Download 43rd AIAA Joint Propulsion Conference Student Design Challenge

43rd AIAA Joint Propulsion Conference Student Design
Challenge
Final Report
Team Members
William Bennett
Jayme Carper
Nicholas Hankinson
Michael Sheridan
Keith Vehorn
Stephen Warrener
Faculty Advisor
Dr. Scott Thomas
TABLE OF CONTENTS
Executive Summary
. . . . . . . . . . . . . . . . . . . . . . . . . . 3
Management Summary . . . . . . . . . . . . . . . . . . . . . . . . . 4
Conceptual Design . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
Airframe
. . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
Propulsion
. . . . . . . . . . . . . . . . . . . . . . . . . . .
5
Power Generation . . . . . . . . . . . . . . . . . . . . . . . .
7
Alternator Drive
8
. . . . . . . . . . . . . . . . . . . . . . . . .
Power Dissipation . . . . . . . . . . . . . . . . . . . . . . . .
9
Power Measurement . . . . . . . . . . . . . . . . . . . . . . .
10
Surveillance
. . . . . . . . . . . . . . . . . . . . . . . . . .
11
Preliminary Design
. . . . . . . . . . . . . . . . . . . . . . . . . .
13
Airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . .
13
Power Generation . . . . . . . . . . . . . . . . . . . . . . . .
14
Alternator Drive . . . . . . . . . . . . . . . . . . . . . . . . .
14
Power Dissipation . . . . . . . . . . . . . . . . . . . . . . . .
15
Surveillance
18
Detail Design
. . . . . . . . . . . . . . . . . . . . . . . . . .
. . . .
Airframe
. . . . . . . . . . . . . . . . . . . . . . . .
21
. . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
Propulsion
. . . . . . . . . . . . . . . . . . . . . . . . . . .
22
Power Generation . . . . . . . . . . . . . . . . . . . . . . . .
22
Alternator Drive
24
Surveillance
. . . . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . . . . . . . . . .
24
Drawing Package . . . . . . . . . . . . . . . . . . . . . . . .
25
Aircraft Geometry, Performance and Weight/Balance Parameters . . . . .
26
Aircraft System Components
. . . . . . . . . . . . . . . . . . .
26
Testing Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
27
Power Generation . . . . . . . . . . . . . . . . . . . . . . . .
27
Static Thrust . . . . . . . . . . . . . . . . . . . . . . . . . .
29
Surveillance . . . . . . . . . . . . . . . . . . . . . . . . . .
30
Telemetry . . . . . . . . . . . . . . . . . . . . . . . . . . .
30
Airframe
. . . . . . . . . . . . . . . . . . . . . . . . . . .
30
Take-Off Performance . . . . . . . . . . . . . . . . . . .
30
Rate of Climb
. . . . . . . . . . . . . . . . . . . . . .
30
Stall Speed . . . . . . . . . . . . . . . . . . . . . . .
30
Maximum Airspeed . . . . . . . . . . . . . . . . . . . .
31
System Tests Performed. . . . . . . . . . . . . . . . . . . . . .
31
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EXECUTIVE SUMMARY
The objective of the 43rd AIAA Joint Propulsion Conference Student Design Challenge is for
university teams to build and fly a radio-controlled aircraft that is able to generate electrical power to be
delivered to a power-consuming device while performing video surveillance. The video and the total
electrical power consumed over 10 minutes will be downloaded to a base station.
The design of the Wright State University airplane started with a model airplane of the Cessna
337, which has a push/pull, twin tail boom configuration. The Cessna model could be easily modified to
allow for a camera in the nose of the aircraft and a pusher propeller mounted to the rear of the fuselage.
This configuration was deemed superior over the Sig Cadet aircraft due to the forward-looking nature of
the surveillance camera as well as eliminating the potential for fouling the camera lens with engine
exhaust.
Electrical power generation choices mainly consisted of a large battery pack, an off-the-shelf
generator system from Sullivan, or a system designed by the WSU team. It was decided to attempt to
design the power generation system ourselves based on weight and cost factors, as well as maximizing
the overall benefit to the team members in terms of learning new things. Starting with an engine driving
an electrical motor which serves as an alternator, it was determined that a rectifier circuit and a voltage
regulator would be needed. The next phase was to decide on the physical arrangement of the
motor/alternator. Originally, the motor was to be connected to the alternator via a long shaft so that the
alternator could be located in the rear of the airplane to mitigate CG issues. While feasible, this system
was thought to be not as reliable as one in which the alternator was driven directly by a pulley system.
This method definitely drives the center of gravity forward due to the motor and alternator being on the
front of the fuselage, which is a negative aspect of the design, but it is by far simpler than a shaft-driven
system.
With the camera located in the rear of the fuselage, the mounting system needed to be revisited.
Initially, a static mount was developed for simplicity. The team then devised a dynamic mounting system
driven by a receiver and two servos that were separate from those of the system for controlling the
airplane. In other words, one person would fly the airplane, and another person would “fly” the camera
with a separate transmitter. This dynamic camera system allows for the camera operator to quickly and
easily locate objects on the ground over a very wide range of operating.
The power consuming device could take on many forms, such as a power resistor, nichrome wire,
or a lamp. Due to considerations such as complexity, weight, reliability and safety, a nichrome wire was
selected, which is suspended between the twin tail booms. This method provides for excellent heat
dissipation characteristics, and also allows for the resistance of the power-consuming device to be
tailored to the requirements of the power generation system.
In summary, the Wright State University entry into the competition consists of a Cessna twinboom aircraft model fitted with a glow-engine/alternator combination mounted to the front firewall. Within
the fuselage lie the rectifier/voltage regulation system, the two receivers for aircraft control and camera
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control, and the data acquisition system used to monitor electrical power dissipation. The dynamic
camera system is mounted on the rear of the fuselage, which is capable of observing objects in front of,
behind and to the sides of the airplane in an active manner. Finally, the power dissipation device is
located between the tail booms, and consists of nichrome wires held in place by ceramic insulators.
MANAGEMENT SUMMARY
The Wright State University AIAA student design team consists of seven members: Dr. Scott
Thomas, William Bennett, Stephen Warrener, Keith Vehorn, Jayme Carper, Michael Sheridan, and
Nicholas Hankinson. Each team member has contributed to the project depending on their skills.
Separating the team into their qualified field provided the most efficient and effective results. The team
members were assigned to the areas as follows:
Airframe
Construction
Power
Propulsion
Generation
Surveillance
3D Drawings
William Bennett
Stephen Warrener
Jayme Carper
Keith Vehorn
Nick Hankinson
Michael Sheridan
Figure 1: Team Member Responsibilities
The team created a time chart plan of when certain areas and objectives needed to be started
and finished. Figure 2 shows the proposed and the actual timing of when design elements occurred.
