RIT Micro Air Vehicle
Preliminary Design Report
February 2005
Brian Gillis Team Leader
Mechanical Engineering
Joshua Baker Mechanical Engineering
Victoria Schoennagel Mechanical Engineering
Aimee Lemieux Mechanical Engineering
Aaron Grilly Mechanical Engineering
Tzu-Chie Fu Computer Engineering
Cuong Le Computer Engineering
David Hein Mechanical Engineering
Atul Phadnis Electrical Engineering
J.E.D. Hess Mechanical Engineering
Dr. Jeffrey Kozak Team Advisor
Mechanical Engineering
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Table of Contents 1 INTRODUCTION AND MOTIVATION ......................................................................................... 9
1.1 MAV ........................................................................................................................................... 9 1.2 MAV USES .................................................................................................................................. 9 1.3 MAVS FOR SURVEILLANCE........................................................................................................ 10 1.4 STATUS OF MAV DEVELOPMENT WORLDWIDE ......................................................................... 10 1.5 CURRENT STATUS OF THE RIT MAV TEAM............................................................................... 11
2 TEAM ORGANIZATION AND WORK BREAKDOWN............................................................ 13
3 LITERATURE REVIEW................................................................................................................. 15
3.1 AIRFRAME.................................................................................................................................. 15 3.2 PROPULSION............................................................................................................................... 16 3.3 ELECTRONICS............................................................................................................................. 20
4 NEEDS ASSESSMENT.................................................................................................................... 23
4.1 PERFORMANCE GOALS............................................................................................................... 23 4.2 VEHICLE GOALS......................................................................................................................... 23 4.3 AIRFRAME.................................................................................................................................. 23 4.4 PROPULSION............................................................................................................................... 24 4.5 ELECTRICAL SYSTEMS ............................................................................................................... 24
5 CONCEPT DEVELOPMENT AND FEASIBILITY..................................................................... 26
5.1 AIRFRAME.................................................................................................................................. 26 5.1.1 Planform Design................................................................................................................... 26 5.1.2 Vehicle Configuration........................................................................................................... 27 5.1.3 Airfoil Geometry ................................................................................................................... 29 5.1.4 Engine and Engine Mounting Configuration........................................................................ 30 5.1.5 Materials and Processes Selection ....................................................................................... 31 5.1.6 Flight Controls ..................................................................................................................... 32
5.2 ELECTRONICS............................................................................................................................. 33 5.2.1 Control System...................................................................................................................... 33 5.2.2 Video System......................................................................................................................... 36 5.2.3 GPS System........................................................................................................................... 38 5.2.4 Servos (Actuators) ................................................................................................................ 41 5.2.5 Antenna................................................................................................................................. 42 5.2.6 Batteries................................................................................................................................ 44
5.3 PROPULSION............................................................................................................................... 45
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5.3.1 Concept Development........................................................................................................... 45 5.3.2 Feasibility Assessment .......................................................................................................... 48
6 DESIGN OBJECTIVES AND SPECIFICATIONS....................................................................... 55
6.1 PERFORMANCE SPECIFICATIONS ................................................................................................ 55 6.2 DESIGN OBJECTIVES .................................................................................................................. 55 6.3 EVALUATION CRITERIA.............................................................................................................. 56
7 ANALYSIS ........................................................................................................................................ 57
7.1 AIRFRAME.................................................................................................................................. 57 7.1.1 Airfoil Testing Design........................................................................................................... 57 7.1.2 Hot Wiring ............................................................................................................................ 58 7.1.3 Fiberglassing........................................................................................................................ 62 7.1.4 Wind Tunnel Testing............................................................................................................. 66
7.2 ELECTRONICS............................................................................................................................. 69 7.3 PROPULSION............................................................................................................................... 71
7.3.1 Propulsion Static Testing...................................................................................................... 71 7.3.2 Propulsion Dynamic Testing ................................................................................................ 75 7.3.3 Future Testing Plans ............................................................................................................ 78
8 REFERENCES.................................................................................................................................. 82
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List of Figures Figure 1.1: Over the Hill Reconnaissance Mission and Chemical Contaminations
Monitoring Figure 5.1: Examples of different planforms, showing trailing vortices Figure 5.2: The BYU Stableyes is an example of a conventional planform with a pusher
motor configuration Figure 5.3: The Black Widow is an example of a modified inverse Zimmerman planform
with a puller motor configuration Figure 5.4: Separation bubble on airfoils at low speeds. Figure 5.5: MIT Lincoln Laboratory MAV Concept Figure 5.6: R4P-JST Receiver Figure 5.7: R-6N/H (Horizontal) Figure 5.8: R-6N/V (Vertical) Figure 5.9: Speed controller YGE3 Figure 5.10: Speed controller Phoenix-10 Figure 5.11: The 200 mw Brown Bag Kit – Transmitter/Receiver Set (Left), 8dBi Patch
Antenna for Base Station (Right) Figure 5.12: Furuno GH-79 Figure 5.13: Sarantel Smart Antenna Figure 5.14: UNAV PICO-GPS-SS Figure 5.15: UNAV OSD-GPS Video Overlay Board Figure 5.16: LS-2.0 Servo Figure 5.17: LS-3.0 Servo Figure 5.18: LS-2.4 Servo Figure 5.19: Antenna Array w/ Radiation Maps Figure 5.20: Batteries – Kokam Figure 5.21: Propeller propulsion (left) and rocket propulsion (right) Figure 5.22: Purchased propeller (left) and mold for propeller fabrication (right) Figure 5.23: Attribute Ranking System Figure 5.24: Feigao 1208430S 12x22mm Brushless Motor Figure 5.25: EP7060 Propeller Figure 5.26: RXC Light Power System (GW/LPS-RXC-A) Figure 5.27: EP7043 Propeller Figure 5.28: Feigao 1208436L 12x30mm Brushless Motor Figure 5.29: EP3020 Propeller Figure 7.1: Blocking out foam Figure 7.2: Attach airfoil template to foam Figure 7.3: Setting up foam and bow for hot wiring Figure 7.4: Counter weight used to create wire movement Figure 7.5: Utilize airfoil templates to cut out rest of airfoil Figure 7.6: Airfoil after hot wiring process Figure 7.7: Layout of supplies Figure 7.8: Sizing fiberglass Figure 7.9: Laying out fiberglass Figure 7.10: Saturating fiberglass with epoxy
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Figure 7.11: Laminating airfoil with fiberglass Figure 7.12: Weight applied to airfoil during curing process Figure 7.13: Airfoil attached to mounting rod Figure 7.14: Airfoil mounted on balance Figure 7.15: SMD S250 Miniature Platform Load Cell Figure 7.16: Calibration electrical schematic Figure 7.17: Calibration setup Figure 7.18: Load cell calibration Figure 7.19: Motor Mount Figure 7.20: Motor test electrical schematic Figure 7.21: Existing wind tunnel setup (left) and new MAV dynamic test setup (right) Figure 7.22: Strain gage mounting locations Figure 7.23: Propeller modification A (left) and B (right) Figure 7.24: Example dynamic test matrix.
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List of Tables Table 3.1: Reynolds Number Range Calculations Table 5.1: Airfoil Normalized Feasibility Chart Table 5.2: RF Receivers Table 5.3: Pugh Chart for RF Receiver Selection Table 5.4: Speed controller alternatives Table 5.5: Pugh Chart for Speed Controller Selection Table 5.6: Camera Specifications Table 5.7: Camera onboard On-board GPS Receiver Table 5.8: Pugh Chart for Video Transmitter Selection Table 5.9: GPS Receivers w/ antenna Table 5.10: Pugh Chart for GPS Receiver Selection Table 5.11: Video Overlay Board from UNAV Table 5.12: Wes-Technik Servos Table 5.13: Pugh Chart for Servo Selection Table 5.14: Video receiver antenna alternatives Table 5.15: Pugh Chart for Antenna Array Selection Table 5.16: Comparison of Battery Cells Table 5.17: Pugh Chart for Battery Selection Table 5.18: Weighted Scale for Propulsion Methods Table 5.19: Thrust Calculation Table 5.20: Revised Thrust Calculation Table 5.20: Relative Weighting of Attributes Table 5.21: Motor Ranking Table 7.1: Data Recording Table for Wind Tunnel Testing of Airfoils Table 7.2: Data Recording Table for Wind Tunnel Testing of Control Surfaces
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Nomenclature CD Coefficient of Drag CL Coefficient of Lift CM Coefficient of Moment C Cost ($) I Current (A) dCL/dα 3-D Lift Curve Slope (Wing) dCl/dα 2-D Lift Curve Slope (Airfoil) MAV Micro Air Vehicle CD0 Parasite Drag P Power (W) e Span Efficiency Factor T Thrust (g) v Velocity (m/s) V Voltage (V) W Weight [Mass in this Case] (g) b Wingspan c Average Chord Length AR Aspect Ratio of the Wing ρ Density ν Viscosity Re Reynolds Number AoA Angle of Attack S Surface Area Fx Applied Load in Dynamic Testing [Drag] (g) Fy Applied Load in Dynamic Testing [Lift] (g) Pcr Critical Load—When exceeded, the part buckles (N) Acs Cross-sectional Area (cm2) h Depth of square tubing (cm) Dm Drag due to the Motor (g) Dp Drag due to the Propeller (g) Dsetup Drag due to the Setup (g) FD Drag Force [from strain gage measurements] (g) c End Condition Constant FS Factor of Safety Fstraingage1 Force calculated by Strain Gage 1 (g) Fstraingage2 Force calculated by Strain Gage 2 (g) L Length of square tubing (cm) E Modulus of Elasticity (GPa) RPM Rotations Per Minute q0 Shear flow in square tubing (N/m) (l/k) Slenderness Ratio t1 Thickness of square tubing (cm) t2 Thickness of location of strain gage mounting (cm) Dt Total Drag (g)
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Sy Yield Stress (MPa) X1 Horizontal distance from applied load to resolved moment at the bottom
corner of the strain gage box (cm) Y1 Vertical distance from applied load to resolved moment at the bottom
corner of the strain gage box (cm)
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1 Introduction and Motivation
1.1 MAV
A Micro Air Vehicle, or more commonly MAV, is an airborne vehicle of a
relatively small size. According to Defense Advanced Research Projects Agency
(DARPA), a MAV is characterized by having a maximum linear dimension of less than
fifteen centimeters or approximately six inches [10]. Over the past decade, research in
MAV technology has increased dramatically due to the intellectual communities’
(Aerospace Corporations, Universities, etc.) involvement. This increase in research has
greatly increased the current knowledge of MAV capabilities, while also maintaining the
overall goal of developing a user-friendly, multipurpose MAV for use by government
agencies by 2010.
1.2 MAV Uses
Most uses for MAVs relate to surveillance operations due to their small size and
video transmission capabilities. These uses have spurred the interest of the government
into the advancing MAV technologies. Not only could MAVs be used for military
operations, but also for intelligence agencies to perform reconnaissance missions. The
MAV technology would allow for soldiers or others to perform surveillance without be
placed in a dangerous position. Figure 1.1 below shows the two typical missions already
being envisioned by the military as possible uses for MAVs. The first image shows an
over the hill reconnaissance mission, where as the second image shows the possibility of
using MAVs for chemical contamination monitoring. This second mission also alludes to
the possible implementation of sensors on MAVs to detect everything from chemical
contamination to forest fires (smoke). Thus, MAVs have other capabilities besides
simply surveillance, and the options are almost limitless in this relatively new and
basically untapped technology. MAV technology has also gained popularity due to
recent success of Unmanned Air Vehicles (UAVs) in combat situations during the
Second Gulf War.
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Figure 1.1: Over the Hill Reconnaissance Mission and Chemical Contamination Monitoring
1.3 MAVs for Surveillance In the future it seems likely that MAVs will become the most basic form of aerial
surveillance. A MAV has significant advantages over other forms of surveillance simply
due to its size and weight. The small size and weight of a MAV make it easier to
transport. A MAV can easily be transported by one person because of its low weight and
size. The small size of a MAV also gives it the advantage of being hard to see, even
when only a few hundred feet in the air. This allows covert surveillance operations to be
made without any endangerment to the operator. Even if the MAV is damaged during a
mission, the relatively low cost of creating a MAV makes it acceptable if it is lost
collecting valuable surveillance information. A typical MAV carrying both video and
GPS only costs a few thousand dollars, and can be mass produced relatively easily. Thus,
MAVs seem to be a very good option for future use during important reconnaissance
missions.
1.4 Status of MAV Development Worldwide
Starting in 1993, DARPA began providing funding support to spark interest in
MAV technology. This initial government funding sparked the involvement of many
universities and aerospace companies into MAV design. This initial researched helped
build a solid foundation for future MAV work. In 1996, DARPA contracted the
aerospace company AeroVironment to create the first Micro Air Vehicle to conform to
DARPA’s standards. AeroVironment succeeded in creating the Black Widow in 1998,
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and this MAV began to be tested by government agencies and the research community as
a whole. The Black Widow was six inches in length, and was capable of transmitting
video, GPS, altitude, velocity, and heading information back to the pilot. This MAV has
become the litmus test for all future MAVs, and has proved to be a good basis for all the
subsequent work in the field.
After the successful creation of the Black Widow MAV, the government stopped
sponsoring large-scale MAV development. Even without the government’s funding,
MAV research continued to flourish at universities around the world. To help continue
this research and advance MAV technology, an annual MAV competition was held at one
of the several universities actively performing MAV research. This year’s competition
marks the ninth annual MAV competition, and also marks the first time it will be a truly
international competition with it being held in Seoul, South Korea. This competition has
proved to be a tremendous tool to advancing the MAV technology due to sharing of ideas
between the many competitors. MAV creation contains many pitfalls, but competition
has helped create a greater knowledge base that can turn these pitfalls into successes.
