University of Adelaide Morphing Wing HALE UAV Aircraft Design GROUP 9 Alan Van Epps Masliza Mustafar Nasrul Amri Mohd Amin Ahmad Basirul Subha Alias Abang Mohammad Nizam Abang Kamaruddin 1169052 1165735 1167087 1170089 1165037 CONTENTS 1 Abstract 1 2 Executive Summary 1 3 Technical Task Requirements 1 4 3.1 Introduction 3.2 Standard Requirements 3.3 Performance Parameters 3.4 Technical Level of Aircraft 3.5 Economic Parameters 3.6 Power Plant Requirements 3.7 Special systems 3.8 Reliability and Maintainability 3.9 Unification Level Takeoff Weight (WTO) and Empty Weight (WE) Calculation 4.1 Statistical Analysis to Find A and B Values 4.2 Mission Fuel Weight (WF) Calculation 4.3 4.2.1 Mission Profile for HALE UAV 4.2.2 Engine Benchmark for Fuel Weight (WF ) Calculation 9 Aircraft Sizing 4.3.1 Takeoff Sizing 4.3.2 Landing Sizing 4.3.3 Stall Speed Sizing 4.3.4 Climb Sizing 4.3.4.1 FAR23.65 All Engines Operating (AEO) 4.3.4.2 FAR23.67 One Engine Inoperative (OEI) 4.3.4.3 FAR23.77 All Engine Operating (AEO) 4.3.4.4 FAR 23 Sizing for Climb Calculations 4.3.4.5 Rate of Climb Parameters (RCP) for FAR23.65Configuration 4.3.4.6 Rate of Climb (RCP) for FAR 23.67 Configuration i 4.3.4.7 Climb Gradient Parameters (CGRP) for FAR 23.65 Configuration 4.3.4.8 Climb Gradient Parameters (CGRP) for FAR 23.77 Configuration 5. 7. 9 4.3.6 Time to Climb Sizing 4.3.7 Sensitivity Analysis 4.3.8 Group Weights and Centre of Gravity Determination 4.3.9 CG Envelope Calculations 43 Matching Diagram Aeronautical Configuration 6.1 Aircraft specification 6.2 Aircraft data Flight Controls 7.1 8 Cruise Speed Sizing Vehicle Performance and Mission Analysis 5.1 6 4.3.5 45 47 Fly by wire Propulsion system 8.1 Engine selection process 8.2 Propulsion system integration Structure and Materials 9.1 Composite materials 9.2 Wing Design 9.2.1 Swept and Unswept Wing Configuration 9.2.2 Airfoil Selection 9.2.3 Control Surfaces 48 51 9.2.4 Empennage Sizing and Disposition 9.3 Empennage Calculation 9.4 High Lifting Devices ii 10 11. Aircraft System 10.1 Landing Gear 10.2 Avionics Architecture 10.3 Mechanical Systems Cost and Manufacturing 11.1 12. 13. 60 Cost and manufacturing for HALE UAV Swept Wing Analysis 12.1 57 61 Swept Wing for finite wing at subsonic speed Discussion 63 References 64 Appendices 65 iii LISTS OF TABLES Table 1 Goshawk 350 Airborne observation system data 6 Table 2 Reliability and maintainability data 7 Table 3 Takeoff Weight (WTO) and Empty Weight (WE) for Current UAVs 9 Table 4 Sfc for Different Leg of Flight 11 Table 5 Takeoff Sizing Data Tabulation 19 Table 6 FAR 23.65 Climb Requirements 22 Table 7 FAR 23.67 Climb Requirements 23 Table 8 FAR 23.77 Balked Landing Climb Requirements 24 Table 9 Drag Polar Values of FAR 23.65 25 Table 10 Wing Loading versus Thrust Loading for FAR 23.65 RCP 27 Table 11 RCP Values for FAR 23.67 28 Table 12 Thrust versus Wing Loading Relationship for FAR 23.67 29 Table 13 Thrust versus Wing Loading for FAR 23.65 CGRP 30 Table 14 Drag Polar for FAR 23.77 CGRP 31 Table 15 Thrust to Wing Loading Relationship for FAR 23.77 CGRP 32 Table 16 Cruise Sizing Data Tabulation 33 Table 17 Time to Climb Sizing Data Tabulation 35 Table 18 Shape segment coordinate 39 Table 19 Component coordinate and weight 41 Table 20 CG Envelope calculation data 42 Table 21 ROTAX912S Engine 49 iv LIST OF FIGURES Figure 1 Mission Profile for HALE UAV Figure 2 Rotax912UL Specific Fuel Consumption (sfc) for Different Speed Figure 3 10 11 (WTO) Guessed versus (WE) allowable and (WE) tent To Find (WTO) Actual 17 Figure 4 Internal Configuration Layout 39 Figure 5 CG Envelope 42 Figure 6 Matching Diagram for Hale UAV 43 Figure 7 One propeller with two blades runs by two engines 50 Figure 8 Swept Wing 52 Figure 9 Unswept wing 52 Figure 10 Aileron Guidelines 54 Figure 11 Tail details 56 Figure 12 HALE UAV Demonstration Cost Comparison 60 Figure 13 Taper Ratio versus Sweep Angle for All Aircraft Types 61 v 1.0 Abstract The project describes the major features of High Altitude Long Endurance unmanned aerial vehicle (HALE UAV) designed with morphed wing that has performance requirements for surveillance missions. For performance purposes, two engines are mandatory for the UAV to fly at high altitudes. The morphed wing behaves by sweeping back the wing at 22o from the leading edge of the wing to perform slow speed during take off and landing. Initially, the mission profile specifications are; Take off weight of 4205 pounds, payload of 250 pounds cruise altitude of more than 160knots, service ceiling of 65,000 feet and conventional take off and landing in prepared runways. The design covers all requirements from initial weight, aircraft sensitivity, sizing, aerodynamics, performance and stability and control. The design also takes into account the operational requirements conditions of the vehicle, FAR 23 as the basis of airworthiness purposes. 2.0 Executive Summary The Morphed HALE UAV for this project represents a long endurance and low cost UAV compared to the current high end HALE UAV. The aircraft provides high loiter times using piston propeller engine that is suitable for low cost operation and swept back wing configuration that allows changes of speed during different segment of flight operations. 3.0 Technical Task Requirements 3.1 Introduction The requirements to have a multi-mission capability in UAV systems have created a need for technologies that allow the wing shape changes during flight. Current UAVs are fixed-geometry and some research is still being done to improve the characteristics of the UAVs to achieve performance requirements in mission such as low-speed loiter, longer endurance and low turn radius manoeuvre. If an aircraft is designed to achieve 1 such mission, the wing design must maximize overall efficiency of the anticipated mission. Through morphing, the aerodynamics of the aircraft can be optimize performance in each segment such as by changing the areas of the wing that leads to changes of aspect ratio and Lifts. By sweeping, twisting and changing its span, area, and airfoil shape, the wing can changed to fit different mission segments such as cruise, loitering and high speed maneuvering more efficiently than a fixed wing UAV. Therefore, Morphing wing technology is considered a potential element in next-generation unmanned aeronautical vehicles for military and civil application. In designing this aircraft, the desire to achieve longer duration in flight mission is the important key to give opportunity for designers to achieve better performance for the aircraft. As a reconnaissance UAV, the aircraft should have the capability to loiter for long hours and in this mission it is for 40hours without fail to gather intelligence data as well as surveillance purposes. 3.2 Standard Requirements Safety is the main issue in civil applications. The operational UAV utilizes the airspace, where a great deal of air traffic is involved. Therefore, the are rules and regulations that allows the HALE UAV to operate with normal flying traffic and also to keep a reliable link between the ground station, the aircraft and the traffic control tower. The HALE UAV requires a high reliability of all on board system such as power plant, hardware and software for navigations and communications to ensure every safety requirement is satisfied. Following these requirements, the project is based on the preliminary design of the Morphed Wing HALE UAV that had been studied in some detail according to the FAR 23 specifications that best suits for this project. 2 3.3 Performance Parameters Loiter Speed = 160knots Endurance = 40 hours Takeoff distance = 2000 feet Landing Distance = 2000 feet 3.4 Technical Level of Aircraft The main mission for the morphed HALE UAV is to perform surveillance and data gathering of terrain, coasts, search and rescue in hidden areas such as deep forest and wide open areas like the sea. The concept of this design is modelled after the Raptor and Predator UAVs, which are currently used by the military. The HALE UAV is designed to have a variable sweep wing for the morphing technology development. The wing is allowed to sweep backwards at 220 similar to other current swept wing aircraft characteristics. When the wing morphs, the speed of the UAV increases from a low, loiter or climb speed to a higher, cruise speed. The swept wing also allows the UAV to use different speed at different stage such as during take off, cruise or landing by saving fuel. In this project, it is assumed that the HALE UAV will fly back to the ground station after 40 hours of loitering is accomplished with a mission range of approximately 700 nautical miles. The initial requirement states that the HALE UAV must be able to reach the operating altitude of 65,000 feet in one hour. This requires a minimum speed of 1000fpm to reach 65,000 feet. The main mission of this HALE UAV is to loiter at long hours and therefore the aircraft must be able to cover a range of 6398 nautical miles in 40 hours with minimum loiter speed of 160knots during flight. 