Figure 2: Anticipated and Actual Timelines for Design Phases to be Accomplished
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CONCEPTUAL DESIGN
The conceptual design phase began immediately after the team was organized and continued
until the written proposal was submitted. This phase consisted of extensive brainstorming sessions to
begin the design process. It was during this phase that each member was assigned as area of the project
based on their interest and skill set.
Airframe
The contest specifications dictated that the airframe must be of an 80” span, 15” chord and a total
flying weight of not more that 15 lbs. These criteria were established with several popular high wing
trainers in mind. The high stability, light wing loading and simplicity of these airframes make them well
suited for the mission. Although practical in some ways, the conventional designs are inherently high
drag, low speed and offer limited camera mounting opportunities. In order to solve some of these issues,
alternative and more innovative airframe possibilities were investigated.
The Nitro Models Cessna 337 is a large scale almost ready to fly replica of the Cessna push-pull
light aircraft. With a wingspan of exactly 80”, and a similar target weight this plane caught the group’s
attention as a viable solution to the camera visibility and performance issues. With the unique engine
configuration, many new optical mounting positions became possible. The intent was to remove the front
engine and tractor propeller and replace this with all of the optics, giving the camera system an
unobstructed forward view. The large fuselage is also designed with a higher power to weight ratio in
mind. By replacing the forward engine with the camera and powering the aircraft with one large engine in
the rear the airframe could facilitate the ideal camera mount, support the large engine, and still have
plenty of room for all onboard systems. The twin boom structure also facilitates the mounting of sensitive
heating elements in the free stream air for cooling and balance purposes. Because the stock wing has a
minor sweep and scale wingtips, a custom wing needed to be fabricated to meet contest specifications.
With a few minor alterations, the Cessna 337 was chosen as the base platform for the project aircraft.
Propulsion
For the propulsion system that will be used on this airplane both electric motors and internal
combustion glow plug engines were considered. The parameters considered in making the comparison
were: total system weight, available power, and the ratio of power to the total system weight.
The first step in this process was to determine basic aircraft propulsion requirements. Using
experience gained from the SAE Aero Design Competition, the team was able to determine that this
airplane would need to be able to generate approximately 5 pounds of thrust to meet the 200-foot takeoff
limit if the plane weighs the competition limit of 15 pounds. Given this information it was then possible,
using a propeller performance program, to calculate the amount of power necessary to produce the
required thrust. Over a large range of propeller options it was found that the power plant will need to
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supply a minimum of 1.4 horsepower in order to meet the takeoff requirement. This enabled the team to
narrow down the potential power plants being considered for this airplane.
The power plants considered for comparison are shown in Table 1. Glow plug engines from OS
Engines and electric motors from AXI were both examined.
Table 1: Power plants Considered for Comparison
Engines
Electric Motors
OS .46 AX
AXI 5330/24
OS .50 SX
AXI 5320/28
OS .61 FX
AXI 4130-20
OS .91 FX
OS .65 LA
OS 1.20 AX
OS 1.4 RX-P
Data was collected for the above power plant options. In addition to the weight of the given motor
or engine, the corresponding hardware and materials required for operation were added into the system
weight. For the electric motors this included, a speed controller, a sufficiently sized battery, and a motor
mount. For the glow plug engines the additional items included the following: fuel tank, twelve ounces of
fuel, muffler, servo to actuate throttle, linkage, and an engine mount. This information was compiled and
the above-mentioned parameters of system weight, power output and power to weight ratio were
computed. The power plant combinations with the greatest power to weight ratios are displayed in Figure
3.
4.000
Weight (lbs)
Power (hp)
Power/Weight(hp/lbs)
3.500
3.000
2.500
2.000
1.500
1.000
0.500
0.000
OS .50 SX
OS .91 FX
OS 1.20 AX
OS 1.4 RX-P
AXI 5320/28
AXI 4130-20
Figure 3: Comparison of Selected Power Plant Options
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The OS 1.20 AX was selected for the design. This engine gives the highest power to weight ratio.
In addition to having the highest power to weight ratio it also supplies a significant amount of surplus
power. This allows for power to be extracted from the engine to drive an alternator used for the power
generation system.
Power Generation
When the question of how to generate power was posed to the group many different ideas
surfaced. These included: turbine driven alternators, an engine driven alternator, solar cells, hydrogen
fuel cells, batteries and thermoelectric devices. Each of these systems were analyzed and conclusions
were drawn as to the best systems for use on this project. Three primary figures of merit were used to
evaluate the systems: total output power, total system weight and the ratio of power to weight.
Turbine driven alternators are becoming more common on modern aircraft due to the increased
need for electrical power. Commercially available systems were found that would produce over 500 Watts
of power at a cruising speed of 50 knots. This appears promising until the weight of the system is taken
into consideration. The system investigated weighs approximately 24 pounds. Given the competition
specification of restricting the maximum aircraft weight to 15 pounds this system was not a viable option.
The idea of designing a custom air driven alternator was considered but when compared to the power
output of the other devices considered it was decided to not pursue this option.
Traditional solar panels would not be practical on this aircraft due to their high weight and
relatively low efficiency. However, a new generation of thin amorphous cells was considered. These cells
come in a very thin strip and could be affixed to the skin of the airplane. The disadvantage of these
devices is the total possible output power. Given the total wing area of the aircraft the highest possible
output power would be approximately 14 Watts. Due to the complexity of adding so many of these
devices to the aircraft and the low output power they were not considered for use in this design.
Small fuel cells represent a new exciting technology in the realm of portable power. They
represent a new option for powering small electronic devices such as cell phones and notebook
computers. The Georgia Institute of Technology recently developed an aircraft powered solely by
hydrogen fuel cells. This was a very sophisticated system capable of delivering 500 Watts of power.
Commercially available systems capable of delivering that amount of power are still very large. A system
investigated yielded 25 Watts of power and weighed 2.7 pounds. Due to their low power to weight ratio
fuel cells were not considered for this project.
Batteries are a viable option for powering both an electrical motor for propulsion and/or for use in
discharging a significant amount of power to the electrical dissipation device. Lightweight lithium-polymer
batteries can be linked both in series and in parallel to achieve an output voltage of 29.6 Volts, which is
very close to the 28 Volt specification required by the contest. Based on previous results from a battery
discharge experiment performed for the 2006 SAE AeroDesign Competition, it was estimated that a pack
of 8 LiPo’s could provide up to 500 W continuously for approximately 10 minutes, with an overall weight of
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1.6 lbs. Using batteries would still require a high power voltage regulator to attain the 28 Volt output
specified. While this option is a strong candidate power system, it was stated in the FAQ’s that the judges
would favor an actual power generation system over a power storage system. Therefore, we opted to not
pursue this avenue.
