the wing could be cut out of foam and then overlaid with carbon
fiber. However, this design choice was eliminated early in the team’s design process for reasons
detailed previously.
The test procedure and results analyzed above were intended to serve as a low order
approximations of material properties associated with materials and construction techniques that could
be used by the aircraft build team. While averaging the theoretical elastic modulii at each data point
was not the most accurate data reduction method available, it was determined to be the most
appropriate in this test due to its ease of implementation. The goal of the testers was simply to establish
a low order quantitative comparison of building methods. The performance of the carbon fiber
laminated balsa sample gave the group confidence that this construction method would be robust
enough to use for major structural elements on the aircraft.
8.3 Systems Testing
Unfortunately, the complete aircraft solution was unable to be tested by the report due date.
Additional testing on the aircraft is planned for when the aircraft construction is complete, as was
detailed in Section 7 of this report. A particular subsystem test of interest to the group is the propulsion
system. It will be critical for the group to establish the static thrust output and endurance of the motor
and batteries respectively. Again, manufacturing and acquisitions delays prevented these tests from
being completed before the report deadline but will be conducted prior to competition.
9 Pre-Crash Construction
The fuselage construction method changed dramatically between the final design stage and the
actual construction of the plane. Originally it was decided to use a rib and spar construction method for
the fuselage, and then cover the ribs in Monokote. It was discovered after the final design iteration that
the design team had access to carbon fiber materials and expertise. During our design stage rib and
spar construction was favored over carbon fiber composite due to the fact that the design team had no
significant experience working with composite materials. With the expertise of the team’s advisor it was
possible to form most of the fuselage out of two ply carbon fiber, using a vacuum bagging technique.
This offered several advantages over the traditional fuselage design. The first advantage was weight.
A carbon fiber composite fuselage could weigh less than the rib and spar construction if great care was
taken during the construction process. The second advantage was survivability. In the case of a crash
a carbon fiber fuselage could be patched, while a rib and spar fuselage would have to have the ribs
replaced and the Monokote skin redone. The third advantage was strength. A carbon fiber fuselage
would have been capable of carrying part of the force load through the skin of the fuselage itself, while
Monokote would not be capable of carrying any load other than the drag resulting from flight. The last
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advantage to the carbon fiber fuselage was aerodynamics. Since a carbon fiber fuselage would be a
single piece cut in half for the cargo hatch, it was possible to sand it to a very smooth and streamlined
surface. A rib and spar design would be covered in Monokote and would have the wrinkles associated
with that covering method. This would cause a higher skin drag coefficient for the fuselage. Due to all
these advantages it was decided to use a carbon fiber composite fuselage.
There were several other additions to the fuselage that occurred during construction. The first
was the main spar. The main fuselage spar consisted of a 3/8 inch by 2 inches rectangular outboard
sheet of balsa coated on either side with carbon fiber. This provided strength and also a place for the
cargo carrier and the wing main spar to attach. At the front of the fuselage a sheet of 3/32 inch balsa
layered with carbon fiber on both sides was used as a firewall for the engine to attach to. The main
spar, cargo section, and the firewall were attached to the fuselage using a mix of resin, 404 structural
epoxy, and short strips of carbon fiber. In addition, a blister to hold the battery pack during the internal
cargo flight was constructed out of two ply carbon fiber composite and was removable from the fuselage
through the use of recessed nuts inside the fuselage and small bolts. The main gear was custom made
using a mold with 12 layers of fiberglass and 1 layer of carbon fiber. This was preferable to an
aluminum gear because the composite gear was lighter and would provide dampening during landing
instead of being rigid.
The next problem that surfaced during construction was the application of the cargo hatch. As
designed with a mid-wing, the cargo hatch would be split along where the wing main spar entered the
fuselage. This was solved by lowering the wing’s attachment point on the fuselage and adding four
degrees of dihedral to retain some roll stability. The wings attached to the fuselage through two points.
The first was the carbon fiber main spar. The main spar for the wing attached to the main spar of the
fuselage through a phenolic tube that was attached using epoxy into the fuselage. The second
attachment point was the wing’s secondary spar. This secondary spar tied into the fuselage through a
hole in each side of the fuselage and was attached via two set screws that slid through two tabs built
into the fuselage itself.
