CReSIS UAV Critical Design Review: The Meridian

Transcription

1 CReSIS UAV Critical Design Review: The Meridian William Donovan University of Kansas 2335 Irving Hill Road Lawrence, KS Technical Report CReSIS TR 123 June 25, 2007 This work was supported by a grant from the National Science Foundation (#ANT ).

2 Executive Summary This report briefly describes the development of the three preliminary configuration designs proposed for the Meridian UAV. This report details the selection of the primary configuration and further, more detailed, analysis including Class II weight and Balance, Class II Stability and Control, Performance Analyses, Systems Design, Class II landing gear, structural arrangement, a manufacturing breakdown and a cost analysis. The design mission for this aircraft is to takeoff from a snow or ice runway, fly to a designated area, then use low frequency radar to perform measurements of ice sheets in Greenland and Antarctica. Three designs were developed: A Monoplane with Structurally Integrated Antennas A Monoplane with Antennas Hanging from the Wing A Biplane with Antennas Structurally Integrated Into the Lower Wing The monoplane with antennas hanging from the wing was selected as the primary configuration for further development. This report describes the Class II design and analysis of that vehicle. i

3 Acknowledgments This material is based upon work supported by the National Science Foundation under Grant No. AST Any opinions, findings, and conclusions or recommendations expressed in this material are those of the author(s) and do not necessarily reflect the views of the National Science Foundation. ii

4 Table of Contents Executive Summary... i Acknowledgments... ii Table of Contents...iii List of Figures... v List of Tables... vi Nomenclature... vii Abbreviations...viii 1 Summary of Preliminary Designs Configuration Selection and Requirement Changes Engine Selection Turboprop Variant Class II Design Class II Weight and Balance The Aircraft V-n Diagram Component Weight Estimations Class II Stability and Control Trim Diagrams Open Loop Dynamics Actuator Size and Rate Requirements Class II Aerodynamics Propulsion Performance Analysis Stall Speed Takeoff Distance Climb Cruise Performance Landing Distance Systems Flight Control System Electrical System Communications/Telemetry System Fuel System Anti-Icing System Class II Landing Gear Tire Selection Strut Sizing Landing Gear Integration Structural Arrangement Wing Structure Fuselage Structural Layout Manufacturing Breakdown Cost Analysis iii

5 Research, Development, Test, and Evaluation Costs Acquisition Cost Cost Estimate Summary Cost Estimate Justification Conclusions References iv

6 List of Figures Figure 1.1: Comparison of Fuel Usage for 3 Fine Scale Missions Figure 1.2: Combined Takeoff Weight Regression Chart Figure Vivaldi Antenna Figure 2.2: CAD Model of Innodyn 165TE Figure The Meridian UAV Figure V-n Diagram for the Meridian Figure Center of Gravity Excursion for the Meridian Figure 3.4 Component C.G. Locations Figure Trim Diagram - Cruise Figure Trim Diagram - Takeoff, Gear Down Figure Trim Diagram - Takeoff, Gear Up Figure Trim Diagram Landing Heavy, Gear Down Figure Trim Diagram - Landing Heavy, Gear Up Figure Trim Diagram - Landing Light, Gear Down Figure Trim Diagram - Landing Light, Gear Up Figure Trim Diagram - OEI Figure Drag Polars for the Meridian without Antennas Figure Lift-to-Drag for the Meridian without Antennas Figure Drag Polars for the Meridian with Antennas Figure Lift-to-Drag for the Meridian with Antennas Figure Relationship of Parasite Area and Wetted Area for Various Single Engine Aircraft [6] Figure Cloud Cap Tech. Piccolo II Autopilot [36] Figure Piccolo II Architecture [36] Figure Piccolo Ground Station and Pilot Controller (Operator Interface Not Shown) [36] Figure Electrical Load Profile for the Meridian UAV Figure Electrical System Layout Figure Fuselage Systems Layout Figure Fuel Tank Integration Figure Lancair Legacy Landing Gear Strut [37] Figure Lancair Legacy Landing Gear Installation [38] Figure Matco Tailwheel Assembly [39] Figure Wing Structural Layout Figure Fuselage Structural Layout Figure Wing-Fuselage Attachment Figure Standard 20 Foot Shipping Container Door [9] Figure Typical Engine Mount for the Innodyn 165TE Figure Manufacturing Breakdown Figure Cost Breakdown by Overall Category Figure UAV Cost in Terms of Payload Weight v

7 Figure UAV Cost Based on System Cost Versus Payload Weight List of Tables Table 1.1: Summary of Preliminary Design Concepts... 9 Table V-n Diagram Parameters Table Design Speeds and Load Factors for the Meridian Table Class II Weight and Balance for the Meridian Table Weight and Balance Summary for the Meridian Table Meridian Flight Conditions Table Dynamic Analysis Flight Conditions Table Stability Deriviatives for the Meridian Table Control Derivatives for the Meridian Table Longitudinal Transfer Functions for Cruise Table Lateral Transfer Functions for Cruise Table Directional Transfer Functions for Cruise Table Meridian Dynamic Stability Parameters Table Roll Control Requirements - Time to Achieve Bank Angle (Seconds). 37 Table Roll Control Results Table 3.15 Wing Geometry for Drag Calculations Table V-Tail Geometry for Drag Calculations Table Fuselage Geometry for Drag Calculations Table Flap and Landing Gear Geometry for Landing Gear Calculations Table Drag Analysis Results Table Resultant Oswald's Efficiency and Parasite Area for the Meridian Table Stall Speed Summary Table Landing Gear Strut Sizing Table Engineering and Manufacturing Rate Estimation Table RDT&E and Acquisition Cost Summary Table Cost Breakdown by Overall Category Table Cost Breakdown by RDT&E and Production Categories Table Current UAV Procurement Cost [40] vi

8 Nomenclature Symbol Description Units AR Aspect Ratio ~ b Wing Span ft, in c Wing chord ft, in C D Drag Coefficient ~ C D0 Zero-Lift Drag Coefficient ~ C L Lift Coefficient ~ C Lα Lift-Curve Slope Rad -1 c p Specific Fuel Consumption Lbs/hp-hr D Drag Lbs D p Propeller Diameter Ft e Oswald s Efficiency ~ f Equivalent Parasite Area Ft 2 L Lift Lbs M Munk s Span Factor ~ n p Number of Propeller Blades ~ P Engine Power hp P bl Blade Power Loading hp/ft 2 R Range Nm S Wing Area Ft 2 S Wet Wetted Area Ft 2 W E Empty Weight Lbs W F Fuel Weight Lbs W pay Payload Weight Lbs W TO Takeoff Weight Lbs Γ Dihedral Angle Deg α Angle of Attack Deg ε Wing Twist Deg η Wing Station ~ η p Propeller Efficiency ~ σ Biplane Interference Factor ~ vii