Thermoelectric devices, also known as Peltier devices, create a temperature differential between
its two surfaces when electricity is supplied to it. The primary application of these devices is in cooling of
components where traditional refrigeration methods are not desirable. These devices can also be used to
generate electricity when a temperature gradient is present between its surfaces. On this project they
were being considered to scavenge power from the electrical systems associated with the regulation of
the power source. The devices are very light weight but are very inefficient when used in this manner. For
the scale of this application the Peltier device operates at approximately 10% efficiency. To produce a
sizable amount of power, the heat being generated by the electric devices would be very large. Due to
this inefficiency the Peltier devices were not selected to be used on this aircraft.
An alternator driven by the glow engine represents the most viable option for use on this aircraft.
By selecting an engine that is larger than what would typically be required to propel an aircraft this size,
the option existed to extract power from that engine to generate electrical power. Two options were
considered for this system. Sullivan Products manufactures an alternator system that mounts directly to
the output shaft of the glow engine. The other option considered was to turn a brushless motor via a belt
drive system that would connect to the shaft of the engine as well as the shaft of the motor. Both
represent viable options for generating substantial power for a given system weight. The Sullivan
alternator has a maximum power output of 500 Watts, whereas the belt driven motor has the potential of
producing up to 700 Watts. While the weight of the belt driven motor is higher, the power to weight ratio is
higher than that of the Sullivan alternator. In addition the competition officials remarked that they would
look more favorably on a student built system as opposed to a commercially available system. A
comparison of output power, system weight, and the power to weight ratio can be seen in Table 2.
Table 2: Comparison of Power, Weight and Power vs. Weight Ratio for Select Power Systems
POWER (W)
WEIGHT (LB)
POWER VS. WEIGHT
(W/LB)
LiPo Batteries
700
3.8
184.21
Solar Cells
14
0.5
28
Air Driven Alternator
590
24
24.58
Thermoelectric Device
30
.046
652.17
Fuel Cells
25
2.7
9.26
Sullivan Alternator
500
2.75
181.82
Custom Alternator
700
3.2
218.75
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Finally the cost became a sizable factor in comparing both a custom alternator and an off the
shelf system. The alternator system from Sullivan was quoted at $3800. This is more than half of the
budget for the entire project. That cost includes a high capacity voltage regulator which separately was
quoted at $1800. The custom built system using a brushless motor from AXI would cost approximately
$400 in addition to the cost of the voltage regulator. This is a significant reduction in cost. Given this fact
and that the custom designed system produces more power, the belt driven alternator system will be
chosen for this aircraft.
Alternator Drive
To generate power, the Wright State University team’s conceptual design was to use two timing
pulleys with a belt set up as direct drive, with both the engine and AXI motor mounted in the rear of the
fuselage. A gear driven and chain driven system were also considered but due to alignment complexity
both of these options were abandoned. With the engine and the motor on the rear of the fuselage, this
would give the camera a clear view directly forward and beneath the plane with no oil residue from the
engine exhaust harming the camera optics. A larger pulley would be connected to the shaft of the engine
while a smaller pulley was connected to the AXI motor shaft, linked with a timing belt. The ratio between
the two pulleys was 2:1 allowing the AXI motor to spin faster than the engine, generating higher voltage at
a lower engine speed.
Power Dissipation
To dissipate the power created by the electric motor, power resistors, nichrome wire, and high
power aircraft lamps were considered. The parameters considered for comparing the materials were:
weight, current capacity and safety.
The first step was to determine the power generation performance and weight budget. By testing
the engine and electric motor used to generate the power, the WSU team discovered that the electric
motor would be able to produce around 700 Watts. This means that the current produced will be around
25 Amperes at 28 Volts. After the airframe was constructed and all necessary components (servos,
engines, receiver, etc.) were installed the aircraft was weighed and it was found that the total weight was
just over 10 pounds. The team then came to the conclusion that the entire power generation system
(electric motor, rectifier, wire, etc.) could not weigh more than 3 pounds. With the electric motor already
taking a third of that budget, the power dissipation device would need to weigh less than a pound, to
accommodate the regulator.
Power resistors were considered because they are compact, designed to handle high power, and
dissipate heat. However, because the amount of current and wattage that the power resistor will need to
handle is so high the resistor will be very large and heavy. Power resistors are also very limited on their
current capacity and cannot be easily adjusted to meet changing resistance needs. A power resistor is
9
much safer than high power lamps and nichrome wire because it is a contained system with no exposed
wire.
High-powered aircraft lamps are made to handle 28 volts at high wattages to make them very
bright. However, these lamps are very heavy and bulky. They are also very limited on their range of
current capacity, and are not easily adjusted to meet the system’s exact needs. They are a safe
alternative in the fact they will reduce the chance for electrical shock, but they are not safe in the fact that
they will also become so bright that it could cause damage to a person’s eyes.
Nichrome wire was considered because it has a very high resistance per unit length and can be
easily adapted to fit any resistance needs. With the amount of power and current the system will be
producing the wire will need to be of a heavy gauge and some kind of system will need to be developed
to insulate the wire from the aircraft. However, two smaller gauge wires in parallel can be used to reduce
the temperature of the wire. This is a big concern because at 25 Amperes 16-gauge wire will reach
temperatures of 1800°F, but two 18-gauge wires in parallel will reduce the wire temperature to around
1200°F degrees. Even though two 18-gauges wires and insulators will be heavier than a single 16-gauge
wire they will not be nearly has heavy as power resistors or high power aircraft lamps. Nichrome wire
does not offer the same level of safety that power resistors, but with the proper switching and fusing the
system can be disarmed when not in flight to ensure safety from electrical shock.
After all the alternatives were compared qualitatively using the decision matrix shown in Table 3,
nichrome wire was chosen. This was chosen because of its ability to be easily adjusted and its low total
weight.
Table 3: Resistor Selection Decision Matrix
Power Dissipation Material
Current Capacity
Weight
Safety
Total
Power Resistors
2
2
3
7
Nichrome Wire
5
3
3
11
High Power Lamps
3
1
4
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Power Measurement
The competition requirements state that a system must be in place that will allow for the live
monitoring of the power generated by the power generation system mentioned above. Two methods were
explored for measuring the power output. The first method would involve measuring the output voltage of
the system as well as the voltage drop across a precision resistor. The voltage drop measured across the
resistor would be used to calculate the current flow through the system. These values would then be used
to calculate the power. The second method involves measuring the output voltage in the same manner
but using a non-contact Hall effect current sensor. This device measures the magnetic field generated by
the current flow through a wire. This value is then translated as the output current of the alternator. An
uncertainty analysis was performed given the manufacturer’s specified accuracy of the measuring
devices. For the precision resistor technique P  V p  V g / R p where Vp is the voltage drop across the
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precision resistor, Vg is the generator output voltage and Rp is the resistance value of the precision
resistor in ohms. ΔP, the RMS uncertainty in the power measurement can be found by the following
equation:

 Vg
  Vp
  V p  Vg

P  
 V p   
 V g   


R
p
R
 R
  R 2

 p
  p
  p

2
2
2
The same analysis was conducted using the voltage and direct current measurement technique.
