Project Erinyes AUVSI Student UAV Competition

Project Erinyes AUVSI Student UAV Competition Overall Design Team Aerospace Team Matt Derginer John Malaney Ryan Siffring Lucas Siron Adam Warden Avionics Team Jeremy Blackburn Cheick Guey Ayman Sheik Computer Programming Team Josh Farmer Saint Louis University Parks College of Engineering and Aviation Department of Aerospace and Mechanical Engineering

Abstract The Project Erinyes team was assembled in late spring of 2003. The mission of the team has been to design, build, and test an Unmanned Combat Aerial Vehicle (UCAV). The conceptual UCAV will travel a total of 2,500 nautical miles (nm); 1,250 nm to location, loiter for approximately six (6) hours, drop a weapons payload of 8,000 pounds, and 1,250 nm return to base. Moving from this conceptual design a 1/8 th scale prototype was then built. The 1/8 th scale prototype was scaled to simulate the aerodynamic and stability and control characteristics of the conceptual. Project Erinyes was completed in conjunction with a capstone design course at Parks College, Saint Louis University. The design course outlines the process and the necessary steps in designing an aircraft. Starting with basic analysis, the process gets more refined as it progresses toward the final design. Basic steps begin with early research and investigation of MILSPECS, along with the fundamental design work. Eventually the team works towards more refined methods and wind tunnel testing. The avionics package for the control of the aircraft was also designed in conjunction with a similar design course in avionics engineering at the university. The utmost time and safety have been put in through construction and into the flight testing phases of development. The prototyped aircraft will be put through a rigid flight test program taking small steps to ensure the flight worthiness of the aircraft and to prove the original mission concepts of the conceptual aircraft. Project Erinyes Parks College of Saint University 1

Table of Contents Section Page Introduction 3 I Background and Conceptual Aircraft Characteristics 4-8 II Features and Performance of Prototype Aircraft 9-10 III Aircraft Construction 11-12 IV Flight Testing of Aircraft 13-14 V Avionics Package 15-17 Conclusion 18 References 19 Appendix A 20 Project Erinyes Parks College of Saint University 2

Introduction The purpose of this report is to outline Project Erinyes and its development. Beginning with a brief introduction to the preliminary research that has been done and then into the construction and testing of the aircraft. The first section contains the background of the project. In the background there is also the conceptual design goals and characteristics for the overall project. Section two then discusses the performance goals of the prototype model aircraft, from intended runway length to fuel usage. The second section also outlines some of the interesting features in the design of the aircraft. Next, the third section goes into how the prototype aircraft was constructed, including safety measures and considerations. The fourth section then outlines the flight testing of the aircraft and all considerations in that process. The last section then discusses avionics package and how the model aircraft will be controlled. Wind Tunnel Testing Project Erinyes Parks College of Saint University 3

I Background and Conceptual Aircraft Characteristics Background In Greek Mythology, Erinyes were a team of three goddesses. These three goddesses were equipped with wings made of brass and they flew around the countryside in search of people that had done evil. When they caught a culprit, they would punish the person for the crimes they had committed. Project Erinyes began in June of 2003. It was then that the preliminary mission objectives of the aircraft were laid out. The conceptual UCAV would be capable of longrange flight, and cruising at high subsonic speed. It would be able to loiter over enemy territory to allow real-time intelligence to be passed back to the base. If no potential targets are found, the aircraft simply returns home. However, if an enemy target is located a weapons package could be delivered. Why is such a mission necessary in modern times? Currently, reconnaissance and attack are two separate phases of an offensive. With modern advances in technology, this is an inefficient process both for time and cost reasons. Two aircraft are required which means two planes to buy, fuel, maintain, and man. Two missions are required, which means extra preparation, flight planning, and time. It is the overall goal of Project Erinyes to eliminate this inefficiency. Research Research began by comparing the mission objectives of the UCAV to existing aircraft. This is not a full size bomber like the B-1, B-2, or B-52 because it is not nearly as large as these aircraft and cannot carry as large of a payload. The current trend in unmanned aircraft is small, extremely long-range, and single mission aircraft. The conceptual aircraft will be larger than and carry a payload more than double any current UCAV. The focus of this project is to achieve similar flying characteristics of current unmanned aircraft, yet incorporate some of the capabilities of the larger bombers. After other aircraft were researched, the next phase was to develop an RFP. It was determined that the desired range would be 2,500 nautical miles, the desired cruising altitude would be 60,000 feet, and the cruise Mach number would be 0.8. The aircraft would be capable of loitering over the target area for six hours. It would carry 8,000 lbs of payload (the equivalent of four J-DAM s) more than double the existing payload of any UAV. The aircraft would be a completely composite structure. Also, it would have twin engines in order to increase aircraft reliability. The preliminary mission profile for Project Erinyes can be seen below in Figure 1.2. Project Erinyes Parks College of Saint University 4

