Camera - Purdue University
AAE 451 – Aircraft Senior Design
Spring 2007
Continuous Area Coverage via Fixed-Wing Unmanned Aerial Systems
Conceptual Design Review
Team 3
Sumitero Darsono
Charles Hagenbush
Keith Higdon
Seung-il Kim
Matt Lewis
Matt Richter
Jeff Tippmann
Alex Zaubi
Table of Contents
1. Table of Content 1
1. Executive Summary 2
2. System Requirement and Definition Review 3-7
1. Business Case 3-4
2. Target Market 4-5
3. Concept of Operation 5
4. Design Mission 5-6
5. Design Requirement 6-7
3. Selected Aircraft Concept 8-11
1. 3 View 8
2. Detailed Design 8-10
3. CG and Other Considerations 10-11
4. Payload Summary 12-16
1. Camera 12-13
2. Avionics 13-15
3. Fuel Cell 15-16
5. Aircraft Sizing 17-18
1. Carpet Plot 18
6. Power Issues 19-20
7. Aerodynamics Detail 21-24
8. Performance Analysis 25-28
9. Propulsion 29-37
1. Motor 29
2. Propeller 29-37
10. Structure 38
11. Weight Statement 39
12. Cost Analysis 40-42
1. Support Equipment 40
2. Aircraft 40-41
3. Pricing 41-42
13. Additional Work 43
14. Reference 44
15. Appendix 45-47
Table of Content 1
16. Executive Summary 2
17. System Requirement and Definition Review 3-7
1. Business Case 3-4
2. Target Market 4-5
3. Concept of Operation 5
4. Design Mission 5-6
5. Design Requirement 6-7
18. Selected Aircraft Concept 8-11
1. 3 View 8
2. Detailed Design 8-10
3. CG and Other Considerations 10-11
19. Payload Summary 12-16
1. Camera 12-13
2. Avionics 13-15
3. Fuel Cell 15-16
20. Aircraft Sizing 17-19
1. Carpet Plot 18-19
21. Power Issues 20-21
22. Aerodynamics Detail 22-25
23. Performance Analysis 26-29
24. Propulsion 30-38
1. Motor 30
2. Propeller 30-38
25. Structure 39-40
26. Weight Statement 41
27. Cost Analysis 42-43
1. Support Equipment 42
2. Aircraft 42-43
3. Pricing 43-44
28. Additional Work 45
29. Reference 46
30. Appendix 47-49
1. Executive Summary
Unmanned Aerial Vehicles (UAVs) are remotely piloted or self piloted aircraft that is are used to carry specific payloads such as cameras, sensors, communications and other equipments during a mission to perform a specific task. This These tasks includes forward reconnaissance and surveillance. The Department of Defense (DoD) has classified the UAV into seven main categories, the Pioneer, Tactical UAV, Joint Tactical UAV, Medium Altitude Endurance UAV, High Altitude Endurance UAV, Tactical Control System and the Micro Unmanned Aerial Vehicles10.
Currently, there are large numbers of UAVs available in the market. However, the availability of a UAV that is small, light, portable, and low costcheap and that is able to provide an endurance of greater than four hours is very limited. This project is aimed to at explore exploring the small UAV market for military and law enforcement, and to provideing an unmanned aerial system that is more capable than those exist currently in the market.
In this conceptual design reviewThe conceptual design review determines, the dimension, weight, and size of the UAV. are determined. These parameters are determinedcome from the sizing codes that were built from the groupthat the group designed. The results from the sizing code showed that the UAV will weight 10.1 lbs withand have an endurance of 5.5 hours, loiter speed of 42 ft/s, stall speed of 33 ft/s, and wing span of 8 ft. The UAV must is also be hand launched to allow for rapid deployment in case of emergency. In addition, a fuel cell will power the UAV will be powered by a fuel cell , and the UAV willd carry a small visual or thermal imaging camera for forward reconnaissance. Furthermore, this stage of the design includes analysis of the stability, aerodynamics, and performance of the UAV were analyzed on this stage of the design.
As the design of the project nears the end, a future study is needed. This includes more detailed structure structural and load analysis, including analysis on the effect of impact at landing, more detailed aerodynamic analysis using CFD, tests to determine propeller efficiency, determine determining the control surface size and trim analysis, and determine determining the manufacturing method. Overall, the design of this UAV provides capability well beyond those UAVs that are available in the market nowadays. With low development and production cost, the concept has the possibility to be very successful.
2. System Requirement and Definition Review
Mission Statement:
To provide continuous aerial coverage using an UAS that is small, light, portable and allows for rapid deployment
The business case of the UAV, target market, and the preliminary design concept and parameters were part of the system requirement review. Based on the system requirement review, the UAV design will feature small size, portability, light-weight, low cost, and rapid deployment as the main criteria in the design mission.
1. Business Case
Currently, there are a large number of UAVs available in the market. However, the idea of a small UAV that is light, portable, cheap, and allows for rapid deployment is something that the market has yet to explore. The current UAVs available in the market either have a limited endurance or larger size. The design of this UAV will solve both problems.
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Figure 1: Evolution XTS – L3 BAI Aerosystems
In order to provide continuous coverage, the unmanned aerial system (UAS) will consist of a system of multiple aircraft and support equipment. They will work in conjunction to provide continuous aerial coverage over the area within a five-mile radius. The aircraft will house a small payload consisting of a video camera, a thermal imaging camera, or a chemical detector. The aircraft will either have a module payload or will carry all of the payload types simultaneously depending upon the final payload weight and the weight of the cameras. The aircraft itself will be a micro unmanned aerial vehicle (UAV) that can be hand launched and carried in a military- style backpack. Two or three people will be capable of transporting the entire systemThe entire system will be transportable by two or three people, depending on the number of aircraft needed. The support equipment is very limited and will consist of a small transmission unit and a laptop to program waypoints and view the incoming video feeds. The aircraft and transmission equipment will both be portable so that theyto facilitate can be used anywhere that surveillance is necessary.