1.5 Current Status of the RIT MAV Team
The MAV team at RIT is currently in its third year, and is hoping to continue to
improve on the MAV design. The first MAV team was an offshoot of the RIT Aero
Design Team, and was formed to build a MAV that would be capable of competing in the
annual competition. Unfortunately, the first MAV team was not able to produce a MAV
that could compete in the competition successfully mostly due to a lack of experience and
knowledge. The following year’s team used the experience it had gained from the
previous teams MAV to produce a more capable MAV. The team again went to
competition, and the MAV was able to achieve flight. Unfortunately, the MAV was not
capable of completing the requirements in the surveillance operation. As the team before
it, last year’s team has helped provide a better foundation for the 2004-2005 MAV team.
The ultimate goal of this year’s team is to produce a MAV that will not only fly well, but
actually complete the surveillance mission. The 2004-2005 team will also try to advance
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on the entire MAV design to continue the advancement of RIT’s Micro Air Vehicle team,
and hopefully make a name for RIT in the MAV research community.
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2 Team Organization and Work Breakdown
The design of a Micro Air Vehicle requires that much of designing and building
occur at the same time. In fact, many of the decisions to be made on one are of the
design relate to many other aspects of the entire design. Therefore, the team was broken
into three subgroups that will handle different aspects of the MAV design. The creation
of subgroups served several purposes. The 2004-2005 MAV team has ten engineers that
will be concentrating on the design. With a team this size, it seemed creating subgroups
would be the most practical and efficient way of creating a successful MAV. Also, last
year’s MAV team utilized subgroups for the design, and seemed to have better success
using this method than all of the members working together on the entire design. The
subgroups were formed by allowing each of the members to decide where they would be
most comfortable working in the design process. This proved to be an adequate method
for forming the groups since each of the members in the respective group had a
background and special interest in the particular design. The three subgroups: Airframe,
Electronics, and Propulsion and their members are shown below.
Each subgroup is responsible for the design of their respective portion of the
project, and also the integration of their portion into the complete MAV design. The
subgroup members must also communicate with the other subgroups, since the design
requires facets of each subgroup influencing the other subgroups design. Each of the
MAV Team 05-001
Airframe Electronics Propulsion
Joshua Baker Aaron Grilly David Hein JED Hess
Cuong Le Tzu-Chie Fu Atul Phadnis
Aimee Lemieux
Victoria Schoennagel
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subgroups will also be responsible for the written portion of their design process, since
their knowledge of the particular design is unparalleled within the entire group. The
report will be created in such a manner that it should make the entire design process as
clear and concise as possible.
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3 Literature Review
The literature review is a very important tool when doing any design work. The
RIT MAV team performed an extensive literature review to become more knowledgeable
on the current state of the MAV technology. This literature review helped give the team
more ideas on what concepts to pursue for our own design. The review also allows the
team to begin where other researchers have left off instead of redoing work that has
already been performed. The majority of the useful information found by the three sub-
groups is described in the following pages.
3.1 Airframe
To perform theoretical airfoil calculations, it is necessary to have certain
aerodynamic properties. Airfoil simulation programs, like XFLR5 [6], utilize Reynolds
number and angle of attack to perform predictions on the aerodynamic properties.
Reynolds number can be easily calculated using Equation 3.1 shown below.
µρVc
=Re (3.1)
Using assumptions for vehicle speed and wing sizing, a range of operating Reynolds
numbers were found. Table 3.1 below shows the calculations and the Reynolds number
range expected for our design. Since the max Reynolds number is less than one million,
these airfoils fall under the term of low speed or low Reynolds number airfoils.
Therefore, all of the airfoils that would be tested in XFLR5 [6] would be airfoils
optimized for low speed.
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Table 3.1: Reynolds Number Range Calculations
Parameters Max Inputs Units Notes Max Chord Length m 0.2032 8 inches Min Chord Length m 0.1524 6 inches Max Density kg/m3 1.225 At sea level Min Density kg/m3 1.21328 At 100 m Max Viscosity n s /m2 1.7890E-05 At sea level Min Viscosity n s /m2 1.7860E-05 At 100 m Max Speed m/s 20 Min Speed m/s 5 Re Max unitless 278745.8 Re Min unitless 51677.997
The results of these theoretical tests must be validated since the programs are still
not completely reliable at low Reynolds numbers. Selig [14-16] performed many tests on
airfoils with respect to 2-D airfoil lift and drag performance. His data will be used to
validate the results of the XFLR5 [6]. The airfoil is only one small art of the overall
airframe. The planform is also quite important to the overall aerodynamics of the
airframe. Torres [21] performed an extensive study of the MAV airframe and in
particular the planform shape. His work along with Selig’s will prove to be quite
valuable in designing the most optimal airframe.
Besides the use of technical papers and dissertations, the websites of other MAV
teams proved to be quite useful when researching other possibilities for airframe design.
This extensive reviewing of the available literature has developed an extensive amount of
knowledge on the type of airframes the 2004-2005 MAV team should investigate. In
particular, use of 2003-2004 RIT MAV team’s work [19,20] has helped the current MAV
team avoid many of the pitfalls the prior team could not, and also has given insight into
the best way to create a success MAV.
3.2 Propulsion
A comprehensive literature review was performed in this area by utilizing the
technical papers from the 2003 and 2004 MAV competitions [2,3] and by performing an
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online search for MAV related information. As these papers show, the trend in
propulsion is heading more and more towards electrically driven propulsion systems as
opposed to internal combustion engines. A summary of the findings from these papers
will be shown in the following pages.
Bringham Young University
In 2003, the BYU MAV team performed a thorough study of motor options by
analyzing there weight, current draw, and voltage draw. This feasibility assessment
helped the team choose their motor with relative ease. After choosing the motor, the
BYU team focused on finding the right propeller for the motor. Instead of using
commercially available propellers, the BYU team utilized a program called JavaProp to
develop propellers that suited the flight conditions and their motor choice. To test their
propulsion systems thrust, the team performed dynamic testing in their wind tunnel.
After the dynamic testing was concluded, the BYU team decided that the Skyhooks and
Rigging KP00 was the ideal motor for their MAV.
The 2004 BYU MAV team performed a similar feasibility assessment as the 2003
team, but this team went for a larger MAV that included autopilot. The team decided
upon the ideal motor of an Astroflight 010 brushless motor because of its ideal power to
weight ratio and good reliability. After choosing the motor, the team then decided on the
ideal propeller. The team chose a MAS 5.5 X 4in propeller because it was decided to be
a perfect match for the motor. The team also chose a Phoenix 010 speed controller to
command the motor’s thrusting due to its compatibility with the motor and its low
weight.
California Polytechnic State University
The 2004 Cal Poly MAV team created two different MAVs to compete in both
the surveillance and endurance mission. Both the surveillance and endurance MAVs
utilized DC electric motors and commercial available propellers that have been modified
for propulsion. The surveillance MAV used a Maxon RE-10 coreless motor with a GWS
MAV propeller with modified tips. The endurance MAV used a Mabuchi M20 that
weighed a mere 3 grams.
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Georgia Institute of Technology
The 2004 Georgia Tech MAV team utilized a computational optimization
program called MAV innovation at the Georgia Institute of Technology (miGIT). This
computer program optimizes the MAV design while also looks into practical design and
construction of the MAV. The team used this software to optimize all aspects of the
MAV design including the motor and the propeller. The propeller program estimates
thrust torque coefficients and propeller efficiency. This program utilizes coding from the
low speed airfoil analysis program PABLO. The program provided information that led
the team to choose the KP-138 propeller for their propulsion requirements. The motor
program used rotational speed and torque required information to provide acceptable
motors. The program helped the team decide upon the Wes-Technik DC5-2.4 electric
motor to meet the rest of the propulsion needs.
Lehigh University
In the 2003, the Lehigh University MAV team utilized the commercially available
Kenway U-80 propeller, and tested it on two electric motors (Maxon RE-10 and Wes-
Technik DC5-2.4). The team performed thrust testing utilizing both motors and lithium
polymer batteries. The team concluded that their ideal propulsion setup would utilize the
Maxon RE-10 electric motor with the U-80 propeller. The 2004 Lehigh MAV team
utilized the same propulsion setup as the 2003 MAV team. The team experimented with
the program MotoCalc to optimize the propulsion system, but this proved to be fruitless
since it could not handle MAV sized aircrafts.
Rochester Institute of Technology
The 2004 RIT MAV team, researched using a combustion engine better later
decided that the electric motor would be a better choice. The team first selected two
electric motors (RE-10 and DC5) and various propellers to perform static thrust testing
upon. Initial thrust testing limited the propeller choices to the EP-0320 and U-80.
Modifications were made to these propellers and testing was continued. The team found
that the modifications to the U-80 made no significant impact, where as the modifications
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on the EP-0320 produced significant advantages overall. The team decided that the
propulsion for the MAV would be best suited by utilizing a Wes-Technik DC5-2.4 motor
with the unmodified U-80 propeller. The U-80 was chosen because it attached to the
motor with a simple press-fit, where as the EP-0320 had to be attached to the motor using
epoxy. Therefore, the team decided that the U-80 was the most obvious choice for the
propeller.
University of Arizona
The 2004 University of Arizona MAV team utilized the RE-10 electric motor and
the U-80 propeller. The team produced both a surveillance and endurance MAV. The
surveillance MAV used the larger 1.5 W RE-10 motor with the unmodified U-80
propeller. The endurance MAV on the other hand utilized the 0.75 W motor with a
modified U-80. The team, though, has been looking at utilizing a three blade propeller,
but have yet to have one ready for competition. The U-80 propeller and RE-10 motor
were chosen because of the success previous University or Arizona MAV teams have had
with that propulsion configuration.
University of Florida
In 2003, the University of Florida MAV team used a Portescap motor on their
endurance MAV, where as the surveillance MAV used a Maxon RE-10. The team used
JavaProp to design the ideal propeller for the flight conditions the vehicle will
experience. The selection method involved using an experimental test matrix which
varied the motor, propeller, airspeed, and voltage settings. The final propeller design was
tested in a wind tunnel using three different motors and thee different diameters for the
propellers. This dynamic testing produced the ideal configuration for each MAV. The
2004 MAV team concentrated on making a more biological inspired MAV. The team
utilized the Maxon RE-10 motor with the commercially available U-80 propeller. The
motor and propeller were dynamically tested in the University of Florida’s wind tunnel.
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Worcester Polytechnic Institute
In 2003, the WPI MAV team selected their motor by evaluating several
commercial available motors with respect to their efficiency. WPI found that the best
motor under these specifications was the Grand Wing Servo Company’s EDP 50XC
electric motor and its paired propeller. The 2004 MAV team chose to utilize a
combustion engine because of its higher thrust output. The combustion engine chosen by
the team was a Cox Tee-Dee 0.010 glow engine. Too throttle the engine, the team chose
a Micro-Flite PET Cox Tee-Dee 0.010 engine and balloon tank for the propulsion system.
The team also performed static and dynamic thrust testing on the combination of the
motor and its paired propeller.
KonKuk University
The 2003 KonKuk MAV team produced a MAV that utilized the Maxon RE-10
motor and the U-80 propeller. The team performed significant amounts of propeller
modifications to find the best possible arrangement. Their modifications included
altering the shape to find the optimal blade shape and overall diameter. The team tested
three different blade shapes (fillet, ‘S’ type, and elliptical) to see if a significant change in
the lift to drag ratio would occur. The team found that the ‘S’ type and elliptical blade
shapes had significantly less drag than the fillet shape. This loss in drag would allow for
a higher RPM, but also a decrease in thrust production.
3.3 Electronics Electronics like most technologies is constantly improving. This is very
important for MAVs which are constantly in need of smaller, lighter, and still acceptable
technologies. Therefore, an extensive search must be performed instead of just relying
on the electrical technology used on other MAV teams in prior years. The greatest tool
for performing these extensive searches is the internet. By using prior technology as a
beginning search and also as the minimum criteria, it is quite easy to find newer, better
technology on-line. The majority of information on prior technology came from
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performing a literature review of the 2003 and 2004 MAV competition papers [2,3]. It
also proved useful to look at the electronics utilized by the 2004 RIT MAV team [19,20].
The literature review showed that battery selection was very important in the
entire electronics design plan. Most of the competition papers showed that a lithium-
polymer battery was the most ideal power source due to its efficiency and lower size and
weight. These batteries can be connected in series to produce the necessary voltage.
Therefore, the main criterion for picking the appropriate battery is based on the maximum
amperage that can be produced by the batteries. An extensive internet search will be
performed to see if there are new batteries that have the same production and lower
weight and size.
The review also gave much insight into control surface actuation. The majority of
MAV teams utilized servo motors produced by Wes-Technik. These servo motors are
utilized because of their very low-weight (2 grams) and their ability to supply a sufficient
amount of force (160 grams) to actuate the control surface during flight conditions. The
review also showed that several MAV teams utilized coil-magnet actuators instead of
servo motors because of their significantly lower weight. These actuators have been
found to be approximately 600% lighter than the Wes-Technik servo motors. An internet
search showed that these actuators are actually better for a typical model airplane design
as opposed to the smaller MAV style aircraft.
Besides the aforementioned electrical components, there are two more
components that are vital to the MAV during surveillance missions. The first component
is the video camera that will be used for real-time video surveillance. The competition
papers have shown that CMOS and CCD cameras are the best for MAVs because of the
small size and weight. Of these two camera types, the CCD cameras are considered to be
the better because of the produce a higher quality image. The second surveillance
component is the Global Positioning System (GPS). Currently, very few MAVs have
utilized this technology due to its weight and size. Improvements have been made in the
technology, so it seems reasonable that there are more GPS components that can be
utilized on a MAV with a larger wingspan. Internet searching was performed to find the
best GPS components that could work for the MAV. Much of the searching also
concentrated on finding transceivers that could handle both GPS and Video
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transmission/receiving. This would save in the overall weight of the electronics, and
make the GPS a greater reality.
