RIT Micro Air Vehicle
Preliminary Design Report
February 2005
Brian Gillis Dr. Jeffrey Kozak Team Leader Team Advisor Mechanical Engineering Mechanical Engineering
Joshua Baker David Hein Aimee Lemieux Mechanical Engineering Mechanical Engineering Mechanical Engineering
Tzu-Chie Fu J.E.D. Hess Atul Phadnis Computer Engineering Mechanical Engineering Electrical Engineering
Aaron Grilly Cuong Le Victoria Schoennagel Mechanical Engineering Computer Engineering Mechanical Engineering 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
2 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
3 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
4 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.
5 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
6 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) 2 Acs Cross-sectional Area (cm ) 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)
7 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)
8 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.
9
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,
10 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
11 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.
12 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.
MAV Team 05-001
Airframe Electronics Propulsion
Joshua Baker Cuong Le Aaron Grilly Tzu-Chie Fu Aimee Lemieux David Hein Atul Phadnis Victoria Schoennagel JED Hess
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
13 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.
14 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.
ρVc Re= (3.1) µ
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.
15 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
16 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.
17
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
18 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.
19 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
20 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
21 transmission/receiving. This would save in the overall weight of the electronics, and make the GPS a greater reality.
22 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
23 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.
24 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.
25 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 Figure 5.1: Examples of different planforms, in other MAV’s. Part of the logic behind these showing trailing vortices [1].
26 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: bb2 AR = = , where Sc
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.
Figure 5.2: The BYU Stableyes is an example Figure 5.3: The Black Widow [6] is an of a conventional planform with a pusher example of a modified inverse motor configuration [3]. Zimmerman planform with a puller fi i
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
27 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
28 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- Figure 5.4: Separation bubble on airfoils at low speeds drag ratios and have the most stability for low Reynolds number flows. The main
29 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.
30 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.
Figure 5.5: MIT Lincoln Laboratory MAV Concept
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:
31
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
32 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.
33 Table 5.2: RF Receivers
Manufacturer Hitec Airtronics Hitec GWS- GWS- GWS- Model HFS- 92515Z Feather PICO NARO NARO 04MG R4P-JST R-6N/V 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.
34 Table 5.3: Pugh Chart for RF Receiver Selection
) ) -
O O -
O O
O
O
s
s r
r R R S T R R S T r
r c
C c C
a
V a V H Z H Z i i F F e A A S e A A I S c I / c e G / / e G / n 5 n 5 h P h P e e H y J H y J o N N o N N 1 N 1 N N N M t M t t
t
- - r - r - i i - - - - 5 6
Factors/Candidates 5 6 6 t a 6 t c a 4 c 4 t t P P S S s s 2 S S - H e 2 S S - e - H e e - r 0 r 0 4 4 i i a t a t 9 F 9 R F R R R W i W i R R W L W W L W A A
( ( G G H H G G G G
Channels 3 4 3 3 5 5 Current Needed 1 1 1 3 5 5 Weight 1 1 3 5 4 3 Volume 2 1 3 4 4 5 Voltage Needed 4 3 4 3 3 3 RX Range 5 1 3 2 3 3
Mean Score 2.66666667 1.8 2.8 3.3 4.0 4.0 Normalized Score 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 Wes-Technik Wes-Technik Phoenix-10 Model YGE-6 YGE-3 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
35 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
k k k k i i ) i i ) 0 0 r n n r n n
1 1 a a h h h h 3 6 3 6 - - e c c e c c - - - - x x e e e e Y Y i i E E
E E
T T n T T Factors/Candidates n t t G G G G e - - e - - s s Y Y s s o Y Y s s o a a e e h e e h L L P ( P ( W W W W
Weight 3 3 1 Current Needed 3 4 1 Compatibility 1 1 5 Dimensions (mm): 3 3 1
Mean Score 2.5 2.8 2.0 Normalized Score 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:
36 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)
37 Table 5.8: Pugh Chart for Video Transmitter Selection
0 0 0 0 0 5 0 5
2 0
2 0 w w 0 0 X X 0 0 X X w w m m 4 4 R R 4 4 R R m m
/ / / / 2 2 2 2
0
Factor/Candidate 0 0 X X V V 0 X X V V 0 0 5 5 T T T T A A A A 2 2 W W W W
B B
B B
Operating Frequency 3 3 Transmission Power 2 5 Number of Channels 3 3 Operating Voltage 3 3 Current Drain 3 2 Weight 3 2 Dimensions 3 3
Mean Score 2.857142857 3.0 Normalized Score 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.