Figure 1.2 Preliminary Mission Profile After the mission profile was determined, the existing aircraft data was revisited this time in an effort to determine typical aspect ratios, wingspans, takeoff lengths, cruise speeds, and fuel consumptions. The estimated aircraft specifications and characteristics to achieve this mission profile will be discussed, as well as takeoff analysis, engine requirements, and cruising conditions. Conceptual Design Performance The preliminary weight sizing and performance characteristics were determined by crude methods. However, this information is absolutely pertinent to begin preliminary design work. Starting with the basics of the mission profile and design goals, a step-bystep process is followed to design the rest of the aircraft. As the process moves along, different aspects of the design are modified with more accurate methods and some parameters change as it is decided they may not be pertinent or attainable. First the maximum takeoff weights were determined under four different primary constraints: the range, loiter endurance, and cruise speed were all kept the same but the type of engine, high bypass versus low bypass, was varied and thus their relative fuel burn rates were also varied for calculations and second varying the engine type and the ultimate payload amount for the weight drop. Once the preliminary weight was determined, the next objective was to set up the constraint analysis for the desired design objectives including the variation of the cruise speed and/or aspect ratio, lift to drag ratio, as well as the runway or landing length. Moving from the preliminary methods into the exact methods other performance characteristics were found. Using a drag polar that broke down the aircraft shell into components, the actual take off and landing field lengths could be found. For the conceptual aircraft they were found to be 5500 ft and 5050 ft respectively. From field lengths into climb the aircraft was found to be able to climb at an average rate of 4800 fpm and reach altitude in just under 12 minutes. The service ceiling overall for the aircraft was 58,000 ft with a one engine inoperable ceiling at 42,000 ft under optimum speed conditions. All original design goals for the conceptual aircraft including range, cruise speeds, and ceiling conditions were met. However in calculations it was found that aircraft could not maintain the six hour loiter because of fuel consumption problems. In turn the loiter time had to be reduced to three hours, which is still better than most UAV s in use today. A typical mission profile can be found in Figure 1.3 below. Project Erinyes Parks College of Saint University 5

Segment Time Distance (nm) Fuel (lbs) Weight (lbs) Taxi Out 6 min - 35.50 27264.50 Take Off 2 min - 151.15 27113.35 Climb 11.60 min 72.5 514.85 26598.5 Cruise 2.72 hr 1177.5 2524.81 24073.69 Loiter 3 hr - 2180.00 21893.69 Payload Drop - - - 13893.69 Climb out 9.19 min 57.47 141.09 13752.60 Return 2.78 hr 1192.53 1321.31 12431.29 Approach 5 min - 98.56 12332.73 Totals 9.06 2603 6967.27 hr nm lbs Reserve 332.73 lbs Figure 1.3 Typical Mission Breakdown and block performance. Aerodynamics The majority of wing design has been discussed on a conceptual and preliminary basis. Based on certain characteristics of the aircraft (i.e. high Mach number, long range, and heavy lift) information can be decided upon on a conceptual level. Many different wing/fuselage setups were discussed in the preliminary phase of this design. Swept back wing at the rear of the aircraft, swept back flying wing, swept forward wing, and wing canard combinations were the heavily discussed configurations. Figure 1.4: Initial wing layouts discussed (From the top left moving clockwise) Aft Swept Flying Wing, Aft Swept Canard, Aft Swept Wing, and Forward Swept Flying Wing Project Erinyes Parks College of Saint University 6