In this case P  V g  I , where I is the output current from the alternator. In this technique ΔP can be
obtained from the following equation:
P 
I  V   V
2
g
 I 
2
g
The analysis showed that for both methods, the uncertainty is less than 1% for a 700 Watt output.
The Eagletree Systems telemetry unit selected for this competition has the ability to display live power
using the non-contact current measuring device. However, this device does not have the ability to record
two voltage measurements and perform the necessary math operations to display power in real time. For
this reason the direct current and voltage measuring technique was chosen over using a precision
resistor.
Surveillance
The model LWA13 camera system was considered for the camera system in the airplane. This
system lacked in picture quality that was necessary for this particular application. The advertised range
of a quarter mile proved to be more than sufficient for our purposes; however, other aspects of the
camera system, such as durability and reliability, caused the team to choose a different camera system.
The decision was made to use the BlackWidowAV.com model KX-141 camera. This camera has
480 lines of resolution and conveniently runs on 5 volts. The resolution of this camera will give the team
an adequate view of the flying field and sufficient video quality to identify the targets.
The team originally decided to design a fixed mount camera on the plane; however, it was soon
apparent that this would not be the optimal configuration. It would be necessary to have a dynamic
camera mount so that the camera controller could sweep for the target, and not force the pilot to
maneuver the aircraft directly over the target. This can be seen when the forward view, time on target,
and side sweep views are compared for a stationary mount and a dynamic mount for the specified
altitude and airspeed. A comparison of the stationary and the dynamic mount for a 64° lens angle can be
seen in Figures 4 and 5.
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Figure 4: Side View Comparison of Dynamic and Static Camera Mounts (Not to Scale)
Figure 5: Top View Comparison of Dynamic and Static Camera Mounts (Not to Scale)
The addition of the dynamic camera mount system gives an additional 36° view both forward and
backward (total 72°), and gives an additional 40° throw for each side view. The chart above shows the
necessity of a dynamic mount. The maximum time on target for the projected airspeed for a stationary
mount is 2.5 seconds; whereas with the dynamic mount, the time on target increases by as much as 10
times. In fact, the time on target increases to the maximum for the 64° lens, 14 seconds, for our airspeed.
This means that for the entire time that it takes for the aircraft to travel from one pylon to the other, the
target can be in the field of view.
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The dynamic mount would require the addition of another radio receiver, wiring, and two servos.
One servo would control each axis of rotation, with the pivot mounted directly onto the end of the servo.
While this increased the camera system weight, it was felt this was justified in terms of greatly improved
performance.
PRELIMINARY DESIGN
The preliminary design phase was initiated based on the assumption that the team’s proposal
would be funded. An aircraft kit was purchased and constructed, and research was conducted to gain
sufficient knowledge in order to make decisions concerning the critical sub-systems on the plane.
Airframe
The first test model was built as a proof of concept in many ways. There were many unknowns
about what the best methods would be for areas such as electrical circuits, control surface linkages,
balance, optics and power. The first task was to build a completely new wing to meet the constant chord
requirements. A Clark-Y airfoil and a dual spar D tube structure were incorporated into a new one piece
wing, making it both light and strong. This wing preserved the 80” span while increasing the wing area
with the 15” constant chord geometry. The modular design of the original was also dropped in favor of a
more structurally sound one piece design. During construction it became apparent that due to severe
balance issues, a forward engine mount would be necessary. In order to mount the optics in the front, a
considerable the amount of nose weight would be required and depending on the moment, this could be
as much as 15% of the total aircraft weight. To alleviate this issue the group decided to run a forward
mounted engine and move the optical systems to the back. This position still provided great visibility, and
cured the balance issues immediately. The aircraft is guided by two separate transmitters: one for flight
controls and the other for the dynamic camera platform. The pilot only focuses on flying the aircraft, while
a separate operator controls the camera. A 5 cell 1000mah NiCad battery powers the flight controls while
all other systems operate from 600mah 4 cell NiCad auxiliary batteries. By separating the power
systems, a greater degree of redundancy and simplicity is achieved. Each of the three power circuits has
a separate switch and charge jack mounted on the side of the fuselage.
Onboard electronics were selected with preference toward light weight. The Futaba 3102 servo
is a mini sized servo with metal gears and high torque. Weighing only 0.7 oz each, these units cut the
total flight control weight by approximately half. For redundancy and simplicity it was elected to use 3102
servos for every function on the entire airframe so that all units are interchangeable and only one type of
spare part would be required. The metal gears made this product ideal for every function from flight
controls to steering. A 5 cell flight pack was selected because the higher voltage would increase servo
torque and the 1000 mAh capacity would meet contest specs but provide light weight power.
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Power Generation
When the decision to design a custom built alternator system was made the next step was to
determine what type of device could be driven off the engine to deliver electrical power. Research
showed that individuals have used brushless motors as alternators on remote controlled aircraft to power
onboard electronics and cameras.
The primary consideration in selecting the size of the motor to be used as the alternator was the
output voltage. Electric aircraft motors are given a rating in KV, which is measured in RPM/Volt and is a
measure of how many RPM a motor will produce for every volt it is supplied. When the motor is being
used to generate power the inverse of the KV value yields the number of volts produced for every RPM
the motor is spun at. The alternator output voltage can be expressed as the following equation:
Vout 

KV
where ω is the engine speed in RPM. For example a motor with a KV value of 700 would need to spin at
19,600 RPM to produce 28 Volts. The OS 1.20AX produces maximum power when operating at 9,000
RPM. For this reason the decision was made to attempt to match that RPM range as closely as possible.
Given an operating speed of 9,000 RPM and an output voltage greater than 28 volts the
necessary KV value was determined to be 320. This led to the selection of the AXI 4130 brushless
electric motor to serve as the alternator on this aircraft. The AXI 4130 has a KV value of 305 which
closely matches the above mentioned value. After further investigation into the additional electrical
equipment that would be required to rectify the 3-phase alternating current produced by the motor into
direct current as well as regulate the output to the required 28 Volts, it was determined that an alternator
output voltage of 32 Volts was desired. Given the operating speed of the engine and the requirement of a
32 Volt output, a gear ratio of 1.5:1 was selected for the drive system connecting the engine to the
alternator.