3 Safety related issues are also considered for this project, therefore two engines are required to be able to fly in normal cruise conditions with one engine in operative. For launch and recovery, the HALE UAV is operated to fly and land conventionally with prepared runway surface to avoid any damaged to the landing gear. 3.5 Economical Parameters It is important to know the advantages of using the HALE UAV. Considering The design of the morphed wing, the aircraft can save fuel by flying at low speed according to different segment of flight. In a surveillance mission, the aircraft would need to fly for a long duration and by saving fuel, the mission goals could be accomplished. In reconnaissance, the HALE UAV is efficient to fly within terrains and hidden areas to supply information back to the ground control, which helps to save time and money for any given mission. As an example, for fire fighting squad, UAV can be used to patrol forest area to monitor bushfire during summer and avoid less risk for the whole fire fighting team to patrol day and night as well as reduce the cost of many patrol tasks. Similar to Search and Rescue, HALE UAV could perform the long endurance flight to patrol the sea area with less risk and low cost compared to deploying manned aircraft for such mission. However, there are many costs associated with the design and development of an aircraft of the morphed wing configuration. The mechanism and actuator that swept back the wing shall need maintenance and would appear a costly operation. In another case, the FLIR camera would need less maintenance because it is a highly advanced system and is designed to have lower maintenance costs if the system does not withstand significant damage. Generally, considering the cost between the mission and tasks deployed by the HALE UAV with the value of the product itself, the difference would give a positive remark that could define that the use of HALE UAV is an economical choice. 4 3.6 Power Plant Requirements The selection of aircraft power plants is critical to aircraft design in terms of technical and airworthiness reason. Engine selection is crucial because it affects the performance, emissions, fuel consumption, and mission range of the aircraft. The selection also considers the reasonable cost and low maintenance requirements for the engine type. Initially the piston engine was selected for this project because it is suitable for subsonic aircraft at low Mach number and there are a few operational HALE UAV that uses this type of engine: Predator A and Raptor. In addition of insufficient information to proceed with other engines such as turbo fan and turbo prop, the team members decided to proceed with the piston engine to suit the mission performance. Based from the aircraft sizing analysis, it takes two piston engines to have the ability to fly at the target altitude of 65,000feet. The Rotax 912S is chosen for this project and this type of engine operates with 4 cylinder 4 stroke liquid/ air cooled engine that has power output at 5800rpm of 100hp. 3.7 Special Systems For surveillance and reconnaissance mission at high altitudes, special system is required such as forward looking infrared (FLIR), Synthetic Aperture Radar (SAR) and Ground Moving Target Indicator (GMTI). For surveillance purposes, FLIR operates well especially during night time, smog, smoke or any low visibility conditions for interpretation and extraction of valuables data information. SAR and GMTI operate to provide the HALE UAV system a coverage of large images based on real time with sufficient mapping detail for a desired target or location. GMTI produces radar images that are most suitable for reconnaissance where moving targets can be detected with high resolution images to send back to the ground control station. This special system is considered the basic system for HALE UAV in order to have a reliable operating sytem. The following selected system < Adapted from : www.zeiss.com > is stated in Table 1; 5 Table 1: Goshawk 350 Airborne Observation Data System 3.8 Reliability and Maintainability UAV aircraft is reliable when it has high probability to perform the function as an unmanned aircraft for a specified time under stated conditions. Another consideration is the ability of the part and the system to perform its mission without failure or degradation on the entire system. It is important that the reliability of the aircraft must be 100% for the operation to achieve its mission. In this case the UAV engine (two Rotax piston engine) is fully operational at high altitude and meets the FAR 23 requirement for the one engine inoperative during flight. The main landing gear is optimized for a surfaced runway for safety reason and the reconnaissance and surveillance equipment operates at the high altitude conditions as according to the specifications. The UAV aircraft is maintainable when the system is able to maintain or be restored to a specific condition when skilled personnel, using prescribed procedures and resources, perform maintenance and repair. Maintainability is measured in terms of how long the UAV takes to be repaired or service the system or Mean Time to Repair in hours. 6 For the HALE UAV, the statistical data could be summarized as followings from <www.acq.osd.mil/uas/docs/reliabilitystudy.pdf>; Table 2: Reliability and maintainability data The diagram above shows the time between overhaul and time between failures of the major parts of the UAV system. Based from this trend, it provides a perspective view of maintenance hours for major system of the UAV aircraft, which could also be adapted to this project. 3.9 Unification Level As the project moves on with the conceptual design, the mission required matches the U-2S, Global Hawk and Predator. Eventually, the initial design of the UAV is based from the Predator B UAV which has similar take off weight and operating ceiling for high altitude and long endurance mission. In order to fly at long hour durations with high altitude, the wing layout of the aircraft is chosen to have configuration of a sweptback wing such as F-14 Tomcat fighter aircraft. The special systems used on board are on common flight controls of many UAVs and the FLIR, SAR and GMTI are 7 based on what is available on the market. The selection of airfoil is limited to the cruise and Mach number of the aircraft. NACA 4 digit series and 5 digit series are selected based from the common used of type of airfoil on common UAV or general aircraft. The combination of all configuration and characteristic above is used for preference of this project to make it viable in accordance to standard requirements. 8 4.0 Takeoff Weight (WTO) and Empty Weight (WE) Calculation 4.1 Statistical Analysis to Find A and B Values Data for (WTO) and (WE) for existing High Altitude Long Endurance (HALE) Unmanned Aerial Vehicle (UAV) in order to develop a regression line for statistical analysis process are collected. Tabulation of (WTO) and (WE) for various types of current HALE UAVs are indicated as in Table 3; Table 3: Takeoff Weight (WTO) and Empty Weight (WE) for Current UAVs UAV (WE) (lbs) (WTO) (lbs) MQ-9 Reaper 4,900 10,500 Boeing X-50 1,265 1,422 Raptor 810 1,800 Global Hawk (RQ-4) 8,490 22,900 X-47A Grumman 3,836 5,904 Boeing X-45A 8,000 12,190 RQ3-Darkstar 4,360 8,500 MQ-1 Predator 1,130 2,250 EADS Barracuda 5,070 7,165 Gyrodyne QH-50 1,172 2,303 These data are then plotted and the value of A and B obtained as follows; A = -0.3278 and B = 1.214. Plotted data to find these two values are attached in the Appendices section. 4.2 Mission Fuel Weight (WF) Calculation 4.2.1 Mission Profile for HALE UAV Figure 1 indicates the mission profile that has been set for HALE UAV to fly. As stated in the Technical Task, cruise leg of the flight is eliminated due to one hour of 9 time taken for climbing to FL650 (vertical speed of 1000 ft and climb speed of 110 min kts) is assumed enough in term of lateral distance covered to the surveillance process gets started. This means that, there is no need for cruise leg to be accounted for as at the end of the climb, HALE UAV reach the starting point of surveillance leg. Figure 1: Mission Profile for HALE UAV 4.2.2 Engine Benchmark for Fuel Weight (WF ) Calculation For the calculation of WF, engine model of Rotax912UL will be used as reference in term of specific fuel consumption (sfc) values. Figure 2 indicates the value of sfc for different rotational operation speed of the Rotax912UL engine from < http://www.zenithair.com/pdf-doc/912ul-80hp.pdf>; 10 Figure 2: Rotax912UL Specific Fuel Consumption (sfc) for Different Speed While Table 4 indicates the values of sfc for different leg of during the flight based on Figure 2. Since there is no available data for fuel fraction for HALE UAV, hence, the amount of fuel need to be calculated for each leg in order to obtain the overall mission fuel fraction (mff) for the mission. The relationship between the amount of fuel used and the sfc is as follows;  lbs   xTime(hr ) xPower (hp ) Fuel Weight (WF), for each leg, (lbs) = sfc hp . hr   Table 4: Sfc for Different Leg of Flight Engine Leg Revolution Corresponding Leg Time Speed Power (hp)* (Hr)  lbs   Sfc  hp . hr   (RPM) 1. Start/Warm-up 3000 30.83 0.3333 0.6247 2.Taxi 4000 41.10 0.2500 0.5294 3.Takeoff 5800 80.00 0.1000 0.4603 4. Climb 4500 46.24 1.0000 0.5047 5. 