38 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 100mA (55mA @3.3V)
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
a a S S n n S S
n
n l l
- e
- e
e e o 9 o t 9 t S V S V t t n n n n 7 7 P P n n A A u u - - A A a a G G r r
Factors/Candidates
N N r r H u H - u - t t a a U U F r F G r G O O S S a a C C m I m I
P P S S
Accuracy 3 3 5 Weight 3 2 1 Dimensions 3 2 2 Operating Voltage 3 3 2 Current Drain 3 1 2
Mean Score 3.0 2.2 2.4 Normalized Score 100.0% 73.3% 80.0%
39 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: z UNAV units work best with R/C radio systems that offer advanced features like "Fail-Safe" but most RC radio systems are compatible. z UNAV can use buffers to reduce RF noise from ignition systems, onboard transmitters and long servo wires. z Connection is simple z 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
40
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 Wes-technik Wes-technik Wes-technik Model LS-2.0 LS-3.0 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
41
Table 5.13: Pugh Chart for Servo Selection
k k k k k k i i i i i i n n n n n n
h h h h h h 4 0 0 4 0 0 c c c c c c ...... e e e e e e 2 3 2 2 3 2 T T T - - - T T T - - Factors/Candidates ------S S S S S S s s s s s s L L L L L L e e e e e e W W W W W W
Max Deflection 3 3 3 Time to Full Deflection 3 3 2 Max Output Force 3 5 4 Operating Voltage 3 3 3 Dimensions 3 3 3 Load Current 3 3 3 Weight 5 2 3
Mean Score 3.3 3.1 3.0 Normalized Score 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:
42
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
43
Table 5.15: Pugh Chart for Antenna Array Selection
P P P P P P G D G D
9 6 4 9 6 4 4 4 4 ) 4 ) t t 0 1 1 r 0 1 1 r 2 1 2 1 s s a a 4 4 4 4 4 4 4 4 4 4 a a
Factors/Candidates e e 2 2 2 2 2 2 2 L 2 2 L 2 ( ( Y Y G G G G G G G G G G
H H H H H H H H H H
Horizontal Beamwidth 4 3 3 2 5 Vertical Beamwidth 4 3 3 2 5 Gain (Directivity) 3 3 4 5 1
Mean Score 3.7 3.0 3.3 3.0 3.7 Normalized Score 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 Kokam Kokam Kokam iRate Model SLPB433452 SLB452128 SLPB523459 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
44
Figure 5.20: Batteries – Kokam
Table 5.17: Pugh Chart for Battery Selection
9 2 9 2 0 0 8 8 5 5 5 5 0 0 2 2
4 4 4 4 5 5 1 1 m m m m m m 3 3 P 3 3 P 2 2 a a a a a a L 2 3 L 2 3 5 5 k k k k k k
5 4 5 Factors/Candidates 4 o 4 o o o 4 o o e e B B B B t t K K K B K K K B P P P P a a L L L L L L R R S S i i S S S S
Capacity 4 2 5 3 Voltage 3 3 3 1 Volume 3 5 1 1 Weight 2 4 1 3
Mean Score 3.0 3.5 2.5 2.0 Normalized Score 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.
45 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 Weight Ability 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%
46
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.
47 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.
dCl dC L = dα dC dα 180 l (5.1) dα 1+ π 2eAR
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
48
2 CL CD = CD0 + (5.2) πeAR
CD = Coefficient of Drag
CD0 = Parasite Drag
CL = Coefficient of Lift
C T = W D C (5.3) L
T = Thrust (g) W = Weight [Mass in this Case] (g)
Table 5.19: Thrust Calculation All Estimated Values Variables 2004 2005
dC l /d α 0.1 0.1 AR 1.422 1.422 Assume same shape for both years so AR remains the same S (m 2 ) 0.048 0.102 For 2004, assume 90% of 8" x 26 cm rectangular wing is remaining b (m) 0.26 0.381 For 2005, assume 15" wing span e 0.9 0.9 Assume 0.9
dC L /d α 0.041 0.041 AOA (deg) 99
C 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 CDR
C D 0.243 0.252
C D0 0.018 0.027 2005 is about 1.5x the size of 2004 so increase CD0 by that much T (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
49 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 Electronics Variables 2004 2005
dC l /d α 0.1 0.1 AR 1.422 1.422 S (m 2 ) 0.048 0.102 b (m) 0.26 0.381 e 0.9 0.9
dC L /d α 0.041 0.041 AOA (deg) 99
C L 0.95 0.95 W (g) 97.9 198.48
C D 0.243 0.252
C D0 0.018 0.027 T (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).
101.97P T = v (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.
50 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 Thrust Cost Weight Current Voltage Additional 1 Additional 2 Row Total Column Total 1+Row + Column Total* Relative Weight Thrust RRRR 40533% Cost CRR 20320% Weight RR 21427% Current R 10213% Voltage 0017% Additional 1 0000% Additional 2 0000%
Column Total 0010000 15100%
*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).