The AXI 4130 is a large motor capable of powering 8 to 10 pound aircraft and it can draw up to
1100 Watts of power. The team chose a target of 700 Watts of power to be produced by the alternator.
Assuming a system efficiency of 80% the engine would be required to give up to 875 Watts of power to
drive the system. This alternator would pull approximately 35% of the OS 1.20AX engine power. This will
still leave enough power to propel the aircraft at a reasonable speed through the flight course.
Alternator Drive
For the preliminary design, the team decided to move the AXI motor towards the front of the
plane for balance purposes, and use a long shaft and multiple pulley setup to drive the motor. For this
design, the team moved the AXI motor 33 inches forward from the original location, and used a long shaft
that would be driven by a pulley off of the engine. This configuration can be seen in Figure 6.
14
Figure 6: Preliminary Alternator Drive System
In order to reduce vibrations, several bearing blocks were placed evenly along the shaft. In case
of any axial movement, thrust bearings were placed at each end of the shaft where they would be against
the bearing blocks, so no movement forward or backward would occur. This design helped the center of
gravity problem to some degree, but a significant amount of weight was still needed in the front. Also, this
design needed several extra mechanical components that could fail and cause problems as the airplane
was in flight.
Power Dissipation
The objective of this analysis is to determine the temperature of the nichrome resistance wire as
a function of the airspeed of the airplane. This is necessary in order to ensure that the maximum
operating temperature of the nichrome wire (2100 F) is not exceeded. The nichrome heater wire is
located behind the airfoil and is stretched between the tail booms. It is held in place with ring-shaped
ceramic electrical insulators that are capable of withstanding high temperatures. The aspect ratio of the
wire (length to diameter) is approximately (18 inches/0.0508 inches) = 354. With this in mind, it was
decided to model the wire as an infinite heated cylinder in cross flow. This analysis breaks down at the
point where the wire passes through the ceramic insulator. For the case where the airspeed is zero it was
observed that the section of the wire in contact with the insulator was at a lower temperature. This is due
to the fact that the conduction through the insulator has a greater affect on heat transfer than the natural
convection into the air. Once airspeed was increased the forced convection began to supersede the
conduction and the highest temperature regions were those on the back side of the insulators where
there was the least airflow. These hot spots will be addressed with small fins to direct more air to the rear
of the insulator. This analysis is also compromised by the fact that the air flow directly behind the airfoil is
not uniform, but is likely to be non-uniform and highly turbulent. However, since turbulence usually
increases heat transfer, it is believed that the analytical model will be conservative in terms of overpredicting the wire temperature, which is considered to be beneficial with respect to safety.
15
Incropera and DeWitt (1990) suggest that the correlation proposed by Zhukauskas (1972) is
appropriate for the situation outlined above:
1/ 4
 Pr 

Nu D  C Re Pr 
 Prs 
0.7  Pr  500
m
D
n
1  Re D  106
where all properties are evaluated at the free stream temperature except Prs, which is evaluated at the
surface temperature of the wire. The values of the constants are given below as a function of the
Reynolds number. For air, with a Prandtl number of approximately Pr = 0.7, the exponent is n = 0.37.
ReD
C
m
1 to 40
0.75
0.4
40 to 1000
0.51
0.5
0.26
0.6
0.076
0.7
1000 to 210
5
2  105 to 106
The properties of air were taken from tables provided by Incropera and DeWitt (1990). A regression
analysis of the data over a temperature range of 250 to 350 K resulted in the following equations:
Pr  2.000  10 4 T  7.690  10 1
( R 2  0.9709)
  1.160  10 10 T 2  2.520  10 8 T  2.110  10 6
k  7.700  10 5 T  3.100  10 3
( R 2  1.0000)
( R 2  0.99949)
where the units of temperature are degrees Kelvin, the kinematic viscosity is in m 2/s, and the thermal
conductivity is in W/(m-K). For the surface Prandtl number, a much wider temperature range was used to
fit the data, since the surface temperature of the wire could go up to 2100 F. The regression equation for
the surface Prandtl number is
Prs  2.04738 10 19 T 6  1.25437 10 15 T 5  2.71086 10 12 T 4  2.2451110 9 T 3
 1.99968 10 7 T 2  4.8828110  4 T  0.830431
( R 2  0.970822)
By neglecting the heat conducted through the ends of the wire where it passes through the ceramic rings,
the electrical power input to the wire is dissipated by convection to the air as well as by thermal radiation.
Qelectrical  Qconvection  Qradiation
The heat convected away by the air is
Qconvection  hAs Ts  T 
16
where
h is the convective heat transfer coefficient, As  DL is the surface area of the wire, Ts is the
surface temperature of the wire, and
is given by

4
Qradiation  As Ts4  Tsur
where
T is the free stream temperature. The heat lost by thermal radiation

 is the surface emissivity of the nichrome wire, 
is the Stephan-Boltzmann constant, and Tsur is
the temperature of the surroundings. The energy equation was solved iteratively for the surface
temperature of the wire using Excel. This analysis tool was used extensively to ensure that the maximum
operating temperature of the nichrome wire was not reached.
Once the expected operating conditions of the power dissipation system were established using
the cylinder in cross flow calculator an experimental test rig was constructed. A picture of this assembly
can be seen in Figure 7.
Figure 7: Resistive Wire Test Stand
The stand consists of parallel walls that replicate the dimensions of the space between the tail
booms on the aircraft. The nichrome is wrapped around ceramic insulators to protect the mounting
hardware from the high temperatures produced. A 28 Volt, 27 Amp power supply was used to simulate
the power generated by the alternator. Variable speed vans were used to produce free stream air
velocities similar to those that are expected to be encountered during flight. The temperature was
measured using a thermocouple in contact with the surface of the nichrome wire. Initially a pyrometer was
used but proved to be difficult to attain consistent measurements. Three power loads were tested at nine
different velocities. The results from these tests can be seen in Figure 8.
17
1200
P=171 Watts Experimental
P=263 Watts Experimental
P=345 Watts Experimental
P=171 Watts Analytical
P=263 Watts Analytical
P=345 Watts Analytical
Wire Surface Temperature(K)
1000
800
600
400
200
0
0
2
4
6
8
10
12
14
16
18
Freestream Velocity(m/s)
Figure 8: Experimental and Analytical Results for the Surface Temperature of Nichrome Wire versus Free
stream velocity
Given the fact that the resistive load will be split into two parallel loads, the maximum power
dissipated by either wire will be approximately 350 Watts. For this reason the tests were conducted up to
a maximum power setting of around 350 Watts. The results show a dramatic decrease in the temperature
as the velocity is increased. Both the experimental and the analytical results show the temperature
approaching some lower limit asymptotically. When comparing the analytic results to the experimental,
the error is never greater than 32%. The analytical results always yielded a higher surface temperature
than those measured experimentally. The wire surface temperature will reach 344 K at a free stream
velocity of 15 m/s. This is well within the safe operating limits of the wire as well as the airframe materials.