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4 Needs Assessment
The 2004-2005 RIT Micro Air Vehicle (MAV) is being developed by a
multidisciplinary Senior Design group. The group will be responsible for producing a
MAV that will be used for surveillance missions. The vehicle will contain real time
video and a Global Positioning System that will be capable of sending the information to
a laptop computer used by the pilot. The team will also research adding sensors and
autopilot to produce a better flying and cutting edge MAV.
4.1 Performance Goals
For the MAV to be a useful commodity, it must be capable of a long stable flight
with an acceptable flight range. The performance objectives for this year’s MAV design
are to fly for a minimum of fifteen minutes with a minimum range of 600 meters.
4.2 Vehicle Goals
The MAV designed last year produced a good MAV, but stability was an issue.
In an effort to solve this stability issue and also produce a MAV that will have a better
ability at completing the surveillance competition, this year’s MAV will have a minimum
dimension of 15 inches. This year’s team will also try and incorporate composite
materials into airframe design to induce both better stability and more resilience to
damage.
4.3 Airframe
The airframe developed must prove to be resilient enough to withstand the impact
during landing, so that it can be used for multiple flights. The airframe must also offer
the necessary protection needed for the electronics since the electronics prove to be the
most expensive part of the entire MAV aircraft.
The computer analysis package XFLR5 [6] shall be used to get theoretical
aerodynamic properties for several airframes found during the literature review and
online research. The airframes that performed the best according to the theoretical
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analysis shall then be tested in the RIT wind tunnel. The wind tunnel results shall give
evidence to which airfoil should be used for the MAV design.
To house all of the electronics needed for the MAV design, a pod will be
designed. This pod must not hinder the MAV’s overall aerodynamic capabilities. The
pod must also incorporate added protection for the electronics. This protection should
ensure both the electronics safety and also constrain the electronics to help maintain the
location of the MAV’s center of gravity.
The control surfaces on the vehicle will need to provide enough control for the
pilot to be able to navigate to the target area. The dynamic surfaces shall be needed for
fundamental control and the static surfaces will assist with stability.
4.4 Propulsion
The motor and propeller must meet the required thrust to propel the MAV during
flight. To find the best motor and propeller, an extensive literary and online search must
be performed. The motors and propellers that meet the thrust requirement, and also meet
the minimums for weight, amperage, and voltage will then be tested dynamically in the
wind tunnel. These results shall be used to choose the motor and propeller configuration
used to power the MAV.
4.5 Electrical Systems
The MAV will be equipped with an on board camera that will be capable of
transmitting real time video back to a laptop. The video feed back to the laptop will prove
to be a useful tool, since it will allow the pilot to perform surveillance of areas or possibly
reconnaissance targets.
A receiver with a range that meets the performance objectives must be used to
receive control information from the pilot’s transmitter. The receiver must be capable of
controlling the servos used for control surface actuation, and throttling the motor. The
motor is throttled by a speed controller that receives commands from the receiver. The
speed controller shall reside electronically between the receiver and the motor.
25
The electronics contained on the MAV will be powered by on board lithium ion
batteries. The batteries must be capable of produce the required voltage and amperage
needed for the electronics. The batteries must also produce these requirements with the
lowest weight possible.
A Global Positioning System (GPS) will also be installed on the MAV to give the
pilot more information on the MAV’s position. The GPS will help the pilot navigate to
known target locations and also report back the location of targets found during flight.
The GPS will also be capable of giving data on the vehicle’s speed and altitude. The
GPS data shall be included in a setup where the operator can view the data along with the
real time video.
26
5 Concept Development and Feasibility
5.1 Airframe
Airframe concept development has been broken down into six main areas of
research. Each one of these areas will have a profound impact on the flight
characteristics of our Micro Air Vehicle (MAV), ease of manufacturing, and overall
usefulness of our design in contributing to the MAV research field. It has been
determined that the MAV design for this year will take a strong research oriented
approach. This will allow the team more freedom in testing different concepts. The six
areas of research are planform design, vehicle configuration, airfoil geometry, engine
mounting configuration, materials and processes selection and flight controls.
5.1.1 Planform Design Within the flying wing category,
several different planforms have been
researched and discussed. The most notable
of these are the Modified Inverse Zimmerman
planform (used by last year’s team [19,20]
and the AeroVironment Black Widow [7],
also see Fig. 3), a triangular planform with a
swept-wing appearance (dubbed the ‘molar’),
the Zimmerman planform and Inverse
Zimmerman planform. At this point, the
Inverse Zimmerman and Modified Inverse
Zimmerman planforms are the strongest
candidates for this year’s MAV planform
design, because they have proven successful
in other MAV’s. Part of the logic behind these Figure 5.1: Examples of different planforms, showing trailing vortices [1].
27
planforms is that they pull the trailing vortices away from the center of the planform, thus
reducing drag. Figure 5.1 shows this in more detail for some various shapes of
planforms.
Because of the applications in which a MAV would typically be used, our goal is
to reduce the max linear dimension while maximizing surface area, as described by the
formula below: 2b bAR
S c= = , where
AR, the aspect ratio, represents the relationship between wing span, b, and overall surface
area, S. This formula can be rewritten to include c, the average chord length. From this
relationship, it follows that to maximize surface area while minimizing the max linear
dimension, the most desirable planform will have an aspect ratio near unity.
5.1.2 Vehicle Configuration
Nearly all research has shown that in the MAV application, a flying wing
configuration, as shown in Figure 5.3, is far more appropriate than a standard
configuration with a wing and tail, as shown in Figure 5.2. The reasons for this are
intuitive. A conventional body design allows for greater flight control characteristics and
by default creates a more dynamically stable aircraft. At the same time however, it
requires a higher thrust-to-weight ratio, because in this configuration so little of the
Figure 5.2: The BYU Stableyes is an example of a conventional planform with a pusher motor configuration [3].
Figure 5.3: The Black Widow [6] is an example of a modified inverse Zimmerman planform with a puller
fi i
28
aircraft is actually contributing to lift. In addition, the industry driving MAV research is
largely military and geared toward covert operations. As this is the case, it will be far
more beneficial to both reduce the planform size of the aircraft and its appearance as an
“airplane”. A flying wing configuration, while making dynamic stability a larger
concern, allows for a higher power-to-weight ratio as well as a higher lift-to-drag ratio.
The prime factors contributing to both of these in a flying wing configuration is the fact
that nearly the entire body of the aircraft will contribute to lift. With a maximum linear
dimension of 15’’ as one of our design parameters, we could choose to use a conventional
planform configuration. However, with the hopes of reducing this dimension in years to
come, it is far more beneficial that we proceed forward with a flying wing design.
In addition to a flying wing configuration, it will be necessary to carry
instrumentation on board the aircraft. Most research shows that airfoils with thicknesses
>10% of the chord length are not good for low Reynolds flows [9], as will be experienced
by a MAV. Thus, storing the necessary instrumentation on-board becomes an issue. A
practical solution to this problem is to mount the instrumentation, actuators, and possibly
even the motor/motors in an independent unit known as a pod. The idea of a pod has
been conceived based on past knowledge of their use on many aircraft including fighter
jets such as the F-16. Much of the thought and development that has gone in to this
portion of the concept assessment has been based on the brainstorming of the team
members. More recently, last year’s senior design team used a pod-like configuration
[19,20], as did WPI [3] on a past MAV design.
There are many advantages and disadvantages of using a pod configuration. The
main advantage in using a pod is that it will allow the instrumentation to be secured
properly, and accurately, in a removable unit that can be easily affixed to different flying
wing configurations. Secure placement of the onboard instrumentation is critical so as to
avoid re-calibration upon crashes or replacement of the flying wing structure. To aid in
this, the pod must be designed to ensure that no shifting of the internal pieces occurs.
Memory foam will most likely be used in this endeavor, as well as smart construction of
the pod itself, as discussed by this year’s team.
With this in mind, many types of pod configurations have been researched or
sketched, and then discussed. The first option is to mount the pod between two airfoils so
29
that the pod effectively becomes the center of the aircraft. This configuration makes
design easy by allowing for the main mass as well as the power plant of the aircraft to be
centered along the x-axis of the MAV—noticeably reducing the pitching moment that the
aircraft would experience if the propeller were to be mounted in a pod offset from the x-
axis.
However, aerodynamics tells us that mounting a pod near the flying wing will
change the lift characteristics of that region of the aircraft. Also, the pod will most likely
be small enough that offsetting it from the flying wing structure will not necessarily
increase the max linear dimension of the MAV. Because of this, other options have been
considered. These involve mounting the pod far enough below or above the aircraft (so
as to not disturb the flow around the flying wing) and either in parallel or at an offset
angle relative to the x-axis. It has been discussed (by the team) that a pod oriented at an
offset angle (relative to the x-axis) and strategically designed as a miniature airfoil could
in fact increase the overall lift of the aircraft as well as contribute to the dynamic stability
of the MAV, respectively.
One further advantage to employing a pod in our design is that though it may shift
the center of gravity (c.g.) above or below the x-axis of our lifting structure, this may be
used to our advantage in enhancing the flight characteristics of our MAV. In addition,
the pod can be smartly adjusted fore and aft in order to compensate for the pitching
moment generated by the wing structure.
5.1.3 Airfoil Geometry Airfoil geometry has been
researched from a variety of past projects
[3] and published papers [9]. The general
conclusions at this point are that thin airfoils
(thickness < 10% of the chord length) with
low camber (camber < 8% chord length)
and some reflex provide the best lift-to-
drag ratios and have the most stability for low Reynolds number flows. The main
Figure 5.4: Separation bubble on airfoils at low speeds
30
problem confronting airfoil at low Reynolds flows is the formation of a laminar
separation bubble that essentially alters the effective airfoil geometry. However, because
this bubble is constantly changing, the flight characteristics change as well. This
phenomenon is shown in Figure 5.4.
With this in mind, several airfoils have been researched adhering to these
guidelines using the program XFLR5 [6], an offshoot of the airfoil analysis program
XFOIL that has been designed for low Reynolds number flows. XFLR5 uses an iterative
technique to analyze a given airfoil. The team has conducted systematic analysis within
this program to find lift, drag and moment coefficients for angles-of-attack in the range [-
6º, +18º] degrees with Reynolds numbers in the range [50000 , 275000]. The analysis
was done in the program using the batch analysis command and the polars for the airfoil
were exported to an Excel spreadsheet for analysis.
The airfoils analyzed include E174, E186, EH2012, EH3012, FX63137,
GOE417a, GOE494, M10, M12, M14, MH46, RAF6, S1210, S2027, S4022, S4083,
S5010, S5020, and the SD7080. The airfoils were chosen on the basis of past merit as
well as on the concepts employed in the design of each airfoil [1, 17]. Airfoil data files
were downloaded from the UIUC Airfoil Data Site [17]. These concepts (such as camber
and thickness ratios) are discussed in the first paragraph of this section. After the initial
analyses, a feasibility assessment was performed using the weighted method to narrow
down the airfoils for wind tunnel testing. The results of this assessment are shown in
Table 5.1 located in the appendices section. Airfoils selected for secondary testing were
the GOE417A, S1210, S4083, and the S4022. Once testing begins, the team will begin to
alter planform geometry as well as airfoil camber, thickness and reflex to arrive at an
optimized wing design to meet our needs.
5.1.4 Engine and Engine Mounting Configuration The first question the MAV team had to face was whether to use one propeller or
multiple propellers in the design. For now it has been decided that a multiple propeller
design would be too complicated to integrate due to alignment issues. Our current
assembly techniques are not refined or reliable enough to overcome this issue at present.
The alignment issue may be researched later on in the project if there is time.
31
The airframe group has also discussed where a single engine should be mounted
on our aircraft. A puller configuration is easier to implement because it does not induce a
large pitching moment on the aircraft; however, it also affects the flow over the lifting
surface of the aircraft in the slipstream area. MIT has conducted theoretical research on a
pusher configuration (Figure 5.5) [10], and BYU [3] implemented a pusher configuration
on a MAV last year (see Figure 5.2 above). However, BYU’s MAV had a larger
maximum linear dimension and a conventional planform design. With this in mind, our
current plan is to start with an un-shrouded puller propeller configuration but to research
pusher configurations as well.
Other ideas that have been considered include the idea of implementing a ducted
fan to increase propeller efficiency and possibly allow for thrust vectoring. This may be
tested later on in a post-research phase after we have a working product. Another idea
that was quickly discarded but may also be considered for research in this post-research
phase is using tilting props to allow for thrust vectoring.
5.1.5 Materials and Processes Selection
The materials considered at the conceptual phase of this project for the airframe
design are:
Figure 5.5: MIT Lincoln Laboratory MAV Concept
32
Balsa wood frame w/ balsa wood leading edge and canvas-like cover (possibly
mylar)
Foam (polystyrene)
Foam with a single layer of epoxy of fiberglass coating
Carbon Fiber
Composite (Kevlar or Fiberglass)
Carbon Fiber frame with lightening holes and a mylar/fabric covering
Many of these options are either labor intensive, imprecise, or present costly
weight penalties that make them unfeasible for implementation. Our most likely choices
at this point in time for an airframe material will be carbon fiber or foam (possibly with a
coating). Carbon fiber foils will require 3D CNC molds to implement. Foam airfoils
simply require two end plates and a hotwire to manufacture, but the resulting product
does not stand up well to crashes and is more likely to break. We will most likely test
both of these options, and possibly proceed to lightening tactics in further test models if
we proceed with carbon fiber. However, lightening will require precision and accuracy
or the resulting, imbalanced planform will affect our flight characteristics drastically.
5.1.6 Flight Controls
Different control surfaces have been considered and are pending testing. The
control surfaces considered are:
Vertical Tail (and rudder)
o Single (high or low orientation possible)
o Double (high or low orientation possible)
o Angled Vertical Tail
Winglets
o Vertical Winglets
o Angled Winglets (creates vector which will contribute to lift)
Canards
33
o Mounted on flying wing
o Mounted on pod
Elevators
Ailerons
Elevons
After some discussion, the control surfaces that will most likely be tested and
receive the most attention with hopes of implementation will be elevons (in place of both
ailerons and elevators) and some sort of vertical tail configuration.