51 T < 80 : 0 T = Thrust I > 1.5 : 1 I = Current 80 ≤ T < 105 : 1 0.8 ≤ I ≤ 1: 2 105 ≤ T ≤ 130 : 2 0.8 > I : 3 130 < T : 3
C > 80 : 1 C = Cost V > 9 : 1 V = Voltage 20 ≤ C ≤ 80 : 2 7 ≤ V ≤ 9 : 2 20 > C : 3 7 > V : 3
W > 30 : 0 W = Weight 20 ≤ W < 30 : 1 15 ≤ W ≤ 20 : 2 15 > W : 3
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 Propeller baseline 2 = same as baseline 3 = better than baseline Thrust Cost Weight Current Voltage 1 Additional 2 Additional Weighted Score Score Normalized 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) EP70350333200 1.93 91% B2C Light Power System (GW/LPS-B2C-C) EP70600332300 1.87 88% B2C Light Power System (GW/LPS-B2C-C) EP70600332200 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%
52 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
53
Figure 5.28: Feigao 1208436L 12x30mm Brushless Motor
Figure 5.29: EP3020 Propeller
54 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
55 • 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.
56 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
57 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: b 2 AR = S = c *b S 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
58 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.
59 Step 1: Blocking out foam.
Transformer
Foam Bow
Figure 7.1: Blocking out foam
Step 2: Attach airfoil bed.
Ledge for Foam bow wire.
Pushpin.
Airfoil bed.
Figure 7.2: Attach airfoil template to foam
Step 3: Weight down foam and place bow in position.
Fiberglass rod.
Aluminum tube.
Wheel.
Figure 7.3: Setting up foam and bow for hot wiring
60 Step 4: Attach electrical leads to bow, route through the pulley, attach to counter weight and turn power on.
Counter weight.
Figure 7.4: Counter weight used to create wire movement
Step 5: Pin airfoil template in the correct position and repeat step 4 except cut over top the airfoil.
Airfoil template pinned in place over top the bed.
Figure 7.5: Utilize airfoil templates to cut out rest of 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.
Figure 7.6: Airfoil after hot wiring process
61 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.
62 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
63 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.
Step 1: Obtaining supplies.
Figure 7.7: Layout of supplies
Step 2: Sizing up the job.
Figure 7.8: Sizing fiberglass
64 Step 3: Laying out Tools and Supplies.
Figure 7.9: Laying out fiberglass
Step 4: Saturating the Fiberglass.
Fiberglass will become clear as saturated.
Figure 7.10: Saturating fiberglass with epoxy
Step 5: Laminating the airfoil.
Fiberglass must be smooth over airfoils, so the plastic squeegee is used to remove air bubbles.
Figure 7.11: Laminating airfoil with fiberglass
65 Step 6: Curing.
Weight is applied to the top of the foam mold with fiberglassed airfoil inside for 24 hours to cure
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.
66 Airfoil
Aluminum Mounting Rod
Figure 7.13: Airfoil attached to mounting rod
Airfoil Mounted in Inverted Position
Aluminum Mounting Rod
Testing Balance
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).
67 Table 7.1: Data Recording Table for Wind Tunnel Testing of Airfoils
Re: 60,000 Moment Vwind AoA Lift Drag (grams- (ft/sec) Re (deg) (grams) (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 Moment Vwind AoA Lift Drag (grams- (ft/sec) Re (deg) (grams) (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
68 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.
69
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.
70
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:
z On-screen video image with GPS data z Servo displacement based on control inputs z Motor acceleration and deceleration z Range test z 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.
71
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.
Multimeter (// to Oscilloscope Oscilloscope) INA114 Instrumentation Amplifier Burr Tetronix Serial Craftsman Serial DC 6V EE Power Brown w/ 100ohm Resistor #CCL02079157 Supply Load Cell #B012473 5V +/- 12V SMD Sensor S250 1kg Serial #EE1015 25Kohm 25Kohm 25Kohm DC 12V EE Power 100ohms Vc Supply 5V +/- 12V 25Kohm Serial #EE1015
Ref 25Kohm 25Kohm
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
72 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
1.8
1.6
1.4
1.2 y = 0.0056x + 0.9795 1 R2 = 1 0.8 Voltage (V) 0.6
0.4
0.2
0 0 20406080100120 Load (grams)
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
73 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.
74 MPJA Multimeter Serial # CCL010412272
R1 R L
Shenzhen Mastech DC Motor T + Power Supply: - HY3003-3
Note: R1 is dependent on the internal resistance (R) of the motor
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
75 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)
76 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.
77 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.
78 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.
79 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.
Propeller EP7060 Drag (g) Thrust (g) RPM Current Draw (A) Thrust/Weight Ratio Thrust/Drag Ratio Voltage (V) Wind Speeds (m/s) AOA (deg) T1 T2 T3 T1 T2 T3 T1 T2 T3 T1 T2 T3 T1 T2 T3 T1 T2 T3 0 8 5 10 4 0 9 5 10 ::: 41510 0 8 5 10 5 0 9 5 10 ::: 51510 ::: 0 9 8 5 10 : :: 9 15 10
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
80 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.
81 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.
82
[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.
83