2. Target Market
The proposed system focuses on a surveillance market, which includes mainly military and law enforcement personnelorganizations. The military will deploy the UAS system UAS out of either a backpack when on foot or out of a Humvee when traveling. The main uses for the UAS by the military will be for surveillance around a temporary base or convoy or for forward reconnaissance. Law enforcement will deploy the UAS out of the back of a squad car. The main uses for law enforcement will be for assessing a hazardous situation before committing personnel or and providing continuous surveillance of large groups.
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Figure 2: Projected Budget for Procurement of Small UAS9
In recent years, the market for small Unmanned Aerial Systems for the military sector has grown dramatically. Due to advances in sensors, materials, and batteries, the mission capabilities of small UAVs are ever continuously increasing. Combined with the changing scope of warfare, current small UA systemsUASs are seeing more and more use in places such as Iraq and Afghanistan, and the United States military has decided to invest substantially in similar systems. In addition, the Department of Defense also planned to spend more than $20M on small UAV over the next three years (Figure 2)9.
3. Concept of Operations
1. Military
Current unmanned vehicles of this size, the Dragon Eye and Raven for example, provide simple “over the hill” type missions where they observe a target location for a few minutes and then return; our system provides the capability to observe a location or multiple locations for hours at a time. The system can be deployed with the infantry at the squadron or platoon level. Similar to other systems of this size, the aircraft is simply launched by handhand-launched and does not require a runway. In addition, the entire system: aircraft, laptop, and supporting equipment would be transported via backpack or a small container in a Humvee (refer to Appendix I).
2. Law Enforcement
Typically, police agencies can use the UAVs to provide overhead surveillance in assessing hazardous situations before committing personnel. Similar to the military, police officers need to gather information on each mission before performing their actions. Usually, the law enforcement personnel carry out these missions. However, placing a police officer in a situation that is relatively unknown or risky may jeopardize the police officer’s safety. Currently, the alternative method is aerial surveillance provided by helicopter. Helicopters are very expensive to buy and operate, require dedicated pilots, and their have availability is limited. Law enforcement agencies can use UAVs as a perfect substitute for a helicopter in the aerial surveillance role.4
4. Design Mission
As expressed in the mission statement, the goal of this aircraft is to provide continuous aerial coverage via a portable UAS. With this in mind, the design mission of this aircraft will consist of a significant loiter segment so as to provide continuous coverage. The pPortability means that a typical mission will require very little cruise time to the area of interest, and, with its hand-launch capabilities, the aircraft can takeoff, as well as land, at virtually any location. This description is summarized below in Figure 3Figure 3 gives a summarized description of a typical mission, with some of the key performance capabilities shown in Table 1.
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Figure 3: Aircraft Design Mission
Table 1: Mission Capabilities
|Parameter |Values |Units |
|Total Weight |10.1 |lbs |
|Endurance |5.5 |hrs |
|Range |10 |miles |
5. Design Requirement
To provide the major design requirements, the team started with the attributes that would be important to the customer. These customer attributes were ranked in order of importance for the military and law enforcement individually. The most important attributes include: continuous coverage, hand launch, ability to pack in a backpack, easy to pilot, and low in cost. These attributes, once ranked, were put into a quality functional diagram (QFD)2. The QFD then related each customer attribute to each engineering requirement. The, and related the engineering requirements were also related to each other. The relations showed the team which engineering requirements affect which customer attributes and other engineering requirements. From these relations, the engineering requirements were ranked in order of importance. The most important engineering requirements include: low take off weight, short wing span, low stall speed, and long endurance. The QFD itself can be seen in the appendix.
3. Selected Aircraft Concept
1. 3-view
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Figure 4: 3-View of UAV
The final concept was chosencame from by an iterative process iterating through involving Pugh’s concept selection method (source: SDR). As more information on component sizing became available, the final design began to take shapematerialized. The final configuration and size of the UAV is roughly 8.5 feet wide, 5 ft long, and 10 inches high.
2. Detailed Design
There are many important engineering requirements built into the design. The carbon fiber construction method reduces the weight of the interior. The skin of the fuselage is not structurally supportive and therefore the weight can be reduced by almost half by only using 1 ply of carbon fiber skin instead of 2 pliesy’s. The skin is designed functions only to protect the inside components and panels from dirt and scrubbing during a landing. The forces from the landing though will beare transferred through the panels and not the skin. The skin is designed to be replaceable when it is becomes too damaged to flyexcessively damaged. The plexi-glass cover is also meant to be replaceable when it becomes too scratched for camera use.
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Figure 5: Break down of UAV
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Figure 6: Component Layout
1. Wing
Each wing is one ply of carbon fiber with a foam core to stiffen the entire wing. Each wing is detachable from the inner structure of the fuselage. Though not modeled, the plan is to have a rod that slides through the fuselage panel and is and through both wings and held in place by pins in each wing. The tail fins are constructed similarly.
2. Tail
The tail boom is a carbon fiber tube roughly 0.02 inches thick. This component will be detachable from the fuselage and tail fins.
3. Fuel Cell
The heaviest component of the UAV, the fuel cell, will sit in the middle of the fuselage, right directly under the wing, thereby reducing the moment on the wing attachments. The volume of the fuel cell was set based on the volume given by Protonex6 (source: Protenex document).
4. Camera
The camera is in the front of UAV to provide the best field of view. The camera will be protected by the plexi-glass cover on the front. The camera will hang from one of the structural panels.
5. Avionics
Above the camera is the avionics and communications board chosen from MicroPilot5. The size of the avionics board is small enough to fit in the small compartment in the fuselage structure.
6. Propeller
In On the front of the UAV is the propeller, which will beis mounted to the bulkhead directly in front of the camera and avionics board.
3. CG and other considerations
The center of gravity (CG) of the aircraft with the components in place is 15.4 inches from the leading edge of the fuselage. The CP is roughly 1.4 inches behind the CG to create a positive static margin of 15% of the wing chord. Because Since the fuel cell weight stays nearly constant during the whole mission, there is no CG travel.