Surveillance
When the camera was purchased, the decision was made to purchase extra lenses since it was
not clear initially what size of a lens was necessary to give an adequate view of the ground and the target.
The team purchased four additional lenses on top of the standard 90° lens that comes with the camera. It
was then necessary to determine which lens would do the best job of identifying a target while at the
same time not sacrificing video clarity. The initial range tests were done indoors using a 3’ by 3’ test
18
target composed of nine black and white squares, each one square foot in area. The tests yielded the
following results:
Table 4: Camera/Lens Performance
Lens Angle(deg)
19°
30°
64°
90°
120°
Clear View of Target (ft.)
240
180
100
80
60
Maximum Identifiable Target (ft.)
300
220
140
110
80
The video transmitter was tested to see the maximum range before the picture begins to
deteriorate. These tests were done in a parking lot where there were several sources of interference,
leading the team to believe that the range would be higher once in the air. This test revealed that the
range of the video transmitter was around 300 feet. Beyond 300 feet, the picture began to deteriorate
rapidly. Once the craft was in the air, the range increased significantly as predicted. The range of the
video system was sufficient for the course that the plane will be flying through. The only issue that was
found was a brief loss of video signal on bank turns where the antenna was on the opposite side of the
camera controller. This resulted in changing the mounting location of the antenna.
The next test that was done was actually flying with the dynamic camera mount attached to the
airplane. This mount was made from ABS plastic, produced by the rapid prototyping machine at Wright
State University. This helped to minimize joints on the individual pieces of the camera mount, thus
increasing its durability. The plastic pieces themselves held up very well, however one issue that was
encountered was joining the plastic pieces to other parts of the mount. Glues did not adhere very well to
the plastic, so it was difficult to fit the pieces together. That aside, a great deal was learned from this
camera mount. The dynamic mount successfully gave the camera controller movement in both axes. It
was also learned that the mount was much larger than it needed to be and that it could be mounted at a
downward angle to maximize the forward viewing angle. Also, the pivots on the axes needed to be
redesigned as the gears were stripped out of both servos. It was decided that gears would be placed on
the servo and used in place of a servo horn.
19
16
14
Time On Target(sec)
12
10
8
6
4
19 Deg
30 Deg
64 Deg
2
0
0
50
100
150
200
250
300
Altitude(feet)
Figure 9: Time on Target as a Function of Altitude for Different Camera Lenses
Figure 9 shows time on target values for the three different camera lenses that were being
considered. The 19° and the 30° lenses show an increase of time on target as altitude increases as
expected; however, this is not the case with the 64° lens. The calculations for time on target were
computed using 700 feet as the absolute maximum for time on target since the only time that counts for
identifying targets is between the pylons. Although the 64° lens has the capability to stay on target for
much longer than 14 seconds, this time is considered the maximum effective time on target. For any
altitude between 150 and 250 feet, the 64° lens can see beyond the 700 feet necessary, thus giving it a
constant time on target value as altitude varies.
The forward view for an altitude of 250 feet and the 64° lens is approximately 2400 feet, which is
more than triple the view of the fixed camera. At 150 feet, the forward view is approximately 1400 feet.
In terms of surface area, the dynamic mount allows the video system the capability to see more than 40
times the ground area than that of the fixed camera.
20
DETAIL DESIGN
The detail design phase started after all experiments with the prototype airplane were completed.
This phase finalized the subsystem designs by using the data and analyses from bench-top experiments.
Airframe
After extensive flight testing with the prototype, many improvements were made to the second
aircraft. In response to major structural problems with the original firewall, a new stronger system was
built for the competition aircraft. The improved design used a 1/4” piece of high quality ply and is bolted
directly to the original firewall with blind nuts. The nose wheel arrangement was strengthened
considerably along with the rudder system. The mounting brackets were spaced further apart for the
nose wheel in an effort to distribute the load over a larger area. The weak balsa at the rudder linkage joint
was also replaced with stronger materials making the rudder system much more durable. After some
problems with poor elevator control at slow speed the new aircraft was fitted with an elevator having a
1/2” larger chord and larger control throws. The mechanical advantage between the servo and surface
was improved by moving the linkage in at the servo horn and out at the elevator linkage. Further
lightening of the fuselage, a stronger main landing gear and an improved firewall were also incorporated
into the final airframe.
The performance of the aircraft was measured using a combination of onboard flight telemetry,
analog measuring of item such as take off distance, and lab testing for items such as prop selection. The
results of the dynamic tests can be seen below in Table 5. The final aircraft configuration can be seen in
Figure 10.
Figure 10: Final Aircraft Configuration
21
Propulsion
After a significant amount of research on the available propulsion systems, an OS 1.20AX glow
plug engine will be utilized to propel the aircraft. This engine was optimal in terms of power to weight
compared to electric motors and other glow engines. It also provided sufficient thrust while simultaneously
driving the alternator for power generation. Extensive bench-top testing showed that this system would
carry the plane at high speeds while supplying surplus power for electrical generation.
Power Generation
The final design of the power generation system consists of an AXI 4130-20 electric motor being
driven by an OS 1.20 AX glow engine via a direct drive pulley system. The pulley drive system consists of
a 1.6:1 gear ratio in order to attain approximately 32 Volts output from the motor at a throttle setting
sufficient to maintain flight stability. By outputting 32 Volts from the alternator the team has accounted for
voltage drops that will occur through the rectifier as well as the regulator. The goal is to have the input to
the regulator be as close to 28 Volts as possible to minimize the energy wasted through the regulator.
The output from the alternator will be connected to a 3-phase full wave rectifier. This device
inverts the negative portion of the alternating current wave yielding an effective direct current. The output
of the rectifier then is connected to the voltage regulator. The regulator is made by Sullivan Products and
is capable of delivering a 28 Volt output at 25 amps. The output of the regulator will pass through a 30
amp main system fuse before being connected to the resistive load. Actuation of the parallel resistive
loads will be done via servo-controlled micro switches. This feature enables the ground operator to
engage the loads individually in order to reduce the likelihood of problems occurring from pulling the
entire 700 W load off the engine at a single instant. Each resistor will be connected with a 15 Amp fuse as
well. A schematic of the electrical power generation system can be seen in Figure 11.