Initially, based on the design objectives and desired flying qualities, it seemed as though a flying wing was the best approach. From there, a great deal of research was put into the decision to sweep the wing forward, or backwards. Instead of making a conceptual decision at this time, it was decided to end the conceptual phase of wing design, and let the calculations drive the rest of the wing parameters. However, for a flying wing, it was known that we would need a reflexive airfoil. From the constraint analysis, a design point was selected which gave the initial values of wing loading, thrust-to-weight ratio, take-off ground roll, and take-off lift coefficient. Using the values that were found during the initial weight sizing, a wing area was set and the actual wing layout could start. From the wing area (326 ft 2 ) and AR (10), a wingspan of 57 ft was calculated. The taper ratio was decided to be 0.33, suited for the Mach number and sweep angle. The NACA 64A410 has a max sectional lift coefficient of 1.6. Also, from the constraint analysis, the aircraft needs a C LTO of 2.25 to have a 6000-foot ground roll. After transforming the two dimensional lift of the airfoil to the three dimensional lift of the wing, it was found that to attain the set take-off lift coefficient, an increment of 2.23 must be achieved. From research on high lift devices, a flap system alone will not suffice, but a combination of flaps and leading-edge slats will be needed. While this combination of high lift devices makes the wing more complicated, the necessity to keep the design point where it is, 6000 ft ground roll, was important so that the aircraft could use as many runways as possible. Thirty degrees of fowler flaps and leading edge slats did not appeal to the designers. To allow the aircraft to not need leading edge slats, the constraint analysis was again consulted. Instead of using C LTO and calculating an appropriate flap setting, it was decided to choose a flap setting and calculate a new C LTO and a new takeoff velocity. This could be done because the engines would have a T/W = 0.50. It was found that with a set flap deflection of 10 degrees with a plain flap at take-off, a take-off speed of 148 knots was needed. This is much higher than the previously determined take-off speed of 110 knots, but obtainable. Also from the constraint analysis a landing lift coefficient was determined for a field length of 6000 feet. The constraint gave the value for landing lift coefficient to be 2.016 which through the analysis of a plain flap gave a deflection of just under 70 degrees. Although high, typical values for flap deflection at landing can fall between 60 and 70 degrees; therefore, this calculated value was acceptable. Structures The mission of Project Erinyes requires a strong structure capable of withstanding high speeds and heavy loads. Also, since this aircraft is being designed primarily for military use, it will need to be durable to withstand a high-cycle routine and harsh flying conditions. The FSW design increases the structural complications of the aircraft. On top of all of that, the key to the structure of the aircraft, is that it needs to be lightweight so that the aircraft will be fuel efficient. In the past few years, advanced composite design technology has grown exponentially. Due to these advancements and by employing the use of a mostly composite structure, the UCAV will in fact be able to attain its design goals. Project Erinyes Parks College of Saint University 7