9 Abbreviations Abbreviation CReSIS FAR NSF UAV Description Center for Remote Sensing of Ice Sheets Federal Aviation Regulations National Science Foundation Uninhabited Air Vehicle viii

10 1 Summary of Preliminary Designs Three preliminary designs have previously been presented to satisfy the requirements for the CReSIS polar research mission: A monoplane with flush-mounted antennas A monoplane with hanging antennas A biplane with flush mounted antennas The three designs are summarized in Table 1.1. The fuel required to complete 3 fine-scale mission is shown in Figure 1.1 for the three designs as well as the Lockheed P-3 and the De Havilland Twin Otter. Figure 1.2 shows the aircraft plotted on the takeoff weight regression chart. Table 1.1: Summary of Preliminary Design Concepts Parameter Units Red Design White Design Blue Design Geometry Wing Area ft Wing Span ft Length Overall ft Height Overall ft Weights Takeoff Weight lbs 760 1, Empty Weight lbs Payload Weight lbs Fuel Weight lbs Performance Range nm 1,750 1,750 1,750 L/D Cr ~ Powerplant Engine ~ Rotax 912-A Rotax 914-F Rotax 914-F Power hp

11 Red Design White Design Blue Design P-3 Orion Twin Otter 1, , Fuel Required for 3 Fine Missions, gallons Figure 1.1: Comparison of Fuel Usage for 3 Fine Scale Missions 10,000 Predator B E-Hunter Empty Weight, lbs 1,000 Shadow 200 Dakota Shadow 600 I-Gnat Predator 100 log 10 (W TO ) = A + B*log 10 (W E ) A = B = ,000 10,000 Takeoff Weight, lbs Figure 1.2: Combined Takeoff Weight Regression Chart 10

12 2 Configuration Selection and Requirement Changes The three preliminary designs presented were compared based on several criteria including weight, aerodynamic performance, and antenna integration. The selection of the configuration for further study was made primarily based on antenna integration issues. While the Red and Blue designs were more efficient than the White design aerodynamically, the antenna integration methods used for these designs would limit the possible bandwidth of the radar, thereby decreasing the total system performance. The success of the CReSIS project is heavily dependent on the level of synergy between the radar and aircraft systems. Design decisions must be made based on high level, systematic concerns. L = 0.5 m H = 0.5 m T = m Figure Vivaldi Antenna 11

13 The White design was therefore selected as the configuration for further development. However, the antenna type must be changed to accommodate higher operating bandwidths. The primary antenna will be a Vivaldi antenna as shown in Figure 2.1. These are essentially flat plates that hang from the wing of the aircraft. While these have been selected as the current antenna for the radar system, it is clear that the antenna type may change again. Therefore, the White design was modified slightly such that eight mounting points will be integrated into the wing structure from which several types of antennas could be mounted. The aircraft will be designed for the Vivaldi antennas, but if another type of antenna of equal or smaller size proves to be more effective, then it may be mounted on the wing without any structural modifications. Essentially, the new design philosophy is that the aircraft is relatively insensitive to the type and size of antennas used within reason. This solution will yield a highly adaptable, high performance vehicle capable of multiple missions Engine Selection Turboprop Variant The engines selections shown for the three preliminary designs were driven by power and specific fuel consumption requirements. The reliability of these engines in a cold-weather environment is questionable. Also, from a logistics standpoint, the Rotax engines are suboptimal as the primary fuel used in Antarctica is Jet-A, not aviation gas. For these reasons, the Innodyn 165TE (Figure 2.2) has been selected for further investigation. The Innodyn is a fairly new, small turbopropeller engine that 12

14 has yet to be fully tested. Therefore, the specific fuel consumption of the engine is somewhat unknown. Figure 2.2: CAD Model of Innodyn 165TE The current specific fuel consumption estimate for the Innodyn engine is 0.70 lbs/hp-hr, which is much higher than the Rotax 912 and 914 value of 0.56 lbs/hp-hr. Nonetheless, the reliability and maintainability issues make this engine very appealing for this mission. The Innodyn has been selected as the primary engine for the Meridian pending testing that will be performed in the Fall of The Rotax 914 will be considered as a backup to the Innodyn in the case of poor test results. 13

15 Figure The Meridian UAV 14

16 3 Class II Design The purpose of this document is to expand upon the chosen aircraft configuration (Monoplane with antennas hanging from the wing) through Class II design. The new Meridian design shown in Figure 2.3 is the result of several design iterations focused on manufacturability and operational constraints. 3.1 Class II Weight and Balance The purpose of this section is to describe the Class II weight and balance performed for the Meridian UAV. This consisted of first calculating and plotting a V-n diagram to determine the limit and ultimate loading for the Meridian. The results of the V-n diagram were then used to create weight estimates for each vehicle component. Finally, a weight and balance analysis is presented to show the aircraft center of gravity travel The Aircraft V-n Diagram A V-n diagram was constructed for the Meridian UAV to help determine the maximum load factors and design speeds that will be used for structural sizing. The V-n diagram was created based on FAR 23 requirements for Normal class aircraft as there are currently no certification requirements for UAVs. The inputs to the V-n diagram creation are shown in Table

17 Table V-n Diagram Parameters Parameter Value Units Altitude 0 ft W gross 1,125 lbs S 69.6 ft 2 W/S 16.2 psf m.g.c ft Cla 3.98 rad -1 C Lmax (+) 1.3 ~ C Lmax (-) ~ C C Lmax (+) ~ C C Lmax (-) ~ The V-n diagram for the Meridian is shown in Figure 3.1. The design speeds and limit load factors are shown in Table 3.2. The positive load factor was set to 3.8 based on FAR 23 requirements [35]. The negative load factor was set to 40 percent of the positive load factor according to FAR 23 requirements [35]. Table Design Speeds and Load Factors for the Meridian Parameter Value Units V s 61 kts V C 133 kts V D 186 kts V A 118 kts V S,neg 66 kts n limit (+) 3.8 ~ n limit (-) -1.5 ~ 16

18 4 Positive g Limit = V C Gust Line +V D Gust Line Load Factor, n 1 0 V S V A V C V D -1 Negative g Limit = V D Gust Line -2 -V C Gust Line Speed, KEAS Figure V-n Diagram for the Meridian 17

