43rd AIAA Joint Propulsion Conference Student Design …
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43rd AIAA Joint Propulsion Conference Student Design Challenge
Final Report
[pic]
Team Members Faculty Advisor
William Bennett Dr. Scott Thomas
Jayme Carper
Nicholas Hankinson
Michael Sheridan
Keith Vehorn
Stephen Warrener
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 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
Airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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
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 twin-boom 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 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|Propulsion |Power 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.
[pic]
Figure 2: Anticipated and Actual Timelines for Design Phases to be Accomplished
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 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.
[pic]
Figure 3: Comparison of Selected Power Plant Options
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 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 |
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 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 |8 |
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 [pic] where Vp is the voltage drop across the 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:
[pic]
The same analysis was conducted using the voltage and direct current measurement technique. In this case [pic] , where [pic] is the output current from the alternator. In this technique ΔP can be obtained from the following equation:
[pic]
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 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.
[pic]
Figure 4: Side View Comparison of Dynamic and Static Camera Mounts (Not to Scale)
[pic]
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.
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.
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:
[pic]
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.
[pic]
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 over-predicting the wire temperature, which is considered to be beneficial with respect to safety.
Incropera and DeWitt (1990) suggest that the correlation proposed by Zhukauskas (1972) is appropriate for the situation outlined above:
[pic]
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 |
|1000 to [pic] |0.26 |0.6 |
|[pic] to [pic] |0.076 |0.7 |
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:
[pic]
[pic]
[pic]
where the units of temperature are degrees Kelvin, the kinematic viscosity is in m2/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
[pic]
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.
[pic]
The heat convected away by the air is
[pic]
where [pic] is the convective heat transfer coefficient, [pic] is the surface area of the wire, [pic] is the surface temperature of the wire, and [pic]is the free stream temperature. The heat lost by thermal radiation is given by
[pic]
where [pic]is the surface emissivity of the nichrome wire, [pic] is the Stephan-Boltzmann constant, and [pic]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.
[pic]
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.
[pic]
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 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.
[pic]
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.
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.
[pic]
Figure 10: Final Aircraft Configuration
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.
[pic]
Figure 11: Power Generation System Electrical Schematic
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.
[pic]
Figure 12: Comparison of Engine Thrust vs. Electrical Power Draw
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.
[pic]
Table 5: Aircraft Geometry, Performance and Weight/Balance Parameters
|Geometry | |
|Length |65.796 in |
|Wing Span |80 in |
|Wing Area |1200 in^2 |
| | |
|Performance | |
|Clmax |1.29 |
|L/D Max |21.5 at α=2° |
|Max Rate of Climb |23 feet/sec |
|Stall Speed |20.5 feet/sec |
|Max Airspeed |111.46 feet/sec |
|Take-Off Field Length |85 feet |
| | |
|Weight and Balance | |
|Airframe |6.22 lb |
|Propulsion System |3.1 lb |
|Control System |0.58 lb |
|Video System |0.71 lb |
|Power Generation System |3.68 lb |
|Gross Weight |14.29 lb |
|CG Location |3.75 from leading edge |
Table 6: Aircraft System Components
|Airframe | |Electronics | |
|Airframe Kit |Nitro Models Cessna Airframe |Servos |Futaba S3102 Micro Servos |