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Figure 7: UAV Center of Gravity and Aerodynamic Center
1. Launching and Landing considerations
Since the launch of the aircraft is by hand, the person launching the aircraft should have a firm grip near the cg CG of the aircraft. The back end of the fuselage will be roughly 4 inches thick at the point where the launcher will hold the aircraft.
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Figure 8: Hand Launch UAV
4. Payload Summary
The payload weight for the UAV incorporates the sensor and communication equipment. Looking at sensors of the low resolution type, these components typically have weights around 2 – 4 lbs. The UAV will carry a system of visual and infrared cameras to provide day and night surveillance. A study was conducted on many visual and infrared cameras to find the best set of cameras, communications package, and fuel cells. The selected components have tomust be light but also must be capable of performing the design mission.
Currently, the selected camera is the Controp D-Stamp Stabilized Camera3. For the avionics, the Advanced Miniature UAV, MP2128LRC5, Autopilots by Micro Pilots was chosen, and the fuel will be the Protonex Procore fuel cell6.
1. Camera
One of the limiting factors in the UAV design is the weight of the payload. Therefore, in selecting the camera, it must be light and have a low power requirement. The selected camera is the CONTROP D-Stamp Camera3. The D-Stamp Camera is a miniature lightweight electro-optical, stabilized airborne sensor that is designed to be carried by a small UAV. This camera is capable for operationof operating at speeds up to 67.5 ft/s and at an altitude of approximately 2000 ft.
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Figure 9: CONTROP D-Stamp Camera
The advantage of the D-Stamp camera compared to others currently available in the market is the its very light low weight of approximately 1.4 lb, with the core only weighting only 1 lb. It is suggested that the enclosure of the camera will be removed for weight saving purpose. In addition, the camera comes with a gyro stabilized stabilizer that maintains the picture orientation independent regardless of the aircraft maneuvers and vibrations. This ensures a continuous high quality display from the camera.
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Figure10: The Results Projected from the Camera at 2000 ft
In addition, the camera has a very low power requirement of only 10 Watts. With such a low requirement, the camera allows continuous operation from the beginning of the climb to landing of the UAV.
2. Avionics5
For the UAV to be able to perform the specified capabilities, it needs requires a controller for the aircraft. Based on the design of the aircraft, the autopilot needs to be light and consume minimum power. Currently, there are several miniature UAV autopilot controllers available on the market. Based on the mission criteria of the UAV, the autopilot must be able to perform both the autonomous flight using GPS and manual control flight.
The autopilot chip comes with an aluminum enclosure. With the enclosure the weight of the component will be 0.727 lbs. With a volume of 38.5 inch3, it is small enough to fit in inside the fuselage of the UAV design. However, it is under study that the aluminum enclosure could possibly be replaced by composite material to reduce the weight of the UAV.
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Figure 11: The Micropilot UAV Autopilot with aluminum enclosure5
The Advanced Miniature UAV, MP2128LRC, Autopilots by Micro Pilots is the world’s smallest UAV autopilot currently available in the market. The chip only weighs 0.06 lbs (includes the GPS receiver) with an extremely low power requirement of 1 Watt. The autopilot chip has the capability to perform GPS waypoint navigation while maintaining constant altitude and airspeed. The autopilot can be controlled using three different modes, autonomous flight using the GPS, manual control flight, and emergency direct servo control. The emergency direct servo control will be activatedactivate when the UAV loses contact with the transmitter, and it will direct the UAV back to the starting location or some other predetermined location.
The autopilot chip comes with an aluminum enclosure. With the enclosure, the weight of the component will be 0.727 lbs. With a volume of 38.5 inch3, it is small enough to fit inside the fuselage of the UAV. However, it is under consideration that the aluminum enclosure could possibly be replaced by a similar composite enclosure to reduce the weight of the UAV.
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Figure 12: The Micropilot UAV Chip Weight at 28 g (0.06lbs)5
In addition,The avionics’ the transmitter has a range of more than 30 miles. This excess range power of the transmitter allows the UAV to operate in the urban environment without the need to worry about the interference created by additional building.
Table 2: Specification for the Avionics5
|Parameters |Value |
|Weight |0.06 lbs |
|Weight with Aluminum Enclosures |0.727 lbs |
|Power Requirement |1 W |
|Volume |38.5 in3 |
3. Fuel Cell
Based on preliminary trade studies performed in the System Requirements Review, the UAV requires a power system with a power density beyond the range of current batteries. This provides the necessary endurance and hand launch capability. Fuel cells have shown promise in providing these high power densities and are just now entering the market. Protonex has developed the ProCore fuel cell system, which is specifically tailored to miniature UAV applications. The fuel cell relies on sodium borohydride as the fuel rather than hydrogen, which could be dangerous in a military UAV application1.
The specifications of the fuel cell are shown below in Table 3 and a picture of the product is shown in Figure 136.
Table 3: Specifications of Protonex ProCore Fuel Cell
|Parameters |Value |
|Output Power |50-200 W |
|Output Voltage |20-30 V |
|Output Current |1-10 Amps |
|Total Available Energy |770 W-h |
|Weight |2000 g (4.41 lbs) |
|Volume |170.8 in3 |
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Figure 13: Protonex ProCore Fuel Cell
As seen in, Table 3, this power system has an energy density of 335 W-h/kg, which is well above the 200 W-h/kg that the best batteries can provide. The power and voltage supplied by the ProCore system also will be sufficient to power the propulsion system as well as the payload and avionics.
5. Aircraft Sizing
The aircraft sizing code was developed independently, since FLOPS, ACS, and other commercially available sizing codes are not suitable for small UAVs with such a low take off weight and an electrical power system.
The sizing code simply calculates the take off weight based on its the UAV’s aspect ratio and wing loading. The initial weight was predicted to be approximately 10 lbs. This prediction is based on the database of the small UAVs currently available in the market. With the weightThe weight then allows for prediction of, the dimension of the wing can be predicted. The weight of the wings are also found by adding the weight of the form core, calculated from the volume of the wing times the density of foam, and the weight of the wing skin, calculated from the wing surface area times the density of carbon fiber composite material. As the airfoil of the UAV was chosen, and the area and circumference of the airfoil could be related to chord length, which provided agave more exact weight estimation. In addition, the fuselage’s maximum width was about approximately 6 inches, so the effective wing dimension yields an additional 6 inches, which results in wing span that is slightly less than 8 feet.