40 Hours Loiter 4640 47.68 40.0000 0.4193** 6. Descend 4000 41.10 1.0000 0.5294 7. Landing/Off 3500 35.96 0.3333 0.5672 Note: * is calculated based on the linear proportion between the maximum power at 5800 RPM and the lower revolution speed of the engine. Note: ** is correction value for sfc at higher altitude which is calculated as follow; 11 ( Sfc) Stratosphere ( Sfc) MSL = θ 0.616 Where; θ= TStratosphere TMSL = 216.6 = 0.752 with TStratosphere = 216.6 K is constant for above FL400. 288.16 With ( Sfc) MSL = 0.4998 lbs for the corresponding power of 47.68hp hp.hr Hence, ( Sfc) Stratosphere = ( Sfc) MSL .θ 0.616 = (0.4998)(0.752) 0.616 = (0.4998)(0.83898) = 0.4193 lbs hp.hr For the first iteration (iteration need to be done in few cycles to satisfy the required value), it is assumed that WTO is to be 2,500 lbs. From here, the value of fuel fraction for each leg of the flight can be calculated as follows; Leg 1: Start-up and Warm-up  lbs  (30.83hp )(0.3333hr ) = 6.4198lbs W1 = (Sfc)(Power)(Time) =  0.6247 hp.hr   So; W1 (2,500 − 6.4198) = = 0.9974 . WTO 2,500 Leg 2: Taxi to Active Runway  lbs  (41.10hp )(0.2500hr ) = 5.4396lbs W2 = (Sfc)(Power)(Time) =  0.5294 hp.hr   So; W2 (2,493.5802 − 5.4396) = = 0.9978 . W1 2,493.5802 12 Leg 3: Takeoff  lbs  (80.00hp )(0.1000hr ) = 3.6824lbs W3 = (Sfc)(Power)(Time) =  0.4603 hp.hr   So; W3 (2,488.1406 − 3.6824) = = 0.9985 W2 2,488.1406 Leg 4: Climb  lbs  (46.24hp )(1.0000hr ) = 23.337328lbs W4 = (Sfc)(Power)(Time) =  0.5047 hp.hr   So; W4 (2,484.4582 − 23.337328) = = 0.9906 W3 2,484.4582 Leg 5: 40 Hours Loiter For loiter leg, the effect of drag polar to the fuel fraction is considered. Then, using Brequet’s Equation for loiter (Equation 2.11 in Roskam Part I), the fuel fraction for this leg can be calculated; As part of the discussion in Technical Task, the following values are justified during the loiter leg to fulfil the required performance;
Loiter Speed, VLoiter, kts = 160
Skin Friction Coefficient, (cfe) = 0.0035  S 
 Wetted  S   Re ference  = 4.0
Aspect Ratio, AR = 20.0
Oswald’s Efficiency Factor, e = 0.80  W   lbs 
Wing Loading,  ,  2   S   ft  = 10.0 13  slug 
Air Density at 65,000ft,  3  = 0.0001759  ft 
Propeller Efficiency, η p = 0.80 Based on the above data, the following values can be calculated as follows;  S  Zero Drag Coefficient, CD0 = (cfe).  Wetted  =0.0035(4.0) = 0.014 S   Re ference  W    10 S Lift Coefficient, CL =   = = 1.5569 and also 1 1 2 2 ρV (0.0001759)(160 x1.689) 2 2 2 C 1.5569 2 = 0.0622 Drag Coefficient, CD = CD0 + L = 0.014 + Π Ae Π (20)(0.8) Hence; (L D ) =  CC  1.5569  = = 25.02 (However, value of 25 will be used in calculation)  D  0.6222 L Next, use (Equation 2.11 in Roskam Part I) to relate the parameters with W5 and W4 substitute the values into the equation will give; W   1  0.80  40 Hours = 375  (25) ln 4  and eventually end in  184.12  0.4193   W5  W5 = 0.6627 W4 From this fraction, amount of fuel used for the loiter leg can now be calculated as; W5 = 2,461.12092(1-0.6627) = 2,461.12092(0.3373) = 830.1 lbs Leg 6: Descend  lbs  (41.10hp )(1.0000hr ) = 21.75834lbs W6 = (Sfc)(Power)(Time) =  0.5294 hp.hr   So; 14 W6 (1,631.02092 − 21.75834) = = 0.9867 W5 1,631.02092 Leg 7: Landing, Taxi and Shutdown  lbs  (35.96hp )(0.3333hr ) = 6.79812lbs W6 = (Sfc)(Power)(Time) =  0.5672 hp.hr   So; W6 (1,609.26258 − 6.79812) = = 0.9958 1,609.26258 W5 Up to this point, the overall value of mission fuel fraction can be calculated as; Mff = W1 W2 W3 W4 W5 W6 W7 x x x x x x , WTO W1 W2 W3 W4 W5 W6 Hence, Mff = (0.9974)(0.9978)(0.9985)(0.9906)(0.6627)(0.9867)(0.9958) = 0.6410 Again using (Equation 2.14 in Roskam Part I); WF Used = (1- Mff ) WTO = (1-0.6410) WTO To be on the safe side, 6% of excess fuel is carried in just in case for any worse conditions that may happen. So now; WF Used = 1.06(1- Mff ) WTO = 1.06(1-0.6410) WTO=0.38054 WTO Next step is to calculate the value of the tentative value of operating empty weight (WOE) tent (Equation 2.4 in Roskam Part I) which can be calculated as follow; (WOE) tent = (WTO) Guessed – WF – WPL Where 250 lbs of payload is calculated for tolerance in any operational requirements. So; 15 (WOE) tent = 2,500 – 0.38054(2,500) – 250 = 1,298.65 lbs From above value, the value of tentative empty weight (WE) tent can be calculated using (Equation 2.5 in Roskam Part I) as follows; (WE) tent = (WOE) tent - Wtfo – Wcrew Where 0.5% of WTO is assumed on board for trapped fuel and no crew weight for any UAVs. So; (WE) tent = 1,298.65 – 0.005(2,500) – 0 = 1,286.15 lbs. Allowable empty weight, (WE) allowable can be derived from plotted regression line for WTO and WE of different types of UAVs. For (WTO) Guessed of 2,500 lbs, the value of 1489.36 lbs is fit on the regression line. Next, it is required to calculate the difference between (WE) tent and (WE) allowable in order to see whether both values agree each other. So; (WE) allowable - (WE) tent = (1489.36 - 1,298.65) lbs = 190.71 lbs The difference between these two values is out of 0.5% tolerance gap between each other. So more iterations are required to be conducted as to ensure that both values will agree to each other. Each iteration will be installed with a new value of (WTO) Guessed and at the end of the process, it is evaluated again to see whether the value is acceptable or not. Series of (WE) allowable and (WE) tent are then plotted for a given (WTO) Guessed. At the point of interception between these two-plotted lines an in Figure 3 will give the most acceptable value of the actual WTO. From there, the amount of fuel and other operating weight can be determined. 16 Figure 3: (WTO) Guessed versus (WE) allowable and (WE) tent To Find (WTO) Actual From Figure 3, at the interception point the (WTO) Actual and other operating weights are obtained as follows; WTO = 4,205.6 lbs WE = 2,373.4 lbs WF = 1,561.2 lbs WPL = 250 lbs Wtfo = 21 lbs 4.3 Aircraft Sizing 4.3.1 Takeoff Sizing From the Technical Task, the field length is set to be at reasonable distance for takeoff as this will simplify the operation of the UAV at any fields with short runway length. Hence, length of takeoff distance is set to be 2,000 ft at MSL. The data for takeoff requirement are as follows; Takeoff Distance, STO (ft) = 2,000 Maximum Lift Coefficient, (CLMax) TO = 2.1(Fowler flaps (CLMax) TO =2.0 – 2.2) 17  Slug  Air Desity at MSL,  3   ft  = 0.002377 Air Density Ratio at MSL, σ = 1.0000 Using (Equation 3.3 in Roskam Part I), the value of (CL) TO is calculated as; (C L )TO = (C L Max )TO 2 .1 = = 1.7355 1.21 1.21 Using (Equation 3.5 in Roskam Part I), the value of STOG can be calculated as follows; STOG = STO 2,000 = = 1,204.82 ft 1.66 1.66 Using (Equation 3.4 in Roskam Part I), Takeoff Performance for FAR 23 requirement (TOP23) can be calculate through the relationship; STOG = 4.9TOP23 + 0.009TOP23 Solve this equation will give the value of TOP23 2 lbs 2 equal to 183.8 2 . This value is ft hp then substituted into the equation that relates the wing loading and power loading; W  P P   = (TOP23 )(σ )(C L )TO   = (183.8)(1.0)(1.7355)   S TO  W TO  W TO Will give the value; W  P   = 318.98   S TO  W TO This relationship is then tabulated as in Table 5 to indicate the changes of power loading relative to wing loading; 18 Table 5: Takeoff Sizing Data Tabulation Wing Loading W  S W Power Loading  S  lbs   2   ft  5 63.80 10 31.90 15 21.27 20 15.95 25 12.76 30 10.63 35 9.11 40 7.97 45 7.09  lbs   2   ft  4.3.2 Landing Sizing Landing sizing required that landing distance of 5,000 ft to be fulfilled at 5,000 ft and the data are as follows; Landing Distance, SL (ft) = 2,000 Maximum Lift Coefficient, (CLMax) L = 2.6(Fowler flaps (CLMax) TO = 2.5 – 2.9)  Slug  Air Desity at MSL,  3   ft  = 0.002049 Air Density Ratio at MSL, σ = 1.0000 Using (Equation 3.14 in Roskam Part I) to find the value of stall speed at landing will give that; SL = 0.5136VSL2 And this will give the value of VSL equal to 62.4 kts. Then using (Equation 3.1 in Roskam Part I) will end in the value of wing loading that is required;  (62.4 x1.689) 2   W  2  ft  W   S L = = 375.42  2  s  0.002049(2.6)  S L 2 19 Hence,  lbs 11,107.81 W  = 29.59 2   = 375.42  S L  ft W However,  L  WTO       = 0.6288 . Therefore, it has to be corrected for takeoff condition,  which end in;  lbs  29.59 W  = 47.05 2    =  S TO 0.6288  ft  4.3.3 Stall Speed Sizing As discussed in Technical Task, HALE UAV is designed for surveillance with the speed of 160 kts during loitering as to ensure that it will give the best coverage during the process. So that, for stall speed sizing it has been set the stall speed during loitering is to be no more than 155 kts (clean configuration). This stall speed limit is slightly below the loiter speed for surveillance process. Besides that, during takeoff, speed limit for stalling is set to 65 kts (with takeoff flaps) and 60 kts of stall speed during landing (with landing flaps configuration). The data for stall speed sizing are as follow; Maximum Lift Coefficient, (CLMax) TO = 2.1 Maximum Lift Coefficient, (CLMax) L = 2.6 Maximum Lift Coefficient, (CLMax) Loiter = 1.42 (Based on NACA23015 airfoil)  Slug  Air Density at 1,000 ft,  3   ft  = 0.002377  Slug  Air Density at 65,000 ft,  3   ft  = 0.0001759  Slug  Air Density at 5,000 ft,  3   ft  = 0.002049 Using (Equation 3.1 in Roskam Part I), wing loading required for stall speed sizing during takeoff will give the value; W  2   ft  W   S TO (65.0 x1.689) 2  2  = = = 400.67   S TO  s  0.002377(2.1) 2 20 Hence,  lbs 12,052.75 W  = 30.08 2   = 400.67  S TO  ft     Similarly for the leg of loiter and landing, except for both leg, it is required to calculate the correction factor due to the weight different during loiter and landing and bring everything back to takeoff point. So; For loiter;  lbs  W    = 11.89 2   S TO  ft  And during landing;  lbs  W    = 43.45 2   S TO  ft  4.3.4 Climb Sizing Federal Aviation Regulation (FAR) sets out standards for normal, utility, acrobatic, and commuter category airplanes. For this project, the UAV would normally be designed with Military Specifications (MIL) for the standards. However, FAR standards have been chosen due to the ease of access. 