The wings were constructed using a rib and spar method with a Monokote skin covering. The
inboard section of the wing main spar was made of a 0.88 inch outer diameter carbon fiber tube for
added strength at the wing root. It was thought that in case of a wing tip strike the wing roots would
need to be stronger to withstand the impact. In addition, the inboard section of the wings carried an
external payload during one of the flight missions. The brackets for the payload mounting were made
out of carbon fiber molded around the phenolic tube on the inboard section of the wing. These
attachments were secured to the wing via four bolts. The wing section outboard of the carbon fiber tube
had a built in dihedral of 4 degrees. The main spar for the outboard section was composed of 5/32 inch
by 1 inch balsa wood. This tied into the phenolic tube in the inboard section that the carbon fiber tube
slid into. Stringers made of thin balsa ran span-wise down the wing to provide more structural support
as well. The ribs for the inboard and outboard sections of the wing were both cut using a laser CNC
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machine. The ailerons were composed of thin balsa covered by Monokote. The hinges for the ailerons
were also made of Monokote. The servos for the ailerons were attached to mounts that tied into the
main spar within the wing.
The vertical stabilizer too was built using rib and spar construction. The ribs were cut using a
CNC laser machine. The main spar was a single piece of balsa wood, 5/32 inch by ½ inch. The rudder
was made of thin balsa wood covered by Monokote. The rudder was attached to the fixed portion of the
vertical stabilizer by a Monokote hinge. The vertical stabilizer connected to the fuselage by inserting the
main spar into a set hole in the fuselage, and then it was attached in place with 404 “Structural
EEEEEEpoxy.” The tail gear was kindly connected to the rudder via tie rods. The servos for both the
vertical and horizontal stabilizers were installed flush mounted to the fuselage.
The horizontal stabilizer was also constructed using the rib and spar method. The ribs were
again cut using a CNC laser cutter. The main spar for the horizontal stabilizer was a thin carbon fiber
tube. The entire horizontal stabilizer was used as the elevator, with the carbon fiber tube as an axle and
a phenolic tube attached inside the fuselage using 404 structural epoxy as its bearing.
10 Crash
Though unfortunate, crashing the plane provided us with the opportunity to make slight
adjustments to the design. The components modified during the rebuild process were the vertical
stabilizer, the wings, and the ball holder. All of the other components were simply repaired.
After our initial building process, we tested our plane under wing tip loading as this was a
requirement to pass technical inspection at the competition. Before the crash, during our first attempt at
this test, we heard cracking near the root of the starboard wing. This led us to add additional balsa
wood reinforcements along the main spar. Also, we added a strip of a thin continuous strip of balsa
wood along the bottom side of the wing. This ensured that the wing would be placed in tension when it
was loaded, bottom side down, and thus would act to carry a significant component of the load along
the wing. At this time we also considered placing an additional strip of carbon fiber along this piece of
balsa wood, but decided not to. We thought that this modification would make the plane overbuilt, and
was therefore not worth the extra weight. After making these modifications, we again performed the
wing-tip loading test and found it to be acceptable and proceeded with the initial flight test.
During the first flight test, the vertical stabilizer spar snapped and de-bonded from the fuselage.
The weakest point of the stabilizer was the main spar due to its incorrect orientation which placed the
aerodynamic loads along the weakest axis of the rectangular balsa spar. In addition to the spar failure,
the stabilizer delaminated from the fuselage. When the stabilizer was bonded to the fuselage, no
bonding agent was used that would permeate past the first layer of balsa wood. Because of this the first
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layer of balsa stuck to the fuselage while the rest of the rib ripped off, leaving the entire vertical stabilizer
attached through the control rod but not attached to the fuselage. This created a moment that forced
the plane into a spin. The plane fell from around 200 feet in the air and landed in brush at the base of
several trees. The wings were almost completely destroyed by the fall through the tree branches and
the nose was badly cracked since it landed nose down. There were several large holes in the carbon
fiber fuselage but it was not damaged irreparably. The crash scene photos can be seen below. The
point of failure on the vertical stabilizer is also shown below.
Figure 10.1 Crash Scene
Figure 10.2 Vertical Stabilizer After Crash
After the crash, the vertical stabilizer was improved by adding several reinforcements. In the
original design, balsa wood was used for the spar, but in the redesigned component we made the
decision to use bass wood as it exhibited a greater modulus of elasticity. The spars were turned so that
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the orientation placed the largest aerodynamic loads on the correct axis to ensure the greatest structural
integrity. Furthermore, we added a secondary spar that helped reduce the torsional load at the joint
between the carbon fiber fuselage and the bass wood main spar; and ensure that the alignment of the
stabilizer would stay fixed.