The size of the tail was dimensioned next. Tail size calculations are discussed in the aerodynamic detail section. By cCalculating more detailed tail sizee, we can findprovides its tail weight in a method similarly to that of thethe wing. The tail was sized based on the tail volume coefficient method described in Raymer8. The horizontal tail was sized using a volume coefficient of 0.7, and the vertical tail was sized using a volume coefficient of 0.04. A vertical tail coefficient of 0.04 was selected because it is common among single engine propeller aircraft, and the horizontal tail coefficient of 0.7 was selected as it represents aircraft of similar configuration. The V-tail was then constructed using these areas as the projected areas of the V-tail, where the tail dihedral angle is given below in Equation 1.
[pic] (1)
As the components of the UAV were pre-selected, the weight of each component was known. Using the volume of each component, a CATIA model of the fuselage was created. Each component was modeled individually, and added together to form a final model. Using the structure and skin area from the CATIA model, the weights of each structural component was found using the density of carbon fiber fabric. Using the initial take off weight, the wing and tail areas can be predicted. These areas were used to find the new weight of the aircraft. This process was iterated until the predicted weight and initial weight converged. Figure 14 shows this process.
As the components of the UAV were pre-selected, the weight of each component was known. Using the structure’s volume and skin fuselage that was found from the CATIA model, the weight of this fuselage with weight of wing and tail can be found, and the sum of the entire component yield the take off weight of the UAV. Compare this to initial guess weight, and the calculation was continued iteratively until the guess and the predicted take off weight converge. Figure 14 shows this process schematically.
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1. Carpet Plot
From tThe sizing code, also produces the carpet plot can be foundcomparing the aircraft characteristics. Since the motor fuel cell was pre-selected, the power available is fixed, so the power to weight ratio is meaningless in the carpet plot. Therefore, instead of plotting the power to weight ratio, a plot of wing loading versus aspect ratio were was studied. From this,This produced the wing loading for the desired stall speed can be obtained. Figure 15 below shows the carpet plot for the UAV.
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Figure 15: Carpet Plot
From the carpet plot, it can be seen that the design point of for the wing loading of is 1.68 lbs/ft2 and with an aspect ratio of 10.5, which yields a take off weight of approximately 10.1 lbs, stall speed of 33.02 ft/s, and effective wing span of 8 ft.
6. Power Issues
The power requirement of the UAV is constrained by the fuel cellThe selected fuel cell places a maximum power available constraint on the UAV. As the pre-selected Protonex Procore has a total energy of 770 W-h available and power output of 200 W, the power requirement of the UAV cannot exceed these parameters.
In calculating the total power consumption of the fuel cell, the UAV flight regime flight segments isare divided into several categoriessegments, : takeoff, climb, accelerate, cruise, loiter, 2-g turn, cruise back, and landing. Using the assumption that the avionics will consume 3 Watts and the camera will consume 10 Watts, with propeller efficiency of 0.5 for take off and 0.7 for other segments, the approximate power requirements can be found. These requirements are given in Table 4.
Table 4: Power Requirement at Each Segement
|Phase |Power |Time (h) |Energy Consumed (W-h) |
| |Required (W) | | |
|Takeoff |191.41 |0.0333 |6.38 |
|Climb |175.80 |0.5000 |87.90 |
|Accelerate |86.52 |0.0333 |5.63 |
|Cruise |76.61 |0.5000 |38.30 |
|Loiter |76.61 |5.5000 |421.33 |
|2g Turn |178.04 |0.5000 |86.65 |
|Cruise back |76.61 |0.5000 |38.30 |
|Land |70.95 |0.0167 |1.18 |
| | |Total |685.69 |
It is important to note that the power required for each segment is below 200 W, which is the maximum power that can be produced by the fuel cellthe fuel cell can produce. With 770 W-h of energy and assuming approximately 10% of reserve for the UAV in case of emergency, the endurance of the UAV is approximately 5.5 hours.
As predicted, take off requires the most power. However, as the amount of time for take off is relatively short, the total energy required is fairly small (~1%). The same situation can also applied applies for to the climb and 2-g turn segments. On the other hand, loiter requires relatively low power at approximately 77 W-h, but . Ddue to the fact that the UAV will be loitering the mostduring most of its flight, the amount of energy required will also be the largest at 421 W-h (~60%). The detailsA basic breakdown of the power required at each segment of the UAV flight can be foundfound inon Figure 16.
Figure 16: Power Budget of the UAV
7. Aerodynamics Detail
From the concept of operations, the aerodynamic qualities of the Saker UAV must serve two major roles: facilitate a hand-launched takeoff and allow for the longest possible endurance. In terms of aerodynamic parameters, these requirements translate into a high CLmax to lower the stall speed and facilitate a hand-launch, and a high L/D to allow for maximum endurance. It is important that these parameters are known as accurately as possible so as to accurately assess the performance of the aircraft and its compliance with the design requirements.
The first step in this process was the selection of the airfoil. When choosing an airfoil, the NACA series is a typical starting point as there is a lot significant amount of information about them, and they perform well for most missions; however, the NACA series of airfoils were designed for operation at a Reynolds number on the order of 1x106 and the highest Reynolds number seen by the Saker UAV is estimated to be on the order of 3x105. Because of this, data from the University of Illinois Low-Speed Airfoil Tests was used to select an airfoil. Shown below in Figure 17 is a graph of the maximum recorded L/D for several low-speed airfoils8.
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Figure 17: Maximum recorded L/D for various airfoils
Using the results in Figure 17 as a guideline for selecting an airfoil from the hundreds tested, the SD7032 appears to be the best performing airfoil of the tests. As mentioned earlier, L/D is not the only parameter of concern, Clmax is also critically important to the performance of the aircraft.