19 3.2 Component Weight Estimations Four methods were used for estimating aircraft component weights: the Cessna, Torenbeek, General Dynamics, and USAF methods. These methods are integrated into the AAA software, which was used for the weight estimations [7]. These methods are designed for conventional, inhabited aircraft, which the Meridian is not. For this reason a certain amount of designer intuition was employed to select the most applicable methods for each component. For example, the Cessna method produced a wing weight of approximately 300 lbs, while the USAF and Torenbeek methods resulted in weights of approximately 100 lbs. The latter results were deemed to be reasonable, therefore the Cessna method was not used for the wing weight estimation. Table 3.3 shows the component weights as well as the methods used. The data shown in Table 3.3 are the result of several iterations. 18

20 Table Class II Weight and Balance for the Meridian Method Class II Weight X CG Y CG Z CG lbs Structure Wing USAF, Torenbeek Empennage Cessna, USAF Fuselage Cessna, USAF Landing Gear Main Gear Cessna Tail Wheel Cessna Main Gear - Retracted Cessna Tail Wheel - Retracted Cessna Propulsion Propeller Torenbeek/GD Engine Manufacturer Fuel System USAF, Torenbeek Engine Systems Torenbeek/GD Fixed Equipment Flight Control System Cessna, Torenbeek Avionics/Electronics Class I/Manufacturer Electrical System Cessna, Torenbeek Icing System USAF, Torenbeek Paint Torenbeek Fuel and Payload Mission Fuel Fuel Reserves Trapped Fuel and Oil Payload Totals Structure - Gear Extended Structure - Gear Retracted Powerplant Fixed Equipment Empty Weight Useful Load Total - Gear Extended Total - Gear Retracted The c.g. locations of each component are shown in Figure 3.3. The c.g. travel due to fuel and payload loading is shown in Table 3.4 and Figure 3.2. Table Weight and Balance Summary for the Meridian Parameter Inches % mgc Most Forward c.g Most Aft c.g Total Excursion Fuel Excursion

21 Wing Chord, %mgc Retract Gear W TO Fuel Weight, lbs Extend Gear - Fuel Trapped Fuel and Oil /- Payload W OE W E Fuselage Station, in Figure Center of Gravity Excursion for the Meridian 20

22 Figure 3.3 Component C.G. Locations 21

23 3.3 Class II Stability and Control The purpose of this section is to describe the Class II stability and control analyses performed for the Meridian UAV. These include: Trim Diagram (Power on and power off) Roll Performance Crosswind Control During Final Approach and While on the Runway Open Loop Dynamic Handling Actuator Size and Rate Requirements Trim Diagrams Trim diagrams were created for the flight conditions listed in Table 3.5. The trim diagrams shown in Figure 3.4 through Figure 3.11 were created using the AAA software [7]. The V-tail incidence was adjusted to i vee = -3.5 deg so that the aircraft could be trimmed at the stall speed with the most forward center of gravity. Table Meridian Flight Conditions Flight Condition Altitude Speed Weight Flaps Gear ft kts lbs deg ~ Clean 5, Up Takeoff Gear Down ,082 0 Down Takeoff Gear Up ,082 0 Up Landing Heavy, Gear Down , Down Landing Heavy, Gear Up , Up Landing Light, Gear Down Down Landing Light, Gear Up Up OEI 5, Up 22

24 Figure Trim Diagram - Cruise 23

25 Figure Trim Diagram - Takeoff, Gear Down 24

26 Figure Trim Diagram - Takeoff, Gear Up 25

27 Figure Trim Diagram Landing Heavy, Gear Down 26

28 Figure Trim Diagram - Landing Heavy, Gear Up 27

29 Figure Trim Diagram - Landing Light, Gear Down 28

30 Figure Trim Diagram - Landing Light, Gear Up 29

31 Figure Trim Diagram - OEI 30

32 The trim diagrams shown in Figure 3.4 through Figure 3.11 show that the aircraft can be trimmed throughout the entire flight envelope requiring no more than 20 degrees of control surface deflection Open Loop Dynamics The open loop dynamics were calculated for the Meridian using the AAA program [7]. The longitudinal and lateral-directional dynamics and flying qualities were calculated for the takeoff, cruise, and approach flight conditions as specified in Table 3.6. These were compared to the flying quality requirements specified in MIL-F- 8785C [6] and MIL-STD-1797A [6] for a Class I aircraft. While it is not necessary for a UAV to meet the military specifications, it is common design practice to use the military flying quality requirements as a basis for the dynamic analysis. Table Dynamic Analysis Flight Conditions Parameter Units Flight Condition Takeoff Cruise Approach Altitude ft ΔT deg F U 1 kts W lbs α deg C L1 ~ n g δ F deg X cg in Z cg in ε vee deg η vee, p. off ~ η vee ~

33 The stability and control derivatives for Meridian are shown in Table 3.7 and Table 3.8, respectively. These values represent the result of several iterations of control surface sizing. 32

34 Table Stability Deriviatives for the Meridian Parameter Units Flight Condition Takeoff Cruise Approach C Tx ~ C MT ~ C Du ~ C Lu ~ C Mu ~ C TXu ~ C MTu ~ C Dα rad C Lα rad C Mα rad C MTα rad C Dα rad C Lα rad C Mα rad C Dq rad C Lq rad C Mq rad C Yβ rad C lβ rad C nβ rad C ntβ rad C Yβ rad C lβ rad C nβ rad C YP rad C lp rad C np rad C Yr rad C lr rad C nr rad C Divee rad C Livee rad C mivee rad

35 Table Control Derivatives for the Meridian Parameter Units Flight Condition Takeoff Cruise Approach C Dδrv rad C Lδrv0 rad C Lδrv rad C Mδrv0 rad C Mδrv rad C hβrv rad C hδrv rad C Yδa rad C lδa rad C nδa rad C hαa rad C hδa rad C L0 ~ C L0 ~ C M0 ~ C m0wf ~ The stability and control derivatives shown in Table 3.7 and Table 3.8 were used to calculate the open loop transfer functions for the Meridian. This was done with the AAA software [7]. The transfer functions for the Cruise condition are shown in Table 3.9 through Table The dynamic stability parameters related to the aircraft flying qualities are shown in Table The Meridian met Level I flying quality requirements for all flight conditions. 34

36 Table Longitudinal Transfer Functions for Cruise Table Lateral Transfer Functions for Cruise 35

37 Table Directional Transfer Functions for Cruise Table Meridian Dynamic Stability Parameters Flight Condition Parameter Units Takeoff Cruise Approach Longitudinal ω sp rad/s ζ sp ~ ω p rad/s ζ p ~ n/α g/rad CAP 1/gsec Lateral-Directional TC Spiral sec TC roll sec ω D rad/s ζ D ~ Note: Level I flying qualities met for all flight conditions. Roll Control Effectiveness The roll control effectiveness is a vital parameter for aircraft controllability, especially during approach and landing. The roll control requirements for a Class I 36