Since the aircraft has a forward swept swing (FSW), aeroelastic calculations will be very important to the design. Forward swept wings have a lower divergence speed than unswept wings. For aft-swept wings, divergence is not a problem, but flutter is. For FSW s, the divergence speed is generally lower than or near the flutter speed; therefore, divergence is the main problem. Divergence is a condition where the aerodynamic moment exceeds the structural restoring moment and causes the wing to fail. As the aircraft s maximum cruise speed is Mach 0.8, structural divergence should not be an issue for the UCAV. However, divergence will still be taken into consideration during the structural design process to ensure that stability and control are not sacrificed. The combination of aeroelastic tailoring of the wing ribs and configuring the matrix of the composite wing skin specific to forces in the direction of divergence, aeroelastic divergence of the wing should not occur and stability should not be affected. Stability and Control A very stable platform is a design characteristic of bombers and surveillance planes. This is because if using an unguided bomb, the vehicle must be at a very stable position so that the necessary calculations can be made to hit the target. In surveillance aircraft, the cameras and other sensitive devices require it to be stable. Vertical tail and rudder are not being used at this time. The goal of a deep strike bomber includes some stealth capabilities. To control yaw, drag rudders (ailerons and spoilers) are used similar to the recent Boeing concepts of crowing. The design of the tail on this aircraft is very unique. First, an all-moving tail is used to reduce the volume coefficient. Next, a computerized active flight control system allows for a further reduction of this value. In the conceptual design of the aircraft, a percentage of the engine bi-pass exhaust will be ducted to blow on the tail surface. This will create a blown flap configuration which research estimated an even greater reduction in the tail volume coefficient. The blown tail configuration restricts the moment arm and tail span due to the fuselage shape. The static margin of the aircraft was then calculated after the component weight build-up of the aircraft. The c.g. position of aircraft structural parts was established from the 3-dimensional drawing. Many parameters were fixed such as engine, tail, wing, fuel in the wing c.g. locations. The exact location of the payload, fuselage fuel tank, and surveillance equipment could be tailored to meet positive static margin. The aircraft was evaluated at from sea level V stall to cruise conditions with different load weights to determine the range of static margin values for the entire flight envelope. The static margin was calculated to be from 9% to 20%. Figure 1.5 illustrates the most forward and aft cg locations with neutral and maneuver points. Figure 1.5 C.G. Travel. Project Erinyes Parks College of Saint University 8

II Features and Performance of Prototype Aircraft Introduction As the conceptual design is seen to be innovative, the model is no different. Several of the characteristics of the aircrafts design were kept for the scaled model so that it could be a close simulation of the actual conceptual aircraft. Along with this simulation it was also helpful to perform theoretical performance calculations on the scaled model aircraft to get an idea of its capabilities. From these the drag polar, max velocity, take off and landing field lengths, as well as the cruise condition characteristics could all be found. Using the most applicable equations, these equations give the basis that will be tested in flight testing. Design Features As noted before in the first section of this paper the control surfaces of the aircraft were discussed in relation to the stability and control characteristics. To again keep within the lines of the original design a split aileron type system is used for lateral control. There is no vertical control surface on the model just as there is none on the conceptual. Instead a spoiler and aileron control system is used to perform a crowing type maneuver just as a vertical tail surface would. Calculations show the scaled spoiler and aileron will have enough control power to ensure the stability and maneuverability of the aircraft. For turning maneuvers if the ailerons fail, the flaps can be used as an inboard aileron to provide roll control. Another feature of the conceptual design that needed to be tested is the blown horizontal tail. In front of the tail there is a 9.6 volt ducted fan system that brings in air from the side of the aircraft and blows it over the tail. As the tail is slightly lower than the upper surface of the aircraft the blown effect is necessary because the airflow of the upper surface would not be enough to provide the proper control power. With the ducted fan system there is more than enough air movement over the tail to provide it with the necessary power to ensure longitudinal stability of the aircraft. This is similar to the ducted bypass air in the conceptual design, and is necessary for the same reasons. Throughout the construction (discussed in section III) and scaling of the model aircraft safety has always been a consideration. As will be discussed in the avionics and flight testing portions of this paper, the proper redundancies are built in to ensure that the aircraft is safe. A complete autonomous system failure will not lead to an aircraft crash due to its redundant control systems. During construction, the same considerations were taken to ensure that the aircraft would be able to survive a hard landing in case of an emergency. By performing each stage of design, construction, and flight testing with safety in mind, it ensures that the aircraft can not only perform as it is expected but that it is also reliable. All applicable AMA regulations were taken into consideration throughout the construction of the prototype, while still keeping the core design intact. Project Erinyes Parks College of Saint University 9