High lift devices, because of their cost and complexity, will not be used on this aircraft, so it is critical that the airfoil alone provide a high Clmax. From the results above, the SD70xx series of airfoils would be the focus of the analysis as they showed promising L/D characteristics. Shown in Figure 18 is a plot of Cl vs. α for two airfoils in the SD70xx series along with the NACA 4414 for a Reynolds number of 3x105 1. Note that this analysis was done using Xfoil.
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Figure 18: Cl vs. α for several airfoils, also shown is a comparison of the geometry of the SD7062 and the NACA 4414
From Figure 18, the SD7062 has a significantly higher Clmax than the SD7032, where the SD7062 is a thicker version of the SD7032. In the lower right of Figure 18, it can be seen that the SD7062 is quite similar in shape to the NACA4414, but the SD7062 clearly outperforms the NACA airfoil for the Reynolds number shown here. Based on its high L/D and high Clmax, the SD7062 was selected for this aircraft.
With the airfoil selected, the drag polar, a useful tool for determining performance, can then be constructed. The drag on the aircraft will be divided into two components, the parasite drag and the induced drag. The parasite drag was estimated using the component build up method, which uses a calculated skin-friction drag coefficient and a form factor to predict the drag for each component. The results of this analysis are shown in Table 5 and Figure 19, with the details shown in the Appendix7.
Table 5: Estimation of CD0 using component build up method
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Figure 19: Breakdown of the contribution of each component to the overall parasite drag
The remaining drag was determined using through the use of Xfoil to predict any separation drag at higher angles of attack, and a lifting line code to predict the drag due to the downwash off the wingtips. The induced drag on the wing is predicted for an untwisted rectangular wing. With an aspect ratio of 10.5, the performance gains from a more efficient wing design were relatively small, and it was decided that the difficulties associated with manufacturing a twisted or tapered foam core wing were not worthdid not justify the increased cost. The drag polar for the entire aircraft is then shown below in Figure 20.
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Figure 20: Drag Polar
Figure 20 also shows the loiter point, which is discussed later, as being at a CL of approximately 0.8. This corresponds to an L/D of nearly 17 and an angle of attack of approximately 5.5 degrees.
8. Performance Analysis
With knowledge of the aerodynamic characteristics, the actual performance and capabilities of the aircraft can then be determined. A typical design mission consists of multiple segments, i.e. takeoff, climb, cruise, loiter, and landing, with each segment having its own characteristics. Because of theSince continuous coverage is a requirement, the loiter segment is arguably the most important segment for this aircraft. Coincidently, the loiter segment is probably the simplest case to analyze, from which weand it can builtd upon to analyze the other flight conditions.
While loitering, the aircraft will be in steady level flight;, that is, the thrust is equal to the drag and the lift is equal to the weight. With the weight known, the thrust required and power required for steady level flight follow directly from the drag polar. Plots of the thrust required and the power required are shown below in Figures 21 and 22. Note that these plots assume sea level standard conditions.
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Figure 21: Thrust required for steady level flight
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Figure 22: Power required for steady level flight
Figures 21 and 22 provide several important pieces of information for determining loiter and cruise performance. Typically, the loiter speed of a propeller aircraft corresponds to the point for the minimum power required so as to maximize the loiter time. From Figure 21, the minimum power required occurs around 38 ft/s, which is too close to the stall speed, so, for safety, a loiter speed of 42 ft/s was chosen in order to be 30% above the stall speed.
From the thrust required curve in Figure 22, the minimum thrust required, which represents the best cruise speed for a propeller aircraft, occurs at approximately 48 ft/s. This is also the point of maximum L/D. Both figures also show the limiting velocities at loiter, with the lower limit being the stall speed at 33 ft/s and the upper limit at 67.5 ft/s due to the limits of the camera payload.
A similar analysis can be performed for climbing and accelerating situations by modifying the steady level flight condition to that shown below in Equations 1 2 and 2738,
[pic] (12)
[pic] (23)
For a given climb rate, dh/dt, and acceleration rate, dv/dt, the results from Equations 1 2 and 2 3 can help in determining the power budget for takeoff and climb where the aircraft is not in steady level flight. The engine throttle settings at takeoff and climb also follow from these results, and this is discussed in the Propulsion section.
In addition to the above conditions, maneuverability is also an important quality, especially if the aircraft is operating in an urban environment. The V-n diagram shown below in Figure 23 shows the maneuvering capabilities of the aircraft during its mission.
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Figure 23: V-n diagram
The limit in Figure 23 at 2-g’s is not a structural limit, but rather is the maximum necessary load factor determined from the initial design requirements and concept of operations. Furthermore, as can be seen in the power budget, a 2-g requires a significant amount of power and maneuvers at much higher loads would require more power than is available from the fuel cell.
9. Propulsion
1. Motor1
The aircraft motor was chosen based on the constraints of most power for the least amount of weight. To improve efficiency, characteristics such as a brushless motor and non-ferrite magnets should be usedwere taken into consideration. Because Based onf the aircraft weight, the motor selection was limited to motors with the ability to lift 10 lbs on take-off as well as low rpm motors so that a gear box was not necessaryto eliminate the need for a gearbox to slow the motor rpm to a value that could be used to turn the chosen propeller.
From the necessary characteristics, a motor study was done from available remote control aircraft motors. Through the study, the best motor for the aircraft was chosen to be the Model Motors AXI 4120/18 Gold Line. This motor is brushless and boasts uses neodymium magnets that produce larger magnetic fields than ferrite magnets and thus more torque. The motor can spin as fast as 9,000 rpm and has a maximum efficiency of 86%. The motor is applicable to aircraft weighing 2 kg to 5 kg, which encompasses the current aircraft weight. The motor can be seen in Figure 24.
[pic]
Figure 24: AXI4120/18 Goldline Engine1
2. Propeller
The propeller affects a number of different aspects of the aircraft. It affects the thrust of the aircraft, the speed of the aircraft, and the amount of power required from the fuel cell in order to fly at a specific speed. The efficiency of the propeller also has a large affect on the range and endurance of an aircraft. Due to the small size of our aircraft, as well as the desire to keep development and acquisition costs as low as possible, existing model aircraft propellers became the focus of the selection process.