5.2 Electronics
The electronic components are the core of an airplane because they must guarantee
the continuous communications between a plane and the base station on the ground. The
constraint of electrical parts is weights and ranges. The weight is a factor to decide to a
force to lift an airplane and the ranges effect to communications between an airplane and
the base station on the ground. These electronic components are concerned that consist of
control system, Camera & GSP system, and motor servos.
5.2.1 Control System
5.2.1.1 RF Receivers
Considering the amount of current drawn, the physical size, the mass, and
receiver range, numerous receiver units were considered. Unfortunately, the receiver
range was not always given for parts researched. But, a design decision was made based
on maximum power transmission from the ground to plane for RF receiver. Those signals
will be used by the actuators to control the plane, so the range of receivers could be
chosen from 150 m-300 m. The major characteristics of receivers are shown in the Table
5.2.
34
Table 5.2: RF Receivers Manufacturer
Model Hitec HFS-04MG
Airtronics92515Z
Hitec Feather
GWS-PICO
R4P-JST
GWS-NARO R-6N/V
GWS-NARO R-6N/H
Channels 4 5 4 4 6 6 Current Needed 9 mA NA 9 mA 5 mA 7 mA 7 mA
Weight 19 g NA 8 g 4.3 g 7.8 g 8.2 g Volume 0.9 in3 1.1 in3 0.416 in3 0.343 in3 0.345 in3 0.317 in3
Voltage Needed 3.6-6.0 V 4.8-6.0 V 3.6-6.0 V 4.8-6.0 V 4.8-6.0 V 4.8-6.0 V Rx Range > 1600 m NA 300 m 150 m 300m 300m
R-6N
Figure 5.6: R4P-JST Receiver Figure 5.7: R-6N/H (Horizontal) Figure 5.8: R-6N/V (Vertical)
During the selection process, the decision to increase the number of channels for a
given receiver was made to allow for additional improvements in the form of additional
feature/control devices. So, the HFS-04MG and 92515Z were rejected based on weight
considerations despite they have a good range characteristics. Due to the fact that there
are not enough channels on the Hitec Feather and GWS-PICO RF receivers, it does not
allow any possible future improvements even though they have a low power consumption
and lower weight. After discussing about the purpose of this project and extension for the
next year project, and also with the previous year’s performance, the GWS-R-6N/H &
GWS-R-6N/V onboard RF receivers have been implemented and operated to
specifications. So the GWS-R-6N/H & GWS-R-6N/V are the leading candidate to be
selected.
35
Table 5.3: Pugh Chart for RF Receiver Selection
Factors/Candidates
HHii ttee cc
HHFF SS
--00 44
MMGG
AAii rrtt rr
oo nnii ccss
99 2255 11
55 ZZ
HHii ttee cc
FF ee
aa tthh ee
rr
GG WW
SS -- PP
II CCOO
RR 44PP --
JJ SS TT
(( LL
aa sstt
yy eeaa rr
))
GG WW
SS -- NN
AARR OO
RR --
66 NN// VV
GG WW
SS -- NN
AARR OO
RR --
66 NN// HH
CChhaannnneellss 3 4 3 3 5 5 CCuurrrreenntt NNeeeeddeedd 1 1 1 3 5 5
WWeeiigghhtt 1 1 3 5 4 3 VVoolluummee 2 1 3 4 4 5
VVoollttaaggee NNeeeeddeedd 4 3 4 3 3 3 RRXX RRaannggee 5 1 3 2 3 3
MMeeaann SSccoorree 2.66666667 1.8 2.8 3.3 4.0 4.0
NNoorrmmaalliizzeedd SSccoorree 66.7% 45% 70% 82.5% 100% 100%
Based on numerical comparison with last year’s component, the GWS-NARO R-
6N RF receivers is the leading candidate for selection.
5.2.1.2 Speed controller
Speed controller is required to increase or to decrease speed of motor as
necessary. The speed controllers are seen at the comparison in the table below
Table 5.4: Speed controller alternatives
Manufacturer
Model Wes-Technik
YGE-6 Wes-Technik
YGE-3 Phoenix-10
Weight 1.3g 1.0g 8.2 g Current Needed 4 A 2 A 10 A
Compatibility NA NA YES Dimensions (mm) 4x6x17 4x8x10 18.5 x 20.3 x 4.06
Figure 5.9: Speed control YGE3 Figure 5.10: Speed controller Phoenix-10
36
Since the speed controller needs to be compatible with the DC motor selected by
the propulsion group, the factor that dictates the selection of the speed controller falls
solely on compatibility. The Phoenix-10 speed controller is the leading candidate to be
chosen due to its compatibility with the DC motor that is being selected.
Table 5.5: Pugh Chart for Speed Controller Selection
Factors/Candidates
WWee ss
-- TTee cc
hh nnii kk
YY GGEE --
66
WWee ss
-- TTee cc
hh nnii kk
YY GGEE --
33 (( LL
aa sstt
YY eeaa rr
))
PP hhoo ee
nn iixx --
11 00
WWeeiigghhtt 3 3 1 CCuurrrreenntt NNeeeeddeedd 3 4 1 CCoommppaattiibbiilliittyy 1 1 5
DDiimmeennssiioonnss ((mmmm)):: 3 3 1
MMeeaann SSccoorree 2.5 2.8 2.0
NNoorrmmaalliizzeedd SSccoorree 90.9% 100.0% 72.7% Based on the numerical numbers, last year’s component turns out to be the best
choice for this project. However, the compatibility issue is the most important factor in
determining the speed controller. Therefore, even though the Phoenix-10 has the lowest
values, it is still considered a first option for its cross-compatibility. Also due to the result
from the Pugh chart, other options for speed controller are still being researched on.
5.2.2 Video System
5.2.2.1 Onboard Camera
Based on last year’s performance, all image acquisition and transmission
equipment will be supplied by Black Widow Audio-Video due to superior image quality,
performance after implementation, and overall unit weight. The specifications for the
selected on-board camera are as follows:
37
Table 5.6: Camera Specifications
Manufacturer Panasonic Model CX-161
Resolution 330 lines Operating Voltage 5.0 V
Current drain 120mA Weight 11.6g
Output Modes NTSC
5.2.2.2 Onboard Video Transmitter
To ensure compatibility and ease of use, Black Widow Audio-Video has
recommended the following Video Transmitter/Receiver:
Table 5.7: Camera onboard On-board GPS Receiver
Manufacturer Black Widow Black Widow
Model BWAV240050 BWAV240200 Operating Frequency 2.4Ghz 2.4Ghz Transmission Power 50mW 200mW Number of Channels 4 4
Operating Voltage 5.0 V 5.0 V Current drain 70mA 240mA
Weight 7g 12.5g Dimensions (mm) 20x20x10 27x24x9
Figure 5.11: The 200 mw Brown Bag Kit – Transmitter/Receiver Set (Left), 8dBi Patch Antenna for
Base Station (Right)
38
Table 5.8: Pugh Chart for Video Transmitter Selection
Factor/Candidate
BB WWAAVV 22 44
00 0055 00
55 00
mmww
TT XX // RR
XX
BB WWAAVV 22 44
00 2200 00
22 00
00 mmww
TT XX // RR
XX
OOppeerraattiinngg FFrreeqquueennccyy 3 3
TTrraannssmmiissssiioonn PPoowweerr 2 5
NNuummbbeerr ooff CChhaannnneellss 3 3
OOppeerraattiinngg VVoollttaaggee 3 3
CCuurrrreenntt DDrraaiinn 3 2
WWeeiigghhtt 3 2
DDiimmeennssiioonnss 3 3
MMeeaann SSccoorree 2.857142857 3.0
NNoorrmmaalliizzeedd SSccoorree 95.2% 100.0%
Based on the numbers from the Pugh chart, the BWAV240200 200mW
transmitter was selected for its increased signal intensity due to an overall higher output
power. At this output power, the listed maximum range was found to be 1.6 km, thus
exceeding design requirements.
5.2.3 GPS System
What is GPS? The Global Positioning System (GPS) is a satellite-based
navigation system made up of a network of 24 satellites placed into orbit by the U.S.
Department of Defense. GPS was originally intended for military applications, but in the
1980s, the government made the system available for civilian use. GPS works in any
weather conditions, anywhere in the world, 24 hours a day.
5.2.3.1 GPS Receivers
The GPS receiver can receive a signal was transmitted by a satellite and
determine the user's position and display it on the unit's electronic map.
39
Table 5.9: GPS Receivers w/ antenna
Manufacturer Furuno Sarantel UNAV Model GH-79 Smart Antenna PICO-GPS-SS
Accuracy 15 m Apprx. 15 m 5 - 25 m Weight 13g w/ antenna 15g w/ antenna 28g
Dimensions (mm) 28x21x10 mm 32x64x13 mm 45.7x31.7x15.2 Operating Voltage 3.1-3.3V 3.1-12 V 3.8 to 8.0vdc
Current Drain 76 mA 180mA (55mA @3.3V)
100mA
Figure 5.12: Furuno Figure 5.13: Sarantel Figure 5.14: UNAV GH-79 Smart Antenna PICO-GPS-SS
Table 5.10: Pugh Chart for GPS Receiver Selection
Factors/Candidates
FF uurr uu
nn oo
GG HH-- 77
99
SS aa rr
aa nntt ee
ll SS mmaa rr
tt AAnn tt
ee nnnn aa
UUNNAAVV
PP IICC OO
-- GGPP SS
-- SSSS
AAccccuurraaccyy 3 3 5 WWeeiigghhtt 3 2 1
DDiimmeennssiioonnss 3 2 2 OOppeerraattiinngg VVoollttaaggee 3 3 2
CCuurrrreenntt DDrraaiinn 3 1 2
MMeeaann SSccoorree 3.0 2.2 2.4
NNoorrmmaalliizzeedd SSccoorree 100.0% 73.3% 80.0%
40
Based on the numerical result, the Furuno GH-79 turns out to be the best choice
for the project. However, the Furuno GPS receiver requires a microprocessor to decode
the signal and manipulate the data in order to obtain the GPS coordinates. This increases
the difficulty of the project and also adds components that could be unnecessary to the
entire system. This causes the need for more power, more weight, and more time.
The next best candidate for the job is the UNAV PICO-GPS-SS receiver. The
UNAV GPS Pico system offers the most complete system in the form of GPS Receiver,
GPS AutoPilot system, and GPS Video Overlay Board. The UNAV GPS was chosen due
to:
UNAV units work best with R/C radio systems that offer advanced features like
"Fail-Safe" but most RC radio systems are compatible.
UNAV can use buffers to reduce RF noise from ignition systems, onboard
transmitters and long servo wires.
Connection is simple
The data downlink from a GPS receiver to a laptop computer running Tracker or
MAPSOURCE is typically RS232 @ 4800
5.2.3.2 Video Overlay Board
Since the UNAV GPS Receiver is selected for the project, a compatible video
overlay board is needed to overlay the GPS data onto the video data to provide GPS
coordinates on the screen. Therefore, the OSD-GPS overlay board from UNAV was
considered due to its compatibility.
Table 5.11: Video Overlay Board from UNAV
Manufacturer UNAV
Model OSD-GPS Accuracy NA
Weight 22.68g Dimensions (mm) 63.5x63.512.7 Operating Voltage 8.0 to 14.0 volts
Current Drain 60 mA
41
Figure 5.15: UNAV OSD-GPS Video Overlay Board
From the specifications, the dimensions and the weight of this video overlay
board are quite large. Even though this particular component is chosen, research is still
being done for other alternatives.
5.2.4 Servos (Actuators)
Servo is an electromechanical device which moves the control surfaces or throttle
of the airplane according to commands from the receiver. So the servo is the one of the
most important parts of electronic components. Some servos and their characteristics are
in the comparing table below: Table 5.12: Wes-Technik Servos
Manufacturer
Model Wes-technik
LS-2.0 Wes-technik
LS-3.0 Wes-technik
LS-2.4 Max Deflection (mm) 14 14 14
Time to Full Deflection (sec) 0.15 0.15 0.2 Max Output Force 160 g 200 g 175 g Operating Voltage 3-5 V 3-5 V 3-5 V Dimensions (mm) 21 x 13 x 9 21 x 13 x 9 21 x 13 x 9
Load Current <100mA <100mA <100mA Weight 2 g 3 g 2.4 g
Figure 5.16: LS-2.0 Figure 5.17: LS-3.0 Figure 5.18: LS-2.4
42
Table 5.13: Pugh Chart for Servo Selection
Factors/Candidates
WWee ss
-- TTee cc
hh nnii kk
LL SS-- 22
.. 00
WWee ss
-- TTee cc
hh nnii kk
LL SS-- 33
.. 00
WWee ss
-- TTee cc
hh nnii kk
LL SS-- 22
.. 44
MMaaxx DDeefflleeccttiioonn 3 3 3 TTiimmee ttoo FFuullll DDeefflleeccttiioonn 3 3 2
MMaaxx OOuuttppuutt FFoorrccee 3 5 4 OOppeerraattiinngg VVoollttaaggee 3 3 3
DDiimmeennssiioonnss 3 3 3 LLooaadd CCuurrrreenntt 3 3 3
WWeeiigghhtt 5 2 3
MMeeaann SSccoorree 3.3 3.1 3.0
NNoorrmmaalliizzeedd SSccoorree 100.0% 95.7% 91.3%
All the servos that are listed have the same technical information with the
exception of the weight and the maximum output force. Based on the numerical values,
the result shows that LS-2.0 is the suitable choice for this project since weight is
considered one of the most important factors.
5.2.5 Antenna
An antenna is a device used to transmit and/or receive radio waves or signals. The
physical design of the antenna determines the frequency range of transmission/reception.
Antennas come in all shapes and sizes, their size and shape depending on the frequency
and use of the signal transmitted. Some antennas can broadcast signals in all directions;
they are called omni-directional antennas. Other antennas can also broadcast signals in a
fine straight line - like a flashlight, they are called directional antennas. Electrical signals
with frequencies higher on the spectrum, for example, are shorter and more directional.