38 aircraft are specified in Table The requirement states that the aircraft must be able to achieve the specified bank angle in the specified amount of time. Table Roll Control Requirements - Time to Achieve Bank Angle (Seconds) Cat A Cat B Cat C Level φ t = 60 deg φ t = 60 deg φ t = 30 deg I II III The roll control analysis results are shown in Table These values were calculated using AAA [7]. As can be seen, the Meridian meets Level I flying qualities for all flight conditions. Table Roll Control Results Takeoff Cruise Approach Cat. C B C φ t t r Actuator Size and Rate Requirements The actuators were sized using the AAA program to calculate the hinge moment derivatives. The actuators were then sized for a factor of safety of 2.0 at maximum deflection. The most critical actuator size was determined to be the flaps, which required a servo with a maximum torque of at least 100 in-lbs. The servo selected is the Model 820 manufactured by Moog Components Group ( This servo has a peak torque of 150 in-lbs and accepts a PWM signal, which is compatible with the selected autopilot. 37

39 3.4 Class II Aerodynamics The Class II drag analysis was performed for the Meridian with the AAA software [7]. The geometries used in the drag calculations are shown in Table 3.15 through Table The drag analysis was performed for the takeoff, cruise, approach, and OEI flight conditions with and without antennas. These flight conditions are described in Table 3.5. Table 3.15 Wing Geometry for Drag Calculations Parameter Units Value S ft AR ~ 10 λ ~ 1 Λ c/4 deg 0 (t/c) r % 18 (t/c) t % 18 LER/c % 1.5 L' ~ 2 x lam /c % 10 Cl α rad f gap ~ 0.97 k sand 10-3 ft Table V-Tail Geometry for Drag Calculations Parameter Units Value S ft AR ~ 4 λ ~ 0.5 Λ c/4 deg 26.3 (t/c) r % 12 (t/c) t % 12 LER/c % 1.58 L' ~ 2 x lam /c % 10 Cl α rad f gap ~ 0.96 k sand 10-3 ft

40 Table Fuselage Geometry for Drag Calculations Parameter Units Value S b ft S wet ft L ft 14.8 S frontal ft 2 3 x lam /L % 0 S wet-lam ft 2 0 S plf ft D f max ft 24 k sand 10-3 ft Table Flap and Landing Gear Geometry for Landing Gear Calculations Parameter Units Value Flaps Flap Type ~ Plain η i f % 27 η o f % 57 c f /c w % 25 Main Gear S ft ft L strut ft 2.5 Tailwheel C Dref ~ 0.5 S ref ft F D ~ 0 39

41 Table Drag Analysis Results Component Takeoff Cruise Approach OEI Zero-Lift Drag Wing Vee Tail Fuselage Flap Retract Fixed Gear Trim Propeller Inlet Nozzle Power Gear Pod Total Zero-Lift Drag Due to Lift Wing Vee Tail Fuselage Total Due to Lift Total Drag Antenna Drag (6) Total Drag w/ Ant The drag results are shown in Table 3.19 and Figure 3.12 through Figure The mid-cruise lift-to-drag ratio was found to be 14.5 as indicated on Figure This is slightly less than the value of 16.0 estimated in the Class I design. However, the performance analysis results show that the range requirement is still met in the current configuration so no further iteration is necessary. 40

42 2 1.8 Takeoff Cruise Approach OEI 1.6 Lift Coefficient, C L Reference Data: S = 68.6 ft 2 AR = 10.0 Takeoff: W = 1,083 lbs e = 0.84 f = 2.84 ft 2 Cruise: W = 963 lbs e = 0.83 f = 1.4 ft 2 Approach: W = 843 e = 0.75 f = 3.7 ft 2 OEI: W = 963 e = 0.81 f = 3.8 ft Drag Coefficient, C D Figure Drag Polars for the Meridian without Antennas 41

43 Lift Coefficient, C L Takeoff Cruise Approach OEI Reference Data: S = 68.6 ft 2 AR = 10.0 Takeoff: W = 1,083 lbs e = 0.84 f = 2.84 ft 2 Cruise: W = 963 lbs e = 0.83 f = 1.4 ft 2 Approach: W = 843 e = 0.75 f = 3.7 ft 2 OEI: W = 963 e = 0.81 f = 3.8 ft Drag Coefficient, C D Figure Lift-to-Drag for the Meridian without Antennas 42

44 2 1.8 Takeoff Cruise Approach OEI 1.6 Lift Coefficient, C L Reference Data: S = 68.6 ft 2 AR = 10.0 Takeoff: W = 1,083 lbs e = 0.84 f = 3.13 ft 2 Cruise: W = 963 lbs e = 0.83 f = 1.65 ft 2 Approach: W = 843 e = 0.75 f = 3.9 ft 2 OEI: W = 963 e = 0.81 f = 4.13 ft Drag Coefficient, C D Figure Drag Polars for the Meridian with Antennas 43

45 Lift Coefficient, C L Takeoff Cruise Approach OEI Reference Data: S = 68.6 ft 2 AR = 10.0 Takeoff: W = 1,083 lbs e = 0.84 f = 3.13 ft 2 Cruise: W = 963 lbs e = 0.83 f = 1.65 ft 2 Approach: W = 843 e = 0.75 f = 3.9 ft 2 OEI: W = 963 e = 0.81 f = 4.13 ft Drag Coefficient, C D Figure Lift-to-Drag for the Meridian with Antennas 44

46 To verify the validity of the drag analysis the Oswald s efficiency and parasite area were calculated for each flight condition using Equation 3.1 and Equation 3.2 respectively. These values are shown on the drag polar plots in Figure 3.12 through Figure 3.15 and Table The cruise values of the parasite area with and without the antennas were plotted against known aircraft in Figure 3.16 from [6] for further verification. As can be seen, Meridian falls somewhere between the lines for overall skin friction coefficient values of C f = to depending on whether the antennas are installed or not. This is a good indication that the Class II drag results are reasonable. e = δc D 1 δc 2 L ea Equation 3.1 f = C D S 0 w Equation 3.2 Table Resultant Oswald's Efficiency and Parasite Area for the Meridian No Antennas With Antennas S wet = 240 ft 2 S wet = 275 ft 2 Flight Condition e f f ~ ft 2 ft 2 Takeoff Cruise Approach OEI