There are a number of different model aircraft propellers available, varying both in geometry and material. There are several different types of materials such as wood, aluminum, fiberglass, nylon, and composite material available in the current market. Since weight is a major consideration, choosing a lightweight propeller is of major importance. The lightest propellers are made of nylon, and are very flexible, which would aid in survivability on landing. However, the efficiency of a nylon propeller is very low, and would not achieve the necessary flight performance in order to operate. Composite propellers are both lightweight and efficient, but they are not very rugged and are more expensive than most other types of propellers. Aluminum propellers are efficient, but very heavy. A fiberglass propeller is the best choice for our aircraft, as it has a good balance of efficiency, low weight, and durability.
A two bladed propeller was chosen over three- or four- bladed propellers because of the low power availability from the fuel cell. While the thrust produced by the propeller is lower, the power required is significantly less. In addition to this, a tractor-type propeller was chosen over a pusher. The reason for this is the increased efficiency of a tractor propeller over a pusher propeller, because the airflow into a tractor propeller is undisturbed, or “clean”. The airflow into a pusher propeller has been disturbed by the wing and fuselage, so the efficiency of a pusher propeller is less than that of a tractor.
In conducting trade studies and analyses of different types of propeller geometry, several variables were considered. These variables were propeller rotation speed, flight velocity, propeller pitch and diameter, thrust provided, power consumed, and propeller efficiency. Certain flight regimes and equipment place limitations on many of these variables, and from these limitations, the propeller best suited to for the UAV can be selected. The flight regimes analyzed are takeoff, climb, cruise/loiter, and a 2-g turn. Takeoff is particularly important due to the fact that the aircraft is hand-launched. Climb is important because the aircraft must increase in altitude and speed during this phase. These two phases place minimum requirements on the initial velocity and thrust needed to successfully maintain flight after the UAV is released. The cruise and loiter regime is important because the flight conditions there determine the endurance of the aircraft, which is of major importance, and the 2-g turn is important because the aircraft must maintain cruise/loiter characteristics while increasing power and thrust while turning.
Since the climb phase has a greater power and thrust requirement than the takeoff phase while maintaining the same speed, it is more restrictive for the given speed. The same is true of the 2-g turn versus straight cruise/loiter. Therefore, these two cases can be used in determining the propeller dimensions. Initially, the selection process used an estimated propeller rotational speed to calculate the dimensions. These dimension then provided a basis for obtaining the operating window for each phase of flight. The operating points then provided new operating speeds for the dimension analysis. The convergence of this iterative process supplied the final dimensions and operating speeds of the propeller.
The climb regime requires a velocity of approximately 33 ft/s, or 22 mph, and the motor operates at a rotational speed of 8300 rpm. This is a reasonable speed at which to expect the person hand-launching the aircraft to throw it. In order to maintain flight while climbing to its operational altitude, the aircraft requires slightly more than 1.8 pounds of thrust. The maximum power available for use from the fuel cell is 200 W, and there are approximately 5 W of power required to run the other onboard systems. This leaves a maximum available 195 W for use by the motor. Using plots of velocity, efficiency, thrust and power, the operating areas are obtained, and the best propeller geometry is chosen. These plots are given in Figure 25, with the design point marked on each graph. The curves on the plot are different pitch/diameter ratios. The diameter is held constant at 10 inches, limited by the geometry of the airplane. If the propeller were larger, it would strike the ground on landing.
[pic]
Figure 25a: Propeller Velocity vs. Efficiency for Climb Regime
[pic]
Figure 25b: Propeller Velocity vs. Power for Climb Regime
[pic]
Figure 25c: Propeller Velocity vs. Thrust for Climb Regime
The 2-g turn regime requires a velocity of approximately 40 ft/s, or 27 mph, and a motor rotation speed of 7500 rpm. According to the constraint analysis, the minimum thrust required to maintain the turn is 1.3 lbs. The maximum available power is the same in this case as for climb, though the aircraft requires more power during climb, so that is the limiting condition. The plots for the turn condition are given in Figure 26, with the operating point being indicated. The propeller size is limited by the takeoff condition, so the operating point simply reflects the point of operation at the previously selected pitch and diameter.
[pic]
Figure 26a: Propeller Velocity vs. Efficiency for Turn Regime
[pic]
Figure 26b: Propeller Velocity vs. Power for Turn Regime
[pic]
Figure 26c: Propeller Velocity vs. Thrust for Turn Regime
As can be seen from Figures 25 and 26, the chosen propeller has a pitch of 7 inches in addition to the 10 inch diameter. The other parameters shown in the plots are listed in Table 6 below. These include the power, thrust, and efficiency of the chosen propeller in each regime, and satisfy the requirements for successful flight. A fiberglass propeller of these dimensions is readily available from many different suppliers, costing approximately $3 – $5. Compared to the overall cost of the aircraft, this is a small amount. The low cost also enables easy replacement of any propeller that may be broken on landing.
Table 6: Propeller Dimension Information
|Propeller Info |
|Diameter |10 |in |
|Pitch |7 |in |
|Taper |0.05 | |
|Root chord |0.75 |in |
|AF |0.045 | |
The propeller operating speeds for each of these four phases was determined by plotting power and thrust versus rpm for each of the two operating speeds (33 ft/s for takeoff and climb, 40 ft/s for cruise and turn). For each phase, limiting conditions for thrust were obtained from the constraint analysis, which, coupled with the maximum available power limit, provided an operating speed window. These plots are shown in Figure 27. The exact operating point for each case corresponds to the minimum thrust required to maintain flight and the corresponding power. These operating points appear in Table 7.