As they get higher on the spectrum, they behave more like light. These must be focused
and thus, require antennas which are shaped like the mirror reflector of a focusing
flashlight. Here are collected antennas in the comparing table below:
43
Table 5.14: Video receiver antenna alternatives
Model HG2414P HG2414D HG2416P HG2424G HG2409P
Horiz.Beamwidth 30 deg. 25 deg. 25 deg 8 deg. 75° Vert.Beamwidth 30 deg. 25 deg. 25 deg 8 deg. 65° Gain (Directivity) 14 dB 14 dB 15.5 dB 24 dB 8,0 dBi
HG2414P HG2414D HG2416P HG2424G HG2409P
Figure 5.19: Antenna Array w/ Radiation Maps
44
Table 5.15: Pugh Chart for Antenna Array Selection
Factors/Candidates
HHGG 22 44
11 44PP
HHGG 22 44
11 44DD
HHGG 22 44
11 66PP
HHGG 22 44
22 44GG
(( LLaa ss
tt YY ee
aa rr))
HHGG 22 44
00 99PP
HHoorriizzoonnttaall BBeeaammwwiiddtthh 4 3 3 2 5
VVeerrttiiccaall BBeeaammwwiiddtthh 4 3 3 2 5
GGaaiinn ((DDiirreeccttiivviittyy)) 3 3 4 5 1
MMeeaann SSccoorree 3.7 3.0 3.3 3.0 3.7
NNoorrmmaalliizzeedd SSccoorree 100.0% 81.8% 90.9% 81.8% 100.0%
Based on the numerical results, the HG2414P and HG2409P seem to be the best
choice for the project. However, the reason for it being the highest ranked is due to the
fact that the HG2409P has the largest coverage area. It is ideal to have an antenna that has
a balance between the coverage area and the gain. Therefore, the HG2414P is chosen for
its fairly good coverage area and an acceptable gain.
5.2.6 Batteries
In the selection of the on-board battery cells, the need to have a list of total power
requirements for all components is critical. Once these values are obtained from the other
sub-groups, a choice of battery pack and quantity can be made. Also, based on last year’s
performance, the Kokam Battery packs performed very well in the field offering
sufficient battery power and an overall discharge rate during the flight time.
The following is an overall listing of the batteries currently being considered for
On-board use: Table 5.16: Comparison of Battery Cells
Manufacturer
Model Kokam
SLPB433452 Kokam
SLB452128 Kokam
SLPB523459 iRate LP500
Capacity 740 mAh 145 mAh 1040 mAh 500 mAh Voltage 3.7 V 3.7 V 3.7 V NA
Dimensions (mm) 52x33.5x4.35 27.5x20.5x4.4 59x33.5x5.25 NA Volume (mm3) 7577.7 2480.5 10376.6 NA
Weight 15 g 3.5 g 20 g 10.8 g
45
Figure 5.20: Batteries – Kokam
Table 5.17: Pugh Chart for Battery Selection
Factors/Candidates KK oo
kk aamm
SS LL PP
BB 4433 33
44 5522
KK ookk aa
mm
SS LL BB
44 5522 11
22 88
KK ookk aa
mm
SS LL PP
BB 5522 33
44 5599
ii RRaa tt
ee LL PP
55 0000
CCaappaacciittyy 4 2 5 3 VVoollttaaggee 3 3 3 1 VVoolluummee 3 5 1 1 WWeeiigghhtt 2 4 1 3
MMeeaann SSccoorree 3.0 3.5 2.5 2.0
NNoorrmmaalliizzeedd SSccoorree 85.7% 100.0% 71.4% 57.1%
After numerical comparisons, the best candidate turns out to be the Kokam
SLB452128 due to its small size and weight. However, the capacity that the battery
provided is too small. Therefore, it is ideal to have a battery that’s acceptable in size and
weight, and provides enough power for the entire system. So the next best choice is the
Kokam SLPB433452. This battery provides a 740 mAh of capacity and weighs around 15
grams. The iRate ended up in the last place due to insufficient operating specifications.
Further research of the iRate operating specifications is underway.
5.3 Propulsion
5.3.1 Concept Development
The propulsion system must be developed to provide enough thrust to power the
MAV so that the endurance and range goals defined in the needs assessment can be
achieved. The most important factor in accomplishing this is the overall thrust produced.
46
In addition, other parameters considered in designing the propulsion system include:
weight, ease of use, and reliability.
Our background search identified two proven ways to power MAVs, either with a
rocketry system or with a propeller driven system [2,3]. Propeller based systems can be
powered by either an electric motor or an internal combustion engine. In addition,
propellers can either be bought as off the shelf items, or specifically designed and
fabricated. All of these options, shown in Figure 5.21 will be considered in the concept
development process for the propulsion system.
Figure 5.21: Propeller propulsion (left) and rocket propulsion (right)
First we considered the use of a rocketry system verses a propeller driven system
in designing the propulsion system. Literature research showed that both methods have
been used in pasts MAV designs [2,3]. It should be noted that the two methods of
propulsion vary significantly in their ability to have multiple uses and in their ease of
control. A weighted scale was developed using an internal combustion engine as the
baseline, to compare an electric motor to rocket propulsion. The results in Table 5.18
show that an electric motor will meet our projects needs the best. With this knowledge,
various electric motors will be researched to find the best fit.
Table 5.18: Weighted Scale for Propulsion Methods
ICE Electric Motor Rocket Propulsion Relative WeightAbility to produce needed thrust 3.0 3 3 22%
Availability 3.0 3 3 8%
Ease of Use 3.0 5 2 11%
Repeatability 3.0 4 2 14%
Ease of Integration 3.0 4 2 14%
Safety 3.0 4 2 11%
Enough student knowledge 3.0 4 2 8%
Weight 3.0 4 5 8%
Average Cost 0.0 4 5 3%
Weighted Score 2.9 3.8 2.6
Normalized Score 76.6% 100.0% 69.3%
47
After the propulsion method of an electric motor was chosen, the options for
propellers were considered. In particular, our team had to decide if a propeller should be
purchased or fabricated for our MAV (Figure 5.22). Background research was done in
both areas and it was found that both have been used successfully on MAVs [2,3]. Based
on this, we explored each method and gave time required to have a working, testable
propeller a significant amount of weight in our decision.
Figure 5.22: Purchased propeller (left) and mold for propeller fabrication (right)
Both propeller options have pros and cons that were considered in making our
decision of what action to take. Regarding fabricating our own propeller, the online
propeller analysis tool, Java Prop, which has been used by other MAV teams [2,3], would
allow us to easily obtain analytical results for various propeller designs. With this in
mind we consulted experts about building a mold for a micro propeller and learned the
process would be fairly time consuming because of the difficulty in making a small,
precise mold. Regarding off the shelf propellers we found many teams have
experimented with modifying purchased propellers, instead of completely designing a
propeller, to obtain better efficiencies for their vehicle. In addition, many motor
manufactures recommend off the shelf propellers for use with their motors.
Taking all of this into consideration, we decided to use off the shelf propellers for
our design. The unknown, and potentially large, amount of time it would take to build a
successful propeller mold is too big of a risk for our team to take in order to reach the
project’s goals on time. Background research has shown that a well selected, and
possibly modified, propeller will be successful on our MAV. Propeller testing will
involve testing the original propellers along with modified versions to find the best match
with a motor.
48
5.3.2 Feasibility Assessment Once it was determined to use an electric motor-propeller combination, the search
for the appropriate motor began. First the amount of thrust required needed to be
determined, as well as an appropriate weight for the motor. Of course, the lower the
weight, the better; but an expectation of what was feasible was desired. With a targeted
thrust to look for, motors were then chosen. Finally, five attributes were chosen to
determine the feasibility of each motor.
Information from the 2004 Micro Air Vehicle (MAV) team [19,20] proved to be
useful for a starting point when estimating the mass of the 2005 MAV. An estimated
mass for this year’s MAV was calculated first by adding up the mass of their
components. A GPS system is also expected to be on the 2005 MAV, so it was added in
as well. Finally the 2005 MAV is projected to be approximately 1.5 times larger than the
2004 MAV, so a factor of 1.5 was multiplied through the components. Therefore a
working mass of 202.35 g (0.446 lbf) was determined.
To determine thrust, several parameters were assumed: the wing on the 2005
MAV would be a scaled up version of the 2004 MAV wing, the plane would fly at 9
degrees angle of attack as it did last year, the span efficiency factor would be 0.9 for both
years, and the parasite drag for the 2005 MAV would be 1.5 times that for the 2004 MAV
due to the size increase. Data for the wing was only given for the airfoil, so it needed to
be converted into information appropriate for use on a wing. Equations 5.1, 5.2, and 5.3
[5] were used in these calculations. A table of values is given in Table 5.19.
(5.1)
dCL/dα = 3-D Lift Curve Slope (Wing)
dCl/dα = 2-D Lift Curve Slope (Airfoil)
e = Span Efficiency Factor
AR = Aspect Ratio of the Wing
eARd
dCddC
ddC
l
l
L
2
1801
πα
αα
+
=
49
(5.2) CD = Coefficient of Drag
CD0 = Parasite Drag
CL = Coefficient of Lift
(5.3)
T = Thrust (g) W = Weight [Mass in this Case] (g)
Table 5.19: Thrust Calculation All Estimated ValuesVariables 2004 2005dC l /d α 0.1 0.1
AR 1.422 1.422 Assume same shape for both years so AR remains the sameS (m 2 ) 0.048 0.102 For 2004, assume 90% of 8" x 26 cm rectangular wing is remainingb (m) 0.26 0.381 For 2005, assume 15" wing span
e 0.9 0.9 Assume 0.9dC L /d α 0.041 0.041
AOA (deg) 9 9C L 0.95 0.95 @ 9 deg AOA, est. from Figure 33 on pg 46 of PDR
W (g) 97.9 202.35 Actually a Mass; 2004 from Table 16 on pg 73 of CDRC D 0.243 0.252C D0 0.018 0.027 2005 is about 1.5x the size of 2004 so increase CD0 by that muchT (g) 25 53.598
The properties of the 2004 MAV were used to calculate a parasite drag by
working backwards. That drag was then multiplied by 1.5 to determine the parasite drag
for the 2005 MAV. From that point, the coefficient of drag, and then the thrust could be
determined using Equations 5.2 and 5.3 [5] respectively. Note that thrust for the 2005
MAV is expected to be 53.6 g (0.118 lbf). However, when choosing a motor, more thrust
than required was desired. Therefore, 80 g (0.176 lbf) was determined as a desirable
amount of thrust.
As components were chosen, actual weights could be added into the Mass
Calculation Table. The new mass of the aircraft (198.48 g [0.438 lbf]) was then entered
into the Thrust Calculation Table. Through this process, a more accurate value for the
eARCCC L
DD π
2
0 +=
L
D
CCWT =
50
required thrust (52.6 g [0.116 lbf]) was obtained, and shown in Table 5.20. After
determining the required thrust, it was decided to continue to use the 80 g (0.176 lbf) as
the desired value of thrust since the initial estimate of 53.6 g (0.118 lbf) is very close to
the revised amount of thrust.
Table 5.20: Revised Thrust Calculation Based on Chosen ElectronicsVariables 2004 2005dC l /d α 0.1 0.1
AR 1.422 1.422S (m 2 ) 0.048 0.102b (m) 0.26 0.381
e 0.9 0.9dC L /d α 0.041 0.041
AOA (deg) 9 9C L 0.95 0.95
W (g) 97.9 198.48C D 0.243 0.252C D0 0.018 0.027T (g) 25 52.573
As the search for motors continued, it became obvious that some gave a thrust
rating and others a power rating. Equation 5.4 [5] was used to determine if those motors
with a power rating would provide enough thrust. The motor needed a power rating of at
least 15.69 W (0.0210 hp).
(5.4) P = Power (W)
v = Velocity (m/s)
Twenty-two motors were then ranked on five attributes against a baseline.
Relative weights of the attributes were determined using the given Pairwise Comparison
Method. Thrust was of most importance (33%) with weight (27%) being also of high
importance. Cost (20%), current (13%), and voltage (7%) also help to evaluate the
motors. Table 5.21 for the comparisons.
vPT 97.101
=
51
Table 5.21: Relative Weighting of Attributes Pairwise Comparison:
Place an "R" if the row is more important. Place a "C" if the column is more
important Thru
st
Cost
Wei
ght
Curr
ent
Volt
age
Add
itio
nal 1
Add
itio
nal 2
Row
Tota
l
Colu
mn
Tota
l
1+Ro
w +
Colu
mn
Tota
l*
Rela
tive
Wei
ght
Thrust R R R R 4 0 5 33%
Cost C R R 2 0 3 20%
Weight R R 2 1 4 27%
Current R 1 0 2 13%
Voltage 0 0 1 7%
Additional 1 0 0 0 0%
Additional 2 0 0 0 0%
Column Total 0 0 1 0 0 0 0 15 100%
*Added 1 to each total to allow each parameter to have some percentage (except for undefined parameters)
A baseline motor was not chosen, but a range of values for each attribute was
chosen for baseline. These ranges are indicated by a rank of 2. The ranking system is
given in Figure 5.23. Note that two attributes (thrust and weight) have a rank of zero for
one range of values. Any motor that received a zero in either of these two categories
would definitely not be chosen because of its inability to perform the task at the
necessary level. Each motor was ranked and a normalized score was determined (Table
5.22).