Performance Characteristics Drag Buildup In determining an estimation of the expected performance characteristics of the scaled aircraft, a drag buildup and drag polar were the first parameters calculated. Keeping the aspect ratio the same as the conceptual design to ensure the aerodynamics could be properly simulated, the wing planform area was found as 5.625 sqft. Using this and an estimation of the body surfaces areas from CAD drawings, a drag polar was found for the aircraft as it would be in a cruise, clean configuration. This method is the same as outlined in Shevell (9). From this, the drag at take off and landing with the certain flap settings and landing gear effects could be estimated. At cruise, the drag was found to be 0.043 and at take off and landing the drag was found to be 0.092 and 0.462 respectively. As the aircraft is small and slow, this drag data is within an acceptable estimation range. Take off and Landing From the drag buildups found at take off and landing, the proper field lengths could then be estimated. This is absolutely necessary so that a proper runway is used through flight testing. The take off field length was found as the length necessary for the aircraft to begin rotation and then transition into a climb. It was found that the aircraft would rotate at just under 300ft and then would be into the climb within 440ft. As well, it was also found that within 770ft the aircraft would be able to clear a 50ft obstacle. For landing the distance was found as the distance after the aircraft is flared and the wheels have touched the runway surface. Once flared it was estimated the aircraft would come to rest in 430ft. If the approach is also included, the distance is closer to 1000ft as the aircraft would need a descent distance to burn off the excess airspeed. Cruise and Fuel Performance From engine wind tunnel testing, it was found that the turbine engines used would provide 12lbs of thrust each and that at full throttle they would use approximately 6oz of fuel per minute. Using this information it was found that at least 3 gallons of fuel would be necessary for flight, to provide at least 30 minutes of flight time at full throttle. Though the aircraft will not always be flown at full throttle this was seen as the worst case scenario and thus, to be able to have more than enough fuel for flight, it was used as the standard. Using the given thrust and assuming a constant cruise at sea level, the max velocity of the aircraft was estimated to be 120 knots or 140 mph. Using equations from Hale (3) the best range and best range velocity were found for the aircraft. The best range velocity was found at 68 knots, which would then allow for a total travel distance of just under 200 nm. This is found to far exceed any expectancies of the aircraft, and will be used as an estimation to be proven in flight testing. Though these equations take into consideration the high thrust specific fuel consumption of the aircraft they are still estimations. In turn the actual performance may be hindered by fuel consumption of the engines when the aircraft is fully loaded with its entire payload. It may also be impossible to fly at best range velocity in level flight due to minimum engine thrust. Project Erinyes Parks College of Saint University 10

III Aircraft Construction Introduction Fabrication of the aircraft was done using primarily composite materials with honeycomb composite substructure. The lay-up methods used were developed last year and slightly modified to suit the complex construction of the aircraft. Comparable to industry plug/mold procedures in many ways, the major difference with the procedure is absence of a heat curing stage. Instead, the single stage wet lay-ups are cured at room temperature in a vacuum bag at approximately 10 psi. The use of a wet lay-up procedure was necessary due mostly in part to the absence of a large-scale autoclave. Construction Procedure Using Rhinoceros CAD software, all the components of the aircraft were modeled and laid out, as they would be during construction. From these drawings, full size templates for the exterior skins of the aircraft were created and used to cut out foam representations of the components. The foam model is inspected and cleared of surface imperfections and then laid-up using bi-directional S-type fiberglass. Using composite fillers, these lay-ups eventually become the male plugs that are used to create the female molds. The plugs are sanded and painted repeatedly and then waxed to create the desired skin texture, in this case as smooth as possible. Once the plug is prepped, it is coated with mold release and using a combination of deck cloth, S-type glass, and glass matting a rigid female mold of the component is created. Both the plug and the mold are cured at room temperature and 10 psi. The primary material used for final component construction is 5.2 oz bi-directional carbon fiber. Carbon fiber was chosen for its lightweight and highly structural material properties. Though each lay-up was tailored differently, all aircraft skins were created from a fiberglass mold, which allows for repeatability in case of a layup error or future need for repair. Fuselage Details The fuselage was created using 3 separate molds. First, fuselage halves were made which overlapped an imaginary centerline drawn from nose to tail. Once the molds were created, the plug was sliced into sections in order to create exact templates for bulkhead formers. The formers were traced onto foam, cut out, and laid-up using a 45 x 45 ply setup. Bulkheads were then bonded to one of the fuselage halves. Once complete, the other half was bonded to the assembly and the center joint was cut, bonded, and filled. The removable bottom panels were created using the third mold. All fuselage skins were manufactured using a 0 x 45 x 90 carbon fiber ply arrangement and reinforced with fiberglass or Kevlar in areas where metal fasteners would be attached. Project Erinyes Parks College of Saint University 11