Because of the crash, we were given the opportunity to completely rebuild our wings. Before
the crash our design was acceptable, but the rebuild allowed us to go ahead with the other structural
reinforcements that we had previously discussed. Therefore, the new wings were built as near to the
originals as we could achieve with the exception of an added carbon fiber reinforced strip of balsa
running along the bottom of both wings. This would act to improve the overall structural integrity of the
wings.
Another component that was modified during the rebuild process was the ball holder. Originally
this component had been made of balsa sheeted with fiberglass. The strength of the fuselage frame had
been compromised due to the holes made by the crash, therefore the ball holder needed to be stronger.
Thus, when the balsa ball holder was rebuilt, it was sheeted with carbon fiber. Also, an extra layer of
carbon fiber was added to the firewall to increase its strength and rigidity. This design exhibited both a
higher ultimate strength and a greater modulus of elasticity than the first design.
11 Testing
11.1 Static Thrust
Static thrust data was collected by for the AXI 4130/20 motor connected to either of the two
battery pack types used during the project. The first pack type used during the competition was a 4000
milliamp-hour pack consisting of 26 sub-C cells. The second pack consisting of 25 sub-C cell 4600
milliamp-hour batteries was ordered as a backup plan incase of delivery or operational problems with
the 4000 milliamp-hour pack. The 4600 milliamp-hour batteries were not used during the competition
due to their higher weight. The objective of the test was to collect thrust, voltage, current, and power
data as a function of time for each pack type. The motor was operated a full throttle for as long as the
packs would allow.
A fish scale was connected to the back of the plane with a rope to measure the thrust in pounds
when the controller was set to full throttle. A power analyzer was connected to the battery pack to take
real time data for the amperage, wattage, and voltage. The trials were run for approximately eight
minutes per pack and data was recorded in thirty second increments. The speed controller and
transmitter settings were identical for each pack.
The forty amp current limit set by the competition rules and the blade fuse was used during
testing. The team’s goal was to draw as close to 40 amps as possible without exceeding the limit.
Figure XX shows the amperage measured during static thrust testing as a function of time. The larger
4600 milliamp-hour battery pack had a higher current draw throughout the entire run than the 4000
milliamp-hour battery.
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35
30
25
Amps
20
4000 mAh
4600 mAh
15
10
5
0
0
50
100
150
200
250
300
350
400
450
500
Time (s)
Figure 11.1 Amperage as a Function of Time
As shown in Figure 11.1, the peak amperage occurred at the first data point, where the stored
energy in the battery was highest and the pack’s temperature was lowest. After approximately 30
seconds, the current in each pack approached a steady state current that remained constant for several
minutes. The 4000 milliamp-hour pack’s performance began to decline at five minutes while the 4600
milliamp-hour battery pack did not severely decline until nearly seven minutes run time.
Figure 11.2 displays the voltage measured during the test with respect to time. The expected
voltage for the 4000 milliamp-hour pack was 31.2 volts and 30 volts for the 4600 milliamp-hour pack. As
shown in the figure, the voltage measured for the 4600 milliamp-hour pack was consistently higher than
that for the 4000 milliamp-hour pack. The measured voltages for both packs were both lower than
those expected based on the number of cells. The difference could be explained by the higher internal
resistance expected for the lower rated batteries. The voltage measurements for each pack behaved
similarly to the current in that they were nearly constant over some range of the elapsed time. However,
the voltage in both batteries began declining earlier than the current. The main difference between the
two packs was that the 4000 milliamp-hour pack’s voltage began to fall much sooner than the 4600
milliamp-hour pack, but this was to be expected due to the difference in capacity.
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29
28.5
28
27.5
Volts
27
26.5
4000 mAh
4600 mAh
26
25.5
25
24.5
24
23.5
0
50
100
150
200
250
300
350
400
450
500
Time (s)
Figure 11.2 Voltage as a Function of Time
The power available to power the motor as measured by the power meter is shown in Figure
11.3. As can be seen in the figure, both batteries supplied nearly constant power over some time
range. The power supplied by the 4000 milliamp-hour batteries began declining after approximately 180
second. The 4600 milliamp hour pack’s power did not begin to decline until nearly 300 seconds. The
figure shows that although the smaller capacity pack had a higher rated voltage based on the number of
cells in the pack, it produced nearly 10 percent less power.