52
3:152:20151:30200:30
3:73:202:972:80201:91:80
3:1303:8.02:1301052:18.01:105801:5.10:80
WWW
WeightWW
VCVC
VoltageVVCostCC
TITIT
CurrentIIThrustTT
>≤≤<≤
=>
>>≤≤≤≤
=>=>
<>≤≤
≤≤<≤=>=<
Figure 5.23: Attribute Ranking System
Table 5.22: Motor Ranking
Evaluate each additional concept against the baseline, score each attribute as: 0 = much worse than baseline concept 1 = worse than
baseline 2 = same as baseline 3 = better than baselinePropeller
Thru
st
Cost
Wei
ght
Curr
ent
Volt
age
Add
itio
nal 1
Add
itio
nal 2
Wei
ghte
d Sc
ore
Nor
mal
ized
Sco
re
T = 105-130 g; C = $20-80; W = 15-20 g; I = 0.8-1 A; V = 7-9 V ----- 2 2 2 2 2 2 2 2.00 94%
GW/EDP-50XC Direct Drive Power System with 2/EP-3020 EP3020 0 3 2 1 2 0 0 1.40 66%
GW/EDP-50XC Direct Drive Power System with 2/EP-3020 EP3020 0 3 2 2 3 0 0 1.60 75%
Feigao 1208430S 12x22mm Brushless Motor EP7060 2 2 3 1 2 0 0 2.13 100%
GW/EDP-150 Motor w/Capacitor & 2-Pin Black Motor Connector EP4530 1 3 0 1 3 0 0 1.27 59%
B2C Light Power System (GW/LPS-B2C-C) EP7035 0 3 3 3 2 0 0 1.93 91%
B2C Light Power System (GW/LPS-B2C-C) EP7060 0 3 3 2 3 0 0 1.87 88%
B2C Light Power System (GW/LPS-B2C-C) EP7060 0 3 3 2 2 0 0 1.80 84%
RXC Light Power System (GW/LPS-RXC-A) EP7060 0 3 1 2 3 0 0 1.33 63%
RXC Light Power System (GW/LPS-RXC-A) EP7060 0 3 1 1 3 0 0 1.20 56%
RXC Light Power System (GW/LPS-RXC-A) EP8043 1 3 1 1 3 0 0 1.53 72%
RXC Light Power System (GW/LPS-RXC-A) EP8060 0 3 1 2 3 0 0 1.33 63%
Astro Mighty Micro Brushless 010 (801 V 14T Direct Drive Motor) 5.5 x 4 MAS 3 1 0 1 2 0 0 1.47 69%
Astro Mighty Micro Brushless 010 (801 V 14T Direct Drive Motor) 5.5 x 4 MAS 3 1 0 1 2 0 0 1.47 69%
Astro Mighty Micro Brushless 010 (801 G 14T Geared Motor) APC 9 x 6E 2 1 0 1 2 0 0 1.13 53%
Astro Mighty Micro Brushless 010 (801 G 14T Geared Motor) APC 9 x 6E 3 1 0 1 2 0 0 1.47 69%
Astro Mighty Micro Brushless 010 (801 G 14T Geared Motor) APC 10 x 7E 3 1 0 1 2 0 0 1.47 69%
Astro Mighty Micro Brushless 010 (801 G 14T Geared Motor) APC 10 x 7E 3 1 0 1 2 0 0 1.47 69%
Sensorless 16 mm dia, Brushless, 15 W (EC 16, 266523) Not Given 0 1 0 1 1 0 0 0.40 19%
Sensorless 22 mm dia, Brushless, 20 W (EC 22, 200858) Not Given 1 1 0 3 1 0 0 1.00 47%
16 mm dia, Graphite Brushes, 4.5 W (RE 16, 118730) Not Given 0 1 0 3 1 0 0 0.67 31%
25 mm dia, Precious Metal Brushes CLL, 10 W (RE 25, 118743) Not Given 0 1 0 2 1 0 0 0.53 25%
Feigao 1208436L 12x30mm Brushless Motor EP3020 1 2 2 1 2 0 0 1.53 72%
Relative Weight 33% 20% 27% 13% 7% 0% 0%
53
Once the raking was completed, only three motors passed all tests. Table 5.22
shows the baseline case (yellow), the motors that will not suffice (gray), and the motors
that will be purchased and dynamically tested for thrust output (white). The 2005 MAV
team will study the Feigao 1208430S 12x22mm Brushless Motor (Figure 5.24) with the
EP7060 propeller (Figure 5.25), the RXC Light Power System (GW/LPS-RXC-A)
(Figure 5.26) with the EP8043 propeller (Figure 5.27), and the Feigao 1208436L
12x30mm Brushless Motor (Figure 5.28) with the EP3020 propeller (Figure 5.29).
Figure 5.24: Feigao 1208430S 12x22mm Brushless Motor
Figure 5.25: EP7060 Propeller
Figure 5.26: RXC Light Power System (GW/LPS-RXC-A)
Figure 5.27: EP7043 Propeller
54
Figure 5.28: Feigao 1208436L 12x30mm Brushless Motor
Figure 5.29: EP3020 Propeller
55
6 Design Objectives and Specifications
6.1 Performance Specifications
After performing the concept development and feasibility assessment, a more
concrete list of specifications were developed for the 2004-2005 Micro Air Vehicle
(MAV) team. The performance specifications are as follows:
• The vehicle shall not exceed 200 grams in mass
• The vehicle shall have a maximum linear dimension between 15 and 18 inches
• The vehicle shall have a minimum flight time of 15 minutes
• The controls, on-board video, and GPS shall have a minimum functional range of
600 meters
• The on-board video and GPS information shall be transmitted to a laptop
computer for viewing and analysis
• The vehicle shall have a cruise speed between 5 and 20 meters per second
• The vehicle shall be propelled by an electric motor and propeller combination
o The powerplant shall generate at least 55 grams of continuous thrust, with
a goal of greater than 80 grams
• Lithium polymer batteries shall be utilized to power the electric systems on-board
the vehicle
• All of the electrical systems must run on 7 volts and have a low continuous
current draw
6.2 Design Objectives
• The airframe shall be created out of composite materials if deemed reasonable
• The vehicle must be of a new design, and not rely on the design created by the
RIT 2003-2004 MAV team
• The vehicle design must be capable of being scaled down in size and still perform
well during flight
56
• All drawings and calculations shall utilize metric standard units
• All CAD works shall be created in Pro-Engineer or Solidworks
• All purchases must be approved by the Team Manager
6.3 Evaluation Criteria
The above specifications must be met prior to the project completion date for this
project to be determined a success. If the specifications are met prior to the completion
date of the project, further research and experimentation will be performed on possible
MAV designs. The completion date for the project is May 20th, 2004. The team will also
compete in the 9th International MAV Competition in Seoul, Korea to demonstrate the
emerging capabilities of the RIT MAV program.
57
7 Analysis
7.1 Airframe
7.1.1 Airfoil Testing Design
The next phase of the development of the airfoil consists of wind tunnel testing.
For this testing, we will be using RIT’s subsonic wind tunnel [4]. Andrew Walter, a
former RIT graduate student, developed a balance for the wind tunnel specifically
designed to measure the small lift, drag, and moment forces produced by MAVs [22]. In
this work, the balance gave accurate lift and drag information, but was unable to provide
accurate moment data. Josh Shreve, a current graduate student at RIT, is working on the
balance to provide accurate moment data. Upon confirmation from Josh Shreve that we
can indeed use the scale for measuring at least lift and drag, the airframe group will begin
their airfoil testing.
The airfoils to be tested include the S1210, S4022, S4083 and GOE417a. The
planform size and shape were determined for research and testing practicality
considerations. Due to the relatively small amount of data in the field of Micro Air
Vehicles, exact correlations between planform shape and aerodynamic characteristics are
not known. With this in mind, the idea of testing airfoils using a planform similar to that
expected for our MAV was rejected because that data would be hard to use by future
MAV teams. The decision was then made to use a rectangular planform for the airfoil
testing.
Practically speaking, wind in the tunnel interacts with every object in the tunnel,
including the walls, balance, and airfoil. For accurate information, the effects of the flow
interaction between the walls and airfoil and the flow interaction between the balance and
airfoil must be kept to a minimum. To reduce the effects of the wall on the flow around
the airfoil, we wanted to keep at least 6 inches of space between the airfoil and walls of
the tunnel. The balance already provides ample spacing between the top and bottom of
the tunnel and the airfoil. Thus, the side walls become the main concern. The wind tunnel
cross-sectional width is 29 inches. To accommodate the six inch buffer zone on the sides,
we decided upon a twelve inch wing span which provides more than a six inch buffer
58
zone on either side. To choose the chord length, we decided to keep an aspect ratio
similar to that which our MAV will have, which is about two. To find our average chord
length we used the following equations:
SbAR
2
= bcS *=
Where AR is the aspect ratio, b is the wing span, S is the wing area, and c is the average
chord length. Knowing our aspect ratio is two and our wing span is twelve inches, our
average chord length is determined to be six inches. Our testing airfoils are a six inch by
twelve inch rectangular planform.
7.1.2 Hot Wiring
In order for the MAV team to expedite construction time of airfoils for the
purpose of wind tunnel testing it was decided to construct the airfoils out of polystyrene
foam. This process, well known to model aircraft hobbyists, is known as “Hot Wiring.”
The process of hot wiring begins by cutting a template of the desired shape out of
a material with a high flash point such that it will not burn easily. The MAV team chose
to use balsa wood light ply because it combines an optimum resistance to burning while
at the same time remaining easy to manufacture. After the desired airfoil is traced onto
the balsa ply from a paper template, it is then rough-cut to shape using an exacto knife.
All edges of the template are then sanded as smooth as possible using high grit sand
paper. Care is taken to remove only as much material as is necessary to produce a
smooth surface finish. Afterward a bed is cut out with three squared edges. The fourth
edge defines the bottom surface of the airfoil. After the bed is cut out all edges are
sanded smooth. Once again care must be taken so that bottom profile of the airfoil
matches the top profile of the bed. Two airfoils and two beds are necessary for each
geometrical configuration to be tested.
Once the airfoils and beds have been manufactured, foam must be cut into
appropriately sized blocks (see step #1 below). This is done by hanging a sheet of foam
off the end of a table and marking off the block to be cut off the sheet. The foam bow is
then hung over the sheet on the line to be cut. Alligator clips are then attached to bow
(the clips attach to the wire spanning the bow). The alligator clips carry the current from
59
a transformer across the bow thereby heating the wire because of its resistance. A power
setting between 3 and 6 on the transformer is used depending on the width of the foam to
be cut. The transformer is then turned on; the wire heats up, and melts its way through
the foam producing a clean trued surface. Care must be taken so that the bow is not
swinging when cutting through the foam in order to keep a tight profile tolerance.
After the foam is cut to shape the foam beds are pinned to opposite surfaces of the
foam block using pushpins (see step #2 below). Care is taken such that the bottom of the
beds are true to the bottom of the foam and that the front edges of the beds are
approximately lined up ~0.25 inches in front of the front face of the foam. This creates a
ledge for the foam bow to rest on. Weight is placed onto the top surface of the foam to
keep it in place (see step #3 below). The foam bow is then placed on the table with the
wire spanning across the ledges of the beds. The electrical leads are attached to the foam
bow wire then run through a system of pulleys leading to a counter weight that is allowed
to drop under its own mass thereby creating tension in the leads (see step #4 below). The
electrical leads continue back to the transformer. The transformer is then turned on to the
appropriate power setting and the how wire is drawn through the foam creating the
bottom surface of our airfoil. Care must be taken such that the wire follows the contour
of the bed. In most cases it is necessary to place some downward force on the wire
(fingers always outside of the alligator clips) such that the hot wire remains true to the
contour. The airfoil is then pinned into place above the bed, the hot wire returned to the
ledge, and the process repeated over the top of the airfoil template (see step #5). Barring
any imperfections the process produces one airfoil and a top and bottom contoured foam
bed that is useful in laying composite over the foam airfoil.
60
Figure 7.1: Blocking out foam
Figure 7.2: Attach airfoil template to foam
Figure 7.3: Setting up foam and bow for hot wiring
Step 1: Blocking out foam.
Step 2: Attach airfoil bed.
Step 3: Weight down foam and place bow in position.
Transformer
Foam Bow
Ledge for Foam bow wire.
Pushpin.
Airfoil bed.
Aluminum tube.
Fiberglass rod.
Wheel.
61
Figure 7.4: Counter weight used to create wire movement
Figure 7.5: Utilize airfoil templates to cut out rest of airfoil
Figure 7.6: Airfoil after hot wiring process
Step 4: Attach electrical leads to bow, route through the pulley, attach to counter weight and turn power on.
Step 5: Pin airfoil template in the correct position and repeat step 4 except cut over top the airfoil.
Step 6: Turn off power; remove weight from foam, foam airfoil, and templates. Inspect for defects and clean up leading and trailing edges using high grit sand paper.
Counter weight.
Airfoil template pinned in place over top the bed.
62
7.1.3 Fiberglassing
After hot wire cutting our four airfoils for testing, it was decided that greater
strength was needed. The foam airfoils do not have the required strength to withstand the
forces of the wind tunnel. Due to the ease of manufacturing and its strengthening
properties, the team decided to cover the foam airfoils with a ply of fiberglass.
Fiberglass work can be summarized in six simple steps. The first step is to get all
the correct materials and tools together. Without the correct tools and materials the final
product will not come out correctly. Suggested materials and tools needed are: Epoxy,
fiberglass, a plastic squeegee, scissors, a single edge razor blade, mixing cups, a yard
stick, rubber gloves, paper towels, and a drop cloth. If the work surface is glass, a drop
cloth is not required, but can be helpful. For the MAV airfoils to be strengthened US
Composites Epoxy was used as well as US Composites 0/90 degree woven fiberglass
cloth. Once all the materials and tools are acquired it is possible to move on to step 2.
The second step is to size up the job. The fiberglass cloth should be cut down to a
size slightly larger than what is required to cover the airfoil. Usually an extra inch and a
half in each direction will provide a sufficient amount of fiberglass to account for
anything being off center when applying the fiberglass to the foam airfoil. In some cases
it is feasible to use strips of fiberglass to work with. Working with multiple strips,
however, usually results in more sanding after the lay up is complete. Therefore, it
seemed only fitting to use a single piece of fiberglass to laminate the airfoils with. Sizing
up the job also helps to provide for the correct amount of epoxy resin that should be used,
cutting down on waste.