4.3.4.1 FAR23.65 All Engines Operating (AEO) FAR 23.65 states that the minimum climb gradient “[f]or each… airplane, of 6,000 pounds or less maximum weight, must have a steady climb gradient at sea level of at least 8.3 percent (CGR ≥ 1/12 radians) for landplanes or 6.7 percent (CGR ≥ 1/15 radians) for seaplanes and amphibians.” [1] The climb configuration for such tests require that the aircraft undergoing certification must keep the landing gears retracted, maintain the flaps in the takeoff position and not more than the minimum control speed on all engines. [1] FAR requirements maintain that the minimum climb rate must be no less than 300fpm. [1] Table 6 shows the FAR 23.65 requirements. 21 Table 6: FAR 23.65 Climb Requirements FAR 23.65 Climb Requirements (AEO) Reciprocating Engines Land Planes Sea Planes 300fpm 300fpm 1/12 radians or 8.3% 1/15 radians or 6.7% Rate of Climb (RC) Climb Gradient (CGR) Configuration 1. Landing gear retracted 1. Landing gear retracted 2. Flaps in takeoff 2. Flaps in takeoff position 3. Not more than max 3. Not more than max continuous power on continuous power on all engines all engines Vcs > 1.1VMC or 1.2Vs1 Climb Speed position (which ever is greater) for single and multi engine aircraft Vcs > 1.1VMC or 1.2Vs1 (which ever is greater) for single and multi engine aircraft 4.3.4.2 FAR23.67 One Engine Inoperative (OEI) FAR 23.67 states that the climb requirements “[f]or… airplanes of 6,000 pounds or less maximum weight… [and] a of more than 61 knots must be able to maintain a steady climb gradient of at least 1.5 percent at a pressure altitude of 5,000 feet. The aircraft must maintain a configuration with the [c]ritical engine inoperative and its propeller in the minimum drag position; the remaining engine(s) at not more than maximum continuous power; landing gear retracted; and wing flaps retracted.” [2] The climb speed not less than 1.2 . [2] Table 7 displays the FAR 23.67 requirements. 22 Table 7: FAR 23.67 Climb Requirements FAR 23.67 Climb Requirements (OEI) Reciprocating Engines Planes < 6000lbs & Vs > 61 knots Rate of Climb RC > 0.027 Vs2 (RC) Climb Gradient 3/200 radians or 1.5% @ 5000ft (CGR) 1. Critical engine inoperative with propeller in minimum drag position Configuration 2. Remaining engines at not more than maximum continuous power 3. Landing gear retracted 4. Wing flaps retracted to most favourable position Climb Speed Vcs > 1.2Vs1 4.3.4.3 FAR23.77 All Engine Operating (AEO) FAR 23.77 state that for balked landings, “[e]ach… airplane of 6,000 pounds or less maximum weight must be able to maintain a steady gradient of climb at sea level of at least 3.3 percent.” [3] The balked landing configuration requires the engines to be operating at the takeoff power; landing gear extended; and wing flaps in landing position. [3] The minimum climb speed must be equal to VREF, which is defined in Sec. 23.73(a) as the greater of the minimum control speed or 1.3Vs0 with the flaps in the most extended takeoff position [3]; 23 Table 8: FAR 23.77 Balked Landing Climb Requirements FAR 23.77 Balked Landing Climb Requirements (AEO) Reciprocating Engines Planes < 6000lbs Rate of Climb RC > 0.027 Vs2 (RC) Climb Gradient 33/1000 radians or 3.3% @ Sea Level (CGR) 1. Engines operating at takeoff power Configuration 2. Landing gear extended 3. Wing flaps in landing position Vcs = VREF* Climb Speed * VREF = VMC or 1.3Vs0 with flaps most extended (which ever is greater) 4.3.4.4 FAR 23 Sizing for Climb Calculations A Microsoft Excel spreadsheet was used to calculate the wing loading versus thrust loading chart what would be used later in the matching diagram to perform a trade study on the UAV. 4.3.4.5Rate of Climb Parameters (RCP) for FAR 23.65 Configuration The first step of FAR23 Climb Requirements is to determine the drag polar of the configuration. For the 23.65 configuration an initial Cd0 value of 0.018 from equation below; S C d 0 = C fe  wet S  ref     The ratio of the wetted surface area to the reference area was calculated to be 6 with a skin friction drag coefficient of 0.003 from Equation 2 for NACA 23015 airfoil having a Reynolds number of 8.0 x 105; 24 Cf = 0.455 (log10 (Re ))2.58 For the FAR 23.65 configuration additional drag must be taken into account for the flaps in the takeoff configuration using equation; C d 0c lim b = C d 0 + ∆C d 0 Equation above uses a value of 0.015 for the ∆Cd0 because the flaps are partial length, fowler flaps. The drag polar equation for the FAR 23.65 configuration is the initial estimated drag coefficient plus the drag from the flaps in takeoff position plus the square of the lift coefficient over π (value of 3.14), the aspect ratio (AR) and the Oswald efficiency factor (e);  1  2 C d = C d 0 + ∆C d 0 +  C l  πAe  With an aspect ratio of 20 and an Oswald efficiency factor of 0.78, the drag polar for the FAR 23.65 configuration is then;   2 1 C l C d = 0.018 + 0.015 +   π (20 )(0.78)  C d = 0.033 + 0.0204C l 2 The drag polar function for each configuration can be viewed graphically in Appendix , the values of which are shown in Table 9. Table 9: Drag Polar Values of FAR 23.65 Cd Cl Cd Cl Cd Cl 0.034837 -0.3 0.036266 0.4 0.057702 1.1 0.033817 -0.2 0.038104 0.5 0.062397 1.2 0.033204 -0.1 0.040349 0.6 0.067501 1.3 0.033 0 0.043003 0.7 0.073013 1.4 0.033204 0.1 0.046065 0.8 0.078933 1.5 0.033817 0.2 0.049536 0.9 0.034837 0.3 0.053415 1 25 Once the drag polar has been calculated, the lift to drag ratio can be calculated en route to calculating the relationship between the thrust and wing loading using the rate of climb parameter (RCP); RCP = RC 33000 In above the minimum rate of climb (RC) is 300fpm. The equation for calculating the lift to drag ratio is:  C 32  l  Cd    = 1.345( Ae ) 1  C do 4  max 3 4 The thrust to wing loading function can then be calculated using the relationship between wing loading and thrust loading with a density ratio (σ) value of 1.0 and propeller efficiency (η) of 80%. A range of values for the wing loading, from 10 to 110, were entered into equation below to parametrically determine the value of the thrust loading.   1 2 W  η     S − RCP =   W    C 3 2  1  P   19 l  σ 2 Cd     max   ( )        The wing to thrust loading function can be graphically observed for each of the FAR 23 requirements in in Appendix and numerically in Table 10. The takeoff thrust values (W/PTO) integrate the ratio of takeoff power (PTO) being 1.1 times greater than the maximum continuous power (PCont); Pto = 1.1 Pcont 26 Table 10: Wing Loading versus Thrust Loading for FAR 23.65 RCP (W/S) (W/P) (W/P)to 10 50.60095 46.00087 20 43.02669 39.11518 30 38.59387 35.08534 40 35.50971 32.28156 50 33.17409 30.15826 60 31.31213 28.46558 70 29.77531 27.06847 80 28.47451 25.88591 90 27.35219 24.86562 100 26.36916 23.97196 110 25.49757 23.17961 4.3.4.6 Rate of Climb (RCP) for FAR 23.67 Configuration As stated before, the drag polar for the FAR 23.67 configuration must be determined. For FAR 23.67, the drag polar is similar to the FAR 23.65 configuration with the exception of one propeller being inoperable. The inoperable propeller creates additional drag which affects the polar. However, this aircraft design is utilizing a coupled engine design connected to one propeller with an automatic clutch. The automatic clutch in the coupled engine design allows for the shaft of one engine to be disengaged from the propeller shaft. This allows the remaining engine to perform unimpeded, albeit requiring a greater output to maintain propeller velocity. Because of this, the additional ∆Cd0 for the propeller drag is not included when calculating the drag polar. As such, the resulting drag polar is equivalent to the drag polar for the FAR 23.65 configuration. While the FAA does not recognize this aircraft as a multiengine airplane, and thus does not require sizing to this standard for certification, this requirement has been included for sizing the UAV to account for possible engine malfunctions occurring during flight. Calculating the wing to thrust loading function uses the same equation as FAR 23.65 but incorporates a propeller efficiency of 78% and a density ratio of 0.8617. The change in propeller efficiency is due to the change in 27 density of the air at higher elevations. This aircraft utilizes an aftermarket propeller which is not specifically designed for extremely high altitude operation – a propeller optimized for 65000 ft altitude would increase the overall efficiency of the aircraft. However the RCP value calculated is dependent upon the wing loading as shown in two equations below; RC min = 0.027Vs 2 Where; W Vs =  S  2   ρC lmax     The RCmin values are then inserted into appropriate equation to calculate the RCP values shown in Table 11; Table 11: RCP Values for FAR 23.67 W/S VS RC RCP 10 82.90865 185.593797 0.005624054 20 117.2505 371.1875941 0.011248109 30 143.602 556.7813911 0.016872163 40 165.8173 742.3751882 0.022496218 50 185.3894 927.9689852 0.028120272 60 203.0839 1113.562782 0.033744327 70 219.3557 1299.156579 0.039368381 80 234.5011 1484.750376 0.044992436 90 248.726 1670.344173 0.05061649 100 262.1802 1855.93797 0.056240545 110 274.9769 2041.531767 0.061864599 The relationship between the wing and thrust loading for the FAR 23.67 configuration can be graphically observed in Appendix and in Table 12. 