[pic]
Figure 27a: Propeller Speed Operating Window for Takeoff and Climb
[pic]
Figure 27b: Propeller Speed Operating Window for Cruise/Loiter and 2-g Turn
Table 7: Propulsion Operating Points
|Propulsion Operating Points |
| |RPM |Thrust |Power (W) |Velocity |
| | |(lbs) | |(ft/s) |
|Take- off |6000 |0.82 |57 |33 |
|Climb |8300 |1.89 |154 |33 |
|Cruise |5800 |0.61 |47 |40 |
|Turn |7500 |1.32 |110 |40 |
10. Structure
Since this UAV needs to be as light as possible, due to the hand-launch design criteria, the aircraft will need to be made of materials with the highest strength to weight ratios. From the chart below, one can see that the carbon fiber and Kevlar composites have the best strength to weight ratios out of the most common aircraft construction materials.
Table 8: Mechanical Properties of Common Aircraft Materials5
|Material |Type |Tensile Strength |Density |Strength to Weight|
| | |(GN/m^2) |(g/cm^3) | |
| |T300/5208 |1.50 |1.55 |0.97 |
|Graphite(Carbon Fiber)/Epoxy |AS/3501 |1.45 |1.55 |0.94 |
| |AS4/3501-6 |2.10 |1.55 |1.35 |
|Boron/Aluminum |B/A1 2024 |1.50 |2.65 |0.57 |
|Fiberglass/Epoxy |Scotchply 1002 |1.06 |1.80 |0.59 |
|Aramid(Kevlar)/Epoxy |Kev 49/Epoxy |1.40 |1.40 |1.00 |
|Aluminum Alloys |- |.14-.62 |2.70 |.05-.23 |
|Steel |- |.34-2.10 |7.80 |.04-.27 |
|Titanium |Ti-6AL-4V |0.92 |4.46 |0.21 |
Because of its high strength to weight ratio AS4/3501-6 was chosen as the main construction material. Carbon fiber was chosen over Kevlar because it’s has the higher strength to weight ratio, it’s easier to work with, and it doesn’t absorb moisture or degrade with exposure to UV light. Carbon fiber is about 50% more expensive than Kevlar, however, the weight savings and ease of manufacturing make up for the added cost. Although carbon fiber is stronger than Kevlar, it doesn’t have the durability or toughness that Kevlar does. Future testing on prototypes will need to be conducted to determine if Kevlar will be needed on the belly of the aircraft in order to withstand the abrasion that occurs on landing.
[pic]
The fuselage structure will consist of carbon fiber honeycomb bulkheads, which are shown in Figure 28, with 2 layers of carbon fiber for a skin. The bulkheads will provide the main support for attaching all the internal payloads and avionics. Hard-points must be incorporated into the honeycomb material in order to use mechanical fasteners to attach the internal components. The skin is there to provide aerodynamics to the aircraft and to protect the internal components from the elements and debris. The wings and tail will consist of 2 layers of carbon fiber and an inner foam core for support. The tail boom will be a pre-fabricated carbon fiber tube.
The carbon fiber skins will be made by molding pre-pregnated AS4/3501-6 carbon fiber fabric over steel molds and cured in an autoclave. The fuselage will be made in 2 sections – a left and right side. The wings and tail will consist of a top and bottom half which will be epoxied together once the servos and other controllers have been installed. Next, a 2-part foam will be poured into the wings. All of the honeycomb bulkheads will be cut to shape and epoxied together. The internal components will then be installed and the removable external skin will then be attached. The removable wings can then be attached to the fuselage using a pinning system.
11. Weight Statement
The weight of the aircraft itself can be distributed into two main groups. These groups are the aircraft components and the aircraft structure. The aircraft components are made up of the avionics, camera core, fuel cell, propeller and motor. The structure is made up of the fuselage skin and structure, the wing, the tail boom and the tail wing. The total structure gives a total empty weight of 3.110 lbs, and the components have a weight of 6.967 lbs. These two weights combine for a total take off weight of 10.1 lbs. The individual weights can be seen in Table 8
Table 8: UAV Components Weight Breakdown
|Components |Weight |
|Payloads | |
|Avionics |0.727 lbs |
|Camera core |1.000 lbs |
|Fuel Cell |4.410 lbs |
|Propulsion | |
|Propeller |0.125 lbs |
|Motor |0.705 lbs |
|Total Components Weight |6.967 lbs |
|Structure | |
|Fuselage (skin and structure only) |1.020 lbs |
|Wing (Include foam and skin) |1.790 lbs |
|Tail Boom |0.100 lbs |
|Tail Wing (both) |0.200 lbs |
|Total Empty Weight |3.110 lbs |
|Total Take Off Weight |10.1 lbs |
12. Cost Analysis
The method used to analyze the cost of the UAS was different from the one offered in the Raymer text book7. The Raymer text book uses past history and fits equations for costs of materials, labor, and development. However, cost data is not readily available and studied for small UAV’s, especially those with weights much smaller than the intended target of these equations.
Because the UAV and support equipment contain a relatively small number of components, it is possible to break down the costs of each of these components to make a rough approximate of the cost of the UAS.
1. Support Equipment
The major components of the support equipment are the laptop, backpack or case, antenna and communications equipment. The prices of these components are detailed below in Table 9..
Table 9 - Pricing of Support Equipment
[pic]
The support equipment will be sold at cost since the support equipment is a one time purchase, and can be used for multiple UAV’s.
2. Aircraft
The aircraft will be the product being sold to make a profit. The UAV was broken down into components of material and labor. The material costs were estimated based on the current market price of carbon fiber. The fuel cell was estimated based on prices of current fuel cells on the market. The camera was an estimated cost. The avionics and communications board was an actual market price of the piece. The component costs are summarized in Table 10below.
Table 10 - Pricing of UAV
[pic]
This is combined with the labor to design and manufacture the UAV. The hourly rate of $86 for an engineer and $73 for manufacturer were taken from the Raymer textbook.
It was estimated the UAV would take 3 engineers working for a 6 months period to complete a full detailed design. It also was estimated it would take nearly 40 labor hours to produce the first UAV, decreasing to a time of 5 labor hours by the 1000th UAV built. This results in nearly 6,500 hours of total manufacturing time.