Figure 11: Power Generation System Electrical Schematic
22
The final large scale static test for the power generation system consisted of determining the
decrease in thrust as the load on the engine increases. Static thrust testing under no load indicated the
16x6 propeller gave the maximum thrust; therefore this propeller was selected for the initial tests under
load. The engine was run to full throttle position and at that point the load was energized via a fuse. It was
shown that even under a 700 Watt load the engine would not stall. Early on in design this was a major
concern. The thrust was then measured for four different power settings ranging from 200 to 670 Watts.
After the initial tests were conducted with a 16x6 propeller it was theorized that the larger, more
aggressive propeller was putting higher load on the engine. For this reason a 15x4 propeller was then
testing. It was theorized that the smaller, less aggressive propeller would draw less power. This proved to
be the case. The results of these tests can be seen in Figure 12. While the initial static thrust of the 15x4
is less than that of the 16x6, as the load is increased the thrust of the 15x4 does not decrease as fast or
as far as the 16x6. Given the fact that the aircraft only requires approximately 5 pounds of thrust for take
off, the 15x4 still provides sufficient performance. This test proved that under the initial design
specification of producing 700 Watts, the engine and alternator system are capable of supplying this
power while maintaining the performance of the aircraft.
9
8
15x4 Propeller
16x6 Propeller
7
Thrust(lbf)
6
5
4
3
2
1
0
0
100
200
300
400
500
600
700
Power(Watts)
Figure 12: Comparison of Engine Thrust vs. Electrical Power Draw
23
Alternator Drive
For the final design, the team decided to go from a pusher style configuration to a traditional
tractor configuration. By moving the engine to the front of the airplane, the team was able to get the
center of gravity in the front of the wing without adding any unnecessary weight. Then to generate power,
the team was able to go back to the direct drive design, with the AXI motor directly under the engine. This
eliminated unwanted weight, large vibrations, and extra mechanical components. Instead of using a 1.5
ratio between the two pulleys, a 1.6 ratio was used to generate the correct output voltage at an engine
speed similar to cruising speed.
Surveillance
The final design of the surveillance system includes the BlackWidow model KX-141 camera fixed
to a custom gear-driven mount. The camera mount uses two servo motors to move the camera through
two axes. Given the position of the mount, the camera operator has the ability to look almost straight
forward. This proved to be incredibly helpful in finding targets on the field, where the entire 700 foot
course is visible during that leg of the flight. Testing also showed that during the opposite leg of the
course the camera can be panned to the side to locate targets. Given the random nature of the target
placement the ability to remove the task of finding targets from the pilot has been a great benefit. The
pilot is going to have a great deal to focus on and removing that additional distraction has proved to be
successful. For the above mentioned reasons the gear driven dynamic camera mount was selected for
the final design on this aircraft.
24
25
Table 5: Aircraft Geometry, Performance and Weight/Balance Parameters
Geometry
Length
Wing Span
Wing Area
65.796 in
80 in
1200 in^2
Performance
Clmax
L/D Max
Max Rate of Climb
Stall Speed
Max Airspeed
Take-Off Field Length
1.29
21.5 at α=2°
23 feet/sec
20.5 feet/sec
111.46 feet/sec
85 feet
Weight and Balance
Airframe
Propulsion System
Control System
Video System
Power Generation System
Gross Weight
CG Location
6.22 lb
3.1 lb
0.58 lb
0.71 lb
3.68 lb
14.29 lb
3.75 from leading edge
Table 6: Aircraft System Components
Airframe
Airframe Kit
Electronics
Servos
Engine
Nitro Models Cessna
Airframe
O.S 120 A.X
Transmitter
Fuel Tank
Great Planes Adjustable
Engine Mount
Dubro 14 oz
Great Planes Standard Fuel
Tubing
Receiver
Landing Gear
TNT Custom Landing Gear
Covering
Material
Building
Balsa
Dave Brown Lite Wheels
Top Flite Monokote
Green/Gold
ordered from Balsa USA and
Lonestar Models
Transmitter
Batteries
Switches
Battery
Monitor
Futaba S3102 Micro Servos
Futaba 20" and Futaba 40"
Heavy Duty Servo Extension
Futaba 9CHPS 9-Channel
Synthesized Tx/Rx 72MHz
Futaba 9C/9CS Synthesized
Transmitter Module 72MHz
Futaba R319DPS
Synthesized 9-Channel PCM
Receiver
Futaba NR4J Receiver NiCd
Flat 4.8V 600mAh
Hobbico Switch Harness
Hobbico Voltwatch 2 4.8V/6V
Rx Battery Monitor
26
Table 6(Cont.): Aircraft System Components
Video
System
Video
Camera
Power
Generation
Black Widow Wireless Video
System
120,64, 30,19 degree Black
Widow Video lenses
Futaba S3102 Micro Servos
Alternator
Video
Conversion
Receiver
Battery
Pinnacle USB Video system
Voltage
Regulator
System
Protection
Rectifier
Futaba NR4J Receiver NiCad
Flat 4.8V 600mAh
Resistive
Load
Transmitter
JR XP7202A Synthesized
72MHz
JR R790 9 Channel
Synthesized
Servos
Receiver
AXI 4130 Brushless Electric
Motor
28Volt/25Amp Sullivan
Regulator
30 Amp Fused System
Crydom B485B-2T
18-Gauge Nichrome
(NiCr60) Resistive Heating
Wire
Data
Acquisition
Data
Recorder
EagleTree Systems FDR Pro
with Seagull Wireless Unit
and Electric Expander
TESTING PLAN
Experimental testing was carried out on a subsystem-by-subsystem basis as much as possible.
This methodology provided the team valuable information on the performance of a subsystem in a timely
manner as each sub-system was procured and brought online. This also allowed individual team
members to work independently, which greatly accelerated the testing process. A summary of the
experiments conducted is shown in Table 7 at the end of this section.
Power Generation
To validate the on-board power measurement data acquisition system, an array of 300mAh
NiCad batteries were mounted to the aircraft and were discharged through a 40  resistive load. The load
consisted of light gauge nichrome wire wrapped around a wooden dowel. The batteries were configured
to deliver 27 volts. This would produce a current of 0.65 Amps for approximately 20 minutes. The noncontact current ring from the Eagletree telemetry system was placed over the wire leading to the resistive
load and the voltage leads were attached. The system was fused and activated in flight using a servo
actuated micro switch once cruising altitude was achieved. The base station, consisting of a Dell laptop
computer running the Eagletree data acquisition software, began displaying the power in real time. This
test proved the capabilities of the wireless telemetry system to monitor the power being generated on
board during flight.