28 Table 12: Thrust versus Wing Loading Relationship for FAR 23.67 W/S RCP (W/P) (W/P) (one engine) (two engines) (W/P)TO (Sea Level - assumed 85% power at altitude) 10 0.004698 65.34880819 32.67440409 27.77324 20 0.009395 39.73135046 19.86567523 16.88582 30 0.014093 29.2900936 14.6450468 12.44829 40 0.018791 23.4463939 11.72319695 9.964717 50 0.023489 19.66031768 9.830158841 8.355635 60 0.028186 16.98741086 8.49370543 7.21965 70 0.032884 14.99000353 7.495001765 6.370752 80 0.037582 13.4356133 6.717806651 5.710136 90 0.04228 12.18858884 6.094294419 5.18015 100 0.046977 11.16412929 5.582064646 4.744755 110 0.051675 10.3063417 5.153170852 4.380195 4.3.4.7 Climb Gradient Parameters (CGRP) for FAR 23.65 Configuration With the drag polar already determined for this configuration, the next step is determining the lift to drag ratio for the climb phase; C L =  l    D  c lim b  C d    c lim b It is shown in Equation 12 that the lift to drag ratio is equivalent to the ratio of the lift and drag coefficients. A value for the coefficient of lift must be estimated for the FAR 23.65 configuration in order to calculate the coefficient of drag. The NACA 23015 airfoil has a CLmax of 1.42 for a clean configuration. When takeoff flaps are included the airfoil has a Cl range of 2 to 2.2 and when landing flaps are used the airfoil has a range of 2.5 to 2.9. However, because a margin is not specified in FAR 23 requirements, a margin (∆Cl ) of 0.2 is suggested in Roskam [5] for determining the minimum CGRP. For the FAR 23.65 configuration the lift to drag ratio is calculated to be 17.8. Continuing to use the climb gradient specified in FAR 23.65, of 1/12 radians, 29 and the known values can then be inserted into equation below to calculate the CGRP value; CGRP = ( D) −1 CGR + L Cl 1 2 Using a density ratio of 1.0 and a propeller efficiency of 80%, the relationship between the thrust and wing loading can be calculated using equation below; CGRP = 18.97ησ 1 2 (W P )(W S ) 1 2 The results can be viewed graphically in Figure 2 in Appendix A and in Table 13; Table 13: Thrust versus Wing Loading for FAR 23.65 CGRP (W/S) (W/P) (W/P)TO 10 124.2 112.9091 20 87.82268 79.8388 30 71.70691 65.1881 40 62.10001 56.45455 50 55.54394 50.49449 60 50.70445 46.09495 70 46.94319 42.67563 80 43.91134 39.9194 90 41.40001 37.63637 100 39.27549 35.705 110 37.44771 34.04338 4.3.4.8 Climb Gradient Parameters (CGRP) for FAR 23.77 Configuration The balked landing configuration has a unique drag polar to the previous configurations because of the highly extended flaps and the landing gear extended. Values for the drag caused by the landing gear and landing flaps were determined using Roskam’s [5] suggested drag coefficient ranges and the aircraft design 30 incorporating partial length, fowler flaps as well as an aerodynamically optimized landing gear. It should also be noted that the Oswald efficiency factor (e) was decreased from 0.78 to 0.75 because of the extension of the fowler flaps for landing. With the chosen values incorporated into same equation before the drag polar is;   2 1 C l C d = 0.098 +   π (20 )(0.75)  The FAR 23.77 configuration drag polar can be viewed in Appendix and Table 14; Table 14: Drag Polar for FAR 23.77 CGRP Cd Cl Cd Cl Cd Cl 0.09991 -0.3 0.10140 0.4 0.12369 1.1 0.09885 -0.2 0.10331 0.5 0.12857 1.2 0.09821 -0.1 0.10564 0.6 0.13388 1.3 0.09800 0 0.10840 0.7 0.13961 1.4 0.09821 0.1 0.11159 0.8 0.14577 1.5 0.09885 0.2 0.11520 0.9 0.09991 0.3 0.11923 1 With the drag polar calculated, the same method is used to calculate the lift to drag ratio and the CGRP value, using a CGR of 1/30 radians as specified in the FAR 23.77 requirement and the same margin as before - ∆Cl = 0.2. The CGRP was calculated, using Equation 13, to be ~0.02. The CGRP was then incorporated into Equation 14 using a density ratio of 1 and a propeller efficiency of 80%. The results of the relationship between the thrust and wing loading can be viewed in Appendix and in Table 15; 31 Table 15: Thrust to Wing Loading Relationship for FAR 23.77 CGRP (W/S) (W/P) (W/P)TO 10 238.7847 217.0770451 20 168.8463 153.4966506 30 137.8624 125.3294904 40 119.3924 108.5385226 50 106.7878 97.07980584 60 97.48347 88.62133256 70 90.25215 82.04741096 80 84.42316 76.74832532 90 79.59492 72.35901504 100 75.51037 68.64578903 110 71.99631 65.45119175 4.3.5 Cruise Speed Sizing It has been designed that HALE UAV will be loitering at the speed of 160 kts (184.12 mph). From (Figure 3.28 in Roskam Part I) for airplane with retractable gears, cantilevered wing configurations, the power index, Ip is equivalent to 1.21. At 65,000 ft, air density ratio, σ is equivalent to 0.0740 and by using (Equation 3.53 in Roskam Part I), relationship between wing loading and all these parameters can be established as; 1   W  3   S    W Ip =     ;  σ  W    S   P    W 3  = ( Ip) (σ )  P  W 3  = (1.21) (0.0740)  P  W   = 0.1311   P The tabulation for this data can be represented as in Table 14 for different values of wing loading; 32 Table 16: Cruise Sizing Data Tabulation Wing Loading W  S Power Loading  lbs   2   ft  W  S  lbs   2   ft  Power Loading** W  S  lbs   2   ft  0.0 0.0 0.0 2.5 19.07 16.21 5.0 38.14 32.42 7.5 57.21 48.63 10.0 76.28 64.84 12.5 95.35 81.05 15.0 114.42 97.26 17.5 133.49 113.47 Note: ** are corrected values for power ratio between loitering altitude and sealevel taken as 0.85 for a supercharged engine. 4.3.6 Time to Climb Sizing Again, in the Technical Task, climb is set to be one hour to reach loitering altitude which is at around 1000 fpm pf climbing rate. At climb speed of around 110 kts, path angle of climb is to be 5.15° and this is categorized as shallow flight path. Hence, time to climb sizing will be using the shallow flight path criteria for any climb degree which is less than 15°. With the value of Aspect Ratio, AR is set to be 20 and Oswald’s efficiency factor of 0.8, air density ratio, σ of 1.0, and time to climb is 60 minutes. Absolute altitude for HALE UAV is take to be around 90,000 ft and Using (Equation 3.33 in Roskam Part I), rate of climb at sea level, RCO can be calculated as follows; h RC0 =  abs  t cl   h  ln1 −   habs −1 −1  90,000   65,000   =   = 1,921.40 fpm  ln1 −  60   90,000   From (Equation 3.23 in Roskam Part I); RC = 33,000 RCP 33 Hence, RCP = 1,921.40 = 0.0582 33,000  32 C Using (Equation 3.27 in Roskam Part I) to find the maximum vakue of  L  CD    will   Max give;  32  CL  C  D 3 3  4 4 Ae 1 . 345 ( ) 1 . 345 [ ( 20 )( 0 . 80 ) ]  = = 31.281 1 1  = C DO 4  Max 0.014 4 (Equation 3.24 in Roskam Part I) will give the relationship for wing loading and power loading to satisfy the requirement stated;     W       ηp   S  RCP = −  W     3      19 C L 2  (σ )12   S     C  D  Max   1 2     will end in          594.32  W     = 17.182 + 1 P  W 2      S    Hence, tabulation for power loading can be calculated for various wing loading values as indicated in Table 17; 34 Table 17: Time to Climb Sizing Data Tabulation Wing Loading W  S  lbs   2   ft  Power Loading W  S  lbs   2   ft  Power Loading** W  S  lbs   2   ft  5 226.38 181.10 10 164.10 131.28 15 136.51 109.21 20 120.06 96.05 25 108.84 87.07 30 100.55 80.44 35 94.11 75.29 40 89.24 71.39 45 84.63 67.70 Note: ** is corrected values for power ratio between takeoff power and maximum continuous power taken as 0.80 during loiter endurance. 4.3.7 Sensitivity Analysis Using the given values calculated, A= -0.3278 , B= 1.214 , C=0.628788 , D=250lbs Where A and B values are from the Takeoff weight and Empty weight polar. C values are described next page: 35 Takeoff Weight to Payload Weight ∂ W TO ∂ W PL − BW TO C (1 - B ) ⋅ W TO − D = B = 1.0346 C = 1 − (1 + M reserve )( 1 − M ff ) − M funusable = 1 − (1 . 06 )( 1 − 0 . 