47 Figure Relationship of Parasite Area and Wetted Area for Various Single Engine Aircraft [6] 3.5 Propulsion The installed thrust of the Innodyn engine was calculated using the AAA program [7]. The Innodyn is rated at 165 SHP. The extracted power is estimated at 5 hp based on the electrical power generation required. The total installed power is 125 hp, which is above what is required for takeoff and climb performance. The AAA program was also used to calculate an estimate for the inlet area. This resulted in an inlet with an area of 0.2 ft 2. 46

48 3.6 Performance Analysis The purpose of this section is to describe the performance requirements imposed on this aircraft design and to verify that these requirements have been met by the Meridian. This includes: Stall speed Takeoff Distance Cruise Performance Landing Distance Stall Speed The stall speed of the Meridian was calculated for heavy and light flight conditions with zero and full flaps as shown in Table The maximum trimmed lift coefficients determined in Section were used for the clean and full flap configurations. Power effects were ignored for the stall speed calculations. Table Stall Speed Summary Clean Full Flaps Parameter Units Light Heavy Light Heavy Weight lbs 843 1, ,082 Flaps deg C Lmax ~ Altitude ft V s ft/s Takeoff Distance The takeoff distance for the Meridian was calculated for conventional tires as well as ski operations using the AAA program [7]. This process uses methods found in [6] 47

49 to calculate the takeoff distance to clear an obstacle of a specified height. The following assumptions were used in the takeoff distance calculations: Ground Friction Coefficient ( μ = for Tires, μ = for skis) The obstacle height is 50 ft Weight = 1,082 lbs Standard sea level conditions Drag is based on Class II drag analysis for Takeoff condition: C = 0. D V 3 /V TO = 1.3 (Based on FAR 23) The takeoff analysis resulted in the following takeoff distances: Standard Tires on Asphalt: S TO = 415 ft Skis on Snow: S TO = 635 ft The takeoff distances with and without skis both exceed the required distance of 1,500 ft by a large amount. This is due to the fact that the selected engine has more power than required by performance matching. G G Climb Cruise Performance The cruise performance calculations for the Meridian were performed with the AAA program [7]. This consisted of estimating the range and endurance assuming constant speed cruise. 48

50 Range The range of the Meridian was calculated using the constant speed range equation found in [6]. The lift-to-drag value was calculated using the mid-cruise Class II drag polar (Section 3.4). The following assumptions were used for the range calculation: W begin = 1,083 lbs W fuel = 240 lbs, W fuel, res = 55 η p = 0.80 c p = 0.90 lbs/hp-hr The range of the Meridian was determined to be: Without Antennas 940 nm (1,735 km) without reserves 1,200 nm (2,220 km) with fuel reserves With 6 Antennas 830 nm (1,530 km) without fuel reserves 1,030 nm (1,900 km) with fuel reserves The range for the Meridian without antennas is acceptably close to the required range of 1,750 km without antennas and the range with antennas exceeds the required 1,500 km. Endurance The endurance of the Meridian was calculated using Equation 3.3. The following assumptions were made for the endurance calculation: Constant speed cruise 49

51 W begin = 1,083 lbs W fuel = 240 lbs, W fuel, res = 55 lbs η p = 0.80 c p = 1.2 lbs/hp-hr (From manufacturer data) U 1 = 80 kts Drag based on mid cruise Class II drag polar E = η p L W Equation 3.3 begin ln c pu1 D Wbegin W fuel The loiter speed was set at 80 kts as this is the speed for maximum L/D. The results from the endurance calculations are: Without Antennas 12.7 hours without reserves 16.1 hours with fuel reserves With 6 Antennas 11.5 hours without fuel reserves 14.8 hours with fuel reserves The Meridian exceeds the specified endurance requirements with and without the antennas and fuel reserves Landing Distance The landing distance for the Meridian was calculated with the AAA program [7]. The landing distance includes the distance from a 50 ft obstacle to the ground and the 50

52 distance from touchdown to a full stop. The following assumptions were used for the landing gear calculation: Weight = 1,082 lbs C Lmax = 1.7 (Based on maximum trimmed lift coefficient) Drag based on Class II drag polar for the Approach flight condition Average ground acceleration = 0.45 g Δn = 0.10 (Correction factor due to pilot technique) The results of the landing distance calculation are: S air = 1,110 ft (Distance in air from obstacle to ground) S LG = 430 ft (Ground run distance) S L = 1,540 ft (Total distance) The total landing distance is acceptably close to the required landing distance of 1,500 ft. This distance is based on conventional tires, which was determined to be the critical requirement as skis actually have a higher coefficient of friction than wheels on asphalt. 3.7 Systems The purpose of this section is to describe the systems both on and off the Meridian that are required for operation. These include: Flight Controls Electrical System Communications 51

53 Fuel Anti-Icing Flight Control System The Meridian will utilize a fly-by-wire control system based around the Piccolo autopilot, which is produced by Cloud Cap Technologies [36]. The Piccolo requires dynamic and static pressure inputs and electrical power. The Piccolo interfaces with the servo actuators using a Pulse Width Modulated (PWM) signal, which is standard for remote control aircraft. The architecture for the Piccolo is shown in Figure Figure Cloud Cap Tech. Piccolo II Autopilot [36] 52

54 Figure Piccolo II Architecture [36] The ground equipment associated with the Piccolo autopilot consists of a ground station, operator interface (PC), and a pilot control unit (Futaba Controller) as shown in Figure Figure Piccolo Ground Station and Pilot Controller (Operator Interface Not Shown) [36] 53

55 3.7.2 Electrical System This section describes the electrical system layout for the Meridian including an electrical load profile for a typical mission. The Meridian will require both 12 and 24VDC power busses. The power system consists of: Electrical Generator Battery Electrical Bus Electrical Wiring The first step in developing the electrical system layout was to generate an electrical load profile for the Meridian. This was done by listing all necessary systems required during each phase of a given flight. The results of the load profile are shown in Figure The total load was estimated assuming the radar system is turned on at takeoff, while the essential load assumes the radar system only requires power during the on-station flight phase. The most critical flight phases are the takeoff and landing segments as the landing gear and flap actuators will be operated in addition to the other systems. The emergency flight phase is representative of an engine flame-out situation. The battery was sized such that all necessary systems could remain operating will the aircraft descends and attempts tot restart the engine. This however, will require the ability to turn some systems off autonomously. The current configuration of the Innodyn engine is with one 600W, 12 V generator and a separate starter. The current generator is a standard off-the-shelf automotive 54

56 alternator, and can be replaced with a larger generator for the Meridian. The electrical load profile indicates that a 1,000 W generator would be sufficient. The electrical system layout is shown in Figure The wiring is not shown in Figure 3.21 for clarity. The aileron and flap servo and landing gear actuator wiring will be located just aft of the aft spar. The antenna wiring will be located just behind the forward spar. A more detailed view of the systems located in the fuselage are shown in Figure