To test the viability of using the brushless motor as an alternator it was decided to build a test
stand that would use another electric motor to drive the alternator. This would enable the alternator to be
27
tested under different loading conditions without running the glow engine. A suitable bench top power
supply was used to power an AXI 4130 motor that drove the alternator. Figure 13 shows the test stand
that was constructed.
Figure 13: Electric Motor Driven Alternator Test Stand
This stand allowed for a large number of resistive loads to be tested. This stand also enabled the
team to determine the no-load draw on the system. This test showed that the power required to overcome
the forces of the magnets within both motors is approximately 80 Watts. The results of these tests can be
seen below in Figure 14.
350
300
y = -0.0006x2 + 1.1724x - 89.407
R2 = 0.9996
Power Output(W)
250
200
150
100
50
0
0
50
100
150
200
250
300
350
400
450
Power Input(W)
28
Figure 14: Display of Power Output vs. Power Input for an Increasing Resistive Load
Static Thrust
Static thrust measurements were taken using a custom designed thrust stand. The stand consists
of an engine mount, L-frame and a digital scale. A picture of the stand can be seen below in Figure 15.
Figure 15: Engine Test Stand
Taking into account the manufacturers recommendations for propeller sizes the team selected
four propellers to test. The results from those tests are shown in Figure 16.
29
14
12
Thrust(lbf)
10
8
6
4
2
0
16x6
16x6-10
15x8
15x4
Figure 16: Static Thrust Results for Various Propellers
Each propeller was tested on three separate runs at full throttle and those values were then
averaged to find the maximum thrust for that propeller. These results show that the 16x6 propeller
provides the greatest static thrust which is of primary concern when determining take-off performance.
Surveillance
The testing procedures undertaken for the video surveillance system have been discussed in
previous sections. The test dates and results acquired can be seen in Table 7.
Telemetry
The data acquisition system was tested first in lab experiments to familiarize the team with its
operation. Measurements of temperature, altitude, power and airspeed were all conducted and compared
with independent measuring devices to confirm the accuracy of the data being collected by the EagleTree
flight data recorder. Once the validity of the EagleTree system was established the unit was placed on the
airplane for flight testing. Issues with data capture rate were initially found but were corrected within the
software. The Seagull Wireless Dashboard can receive information at a maximum rate of 10 frames/sec.
This proved to be sufficient for all tests being conducted.
Airframe
Throughout the duration of our flight testing, ample monitoring of many flight details took place
including take off, rate of climb, max speed, and stall speed. Using both the pilot and live/recorded on-
30
board Eagle Tree telemetry the necessary flight data was collected. The following represents the
procedures taken during the data collection. The results of these tests can be seen above in Table 5.
Take Off Performance
When testing the take off performance of the aircraft the total distance from rest to initial liftoff was
measured. The data is both the result of visual flag markers and the on-board telemetry. The flags were
set in intervals of 10 feet across the entire length of the runway and allowed a visual take off distance to
be calculated with the actual distance recorded by the on-board telemetry.
Rate of Climb
Once the aircraft established flight, the aircraft was then immediately positioned into the highest rate of
climb achievable without the potential of stalling the aircraft. A number of these flight patterns were flown
with similar consistency so an average rate of climb would be obtainable after reviewing the on-board
telemetry.
Stall Speed
During the stall speed tests the aircraft was first positioned to the competition flight ceiling of 250 feet.
With flight altitude established, the throttle was then slowly decreased until the time the pilot recognized
he had lost flight sustaining lift. Using the live on-board telemetry for airspeed, the correct speed was
recorded for the time in which the aircraft stalled.
Maximum Airspeed
The maximum airspeed was determined after exiting each 180 degree turn. Upon exiting each turn, full
throttle was applied and speed was monitored through both live and recorded telemetry. Circuits were
repeated depending on the weather each day to determine the highest maximum airspeed.
The following table indicates when certain tests were conducted for the various systems on the
aircraft and what conclusions were reached upon completion of that test.
Table 7: System Tests Performed
TEST DESCRIPTION
Surveillance
DURATION
DATE
RESULTS
Indoor Stationary Optical
Range Testing of
Surveillance System
2 Days
2/6/2007
Using a 3 square foot target determined the
maximum range that the 1 foot black and
white squares could be identified.
Outdoor Stationary Optical
Range Test
1 Day
3/29/2007
Same objective as Indoor test but at greater
maximum range.
3/29/2007
Determine the point at which the signal
strength from the camera transmitter fails to
yield a clear image
4/3/2007
Proved the ability of a remote operator to
control the camera mount system while
viewing a monitor
Outdoor Stationary Wireless
Signal Range Test
In-Flight Dynamic Camera
Mount Testing
1 Day
3 Days
31
Table 7 (Cont.): System Tests Performed
In-Flight Camera Lens Angle
Testing
Determined 64° Lens provides the best
compromise of Time on Target and Target
Size
1 Day
4/22/2007
Initial Rectifier Testing
1 Day
3/15/2007
Electric Motor Driven
Alternator Testing
2 Days
4/13/2007
Stationary Power Dissipation
Testing
4 Days
5/10/2007
No Load Static Thrust
Testing
1 Day
5/14/2007
Engine Driven Alternator
Testing
2 Days
5/18/2007
Showed the expected DC voltage output was
acquired given the speed of alternator
Determined the Minimum Power Input Before
Power can be generated. Proved the
concept of generating power using a
brushless motor.
Showed the affects of airflow on the surface
temperature of the resistive wire. Established
the safe operating conditions for the wire in
flight.
Measure the Static Thrust produced by
different propellers. Determined which would
produce the greatest thrust for use on the
airplane.
Determined the effects of drawing the
electrical load off the engine. Tests indicated
that the engine will provide sufficient thrust
under full load.
In-Flight Telemetry Testing
1 Day
4/1/2007
Determined Altitude, Airspeed, Distance to
Operator
In-Flight Power
Measurement Testing
1 Day
4/3/2007
Verified the Functionality of using EagleTree
Telemetry to Measure the Power Dissipated
5/19/2007
Using Both Onboard Telemetry and markers
on the Field Measured Distance from
Starting Roll to Wheels Leaving the Ground
5/19/2007
Measure Time from Start of Roll to the time
the Aircraft has reached 250 Feet using
onboard telemetry
5/19/2007
With the aircraft at altitude monitor airspeed
via telemetry system and slow the aircraft
until a stall occurs.
5/19/2007
While monitoring airspeed via telemetry
system make 3 high speed passes from both
directions.
Power Generation
Airframe and Data
Acquisition
Determine Take-Off Distance
Determine Time to Altitude
Determine Stall Speed
Determine Maximum
Airspeed
1 Day
1 Day
1 Day
1 Day
32