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900
800
700
Power (W)
600
500
4000 mAh
4600 mAh
400
300
200
100
0
0
50
100
150
200
250
300
350
400
450
500
Time (s)
Figure 11.3 Power Available from vs. Time
Figure 11.4 shows the force recorded by the fish scale as a function of time for both battery packs. The
force measurements showed that the values between 90 and 150 seconds were approximately equal
for both packs. This was despite the power supplied to the motor being higher for the 4600 milliamphour pack. It would seem counterintuitive that the motor would produce the same thrust despite the
higher power setting. This could be explained by the propeller efficiency at zero velocity for each case.
The electric motor is a nonlinear device, and its revolution rate is sensitive to current and voltage
changes. This speed of revolution is a key factor in propeller efficiency. Thus, it could be possible that
the lower power setting is running at a more efficient condition, thus generating a similar static thrust to
the higher powered condition. The static thrust was constant up to approximately 150 seconds for the
4000 milliamp-hour pack and 300 seconds for the 4600 milliamp-hour pack.
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7
6
Force (lbs)
5
4
4000 mAh
4600 mAh
3
2
1
0
0
50
100
150
200
250
300
350
400
450
500
Time (s)
Figure 11.4 Force vs. Time
11.2 Rolling Resistance
The rolling resistance was measured by two different methods. The first used a fish scale and
the second was by using Cen-Tech compression scale. In the fish scale method, the aircraft was pulled
by a rope attached to the scale. In order to establish repeatability of the results, the compression scale
was used to push the aircraft by the tail. Four trials were done with each scale for each surface tested
for a total of eight trails per surface. The aircraft’s rolling resistance was tested on asphalt and concrete.
These values are shown below in Table 11.1. These results show that the concrete produced a lower
rolling resistance coefficient, as would be expected due to its finer surface.
Table 11.1 Rolling Resistance Predictions
Fish Scale
Cen-Tech Scale
Average
Concrete
0.071
0.064
0.067
Asphalt
0.075
0.077
0.076
12 Revised Performance Predictions
The aircraft’s performance predictions were recalculated using parameters measured from the
finished aircraft. These results are presented in Table 12.1.
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Table 12.1 Recalculated Performance Parameters
Parameter
Empty
10 Balls
2 Bats
Weight (lbs)
10.4
14.4
14.2
Stall Velocity (ft/s)
34.8
40.9
40.6
Maximum Velocity (ft/s)
119
119
109
Maximum Rate of Climb (ft/min)
1200
810
730
Velocity for Maximum Rate of Climb (ft/s)
78.9
87.3
80.1
Takeoff Distance (ft)
57.9
129
126
2
Wing Loading (lb/ft )
1.73
2.40
2.37
Power Loading (shaft W / lbf)
53.1
38.3
38.9
48
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13.2
13.2
11.3
Maneuver Speed (ft/s)
Maximum Lift to Drag Ratio
The takeoff and maximum velocity prediction codes for the aircraft were verified via flight test.
The aircraft’s takeoff distance was established by marking the runway in five foot increments and
reviewing a video recording of the takeoff roll. The maximum velocity of the aircraft was measured
using a radar gun. The measured and predicted takeoff and maximum velocities for the aircraft are
shown in Table 12.2.
Table 12.2 Comparison of Predicted and Measured Parameters
Parameter
Measured
Predicted
117.4
119
1.4
Takeoff Distance (ft)
45
57.9
28.7
Takeoff Time (s)
2.5
2.85
14.0
Maximum Speed (ft/s)
% Difference
As can be seen in the table, the prediction codes over predicted maximum speed of the aircraft
and its takeoff distance. The maximum velocity predicted was within five percent of the value measured
with the radar gun. The takeoff distance displayed a substantially greater error of nearly 30 percent.
This suggested that the drag and thrust predictions for the aircraft at high speeds were more accurate
than those at lower velocities. The validity of the thrust and drag predictions were confirmed by
calculating the drag coefficient for the actual flight test using measured maximum velocity and the
theoretical thrust obtained from propeller efficiency curves and manufacturer data. These results are
shown in Table 12.3.