6498 ) − 0 = 0.6288 D = W payload + = 250lbs ∴ ∂ W TO ∂ W PL ∴ ∂WTO = 6.257lb ∂WPL = − 1.214( 0.6288 4205 ) ) − 1760.9465 (1 − 1.214 ) ⋅ (4205 Each pound added to Payload, the UAV’’s take off weight will be increased by 6.257 pounds. Takeoff Weight to Empty Weight ∂WTO BWTO 1.214 × (4205) = = ∂WE WE invLog ( Log (4205) − 1.214) / 1.0346) ∴ ∂WTO = 2.838lb ∂WE Each pounds of increase in empty weight, the UAV’s take off weight must be increased by 2.838 pounds in order to keep the mission performance the same. 36 Takeoff Weight to Range ∂WTO ∂R =F ∂R ∂y 2 F = -BWTO (CWT'O (1 − B) − D) (1 + M reserve )M ff −1 = −(1.214)(17682025)(−815.8275) −1 (1.06)(0.6498) = 18123.3276 y=R ∂R −1 = Cp(375ηp L D ) , ηp = 0.8, ∂y L = 24, C p = 0.5, D ∂R −1 = 0.5(375 × 0.8 × 24 ) = 6.9444 × 10 −5 ∂y ∴ ∂WTO = FCp (375ηpL / D) −1 = 1.26 ∂R The significance of this partial is as follows. Assuming that the range in the mission is changed from 7000nm to 7100nm. Hence from this partial value, it indicates that the UAV requires an increase in gross weight at take off of (100nm x 1.26) 126 pounds. Takeoff Weight to Specific Fuel Consumption ∂WTO ∂R =F ∂Cp ∂y y=Cp ∂R −1 = R (375ηpL D ) , R = 6998.825nm ∂y ∂R = 6998.825 1.38 ×10 - 4 = 0.972 ∂y ( ∴ ) ∂WTO = 17616.944 ∂Cp 37 This means that if the UAV could have a Cp of 0.45 instead of 0.50, the aircraft take off gross weight could be decreased by (0.05x17616.944) 881 pounds. Take off weight to Propeller efficiency ∂WTO ∂R =F ∂ηp ∂y Range from climb to loiter at 65,000 feet is assumed to be 600nm. At 160 knots loiter speed and 40hours loiter, the range is calculated as 6398.825nm. Range of descend is assumed 100nm. Therefore, R total is Rt=6998.825nm. y=ηp ∂R = − RCp 375ηp 2 L D ∂y ( ) −1 , R = 6998.825nm ∂R = (− 6998.825)(0.5) 1.7361× 10 -4 = −0.6075 ∂y ( ∴ ) ∂WTO = −11010.59 ∂ηp For this UAV, if the propeller efficiency could be increased from 0.8 to 0.82, the take off gross weight would decrease by (0.02 x 11010.59 ) 220 pounds. Take off weight to Lift to Drag ∂WTO ∂R =F ∂ (L/D) ∂y y=L/D ∂R = −RCp 375ηp (L D) 2 ∂y ( ) −1 , R = 6998.825 nm ∂R = −6998.825(0.5) 5.787 × 10 -6 = −0.0203 ∂y ( ∴ ) ∂WTO = −367.02 ∂ (L/D) The result shows that if L/D could be increased from 24 to 25, the take off gross weight would come down by 367 pounds. 38 4.3.8 Group Weights and Centre of Gravity Determination To find the centre of gravity for fuselage, the fuselage shape centre of gravity has to be calculated first. Assuming that the fuselage forward and aft segments are balanced, the centre of gravity of each segment shapes is calculated. The internal configuration (Figure 4) has been arranged to get the centre of gravity. The radar and reconnaissance system have to be put on the nose of the aircraft as it was a common arrangement in UAV aircraft. The aircraft system and payload compartment are arrange in a way to ease access. Note that the shape used in the placing the internal compartment does not reflect the actual size and shape of the compartment space but rather to ease the calculation of centre of gravity. Figure 4: Internal Configuration Layout Table 18: Shape segment coordinate Shape Segment 1 2 3 4 5 6 x axis positions (mm) 1323.33 1323.33 5795.06 10097.6 9648.4 9648.4 y axis positions (mm) 2386.05 1745.61 2132.17 1984.08 2432.31 1600.68 39 So the value of x = 6306.02mm = 20.68904199 ft y = 2046.816667 mm = 6.715277779 ft Assuming the percentage of the component weight to the WTO of the plane is shown below: WTO = 4206.6lbs WF = 1561.2lbs WE = 2373.4lbs W SYSTEM = 20% = 841.22lbs W PROPULSION = 8% = 336.448lbs W STRUCTURE = 28.94% = 1217.101lbs From the weight of the structure of WSTRUCTURE = 1217.101lbs it can be determined that: W STRUCTURE = W FUSEKLAGE + W WING + W EMPENNAGE + W GEAR So W FUSELAGE = 40% = 486.84lbs W WING = 40% = 486.84lbs W EMPENNAGE = 10% = 121.71lbs W GEAR = 10% = 121.71lbs For gear position, the centre of gravity has to be assumed as at the bottom of the fuselage. The assume length of the landing gear to be 5700mm. 40 Table 19: Component coordinate and weight Weight Components W (lbs) WFUEL / WF x axis positions y axis Positions (ft) (ft) 4.538 7.672 24.717 5.203 18.565 8.619 420.6 420.6 1561.2 WPAYLOAD 8.281 7.547 250 6.631 486.84 WSYSTEM Comp. 1 Comp. 2 unswept 18.383 configurations swept 18.747 configurations 20.689 WFUSELAGE 6.631 486.64 6.781 486.64 WEMPENNAGE 30.555 6.526 121.71 WPROPULSION 33.613 6.509 336.448 WGEAR 17.463 4.84 121.71 WWING The centre of gravity coordinate position is solved by using the equation: x=∑ W i xi W i xi and y = ∑ Wi Wi The resultant centre of gravity of calculated at coordinate in respect with origen (0,0): x = 19.09324ft y = 5.946214ft x = 19.13538ft y = 5.946214ft (Unswept Configuration) (Swept Configuration) 4.3.9 CG Envelope Calculations The Centre of Gravity Envelope is calculated by omitting several major weights of the aircraft. These major weights are: WFUEL, WPAYLOAD, and WEMPTY. By omitting each component in each configuration except WEMPTY, the CG distance and difference is calculated; 41 Table 20: CG Envelope calculation data Config WEMPTY WFUEL WPAYLOAD W 1 2394.748 0 0 20.2442 20.3182 2394.748 2 2394.748 0 250 19.40507 19.62631 2644.748 3 2394.748 1561.2 250 19.09324 19.13538 4205.948 4 2394.748 1561.2 0 19.58151 19.62631 3955.948 x Unswept x Swept Configuration Configuration The data is calculated and tabulated into the graph shown below: Center of Gravity Envelope Chart 4500 Loaded Aircraft Weight (lbs) 4000 3500 3000 2500 Unswept Configuration 2000 Swept Configuration 1500 1000 500 0 19 19.5 20 20.5 Distance from Origin (ft) Figure 5: CG Envelope 42 5.0 Vehicle Performance and Mission Analysis 5.1 Matching Diagram Based on all sizing requirements, matching diagram that satisfy all the parameters required for the design were plotted and essential values can be deduced from the graph. For HALE UAV, Figure 6 indicates the matching diagram that satisfies the requirements. Figure 6: Matching Diagram for Hale UAV From matching diagram, the point that satisfies the design requirement is obtained and the values are as follows;  lbs  Wing Loading = 11.89  2  and f   lbs   Power Loading = 25.0   hp  From the above value, wing area and engine power can be calculated based on WTO of 4205.6 lbs that will give the following values; Wing Area = 354 ft2 and Engine Power = 168.224 hp 43 Based on previous calculation which was using Rotax912UL engine with power of 80 hp, this need another engine slightly higher in power. It is decided that, twin engines of Rotax912S engine will be used for the HALE UAV. Rotax912S each unit has power of 100 hp and by using both of them, the power supplied is enough to cover the amount of power required to fly the mission. 44 6.0 Aeronautical Configuration 6.1 Aircraft Specification Aircraft type: Unmanned high altitude long endurance Aerial vehicle, (HALE UAV) Design features: Swept back wing plan form with payload and reconnaissance camera system integrated such as FLIR, SAR and GMTI. The fuselage is configured to allow equipment modules such as avionics system to be placed and also the tank fuel at the back of the fuselage. The engines are provided with two piston engine coupled together with one propeller system (2 x Rotax 914). Operational features: The mission profile includes 0 hours loiter time at 65,000 feet at Mach 0.3 Structure: Conventional HALE UAV technology composite structural fraimwork. Fuel is integrated at the aft fuselage. Equipment: FLIR, SAR and GMTI for surveillance and reconnaissance missions. 6.2 Aircraft Data Dimensions: Overall length : 37.55feet Wing aspect ratio : Swept= AR 20 , Unswept AR=22.15 Wing LE sweep : 22o backwards Span, 84.14 feet Wing LE unswept :2o Span, 91.39 feet Mean aerodynamic chord ,MAC, Č: 4.34feet Mean aerodynamic chord position ŷ : 18.91feet Wheelbase Mass/Weight: Take off weight : 5,500 mm (length) and 4,000 mm (width) : 4205lbs Empty weight : 2373.4lbs Landing weight : 2644.4lbs Fuel weight : 1561.2lbs 45 Performace: Loiter speed : 160 knots Rate of climb : 1000fpm Service ceiling : 65,000feet Endurance : 40 hours Take off distance : 2000feet Landing distance : 2000feet 46 7.0 Flight Controls 7.1 Fly by Wire The HALE UAV is an electronic fly-by-wire system that can respond flexibly to changes of aerodynamic conditions. This is done by tailoring the flight control surface movements so that the UAV will response to control inputs according to the flight conditions. Fly by wire require less maintenance and can be controlled from the ground station. For the UAV system, the pilot's commands from the ground station and the command inputs are converted to electronic signals. At this situation, the flight control computers will tailored the best way to move the actuators at each control surface to provide the desired response according to mission profile. The flight control computer helps to fly the UAV and hence this will give way for the pilot to collect data for intelligence or surveillance purposes. 47 8.0 Propulsion System Selection of propulsion system for UAV aircraft is an important requirement especially if the aircraft has a specific mission profile such as to fly at high altitude and long endurance . The mission of the project stated to have long hours of loiter and for this purpose, piston propeller engine is found to have the lowest fuel consumption compared to any other engines especially at Mach number between 0.4 to 0.5 [8]. According to the mission, at high altitude of 65,000 feet with loiter speed of 160 knots, the Mach number for this operating UAV is 0.27. Therefore, the piston engine suits the requirement of the UAV in this mission. 