57 Min 5 Min 5 Min 5 Min 20 Min 105 Min 540 Min 105 Min 15 Min 5 Min 10 Min Total Load Essential Load Electrical Power, W Loading Start Taxi Takeoff Climb Cruise Out On Station Cruise In Descent Land Emergency Flight Segment Figure Electrical Load Profile for the Meridian UAV 56

58 Figure Electrical System Layout 57

59 Figure Fuselage Systems Layout Communications/Telemetry System The Meridian will utilize dual line-of-site communication links: the piccolo communications will be used for command and control and a secondary communications link will be used for vehicle health monitoring/telemetry. For beyond line-of-site (BLOS) communications, an Iridium satellite communication link will be utilized. The Piccolo autopilot is configured to transmit and receive data over an Iridium link. This communications link will be used for low-bandwidth health monitoring and limited control. 58

60 3.7.4 Fuel System The fuel tank integration was difficult for the Meridian due to the removable wing design. Several options were investigated including a hinged wing joint such that the wing pivots rearward but is not removed. This would theoretically allow for fuel to be placed in the outboard wing section, but this type of fuel system would have an extremely high probability of leaking. Another option is to use quick fuel line connectors at the wing split. Again, this type of integration poses serious leaking problems. The design was iterated such that the fuel could be stored inboard of the wing split. This involved increasing the wing thickness to an 18 percent thick airfoil and adding a tank in the fuselage. The latter decision required the fuselage height to grow. The required fuel volume is 43.7 gallons or 5.84 ft 3. Approximately 45 gallons of fuel fits in the fuel tanks in the inboard wing and fuselage sections as shown in Figure Fuel bladders will be utilized for the wing and fuselage tanks. These bladders are commercially available and include all of the pickups, lines, and baffling as required. The fuel tank will be split into 3 separate bladders in the wing (1 center, and two outboard of the inboard rib), and 1 bladder in the fuselage. The center wing bladder will serve as the fuel collection point. 59

61 Fuel Storage Figure Fuel Tank Integration Anti-Icing System A combination of muffed engine exhaust and electrically heated elements will be utilized for the anti-icing system. The location of the wing is such that the leading edge of the wing is forward of the firewall. Similar engine installations have been performed ( utilizing NACA inlets to pressurize the engine cowl volume. This air will be pushed through a valve into the leading edge of the wing. The temperature of the muffed exhaust air has been measured at 180 o F. Much attention will be given to the thermal effects on material properties and stress states in the detail design and analysis phases. 60

62 3.8 Class II Landing Gear This section discusses the design of landing gear in terms of stroke length, tire diameter, and strut diameter. The landing gear must be retractable so as not to interfere with the radar. More importantly, the landing gear must be retractable with skis or conventional wheels as the Meridian will be operated from snow and paved runways. The Red, White, and Blue designs all incorporated tricycle type landing gear that retracted on a tilted pivot into the fuselage. While this is feasible for conventional tires, this retraction scheme does not work with skis. The tricycle gear had several other design problems: The nose gear had to be mounted far enough from the propeller to leave room for the nose ski. This required a very wide gear to meet lateral tipover. The Meridian should be able to be shipped in a 20 foot container, which is approximately 90 inches wide. Lateral tipover requirements called for a wider gear than this, so the gear would have to be removed for shipping. There were no commercially available landing gear similar to the previous design. All of these problems lead to the development of a new landing gear integration scheme. The gear disposition was changed to a tail dragger to solve the nose ski integration and lateral tip-over problems. The landing gear were then moved to pods mounted to the wing. This had two benefits: 61

63 The landing gear could be purchased commercially The landing gear retract straight aft, which allows for ski retraction The decision to put the landing gear on the wing calls for the use of either an oleo, pneumatic, or rubber damped type strut. The oleo gear was chosen as this type of gear is commercially available. The landing gear selection will be discussed further in Section The following assumptions were made for the landing gear design: The main gear shall be able to sustain 100% of the static load. (This is due to the tail dragger configuration.) The gear will be sized for a maximum touchdown rate of 10 ft/s. The stroke length will be sized such that a 10 ft/s decent rate imparts 1g on the airframe. The strut will have an energy absorption efficiency of 80%. The strut will be sized for skis, thereby setting the tire deflection to zero Tire Selection The main gear tires will be 3.00 x 4 tires. These tires have an outer diameter of 10.0 inches, a width of 3.2 inches, a maximum pressure of 50 psi, and weigh 3.5 lbs each. The tail wheel tire will be a 6.0 inch diameter solid rubber tire, which weighs 4.75 lbs. As will be discussed in Section 3.8.3, the main and tail gears are commercially available parts currently used on homebuilt aircraft. 62

64 3.8.2 Strut Sizing The strut stroke length and diameter sizing were performed using the methods described in [6]. The stroke length is calculated by determining the touchdown energy using Equation 3.4. The gear absorption energy equation is then used to determine the appropriate stroke length using Equation 3.5. The results are shown in Table E T 1 = W 2 L wt g 2 Equation 3.4 S s = ET n P N s m g η s η s t t Equation 3.5 Where: S s = Strut stroke length E t = Touchdown energy n s = Number of struts P m = Max static load per gear N g = Ratio of max load to static load η t = Tire energy absorption efficiency s t = Tire deflection η s = Strut energy absorption efficiency 63

65 Table Landing Gear Strut Sizing Parameter Units Value W L lbs 1082 w t fps 10 n s ~ 2 P m lbs 541 N g g 1 η t ~ 0 s t in 0 η s ~ 0.6 Diameter in 0.1 Stroke in Landing Gear Integration The landing gear placement, integration, and sizing were iterated such that a commercially available landing gear could be integrated with the Meridian. This greatly increases the feasibility of manufacturing the Meridian in the time allotted as landing gear development is a fairly complicated process. The landing gear strut produced for the Lancair Legacy homebuilt aircraft [37] will be used for the main gear (Figure 3.24 and Figure 3.25). Figure Lancair Legacy Landing Gear Strut [37] 64

66 Figure Lancair Legacy Landing Gear Installation [38] The tail wheel will also be purchased commercially. The tail wheel assembly is manufactured by Matco [39] and is commonly used on homebuilt aircraft. Figure Matco Tailwheel Assembly [39] 65