Wing Details Advanced structural tailoring was used in designing the structure for the wing in order to withstand specific forces associated with swept forward wings. The rib pattern is not a standard setup in which the ribs are aligned with the air flow; instead they cross at 90 angles with the joints aligned on the elastic axis of the wing. This is the axis of rotation at which any type of torsional divergence would take place. By arranging the ribs in this manor, the wing is most ridged at the elastic axis, which is favorable to withstand divergence. The main bending structure in the wing is the ¼ chord spar. The spar is an H-boom design created using 4 layers of carbon fiber and two layers of Kevlar. The carbon plies are arranged in a 0 x 45 x 45 x 90 pattern with the two layers of Kevlar between the 45 layers of carbon. The carbon fiber provides both the desired bending and torsional stiffness values required for the swept forward design. Necessary bending and torsional stiffness values were attained using a method developed by Dr. Marty Ferman of Parks College. The method outputs stiffness values based on wing geometry, atmospheric conditions, and flight speed envelope. The addition of Kevlar to the spar provides the spar with the ability to flex slightly without failing. At the root of the wing, the spar is laminated to an aluminum C-channel with a 1/8 th inch wall thickness. This was done as a means of fastening the wing to the torque box using standard bolts so that the wings would be easily removable for aircraft transportation. Wing skins were manufactured using two molds, top and bottom. The bottom skin extends around the leading edge and joins the top skin at the ¼ chord spar. Each wing skin has two layers of carbon fiber tailored such that they are arranged in a 0 x 90 relative to the airflow over the wing. The combination of the ply tailoring in the wing skins, the arrangement of the ribs, and the rigid spar, the wing is able to withstand the complete flight envelope of the aircrafts capabilities. Figure 3.1 Aircraft Wing and Body Skins Project Erinyes Parks College of Saint University 12

IV Flight Testing of Aircraft Introduction Due to the complexity of the avionics package and the unconventional design of the aircraft, a detailed flight test program was setup in order to verify all the flight characteristics of the aircraft and the ability of the avionics to control the aircraft. The flight test program was setup in three sections, each progressing to the final goal of a complete digital flight control package. In each stage a set of maneuvers were performed to better verify the ability of the aircraft and flight systems. Flight Maneuvers Before actually flying the aircraft, it was decided to take the flight portion of the testing one step at a time. This means that the aircraft will not be advanced in the program until it successfully passes the previous maneuver. By setting the program up this way it is possible to work out any sources of error prior to allowing the aircraft to perform an entire flight. The break down of the flight maneuvers is nearly identical to those of any full-scale aircraft. The first maneuver is a simple high-speed taxi to determine the distance needed for takeoff compared to the theoretical values found in the performance section. The second maneuver is to accelerated the aircraft to takeoff speed, lift off the ground, and immediately set back down. This maneuver accurately outputs the stall speed of the aircraft which is very crucial because it instantly determines if the aircraft will actually fly. Once the takeoff maneuver is successful, the aircraft will then be allowed to takeoff and fly straight and level for a given time before landing. Once the aircraft is determined to be flight worthy, the next maneuvers will incorporate turns. The first turn maneuver will be done by performing slow/gentle S-turns to further the understanding the lateral control of the aircraft. The second turn maneuver and final flight maneuver is to put the aircraft through a full 360 turn before landing. This will entail takeoff, climbing to a given altitude, making several turns in both directions, and then landing. By taking these maneuvers one at a time, the aircraft can be evaluated multiple times for each maneuver and minor changes made to improve the flight characteristics. Once the aircraft is proven flight worthy, numerous flight test procedures will be used to verify aircraft performance characteristics. Project Erinyes Parks College of Saint University 13