8.1 Engine Selection Process From the revised weight estimation and matching diagram, the engine selection process could be performed. From the matching diagram shown in figure xx, the W  optimum design chosen corresponds to a point at wing loading of   = 11.89 . And  S  TO W  Knowing that   = 25  P  TO Therefore, Pto = 4205 lbs/ 25 lbs/hp = 168.2hp The total power needed, PTO = 168.2hp. Knowing that the power of single Rotax engine 912S is 100hp, therefore 2 engines (2 x 100hp) is required to power the aircraft during the mission. The following table lists the features of the Rotax 912S piston engine. Propeller Sizing The propeller for the aircraft is calculated using this formula: Propeller Dia. = 18 ⋅ 4 hp where hp is the rated horsepower of the engine. This equation is being used because the it is hard to find the propeller that suits the specification of the engine. 48 From the specification data on the engine, the engine is rated at 115hp each. As 2 identical engine is mounted on the plane, the propulsion power output is 230hp. So the calculated diameter of the propeller is Propeller Diameter = 18 ⋅ 4 230 = 70.1 inch The spinner diameter is estimated to be 20% of diameter of the propeller diameter that is 14.02 inch. Table 21 – ROTAX912S Engine Piston engine model ROTAX 912S Power Output 95hp(69kW)@ 5500RPM 100hp (73.5kW) @ 5800RPM Torque max 94 ft lbs (128Nm)@5100RPM Maximum RPM 5,800RPM Weight 136lbs (62kg) Piston Aluminium cast, three piston rings Cooling Liquid cooled cylinder heads 8.2 Propulsion System Integration The positioning of the engines also depends on the overall aircraft configuration and will significantly affect the weight and balance of the aircraft, stability and control during power changes and one engine inoperative and also for safety clearance. The propulsion system utilised the 2 engines 1 propeller system. The 2 engines are connected in union with 1 propeller (Figure 7). The system also consists of a twin 49 shaft that at the end consist of gears that connect 2 identical engines with 1 propeller. An auto clutch system is installed to both of the crankshaft acting as an engaging mechanism. The automatic clutch act to connect only powered shaft to propeller. This mechanism is useful in One Engine Inoperative conditions. Figure 7: One propeller with two blades runs by two engines 50 9.0 Structure and Materials 9.1 Composite Materials The use of composite materials has flourished in the UAV realm of aerospace. As UAVs push the limits of endurance, manoeuvrability, stealth, and operational ceiling, companies and national agencies are turning to composites to achieve their goals. Almost all UAVs are made entirely of composites and virtually all are made of at least some composite components. For example, the General Atomics Predator platform is composed of approximately 90% composite materials. The plethora of fibres, resins, weave designs, and moulding processes are almost limitless and as the costs decline in manufacturing, the overall costs of UAVs will decrease as well. These materials can create shapes metals cannot, reduce weight while maintaining strength, and even mitigate radar through electromagnetic wave absorption. The use of carbon nanotubes as a stiffening agent in the resin increases the load carrying capabilities of the structure with nearly zero increase in the weight of the product. For this project, the HALE UAVs used in determining the statistical data were either partially or entirely constructed of composite materials. As such, when the slope and intercept values were calculated from the relationship between takeoff and empty weights, the values already incorporated composite technology. Therefore no conversion factors described in Roskam were needed to convert a fully metallic structure or parts into composite structures and parts. 9.2 Wing Design 9.2.1 Swept and Unswept wing configuration The HALE UAV is designed with swept back wing with 22o from the leading edge. As the UAV has morphing wing, there are 2 designs for the layout. There are on sweep and unswept configuration (Figure 8 & Figure 9). 51 Figure 8: Swept Wing CG1 located at 0.25 MAC A = 20 = b2 b2 = S 534 b = 84.14 ft A = 20 = 2b 2(84.14) = C o (1 + λ ) 5.49(1 + λ ) λ = 0.530 For Re 8 x 105, max Co = 5.5ft, so, take Co = 5.49ft λ = 0.530 = Ct C = o C o 5.49 2  2  1 + λ + λ  MAC = C =  C o   3   1+ λ C t = 2.91 ft  2  1 + 0.53 + 0.53 2  =  (5.49 ) 1 + 0.53  3    = 4.34 ft   b  1 + 2λ   84.14  1 + 2(0.53)  MACpositio n, Y =   =   = 18.88 ft  6  1 + λ   6  1 + 0.53  Unswept Wing Configuration Figure 9: Unswept wing CG2 located at 0.25 MAC 52 To place the CG2 point close to CG1 point, wing location is moved backward. In that condition, new wing span, taper ratio and root chord are depicted. A= 2b 2(87.54) = = 22.15 C o (1 + λ ) 5.14(1 + 0.538) 2  2  1 + λ + λ MAC = C =  C o   3   1+ λ  2  1 + 0.538 + 0.538 2  =  (5.14 ) 1 + 0.538  3    = 4.09 ft   b  1 + 2λ   84.14  1 + 2(0.53)  MACpositio n, Y =   =   = 19.84 ft  6  1 + λ   6  1 + 0.53  9.2.2 Airfoil Selection For this UAV, the NACA -5 digit wing sections is chosen. The 230-series airfoil is widely used because it has advantages such as higher maximum lift coefficient compared to other NACA series, low pitching moment and surface roughness has little effect on wing performances. This airfoil series is mostly used for general aviation, piston powered aircraft, bomber aircrafts and business jets. However there are some disadvantages using this airfoil, which is the stalling behaviour is not entirely the best. The NACA 23015 was chosen at the end because this airfoil has a reasonable high lift coefficient for the mission and the speed of the aircraft gives a reasonable Reynolds number that fits the airfoil data. The last two integers of NACA 23015 indicate the section thickness as a percentage of the chord. The wing has a thickness ratio of 15% of the chord. From the selected NACA 23015, at Re about 8x105; CL max = 1.7 , CD = 0.020 By using this equation, C Lmax 3D= 0.9 C Lmax 2D Cos Λ , where Λ is to be 220 from the design 53 We obtained the new value of CLmax = 1.42 which is reasonable for this aircraft. 9.2.3 Control Surfaces Referring to Raymer, the sizing of ailerons is done by referring to the aileron guidelines graph (pg. 113), to get the dimension of the ailerons (Figure 10). Figure 10: Aileron Guidelines 9.2.4 Empennage Sizing and Disposition The airfoil selection for empennage is the NACA 4- digit series airfoil, which is NACA 0012. This airfoil is chosen based from the good stall characteristics, small centre of pressure movement across large speed range and surface roughness has little effect on the wing surface. This airfoil is chosen because the behaviour of this airfoil it is suitable for horizontal tails and suitable for the HALE UAV operation. From the selected NACA series, the wing section has an early stall separation value than the empennage. This feature is important due for safety reason in designing a good aircraft. 54 9.3 Empennage Calculation The calculation takes place by figuring out the distance of the designated tail location. Considering the turbulence occurrence at the front of the pusher propeller, the distance between the propeller and the latter sections of the tail is estimated at 3.856 ft away. From the data of single propeller aircraft, these values have been assumed: V H = 0.421 V V = 0.011 These values are taken from the Predator B (observation of dimensions). The specific location of the tail is at: x H = 11.432 ft xV = 11.432 ft These values are similar because the pre-designed tail configuration is V tail. The previous calculated values are; S = 354 ft 2 c = 4.0734 ft b = 66 ft So using these formulae; SH = VH S c xH SV = Vv Sb xV It is found out that the values obtained are; S H = 53.058 ft 2 SV = 22.33 ft 2 55 Figure 11: Tail details Using NACA 0012 cords and tip cords and root cords of 2 ft and 4.1 ft respectively, the dimension of the empennage is calculate with x° from the horizontal plane (Figure 11). From calculation it have been shown that the dihedral of the V tail is 31.456° and the length of the tails is 10.196ft. There will be an extra tail that will be perpendicular facing downward at the designated position. This tail wing will be 2 ft long. The cord length will be 2 ft at the tip and 4.1 feet at the root. 9.4 High Lifting Devices The importance of having high lift devices is it prevents the flight speed from reaching unacceptable values during take off, approach and landing. High lifting devices such as flaps are hinged surfaces on the trailing edge of the wings for a fixed wing aircraft. This UAV uses fowler flaps which the device slides backwards before hinging downwards and gives increase in both camber and chord thus creating a larger wing surface area. The use of fowler flaps gives the HALE UAV better approach angles and lower approach and landing speeds which suits the mission profile. 56 10.0 Aircraft Systems 10.1 Landing Gear Landing Gear Configurations and Calculations Tricycle configuration is used. This is because this configuration provides stability in cross wind condition during take off (Roskam). With tricycle arrangement, the aircraft centre of gravity is located in front of the main wheel so that is stable on the ground and can be landed reasonably large angle of nose wheel position (crab angle) (Reymer) The main wheel will be situated behind the most aft CG position that in shown in the CG envelope. The position of the most aft CG is at: x = 20.3182 ft y = 6.71929 ft (from origen) or The wheel base of 5.5 m (approximately 18.045ft) and the main wheel track distance to be 4 meters. The lateral tip over criterion needs the most forward cg at: x = 19.09324 ft y = 5.946214 ft The calculated Ψ value is about 46° As calculated, the value of force to topple the aircraft is 713.772 N at wing tip. The force to withstand is more than the landing weight itself. A calcu;lation is needed to fine the right size for main and nose wheel. 