67 3.9 Structural Arrangement The purpose of this section is to discuss the proposed structural arrangement for the Meridian UAV. This includes material selection, structural layouts for the wing, v- tail, and fuselage, as well as preliminary structural sizing. The mission of the Meridian is considered to be extreme in that the locations of operation, Greenland and Antarctica, are known for extremely cold weather. While this must be considered during the structural arrangement and material selection, it must be noted that the temperatures the Meridian will experience are not much different from High Altitude Long Endurance (HALE) UAVs. In fact, the Meridian mission may be less extreme than a HALE UAV because it will not experience large changes in temperature throughout a flight. This is mentioned only to emphasize the fact that the material selections should not be arbitrarily limited due to cold weather operations. Rather, the changes in material properties due to temperature should be acknowledged and accounted for in the design process such that the final product is an optimized solution in terms of weight, manufacturability, and service life. One of the primary drivers of the material selection and structural layout is the advanced development time requirement. For the Meridian to be a successful project, manufacturability has to play a big role in the structural design process. In addition, many of the structures will be manufactured and assembled by graduate students with limited manufacturing experience. Therefore, the aircraft should be designed in such a way that limits the manufacturing skills and facilities required as is often done with homebuilt aircraft. These two concerns warrant the need for a limited part count as 66

68 well as a high level of automated processes such as computer numerically controlled machining Wing Structure The structural arrangement of the wing for the Meridian was driven by the following concerns and requirements: Shipping requirement Hard point mounting requirement for antennas Fuel system integration Cost Manufacturability Weight On site storage facility limitations The shipping, storage, and hard point requirements were determined to be the most critical and therefore had the biggest effect on the wing structural arrangement. The shipping and storage limitations [5] were such that the wing had to be designed in at least three pieces. In terms of structural optimization, the best solution would be to make the wing in two pieces. This however, does not consider the other system requirements and limitations such as landing gear and fuel tank integration nor does it consider manufacturability. The structural layout of the wing was determined by integrating the landing gear placement, fuel tank sizing, control surface sizing, shipping requirements, and manufacturing limitations. The final solution is a three-piece wing: the inboard 67

69 section contains the fuel tanks and landing gear and the outboard sections contain all of the control surfaces. The outboard sections are removable for shipping and storage in small field hangars. In addition the length of the longest part in the wing is less than that of the current composite curing facilities at the University of Kansas (~10 ft). Several possible arrangements were investigated for the wing structure including: Single spar Two Spar Three Spar Tube Spar The single spar concept was eliminated as the control surfaces will require some sort of closeout mechanism. The three spar concept was eliminated based on preliminary structural sizing analyses. The two spar concept was selected as the primary configuration with the tube spar as a secondary option. The wing is designed with a rectangular forward spar and a c-channel rear spar. The spar of the outboard wing slides into the inboard spar and is held by fasteners on top and bottom. This allows the outer portion of the wing to be removable without adding a great deal of complexity of weight. 68

70 Figure Wing Structural Layout The material selection for the wing was influenced primary by manufacturability, load types, and thermal considerations. In terms of manufacturability, composite skins allow for a high level of automation in the tooling manufacturing and provide excellent surface finish. In terms of the substructure, there are several locations where loaded fasteners are required such as the landing gear attachment, wing joint, and antenna hard points. This warrants the use of aluminum in several of the structural components such as the forward and aft spar as well as several of the ribs. The combination of different materials in the wing has implications in terms of thermal expansion. These will be investigated further in the detailed design and analysis of the structure. 69

71 3.9.2 Fuselage Structural Layout The structural layout of the fuselage was driven by the following: Wing-Fuselage Integration Manufacturability Weight Engine Installation Payload Integration Accessibility Requirements The structural layout of the fuselage was integrated with the configuration design in terms of wing and payload placement. The wing placement, which is driven by the aircraft center of gravity was iterated until the main spar of the wing was collocated with the firewall. To produce structurally efficient aircraft designs, this type of synergy must be implemented between the design aspects such that the amount of structural members required is decreased. By locating the landing gear, wing main spar, and fuselage firewall at the same fuselage station, the amount of heavy structural members has been decreased, which provides weight savings and improves the manufacturability of the vehicle. The primary frames in the fuselage were placed at the locations of the wing spars, payload hatch closure, payload rack mount, fuselage split, and v-tail spars. The remainder of the fuselage frames were spaced according to preliminary buckling calculations. The upper longerons were placed in line with the top engine mounting bracket as well as the payload hatch opening. The lower longerons were located at 70

72 the upper surface of the wing and coincide with the lower engine mounting brackets. Two frames were located at the forward and aft v-tail spar locations. These frames were also used to mount the tailwheel assembly. Figure Fuselage Structural Layout 71

73 Figure Wing-Fuselage Attachment The aircraft structure was iterated several times such that the fuselage and center wing section would fit in a standard 20 foot container [9]. The goal was to minimize the amount assembly that would have to be performed on-site. This is important for shipping, but also for on-site storage. The projected hangar size is approximately 15 feet wide, which means the wings must be removed after every flight. The aircraft is shown in a 20 foot container in Figure

74 Figure Standard 20 Foot Shipping Container Door [9] The engine mount will be procured from the engine manufacturer and will be very similar to the mount shown in Figure

75 Figure Typical Engine Mount for the Innodyn 165TE 3.10 Manufacturing Breakdown The aircraft skin is divided into 6 pieces: 2 for the cowl, left and right sections for the forward fuselage, and top and bottom sections for the aft fuselage sections. Again, manufacturability was a primary driver in the fuselage design as the leading edge of the v-tail was placed such that it coincides with the mold line of the fuselage. This allows for the aft fuselage and v-tail skin to be continuous, which improves structural rigidity and reduces the parts count. 74

76 Figure Manufacturing Breakdown 3.11 Cost Analysis The cost estimate for the Meridian UAV was created using the AAA software [7] and the methods described in [6]. The vehicle cost is broken down into research, development, test, and evaluation or RDT&E costs; and acquisition cost. The methods of [6] are typically used for production type aircraft that will be sold for some profit. This Meridian is strictly a research aircraft developed for a specific mission. The marketability of the Meridian, while may be exploited at a later date, is not part of this cost estimate. For this reason, the cost estimates of [6] were augmented with quotes from vendors for items such as avionics, tooling, and engines. 75