Stage One The first phase of the program is to verify the flight characteristics of the actual aircraft. This testing will be done using analog RC servos, receiver, and transmitter. This will be the only flight test stage where a human has full control of the aircraft from takeoff to landing. The procedure of maneuvers for this stage were discussed earlier. Stage Two The second stage of the flight test program will be to integrate an off-the-shelf flight control system to the previously used RC system. A Micro-Pilot autopilot will be used in order to determine aircraft response to autonomous control. With the installation of this system s components into the aircraft, the same flight procedure will be used as outlined above and it is desired that the aircraft responds similar to stage one results. This stage is vital to the program because it will study and evaluate the response of the aircraft to nonhuman inputs. Stage Three The third and final stage is to install a custom designed and programmed flight system. This system is controlled by a laptop motherboard and a set of six digital step motors. Three students in the Aerospace Technology department and one student from the Computer Science department at Parks College undertook the design of this system. The details of this system can be seen in the avionics section of this report. Conclusion At the time of submission of this document, the flight test procedure has not been completed thus no quantitative results can be given. The testing is set to begin in mid May and the schedule will be determined based on the success of the testing one step at a time. Each of the three systems have been bench tested and have performed to their designed specs. The final phase of Project Erinyes is to integrate all of these systems into the aircraft and prepare a full presentation of the flight test results. Project Erinyes Parks College of Saint University 14

V Avionics Package The mission of Project Erinyes is to create a prototype of the conceptual design, which models the aircrafts aerodynamic characteristics and flies autonomously. In order to achieve autonomous flight, the design and construction of a flight control package is necessary. This was done by the Avionics Engineering and Computer Science divisions of Project Erinyes. Just as the aircraft was designed as part of a Capstone design course, the avionics package was designed and assembled the same way. This PC based digital system will be able to control the aircraft throughout its entire flight regime. The turbines being used are not compatible with an alternator thus the system must draw power from an on-board battery. The battery used is a lithium battery similar to that used in a modern laptop. The motherboard receives signals from both an onboard Global Positioning System (GPS) and a gyro. The GPS, if triangulated, will supply latitude, longitude, altitude, and airspeed and the gyro will be used to provide longitudinal and lateral attitude (pitch, roll, and yaw angles). Compared to the desired information in the flight code, the motherboard will be used to process the amount of error in the flight path and trajectory. From the amount of error between actual and desired flight path, standard equations of motion are programmed to determine the necessary control output to correct the error. The motherboard outputs the control movement to one of six digital control boards, which converts the control signal to a digital step command, which is then relayed to a digital step motor that performs the mechanical control movement. The block diagram for this system can be seen in Figure 5.1. All of the flight coding was done using C programming. As stated earlier, the program takes information from the GPS and gyro at a rate of 60 Hz. Every time new information is received, the program strings out the information into respective variables and determines the error between the desired flight path, which is pre programmed into the code, and the actual flight path based on the gathered information. Control outputs are then sent to one of six digital control boards, which control the nose wheel, horizontal tail, left aileron, right aileron, left spoiler, and right spoiler. Control outputs are also sent to five analog systems on board the plane. The flaps are controlled by analog servos because there deployment is based solely on airspeed and altitude. Also, in the event of a system failure, the flaps become the backup pilots lateral control devices. The engines each have an analog Engine Control Unit (ECU) which will receive signals from the motherboard. A digital to analog converter transforms the digital signal to an analog current. The final analog system is the blown tail device. This is controlled in a simple on and off manor based on airspeed. At slow airspeeds the fan will be on to provide tail power, at high speeds the fan will be off because tail power from exterior airflow is sufficient. Project Erinyes Parks College of Saint University 15