The third step is to lay out all of your tools in the order in which they will be used,
or at least within an arm’s reach so that time is not spent searching for tools while the
epoxy resin is curing. It is important that the drop cloth, or glass surface is at least two
and a half times the size of the fiberglass cloth being used. Following this standard will
provide that there is enough room, not only to work with the fiberglass, but also for all of
the tools and supplies to be within quick reach throughout the lay-up process. For the
MAV team a fiberglass station was set up in the wind tunnel area, away from all of the
other tables and stations. A plexiglass sheet was placed on a table top to provide the
smooth flat surface to lay up the fiberglass on.
63
The fourth step begins the actual fiberglass lay up process. Before starting this
step it is very important that the fiberglass cloth is on a smooth flat surface and all the
tools are in order. With rubber gloves on the epoxy resin can be mixed. It is essential
that the 3:1 ratio is kept between the resin and hardener. The RIT Aero Team and MAV
team have observed that the US Composites Epoxy can be very temperamental and if the
ratio of resin to harder is off slightly, the final product will not turn out as desired. To
achieve an optimum mix of hardener and resin the MAV team will use calibrated ratio’d
pumps or a scale. Once pumped into the mixing cup the resin and hardener should be
mixed for two minutes, making sure to scrape the sides of the cup to ensure all the resin
is in solution. Once the epoxy is thoroughly mixed, the fiberglass cloth can be saturated
with the epoxy. The plastic squeegee should be used to ensure that the fiberglass cloth is
thoroughly and evenly saturated with epoxy. The fiberglass cloth will begin to turn clear
when it’s completely saturated with epoxy.
During step five the saturated fiberglass is placed on the foam airfoil. Before
applying the fiberglass to the foam airfoil, if any epoxy is remaining after saturating the
fiberglass, it can be applied to the surface of the foam airfoil. Through the MAV team’s
experience, however, this does not improve the quality of the final product, but can add
extra variable to the process, therefore extra epoxy will be used for saturating more
fiberglass, or disposed of. The saturated fiberglass sheet can be applied to the airfoil and
air bubbles or imperfections can be smoothed using the plastic squeegee.
Step six is when the fiberglass laminated airfoil should be stored so that it can be
left to cure for twenty-four hours. The extra fiberglass around the edges of the airfoil can
be trimmed. Once the airfoil looks satisfactory it can be placed back into the foam block
it was cut from. The foam block provides a structure that helps the fiberglass bond to the
foam airfoil as weight is applied to compress the foam mold around the fiberglass
laminated airfoil. After twenty-four hours the epoxy is fully cured and the airfoil can be
removed. During curing, the foam mold must be covered with something to keep the
epoxy saturated fiberglass from adhering to the mold. For the MAV team’s first trials
cellophane was used to cover the foam block. The surface roughness of the airfoils using
this method was not satisfactory and we approached the Aero Team for suggestions. It
was suggested that the MAV team use Mylar in between the airfoil and foam block. The
64
first trial with Mylar resulted in a disaster. The Mylar stuck to the fiberglass at some
areas on the airfoil. In searching for a reason, the MAV team members concluded that
the Mylar stuck due to the fact that as epoxy cures it experiences an exothermal reaction
and heat is expelled. To combat this, the Mylar was waxed, and all trials utilizing the
waxed Mylar have produced promising results. Wet sanding can smooth out any
imperfections and help produce a smooth surface for the fiberglass laminated airfoil.
Laminating the foam airfoils with a ply of fiberglass gives each airfoil a
substantial increase in strength, with minimal weight gain. Depending on the airfoil, the
fiberglass laminate adds twenty to thirty grams to the weight of the airfoil. For future lay
ups, the MAV team hopes to make use of the vacuum bagging equipment available to
hopefully improve the uniformity of our airfoil manufacturing.
Figure 7.7: Layout of supplies
Figure 7.8: Sizing fiberglass
Step 1: Obtaining supplies.
Step 2: Sizing up the job.
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Figure 7.9: Laying out fiberglass
Figure 7.10: Saturating fiberglass with epoxy
Figure 7.11: Laminating airfoil with fiberglass
Step 3: Laying out Tools and Supplies.
Step 4: Saturating the Fiberglass.
Step 5: Laminating the airfoil.
Fiberglass will become clear as saturated.
Fiberglass must be smooth over airfoils, so the plastic squeegee is used to remove air bubbles.
66
Figure 7.12: Weight applied to airfoil during curing process
7.1.4 Wind Tunnel Testing With the planform and airfoils selected, we hotwired the airfoils to the proper
dimensions. The pure foam wing was determined to be too weak, so a layer of fiberglass
was applied to the airfoils to provide the additional strength required for the testing.
Currently, we have two airfoils ready to be tested and we need to add the final touches to
the other two airfoils before they can be tested in the wind tunnel. At the present time, we
are waiting for confirmation from Josh Shreve that the balance is ready for use before we
start testing.
Each airfoil and subsequent control surface configuration will be mounted in the
wind tunnel via an aluminum rod. The rod attaches to the airfoil at the trailing edge of the
airfoil along the centerline. Two mounting screws, an inch apart, hold the airfoil to the
balance. The airfoils will be mounted in an inverted position. The airfoil will be tested
upside down for one main reason. This reason is that our design idea for control surfaces
includes a V-tail protruding from the pod below the wings. With this orientation,
mounting the airfoil upright on the balance will cause a great deal of flow disturbance
directly due to the balance. To alleviate most of the disturbances due to the balance as
possible, the airfoil will be mounted upside down. Figures 7.13 and 7.14 show the
mounting of an airfoil.
Step 6: Curing.
Weight is applied to the top of the foam mold with fiberglassed airfoil inside for 24 hours to cure
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Figure 7.13: Airfoil attached to mounting rod
Figure 7.14: Airfoil mounted on balance
We will be testing each airfoil at four different Reynolds numbers including
60,000, 100,000, 150,000 and 200,000. For each Reynolds number, ten angles of attack
will be tested. The table below shows the data we expect to collect for one test (moment
data is hopeful, but not necessary at this time).
Airfoil
Aluminum Mounting Rod
Airfoil Mounted in Inverted Position
Aluminum Mounting Rod
Testing Balance
68
Table 7.1: Data Recording Table for Wind Tunnel Testing of Airfoils Re: 60,000
Vwind (ft/sec) Re
AoA (deg)
Lift (grams)
Drag (grams
Moment (grams-
inch) CL CD Cm -6 -3 0 3 6 9 12 15 18
With the successful completion of these tests, we will then test the control surface
configurations in a similar manner. Using a planform that is decided up, we will attach
various configurations of control surfaces. Only one airfoil, the airfoil we decide upon,
will be used for the control surface testing. Another dimension will be added to the
testing matrix however, and that is the deflection of the control surfaces. The control
surface deflections that will be looked at include -60o to 60o, incremented by 15o. The
following table shows the data we expect to collect.
Table 7.2: Data Recording Table for Wind Tunnel Testing of Control Surfaces
Re: 60,000 Deflection: -60o
Vwind (ft/sec) Re
AoA (deg)
Lift (grams)
Drag (grams
Moment (grams-
inch) CL CD Cm -6 -3 0 3 6 9 12 15 18
After successful completion of the testing of control surfaces, we will then design
and build a full scale MAV implementing the airfoil and control surface designs. Wind
69
Tunnel Testing will then be conducted for the same four Reynolds numbers in the same
testing matrix was used for one airfoil. Modifications will be made as needed. Finally, a
full-scale working MAV will be constructed and flown. Modifications to the plane will
then be made on an as-needed basis.
7.2 Electronics Over the next ten weeks, the following list of experiments has been proposed: 1. Capture/verify the waveforms coming out of the RF Receiver.
To perform this test, the RF receiver is placed on a breadboard with the
oscilloscope capturing the signal from each pin for various input states. The input states
are generated using the RF controller. Once a waveform appears on the oscilloscope, this
means that the channel is reacting to the input from the RF controller.
2. Range test using the RF Receiver interfacing with a LED test circuit.
Using the same circuit from the previous experiment, six LEDs are connected to
each channel to represent that the signal has been received. This will determine the
signal strength at different ranges using the RF controller as the input. The ranges are
from 100 meters to 800 meters.
3. Find out the amount of time for the battery to discharge.
Connect the battery to a given set of resistive loads and obtain the time required
for a full discharge. The set of resistive loads can vary from 5 kΩ to 500 kΩ. The MAV,
when fully constructed, should be similar to a given discharge pattern. Thus a predicted
flight time can be calculated.
4. Testing of the speed controller.
Connect the RF receiver to the speed controller and verify the output of the speed
controller corresponds to acceleration and deceleration of the motor. This can be
determined using an oscilloscope connected at the output of the speed controller to view
the change in amplitude of the waveform.
70
5. Servo test to verify the functionality.
Connect the RF receiver to servo #1 and verify its full displacement using input
from the RF controller. Once the full displacement of the servo has been verified through
visual inspection, repeat the procedure with servos #2 and #3. It is necessary to use the
same pin on the RF receiver in conjunction with the same set of input for all three servos.
6. Control testing with the mixing chip.
Referring to the RF receiver pin-out diagram, connect servos #2 and #3 in
conjunction with the mixing chip to verify the simultaneous displacement of the servos.
The expected pattern of displacement is based on the control surface configuration
suggested by the airfoil subgroup.
7. Voltage regulator test circuit to verify its functionality.
For each onboard electrical component, specific operating condition must be
satisfied. Individual voltage regulation circuits must be constructed specifically for each
component. Using resistive loads to represent onboard components, the various voltage
regulation circuits will be constructed and proper functionality over time will be verified.
Voltage regulation waveforms will be captured for reference.
8. Testing the camera, video transmitter, and receiver to verify the functionality.
Apply power to the camera and the video transmitter. Then connect the video
receiver to the laptop. Once the connections are made, verify the transmission of signals
via the real-time video image displayed on the laptop.
9. Testing of the antenna array to verify functionality, coverage area, and signal
strength.
Using a similar setup from experiment 2 and the fully-built onboard video system,
a range test using the antenna array at the base station will be conducted. From a range of
100 meters to 800 meters, the signal strength will be verified from the real-time image on
the laptop. A similar test will be conducted for the coverage area test.
71
10. Testing of the GPS receiver and the Video overlay board to verify its functionality.
Apply power to the GPS receiver and the video overlay board, and connect the
onboard video system to the video overlay board. Repeat experiment 9 with the
video/GPS circuit and verify successful image transmission from laptop.
11. Full circuit wiring test board.
Wire all the components together on one circuit board to represent the fully wired
MAV. Also included in the wiring, is the onboard battery cells.
12. Run simulated flight path. Verify full system operation using an RF controller and visual inspection of:
On-screen video image with GPS data
Servo displacement based on control inputs
Motor acceleration and deceleration
Range test
System fallout test
13. Actual construction and implementation on the MAV. 14. Repeat experiment 12 for the implementation.
7.3 Propulsion
7.3.1 Propulsion Static Testing
A goal of the 2004 MAV team was to leave behind a static test setup that future
MAV teams could use to test propulsion components. We reviewed the test setup, and
have decided to use the basic idea for our testing. To measure static thrust, a
commercially available load cell is used to sense an applied thrust and provide an output
as voltage. The S250 Miniature Platform SMD load cell, with a 2mV/V output has been
selected (see appendix for specifications). Figure 7.15 shows the load cell chosen for
testing.
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Figure 7.15: SMD S250 Miniature Platform Load Cell
The entire static test setup includes: a load cell, power supply, instrumentation
amplifier, multi-meter, and an oscilloscope. The basic principle of the setup is for the
load cell to be supplied with proper excitation, and for the amplifier to increase the output
so that the multimeter can obtain accurate voltage measurements. The oscilloscope will
be used to monitor the signal for any interference.
7.3.1.1 Calibration
The static test setup has been calibrated following the report of the 2004 MAV team
[19]. The electrical schematic, taken from the previous team, is shown in Figure 7.16.
Figure 7.16: Calibration electrical schematic
The power supply provided 5 volts to the load cell and 12 volts to the amplifier.
A 100 ohm resister was used to amplify the load cell’s voltage output (see appendix for
amplifier data). This allowed the output voltage to be accurately measured and recorded.
Calibration of the load cell was completed by using a set of calibrated weights ranging
from 0–100 grams (0-0.2205 lbf). The calibration was run with the weights hanging from
DC 6V EE Power Supply 5V +/- 12V Serial #EE1015
Load Cell SMD Sensor S250 1kg
100ohms
INA114 Instrumentation Amplifier Burr Brown w/ 100ohm Resistor
Oscilloscope Tetronix Serial #B012473
Vc
Ref
Multimeter (// to Oscilloscope) Craftsman Serial #CCL02079157
DC 12V EE Power Supply 5V +/- 12V Serial #EE1015
25Kohm 25Kohm 25Kohm
25Kohm 25Kohm
25Kohm
73
the motor mount to imitate actual thrust testing. Figure 7.17 shows the calibration setup
in the lab.
Figure 7.17: Calibration setup
Several trials showed consistent results for calibration, which are presented in Figure 7.18.
Load Cell Calibration
y = 0.0056x + 0.9795R2 = 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
0 20 40 60 80 100 120
Load (grams)
Voltage (
V)
Figure 7.18: Load cell calibration
From the calibration results, we have obtained a linear relationship between
applied load and output voltage, which allows thrust, T, to be calculated in grams as:
T(V) = 178.36V - 174.71
Using this result, we will be able to calculate the thrust produced by each motor and
propeller tested. Throughout testing, calibration will be repeated to ensure results are
accurate.
7.3.1.2 Test Setup
The motor mount left behind by the 2004 MAV team will be used for static
testing. The mount is made out of PVC tubing. It is used to attach the motor to the load
74
cell with two screws that thread into holes built into the load cell design. Figure 7.19
shows the individual motor mount, which will attach to the load cell [19].