77 The first step in the cost estimation is to determine the Aeronautical Manufacturer s Planning Report (AMPR) weight. This is defined as the vehicle empty weight less all of the items that will be purchased from vendors such as the engine, actuators, avionics, wheels, etc. The AMPR weight of the Meridian was estimated to be: W AMPR = 260 lbs The next step in the cost estimation was to develop the hourly rates to apply to each cost estimate. This project is different from a typical aircraft manufacturing process as much of the work will be performed by students, whom work at much lower rates than typical industry standards. Average rates for manufacturing and engineering time were developed based on the rates for undergraduate students, graduate students, professors, and industry labor as shown in Table The expected breakdown of time is also shown in Table 3.23, which was used to create a time-weighted average. The industry rate was included in the wage calculations as some of the part manufacturing will be outsourced. The rates shown in Table 3.23 include typical overhead rates. Table Engineering and Manufacturing Rate Estimation Engineering Labor Manufacturing Labor % of Total Time Hourly Rate % of Total Time Hourly Rate % $/hr (2006) % $/hr (2006) Undergraduate 15 $ $16.00 Graduate 60 $ $24.00 Professor 15 $ $96.00 Industry 10 $ $60.00 Total (Averged) $37.20 $

78 Research, Development, Test, and Evaluation Costs The total RDT&E cost for an aircraft is defined as the sum of the airframe engineering and design cost; development, support and testing cost; flight test airplane cost; and flight test operations cost. This cost is then adjusted by factors accounting for test facilities, profit, and financing. This was not done for the Meridian, however, as there will be no profit or financing, and the manufacturing facilities will be paid for by other university funding. The following assumptions were made in the RDT&E cost estimation: The number of aircraft built during the RDT&E phase is 1. The workforce is assumed to be relatively skilled in Computer Aided Design The engine cost was estimated at $30,000 per the manufacturer. The propeller cost was assumed to be $5,000 per the manufacturer. The avionics cost was estimated at $100,000 per the manufacturers. No profit or financing were included in the RDT&E phase. The total cost for the RDT&E phase was determined to be $2.018 million. The cost breakdown is shown in Table 3.24 on page Acquisition Cost The acquisition cost of an aircraft is defined as the sum of the manufacturing cost and the profit. As there will be no profit for the Hawkeye, this reduces to simply the manufacturing cost, which is comprised of the airframe engineering and design cost for the production phase; the airplane program production cost; and the production 77

79 flight test operations cost. The following assumptions were made for the acquisition cost estimation: The total number of aircraft produced for the production phase is 1. The manufacturing rate is assumed to be 0.1 aircraft. The interior costs were set as $0. 40 hours of flight testing at $500/hr with an overhead factor of 4.0 were assumed for the production vehicle. No profit or financing were included in the production cost estimate. The total acquisition cost was estimated as $1.009 million. The cost breakdown for the acquisition phase is shown in Table Cost Estimate Summary The RDT&E and acquisition cost estimates are summarized in Table The total costs are broken down into overall categories in Table 3.25 and by RDT&E and production categories in Table

80 Table RDT&E and Acquisition Cost Summary Item Cost 10 $ RDTE Engineering and Design Development, Support, and Testing RDTE Labor Costs Material Costs Avionics Equipment 0.1 Tooling 0.15 Quality Control Engine Flight Test Operations Production Cost Airframe Engineering and Design Labor Production Materials Production Avionics 0.15 Manufacturing Tooling 0 Manufacturing Quality Control 0.05 Engines Flight Test Operations Acquisition Cost (2 Vehicles) Table Cost Breakdown by Overall Category Item Cost 10 $ Labor Materials Avionics 0.25 Tooling 0.15 Quality Control Engines 0.07 Flight Test Operations 0.13 Total

81 Table Cost Breakdown by RDT&E and Production Categories Item Cost 10 $ RDTE Engineering and Design Development, Support, and Testing Test Aircraft Flight Test Operations Production Cost Airframe Engineering and Design Production Manufacturing Flight Test Operations Acquisition Cost (2 Vehicles) Quality Control 5% Tooling 5% Flight Test Operations 4% Engines 2% Avionics 8% Labor 54% Materials 22% Figure Cost Breakdown by Overall Category 80

82 Cost Estimate Justification The cost estimate performed is based on conventional aircraft production costs. The viability of using these methods for a UAV is questionable. Therefore, the estimated cost was tabulated against several current UAV systems for comparison as shown in Table It is important to note that the cost of these systems is listed in terms of the vehicle costs and the system costs, which include ground support equipment and a certain number of vehicles. It is difficult to estimate the cost of one vehicle with ground support equipment, therefore the system cost was divided by the number of aircraft per system. This gives a more reasonable estimate as to the actual cost of functional UAV. The aircraft cost and the cost per aircraft based on system cost were plotted versus payload weight in Figure 3.34 and Figure 3.35 respectively. Figure 3.34 shows that the estimate vehicle cost of the Meridian is almost exactly on the linear regression. This indicates that the vehicle cost estimate is reasonable. The cost per aircraft based on system cost plotted in Figure 3.35 shows how the Meridian is a more cost-effective system because the ground support equipment (ground station, charging system, etc) is already included in the vehicle cost estimate. (The ground station costs are included in the avionics cost estimates.) Table Current UAV Procurement Cost [40] System Aircraft Weight Payload Aircraft Cost System Cost Number Cost Per Aircraft lbs lbs FY06$ mil FY06$ mil Acft/System In System FY06$ mil Dragon Eye RQ-7A Shadow RQ-2B Pioneer RQ-8B Fire Scout 1, RQ-5A Hunter 1, MQ-1B Predator 1, MQ-9A Predator 3, RQ-4 (Block 10) Global Hawk 9,200 1, RQ-4 (Block 20) Global Hawk 15,400 3, Meridian 1,

83 Global Hawk Global Hawk Meridian Predator Fire Scout Predator Cost (FY06$ 10^6) Cost(FY06$ 10^6) = (0.0089)W PL Pioneer Shadow Hunter Dragon Eye ,000 10,000 Payload Weight, lbs Figure UAV Cost in Terms of Payload Weight 82

84 Global Hawk Global Hawk Predator Predator Shadow Pioneer Hunter Fire Scout Cost (FY06$ 10^6) 1.00 Cost (FY06$ 10^6) = (0.0225)W PL Meridian 0.10 Dragon Eye ,000 10,000 Payload Weight, lbs Figure UAV Cost Based on System Cost Versus Payload Weight 83

85 4 Conclusions This document summarizes the redesign of the Meridian UAV based on the response from the Preliminary Design Review. The antenna type has been changed from a bow-tie to a Vivaldi or exponential antenna. In addition, the aircraft has been modified to incorporate several commercially available off-the-shelf parts. The goal with the redesign of the Meridian is to produce a design that is not only novel, but is feasible considering the extremely short development time. This meant integrating manufacturability, performance, and operational constraints into the design process. The product is a vehicle that can be shipped in a standard 20 foot container and quickly assembled and disassembled with minimal tools. The Meridian is the smallest turboprop powered UAV in the world. It is also one of the only UAVs with retractable ski landing gear. The purpose of this continued development of the Meridian is to completely flush out the best new UAV design based on the mission specification. 84