|Engine |O.S 120 A.X | |Futaba 20" and Futaba 40" Heavy Duty |
| | | |Servo Extension |
| |Great Planes Adjustable Engine Mount |Transmitter |Futaba 9CHPS 9-Channel Synthesized Tx/Rx |
| | | |72MHz |
|Fuel Tank |Dubro 14 oz | |Futaba 9C/9CS Synthesized Transmitter |
| | | |Module 72MHz |
| |Great Planes Standard Fuel Tubing |Receiver |Futaba R319DPS Synthesized 9-Channel PCM |
| | | |Receiver |
|Landing Gear |TNT Custom Landing Gear |Transmitter Batteries |Futaba NR4J Receiver NiCd Flat 4.8V |
| | | |600mAh |
| |Dave Brown Lite Wheels |Switches |Hobbico Switch Harness |
|Covering Material |Top Flite Monokote Green/Gold |Battery Monitor |Hobbico Voltwatch 2 4.8V/6V Rx Battery |
| | | |Monitor |
|Building Balsa |ordered from Balsa USA and Lonestar Models| | |
Table 6(Cont.): Aircraft System Components
|Video System | |Power Generation | |
|Video Camera |Black Widow Wireless Video System |Alternator |AXI 4130 Brushless Electric Motor |
| |120,64, 30,19 degree Black Widow Video |Voltage Regulator |28Volt/25Amp Sullivan Regulator |
| |lenses | | |
|Servos |Futaba S3102 Micro Servos |System Protection |30 Amp Fused System |
|Video Conversion |Pinnacle USB Video system |Rectifier |Crydom B485B-2T |
|Receiver Battery |Futaba NR4J Receiver NiCad Flat 4.8V |Resistive Load |18-Gauge Nichrome (NiCr60) Resistive |
| |600mAh | |Heating Wire |
|Transmitter |JR XP7202A Synthesized 72MHz | | |
|Receiver |JR R790 9 Channel Synthesized | | |
|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 non-contact 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 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.
[pic][pic]
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.
[pic]
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.
[pic]
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.
[pic]
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-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 |DURATION |DATE |RESULTS |
|Surveillance | | | |
|Indoor Stationary Optical Range Testing |2 Days |2/6/2007 |Using a 3 square foot target determined the maximum range that|
|of Surveillance System | | |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. |
|Outdoor Stationary Wireless Signal Range |1 Day |3/29/2007 |Determine the point at which the signal strength from the |
|Test | | |camera transmitter fails to yield a clear image |
|In-Flight Dynamic Camera Mount Testing |3 Days |4/3/2007 |Proved the ability of a remote operator to control the camera |
| | | |mount system while viewing a monitor |
Table 7 (Cont.): System Tests Performed
|In-Flight Camera Lens Angle Testing |1 Day |4/22/2007 |Determined 64° Lens provides the best compromise of Time on |
| | | |Target and Target Size |
|Power Generation | | | |
|Initial Rectifier Testing |1 Day |3/15/2007 |Showed the expected DC voltage output was acquired given the |
| | | |speed of alternator |
|Electric Motor Driven Alternator Testing |2 Days |4/13/2007 |Determined the Minimum Power Input Before Power can be |
| | | |generated. Proved the concept of generating power using a |
| | | |brushless motor. |
|Stationary Power Dissipation Testing |4 Days |5/10/2007 |Showed the affects of airflow on the surface temperature of |
| | | |the resistive wire. Established the safe operating conditions |
| | | |for the wire in flight. |
|No Load Static Thrust Testing |1 Day |5/14/2007 |Measure the Static Thrust produced by different propellers. |
| | | |Determined which would produce the greatest thrust for use on |
| | | |the airplane. |
|Engine Driven Alternator Testing |2 Days |5/18/2007 |Determined the effects of drawing the electrical load off the |
| | | |engine. Tests indicated that the engine will provide |
| | | |sufficient thrust under full load. |
|Airframe and Data Acquisition | | | |
|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 |
|Determine Take-Off Distance |1 Day |5/19/2007 |Using Both Onboard Telemetry and markers on the Field Measured|
| | | |Distance from Starting Roll to Wheels Leaving the Ground |
|Determine Time to Altitude |1 Day |5/19/2007 |Measure Time from Start of Roll to the time the Aircraft has |
| | | |reached 250 Feet using onboard telemetry |
|Determine Stall Speed |1 Day |5/19/2007 |With the aircraft at altitude monitor airspeed via telemetry |
| | | |system and slow the aircraft until a stall occurs. |
|Determine Maximum Airspeed |1 Day |5/19/2007 |While monitoring airspeed via telemetry system make 3 high |
| | | |speed passes from both directions. |
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