Figure 5.1 Avionics Block Diagram Project Erinyes Parks College of Saint University 16

The avionics system also features a video camera, hard drive, and a four-channel transmitter. The camera is mounted in the nose of the fuselage and used to record flight information and target locations. This information is stored on the hard drive along with the flight code. It is also desired to transmit information while the aircraft is in flight. With the four channels, live video, gyro information, GPS information, and system stability will be sent. The live video will aid in determining if the aircraft is performing as planned. The backup pilot will know to take control of the aircraft if the system stability warning goes off. Basically, if the system is set up to determine the amount of error between actual and desired position; if this error reaches a certain value, the system shuts down and transmits a system failure warning at which time the pilot assumes control until landing. Conclusion As stated in the flight test section, this is the final phase of flight testing. The success of this system largely depends on the success of prior flight testing. If the aircraft responds as calculated, very few adjustments will need to be made to the system. If undesirable aircraft characteristics are found to exist, then modifications may set back the flight testing of the custom system. In line with the step-by-step integration of system, this system will be added piece by piece to the flight controls. First, the control surface/digital system will be integrated. Once the aircraft can perform level flight and steady turns using the custom system, the analog subsystems will be added and eventually the aircraft will be aloud to fly completely autonomously. Project Erinyes Parks College of Saint University 17

Conclusion Progressing from one stage to the next through the design, into the construction, and eventually into flight testing the core basics of Project Erinyes have always stayed true. The Project Erinyes team set out to design and Unmanned Combat Aerial Vehicle that could meet and/or surpass any aircraft currently used in the combat arena today. Looking at the conceptual design characteristics this aircraft can do exactly that. The next goal of the team was to prove this idea beyond the paper. With the construction of the eighth scale prototype and through wind tunnel testing this is exactly what should be done. The scaled prototype currently is an exact replicated model of the conceptual design. The current performance estimates also show that the aircraft should be able to perform within its designed region. As well, calculations currently show that the model should be able to adequately prove the aerodynamic and stability characteristics of the conceptual aircraft. The avionics package has paralleled the progress of the model construction. Issues are expected to arise in the integration of the digital flight controls and subsystems, however flight testing will prove the mission success of Project Erinyes. Project Erinyes Parks College of Saint University 18

References 1. Abbott, Von Doenhoff. Theory of Wing Sections. Appendix IV. 1949 2. Broeren, Giruere, Fopalarathnam, Lyon, Selig. Summary of Low-Speed Airfoil Data. Volume 1, 2, 3. 1995, 1996, 1998. 3. Hale, Francis J. Aircraft Performance, Selection, and Design. 1984 4. Kuethe, Arnold M, and Chuen-Yen Chow. Foundations of Aerodynamics Fifth Edition. 1998. 5. Mattingly, Jack D, Elements of Gas Turbine Propulsion. 1996. 6. Nelson, Robert C, Flight Stability and Automatic Control Second Edition. 1998 7. Raymer, Daniel P. Aircraft Design: A Conceptual Approach. 3 rd Edition. 1999. 8. Roskam, Dr. Jan, and Dr. Chuan-Tau Edward Lan, Airplane Aerodynamics and Performance. Volume I VII. 1997 9. Shevell. Fundamentals of Flight. 2 nd Edition. 1989 10. Tau Edward Lan, Roskam Airplane Design Part II: Prelimenary Configuration Desgin and Integration of the Propulsion System. Project Erinyes Parks College of Saint University 19

Appendix A Aircraft Photos Figure A.1 Model Side View Figure A.2 Conceptual Rhinoceros Drawing of Aircraft Project Erinyes Parks College of Saint University 20