Table 12.3 Predicted vs. Measured Drag Coefficient
Drag Coefficient
Measured
Predicted
0.028
0.027
61
% Difference
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As shown in the table, the drag coefficient from the measured maximum velocity calculated
using theoretical thrust compared well with the drag coefficient predicted by the team using empirical
methods. The lift coefficient required by the empty aircraft flying at maximum speed was 0.12. At this
low lift coefficient, the majority of drag was parasite drag and was a weak function of lift. This suggested
that the group’s zero lift drag predictions were accurate to within approximately five percent. The close
agreement of the measured and predicted drag coefficients also gave credence to the group’s thrust
predictions at high velocities. Had either the drag coefficient or the thrust predicted been significantly
wrong, the results shown in Table 12.3 would have varied greatly.
The large discrepancy in the takeoff code could be the result of several errors. First, the low
speed thrust estimates were suspect. Static thrust tests were performed, but these measurements did
not indicate the change in the thrust as a function of velocity. Additionally, the performance of the
motor, especially the revolutions per second of the propeller, was a major contributor to thrust
predictions. Manufacturer supplied data was used to estimate the efficiency of the motor and its
revolution rate at the voltage and current settings measured during static testing. However, the
performance of the motor at non-static conditions was unknown and thus a potential contributor of large
errors. Low thrust predictions could be to blame for the exceedingly large takeoff distance predicted by
the code.
Another potential pitfall in the takeoff code could have been the low speed drag predictions.
The aircraft was set at a large angle of attack by virtue of its tail dragger configuration. Thus, the wings
began developing lift and, therefore, induced drag as soon as the aircraft began rolling. The takeoff
code did not account for any reduction of induced drag due to ground effect. The code also failed to
account for the ground effect’s tendency to increase lift near the surface of the ground due to its
destruction of the wing’s downwash. Both of these factors would tend to reduce takeoff distance from
that predicted.
The static stability margins for the aircraft were recalculated using measured center of gravity
and aircraft geometry. The aircraft’s payload loading system was designed to maintain a static margin
of approximately 15% by sliding the location of the engine’s batteries. These results are shown in Table
12.4.
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Table 12.4 Static Stability Margins
Static Margin (Percent)
Small balls
#
0
1
2
3
4
5
6
7
8
9
10
0.18
0.18
0.19
0.19
0.20
0.18
0.19
0.19
0.20
0.20
0.18
0.19
0.19
0.20
0.20
0.18
0.19
0.19
0.20
0.20
0.18
0.19
0.19
0.20
0.20
0.18
0.19
0.19
0.20
0.20
0.20
0
1
2
Big Balls
3
4
5
6
0.18
0.19
0.19
0.20
7
0.19
0.19
0.20
0.20
8
0.19
0.20
0.20
9
0.20
0.20
10
0.20
63
Team
VOLocity
13 Works Cited
1. Abbott, I.H and Von Doenhoff, A.E., Theory of Wing Sections. New York: Dover Publications,
1959
2. “2009/10 Rules and Vehicle Design.” 31 Oct, 2009
<http://www.aiaadbf.org/2010_files/2010_rules.htm>
3. Anderson, John D. Aircraft Performance and Design. United States: McGraw-Hill, 1999.
4. “AXI Gold 4130/20 Brushless Outrunner Motor.” BP Hobbies. 14 Oct, 2009
<http://www.bphobbies.com/view.asp?id=B0658308&pid=AXI111>.
5. “AXI 4130/20 Gold Line.” Model Motors. 10 Jan, 2010
<http://www.modelmotors.cz/index.php?page=61&product=4130&serie=20&line=GOLD.>
6. Raymer, Daniel P. Aircraft Design: A Conceptual Approach. Third Ed. Reston, Virginia: AIAA,
1999
7. Rice, M.S., Airfoil Sections for Lighter Aircraft. Milwaukee, Wisconsin: Aviation Publications:
1971
8. Roskam, Jan. Airplane Design Part VI. Lawrence, KS: DARcorporation, 2000.
9. “Sub C NiMh Rechargeable Batteries.” Battery Junction. 20 Jan, 2010
<http://www.batteryjunction.com/red-subc4500-tab.html.>
64
Team
VOLocity
Appendix A: Drawing Package
65