Figure 7.19: Motor Mount
The setup varies slightly from the 2004 MAV team, in that the propeller will be
mounted horizontally instead of vertically. This change in setup has been made so that
there will be no variations in setups when dynamic testing is completed. For each test,
the propeller will be secured to the motor with epoxy to ensure it remains on the motor
shaft while allowing for its easy removal after testing.
The electrical schematic for thrust testing is very similar to the calibration setup.
In addition to the calibration circuit, a motor circuit would be setup using a power supply
(Shenzhen Mastech DC Power Supply: HY3003-3), the motor, an additional resistor, and
a multimeter (MPJA Multimeter Serial # CCL010412272). The setup will power the
motor and allow for the current draw of the motor to be calculated from the additional
resistor and voltage measured across the multimeter. Figure 7.20 shows a schematic of
the additional motor circuit.
75
+ -
R L
Motor
MPJA Multimeter Serial # CCL010412272
R1
Shenzhen Mastech DC Power Supply: HY3003-3
Note: R1 is dependent on the internal resistance (R) of the motor
T
Figure 7.20: Motor test electrical schematic
7.3.2 Propulsion Dynamic Testing
To help choose the best motor and propeller combination for our MAV, we have
decided to complete dynamic testing of propulsion items. There was no existing test
setup that would meet our needs in the wind tunnel, so a testing method needed to be
developed. The two design ideas which were considered for dynamic testing are the
following: 1) to mount the static test setup in the wind tunnel on a block which could be
rotated to predetermined angles of attack, or 2) to design an inexpensive setup similar to
the existing setup, by replacing the current expensive load cells with strain gages, which
would allow measurements of drag and lift at numerous angles of attack. In order to
decide on the best method for our needs a listing of pros and cons for each design was
compiled.
1) Mount Static Test Setup to Rotating Block
Pros:
• Minimal time designing setup
• Only one part required to be made
76
Cons:
• Only a few angles of attack can be tested in a short time frame
• Each angle of attack tested would require precise geometric calculations and
measurements to be made
• Maintaining angel of attack during testing will be difficult
• Can only be used for one specific application
2) Design Strain Gage Setup
Pros:
• Will allow for quick and accurate changes to any angle of attack using wind
tunnel software
• Once set up, testing can run quickly
• Universal RIT application for future projects
Cons:
• Large amount of design calculations needed
• Strain gages need to be mounted (time consuming, less accurate than load cells)
Taking into consideration the above pros and cons it was decided to proceed with
the second option of designing an inexpensive setup which will hook up with the wind
tunnel’s current hardware and software. This method involves an initial time investment,
but once the setup is in place wind tunnel testing will be accurate and fairly quick to carry
out. In addition, future MAV teams, and other RIT teams or classes, will be able to use
the test setup. Figure 7.21 shows the existing wind tunnel setup and how it has been
designed to be modified for MAV testing.
Figure 7.21: Existing wind tunnel setup (left) and new MAV dynamic test setup (right)
77
The main principle of the dynamic test setup is the strain gage block. The block
is designed so that the strain gages will pick up forces due to drag. Buckling calculations
were completed to find an appropriate thickness to mount the strain gages to, based on a
maximum force of 500 grams (1.1 lbf). Calculations were done with drag as 200 grams
(0.44 lbf), (based on a maximum expected thrust of 120 grams [0.265 lbf]) and with lift as
458 grams (1.01 lbf). The appropriate thickness for strain gage mounting was found to be
0.0508 cm (0.02 inches). Further details of the analysis and drawings of parts that will be
fabricated are in the appendix.
The strain gages will be mounted where the drag force is most prominent in order
to obtain accurate measurements. Figure 7.22 shows the chosen location on the strain
gage box to mount the strain gages.
Figure 7.22: Strain gage mounting locations
The stain gages will be calibrated before testing to obtain the relationship between
the output voltage of the strain gage and an applied force. This calibration will be done
similar to the load cell calibration. Using the outputs of strain gage 1 and strain gage 2,
the force due to drag can be calculated as:
FD = Fstraingage2 - Fstraingage1
During testing, strain gage 2 will be in tension and strain gage 1 will be in compression.
During dynamic testing, the RPM of the motor shaft will also be recorded. This
will be done by painting a white vertical line on the back of the propeller and using a
strobe light to match the frequency of the propeller [8]. The frequency at which the strobe
light matches the propeller rotation will be used to calculate the RPM of the motor shaft.
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7.3.3 Future Testing Plans
Using the static and dynamic test setup described above, the best motor and
propeller combination will be selected for the MAV. From our literature research we
found that if an off the shelf propeller was used, the propeller should be chosen first and
then a motor matched to the propeller. Using this, we will test in the following order:
static propeller, static motor, dynamic propeller, and dynamic motor. If the conclusion of
dynamic motor testing does not identify the proper motor and propeller combination,
selected combinations of motors and propellers will be dynamically tested.
Propellers will be tested with both modified tips and in original shape. Based on
the work done by the 2004 MAV team the two most promising modifications A and B
(see Figure 7.23) will be used. For modification A both the leading and trailing edge of
the propeller will be rounded, and for modification B only the leading edge will be
rounded [19]. In addition, as testing is carried out, modifications to the propeller
diameter may be made in order to achieve the needed thrust.
Figure 7.23: Propeller modification A (left) and B (right)
7.3.3.1 Static Propeller Testing
Using the static setup described previously, each propeller will be tested on the same
motor so results will be comparable. Each propeller will be tested with an input voltage
to the motor of 4-9 volts in 1 volt increments. In addition, each propeller will be tested at
the manufacture’s recommended nominal voltage of 7.4 volts. Both the current draw of
the motor and the output voltage of the load cell will be recorded. Using the thrust
equation determined from calibration thrust will be calculated. With this, data plots of
current versus thrust will be made for each propeller. This will identify which propeller
has the best static thrust to current ratio.
79
7.3.3.2 Static Motor Testing
Each motor will be tested using the propeller that had the best static thrust to
current ratio. Similar to static propeller testing, each motor will be tested with an input
voltage to the motor of 4-9 volts in 1 volt increments. Again, the current draw of the
motor and the output voltage of the load cell will be recorded. A thermocouple will be
mounted to the motor bodies to record temperature during testing. This will be done to
obtain the relationship between temperature reached by the motor and current draw.
Similar to the propeller, current versus thrust will be plotted for each motor, and the ratio
of thrust to current calculated. In addition, a thrust to weight ratio will be calculated for
each motor at the manufacture’s recommended voltage of 7.4 volts.
7.3.3.3 Drag of the Dynamic Test Setup
Before dynamic testing can begin, the drag of the setup must be determined so it
can be taken out of the test results. Drag will be measured on the dynamic test setup
described before as a function of airspeed from 8-15 m/s (26.2-49.2 ft/s) in 1 m/s (3.28
ft/s) increments. Using the equation for force due to drag explained earlier, the output
voltage from the strain gages will allow the drag of the setup, Dsetup, to be calculated.
This drag measurement will be used to determine the actual propeller and motor drag
from testing.
7.3.3.4 Dynamic Propeller Testing
Based on static testing results, all propellers may not be selected for dynamic
testing if they do not appear to be worthy candidates, which will simplify testing. Using
the dynamic setup, each propeller will be tested in the wind tunnel, with the same motor
that was used for static testing. Propeller drag, Dp, can be calculated as:
Dp = Dt – Dm – Dsetup
Where Dt is total drag recorded during propeller testing, Dm is the drag of the
motor, and Dsetup is the drag of the dynamic test setup. The drag due to the motor can be
found in the same way that the drag due to the test setup is found. Both Dm and Dsetup
will be constant, with respect to air speed, for all dynamic propeller tests.
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By testing in the wind tunnel, new independent variables of air speed and angle of
attack are introduced. The independent variables will be tested in the following ranges:
input voltage to the motor of 4-9 volts in 1 volt increments, air speed of 8-15 m/s (26.2-
49.2 ft/s) in 1 m/s (3.28 ft/s) increments, and angle of attack of 0-10 degrees in 5 degree
increments. During each run drag, thrust, motor current draw, and the motor shaft RPM
will be recorded. To compare the propellers, a thrust to weight ratio, and a thrust to drag
ratio (both at 7.4 volts) will be calculated. In addition, a thrust to current ratio will be
calculated which will be compared to the static thrust to current ratio to show the
variation between static and dynamic testing.
Figure 7.24 shows a portion of the test matrix that has been developed for
propeller dynamic testing. Each propeller and modified versions will have the same
testing sequence.
T1 T2 T3 T1 T2 T3 T1 T2 T3 T1 T2 T3 T1 T2 T3 T1 T2 T305
1005
10: : :4 15 10
05
1005
10: : :5 15 10
: : :05
10: : :9 15 10
Voltage (V) Wind Speeds (m/s)
Propeller EP7060
4
Thrust/Weight Ratio Thrust/Drag RatioAOA (deg) Drag (g) Thrust (g) RPM
8
9
8
Current Draw (A)
9
5
89
Figure 7.24: Example dynamic test matrix.
7.3.3.5 Dynamic Motor Testing
A similar method to the dynamic propeller tests will be used for dynamic motor
tests. The propeller with the best thrust to drag ratio from dynamic testing will be used
for all motor tests run. The same independent variables and ranges will be tested for the
motors: input voltage to the motor of 4-9 volts in 1 volt increments, air speed of 8-15 m/s
(26.2-49.2 ft/s) in 1 m/s (3.28 ft/s) increments, and angle of attack of 0-10 degrees in 5
degree increments. In addition, a thermocouple will be mounted to the motor bodies, as
81
was done during static testing, to record the motor’s temperature during testing.
Temperature information may be useful when selecting motor position for the final
design. The same data collection and calculations will be made for the motors as will be
done for the propellers.
To calculate the drag of each motor, the output voltages of the strain gages will be
used again. The drag of the propeller will have already been found during dynamic
propeller testing so that the drag due to the motor, Dm, can be calculated:
Dm = Dt – Dp – Dsetup
Where Dt is total drag recorded during motor testing, Dp is the drag of the propeller, and
Dsetup is the drag of the dynamic test setup. Both Dp and Dsetup will be constant, with
respect to air speed, for all dynamic motor tests.
Once both static and dynamic testing has been completed the best motor and
propeller combination will be chosen. When making the final decision between
combinations which meet the thrust requirement of 80 grams (0.176 lbf), the weight to
thrust ratio and the thrust to current ratio will be given the most consideration.
Minimizing weight is important in all aspects of an MAV design, and also the propulsion
system must not exceed the current provided by the electrical system. The final motor
and propeller combination will be further tested during flight tests.
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8 References [1] “Airfoil Database Tailless and Flying Wings,” http://www.aerodesign.de/english/profile/profile_s.htm#hs520, Hartmut Siegmann, (1998-2004). [2] “Micro Air Vehicle Design Papers”, 7th Annual MAV Competition. University of Florida, April, 2003. [3] “Micro Air Vehicle Design Papers”, 8th Annual MAV Competition. University of Arizona, April, 2004. [4] “Rochester Institute of Technology Wind Tunnel,” http://www.rit.edu/~ritaero/windtunnel/ [5] Anderson, John D., Jr., Introduction to Flight. 4th Edition, McGraw Hill, New York, 2000. [6] Deperrois, André. XFLR5, A tool for the design of Airfoils for Model Aircraft and Sailplanes operating at low Reynolds numbers, (2003). [7] Grasmeyer, J.M. and Kennon, M., Development of the Black Widow Micro Air Vehicle, Progress in Astronautics and Aeronautics,Vol 195, 2001 [8] Kotwani, Kailash et. al, “Experimental Characterization of Propulsion System for Mini Aerial Vehicle,” 31st National Conference on FMFP, Jadavpur University, Kolkata, December 16-18 2004. [9] Kunz, Peter and Kroo, Ilan, “Analysis and Design of Airfoil for use at Ultra-Low Reynolds Numbers,” published in American Institute of Aeronautics and Astronautics, (2001). [10] McMichael, J., Francis, M.: Micro Air Vehicles - Toward a New Dimension in Flight, Defense Advanced Research Projects Agency, 1997. [11] Mueller, T. J. , Fixed and Flapping Wing Aerodynamics for Micro Air Vehicle Applications, AIAA, Virginia, (2001). [12] Nelson, Robert C., Flight Stability and Automatic Control, 2nd ed., McGraw-Hill, Boston, 1998. [13] Niu, C. Y. Michael, Airframe Stress Analysis and Sizing, Conmilit Press Ltd., 2nd ed., 2001. [14] Selig, M. S. et al, Summary of Low-Speed Airfoil Data, Vol. 1, University of Illinois at Urbana-Champaign, SoarTech Publications, Virginia Beach Virginia, 1997. [15] Selig, M. S. et al, Summary of Low-Speed Airfoil Data, Vol. 2, University of Illinois at Urbana-Champaign, SoarTech Publications, Virginia Beach Virginia, 1997. [16] Selig, M. S. et al, Summary of Low-Speed Airfoil Data, Vol. 3, University of Illinois at Urbana-Champaign, SoarTech Publications, Virginia Beach Virginia, 1997. [17] Selig, M.S., “UIUC Airfoil Data Site,” http://www.aae.uiuc.edu/m-selig/ads.html. [18] Shigley, E. Joseph and Mischke, R. Charles, Mechanical Engineering Design, 5th Ed., 2002.
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[19] Szachta, C. et al, Critical Design Report, Rochester Institute of Technology, Rochester, New York, 2004 [20] Szachta, C. et al, Preliminary Design Report, Rochester Institute of Technology, Rochester, New York, 2004 [21] Torres, G. E., Aerodynamics of Low Aspect Ratio Wings at Low Reynolds Numbers with Applications to Micro Air Vehicle Design, University of Notre Dame, Notre Dame, Indiana, 2002. [22] Walter, Andrew, Design, Fabrication and Testing of Longitudinal Wind Tunnel Balances for Micro Air Vehicle Applications, Rochester Institute of Technology, Rochester, New York, 2004.
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