The estimation for the purchase of molds and tools to purchase at start up is nearly $20,000. There will also be support of 4000 hours over the course of the life span of the UAV to provide to the customers in case of failure.
Any additional replacements will be billed at costs and is not included in this cost analysis.
3. Pricing
T The above total costs for 1000 UAV’s and the labor to design and build them is roughly $12.2 million. If each UAV costs $18,000, the total profit for the sale of 1000 UAV would result in a total profit of $4.8 million.
BecauseSince the manufacturing costs are accumulated over time, the quantity sold over time was estimated. The first year was estimated to have a sale of 50 UAV’s, picking up to 250 in the fourth year, before steadying at 100 per year for the last year and sixth year of production. A plot of the aircraft sold and profit over time is given in Figure 28.
[pic]
Figure 28- Aircraft Sold and Profit over Time
13. Conclusion and Additional Work and Conclusion
The design of the a UAV that is small, light, low costcheap, and allows for rapid deployment will provide the military and law enforcement with far greater reconnaissance and surveillance capability than what is currently available. Currently, the general concept of the UAV has been determined developed along with the components of the UAV, . This including includes the camera, avionics, fuel cell, engine and the propellers. In addition, the dimensions of the UAV and the weight has have also been determined.
The current design of the UAV does satisfy with the target that was set earlier during the System Review Requirement phase. The current weight of the UAV is approximately 10.1 lbs with an endurance of greater than 5.5 hours. This high endurance is accomplished possible bythrough the use ofing fuel technology, which adds significantly toing the amount of power density available for to the UAV. The stall speed of approximately 33 ft/s is also achieved and this alloweds the UAV to be hand launched.
As for next stepfuture work, the team will further detail the structure and load analysis, including the effect of an impact at landing, perform a more detailed aerodynamic analysis using CFD, tests to determine propeller efficiency, determine the control surface size and trim analysisconditions, and determine the manufacturing method.
For the prospect of the. In comparison to small-sized surveillance UAVs, this design provides endurance greater than the smallany similar-sized UAVs currently available in the market. With extremely low development and production cost, the concept of the this UAV is feasible and the concept should be studied further.
14. Reference
1. AXI4120/xx. Model motors S.R.O.
2. Clausing, Don. Total Quality Development. New York: ASME Press, 1994
3. CONTROP Precision Technologies. D-Stamp. February 28, 2007.
4. “Law Enforcement UAVs”. Aeronautics Defense System Ltd. January 25, 2007.
280&Page=1
5. Mechanics of Composite Materials and Laminates: Lecture Notes for A&AE 555, C.T. Sun, Spring 2007.
6. MicroPilot – World Leader in Small UAV Autopilot.
7. Protonex Technology Corporation. .
8. Raymer P, Daniel. Aircraft Design: A Conceptual Approach.Blacksburg, Virginaia: AIAA Education Series. 2006
9. Selig, M. “UIUC Airfoil Data Site,” [retrieved April 2007].
10. “Unmanned Aerial Systems Roadmap 2005-2030,” Department of Defense, August
2005, pp 39-51.
11. Unmanned Aerial Vehicle (UAVs) – Military Aircraft . March 1, 2007. March 01, 2007
12. US Composites. Carbon Fiber. . March 20, 2007
1. AXI4120/xx. Model motors S.R.O.
2. Clausing, Don. Total Quality Development. New York: ASME Press, 1994
3. CONTROP Precision Technologies. D-Stamp. February 28, 2007.
4. “Law Enforcement UAVs”. Aeronautics Defense System Ltd. January 25, 2007.
280&Page=1
5. MicroPilot – World Leader in Small UAV Autopilot.
6. Protonex Technology Corporation. .
7. Raymer P, Daniel. Aircraft Design: A Conceptual Approach.Blacksburg, Virginaia: AIAA Education Series. 2006
8. Selig, M. “UIUC Airfoil Data Site,” [retrieved April 2007].
9. “Unmanned Aerial Systems Roadmap 2005-2030,” Department of Defense, August
2005, pp 39-51.
10. Unmanned Aerial Vehicle (UAVs) – Military Aircraft . March 1, 2007. March 01, 2007
15. Appendix
Appendix I: QFD
Appendix II: UAV Database
Appendix III: Prediction of CD0 using the component build up method2:
[pic]
where the subscribt c refers to each component.
The Cfc is determined from the following equations:
Laminar: [pic]
Turbulent: [pic]
In addition, define the parameter xc which describes the percentage of the component that experiences laminar flow.
The form factor (FF) is defined for each component as follows:
Wing, tail, strut, and pylon
[pic]
Fuselage
[pic]
where [pic]
These parameters are summarized below for each component. Note that a factor of 25% was added to the fuselage form factor to account for its somewhat unconventional shape. It was also assumed that the flow was turbulent over each component due to the propeller wake.
| |xc |Cfc |FFc |Swetc (ft^2) |
|Tail |0 |0.0063 |0.8016 |3.45 |
|Fuselage |0 |0.0047 |1.921 |4.0 |
|Tail Boom |0 |0.0047 |1.18 |0.40 |
The total parasite drag is then summarized in Table 5, where the wing CD0 was found from Xfoil. CDmisc and CDL&P were assumed to be zero.[pic]
-----------------------
Fuel Cell
CONTROP UAV Camera
Wing
Goldline Motor
10"x7" pitch Propeller
Tail Boom
Tail Surfaces
GPS and Com. board
Descend and Land
Return
Loiter
Cruise up to 5 miles
Climb to 1000 ft AGL
Hand-Launched Takeoff
Parasite Drag
Induced Drag
Loiter Point
Cruise Speed
Payload Limit
Stall Speed
Payload Limit
Stall Speed
Loiter Speed
Payload Limit
[?] |45tvyŠ’“š› ! 2 4 öíØöÌű© ”©‰±¹…zV99??hà=‡hsJƒhblLoiter Speed
Stall Line
Figure 14: Flowchart of the Aircraft Sizing
Figure 28: Internal Structural Layout
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