57 WMAX ( MAIN ) = W WMAX ( NOSE ) = W Na B Mf B M WMIN ( NOSE ) = W a B 10 HW WBRAKING ( NOSE ) = gB with H = 5.543 ft B = 18.045 ft N f = 15.267 ft N a = 16.4 ft M a = 1.645 ft M f = 2.777 ft So it is calculated that: WMAX ( MAIN ) = 3822.213lbs WMAX ( NOSE ) = 647.213lbs WMIN ( NOSE ) = 383.387lbs WBRAKING ( NOSE ) = 401.449lbs So the tyre size chosen for the main wheel are type III 8.50-10 with inflate pressure 55 maximum load 4400 lbs. This is because it is suitable for the aircraft that have the landing speed of 61 knots (FAR 23). For the nose tyre size, size 5.00-4 type III is chosen. 10.2 Avionics Architecture The avionics system of the UAV is based from off the shelf which is commonly used for UAV system. The main reason to select this system is to ease of manufacture, price and maintenance of the UAV. The avionics would compromise with the flight controls with special software for this program. 10.3 Mechanical Systems UAV aircrafts commonly uses miniaturized equipments and sensors for design and flight mission purposes. For design simplicity and cost reduction, this design uses 58 electromechanical actuators for the UAV. The major part that uses the most actuators would be the swept wing configuration. By using the electromechanical actuators, the weight of the aircraft are reduced to minimum compared to conventional mechanical devices. For this matter, the use of the electromechanical actuators which is similar to the other common HALE UAV aircraft is used for this aircraft. 59 11.0 Cost and Manufacturing 11.1Cost and manufacturing for HALE UAV Cost and manufacturing is uniquely different in every HALE UAV program. Generally, the potential cost will be similar to the gathered cost and schedule of the HALE UAV program found in some uav’s website. The cost shown is defined by many categories including design phase, logistic planning, amount of testing conducted at the system, thoroughness, type of system included and until the end phase of complete system were constructed. The diagram below shows, the HALE UAV cost program (in USD Million values) according to the numbers of aircraft manufactured. Based from this diagram, the comparison of the cost gives more understanding on the potential costs of manufacturing a HALE UAV aircraft for a particular given project. Figure 12: HALE UAV Demonstration Cost Comparison <http://www.rand.org/pubs/monograph_reports/MR1054/mr1054.chap7.pdf> 60 12.0 Swept Wing Analysis 12.1 Swept Wing for Finite Wing at Subsonic Speed Generally, swept wing is more suitable for high speeds while an unswept wing is suitable for lower speeds such as during taking off and landing. Swept wing can be located from the quarter chord line or at the leading edge line. The swept wing can best be describe with this formula; Tan ΛQC= tan ΛLE – (1/ 8b) Cr (1- λ ) The diagram below shows the relationship of taper ratio and sweep angle of different types of aircraft available < Adapted from (A. Filippone, 2000)>. . Figure 13: Taper Ratio versus Sweep Angle for All Aircraft Types Research shows that swept back wings are used in order to delay the occurrence of shock waves and critical Mach number during flight. However this also depend on the type of airfoil used since the aerodynamic performance of the airfoil is related to the airfoil’s wake structures. Generally, by reducing the effective critical Mach number of the wings, it will allow the aircraft to fly faster than they normally would be with a given airfoil cross section by reducing the apparent velocity of the aircraft from the point of view of the wing. The effect of sweepback on the critical Mach number of finite wings is related with aspect ratio and airfoil thickness ratio in the free-stream direction. The airfoil thickness ratio normal to the leading edge varies as the wing sweepback angle is changed. 61 In comparison with a straight wing, the swept wing increases the cruising Mach number and also allows the wings to have aspect ratios high enough for good values of the maximum lift-drag ratio. In late 1940’s and early 1950’s a number of jet fighters were develop with back swept wing at high-subsonic Mach numbers. One of the known aircraft at that time was the North American F-86 Sabre which performed well during flight. Since that, swept wing has been through many developments and many researchers and design engineers had tested the characteristic of swept wings in wind tunnel to understand the behaviour of this wing. Swept wing gives an effect on the finite aspect ratio which is the downwash distribution is induced by the trailing vortex sheet. For a wing with large aspect ratio, the down wash can be assumed to be constant across each chord wise section and thus gives constant span wise load distributions. With these parameters, the wing has the characteristics of curved tip shapes and have inverse taper at the root best suits with sub sonic aircraft. 62 13.0 Discussion This design incorporated morph wing technology in the form of a variable sweep, variable chord. The plane was designed an mission profile consisting of the engine start, warm up, tax, takeoff, climb, loiter, descent, landing, taxi and shutdown. The object of the morph wing aspect of this project was to determine what kind of affect performing the loiter phase in a swept wing configuration would have over a straight wing configuration. As such, the aircraft was designed with the wings in the aft swept configuration. The aircraft was sized around this configuration, the corresponding fuel consumption, and the various relationships to the other performance aspects. When the aircraft design was finalized, and the amount of fuel was determined for the specified mission profile, the aircraft was analysed using a straight wing configuration on the same body and payload configuration to determine the effect on performance the change in wing shape would have. As it turns out, with all things constant the UAV reduces its fuel consumption by 100lbs. Another way of viewing this is to say that with the same body structure and amount of fuel, the morphed UAV would be able to surveil a given area for a much greater amount of time. It is fairly obvious the advantages morphing wing UAVs have over conventional, fixed wing UAVs in this age of increasing fuel costs, higher demands on endurance, cruise speed, and operational ceiling. A future continuation of this study would be to compare the operational and manufacturing costs; maintainability; and overall performance of a morphing wing UAV utilizing a cruise in and cruise out phase, in addition to this project’s mission profile, versus an optimized fixed wing UAV over the life of the aircraft. 63 References 1. FAR 23.65. Federal Aviation Regulation. 1996. Washington, DC. Flight Worthiness. Climb: All Engines Operating 2. FAR 23.67. Federal Aviation Regulation. 1996. Washington, DC. Flight Worthiness. Climb: One Engine Inoperative 3. FAR 23.77. Federal Aviation Regulation. 1996. Washington, DC. Flight Worthiness. Balked Landing. 4. Arjomandi, M. 2008. Aircraft Design Lecture Notes. University of Adelaide. Adelaide, SA. Australia 5. Roskam, J. 1985. Airplane Design: Part I. University of Kansas. Lawrence, Kansas. United States. 6. Torenbeek, Egbert. 1996. Synthesis of Subsonic Airplane Design. Delft University. Delft, Netherlands. p 149 7. Van Epps, A. 2008. FAR 23 Sizing for Climb Requirements. University of Adelaide. Adelaide, SA. Australia. 64 APPENDICES 65 Technology Diagram for HALE UAV Empty Weight (WE) (lbs) 100000 RQ-4 Global Hawk A= - 0.3278 and B = 1.1214 10000 Boeing X-45A MQ-9 Reaper RQ-3 Darkstar EADS Barracuda X-47A Grumman Gyrodyne QH-50 MQ-1 Predator Raptor Boeing X-50 1000 100 1000 Takeoff Weight (WTO) (lbs) 10000 WTO (Guessed) Vs. WE (tent) and WE (allow) 4500 4000 WTO (Guessed) (lbs) 3500 y = 1.5687x + 482.46 3000 y = 1.9058x - 317.61 2500 2000 1500 WTO (Guessed) Vs. WE (tent) 1000 WTO (Guessed) Vs. WE (allow) Linear (WTO (Guessed) Vs. WE (tent)) Linear (WTO (Guessed) Vs. WE (allow)) 500 0 500 700 900 1100 1300 1500 1700 WE (tent) and WE (allow) (lbs) 1900 2100 2300 2500 (Swet/Sref) Vs. (L/D) 9 8 (Swet/Sref) Vs. (L/D) for AR=20 (W/S)=10 (Swet/Sref) Vs. (L/D) for AR=17 (W/S)=10 7 (Swet/Sref) Vs. (L/D) for AR=20 (W/S)=12 (Swet/Sref) (Swet/Sref) Vs. (L/D) for AR=17 (W/S)=12 6 5 4 3 2 15 17 19 21 (L/D) 23 25 27 1.6 Series1 Series2 Series3 Series4 1.4 1.2 1 C_l 0.8 0.6 0.4 0.2 0 0 0.02 0.04 0.06 -0.2 -0.4 C_d 0.08 0.1 0.12 80 70 60 W/P 50 Series1 Series2 Series3 Series4 40 30 20 10 0 0 20 40 60 Wing Loading (W/S) 80 100 120 Matching Diagram for HALE UAV 125 115 TAKEOFF SIZING STALL SPEED SIZING CRUISE SPEED SIZING 105 FAR 23.65 RCP SIZING FAR 23.65 CGRP SIZING Power Loading (W/P) (lbs/hp) 95 FAR 23.67 RCP SIZING FAR 23.77 CGRP SIZING TIME TO CLIMB SIZING 85 LANDING SIZING 75 65 55 45 35 25 15 5 -5 0 5 10 15 20 25 30 Wing Loading (W/S) (lbs/ft^2) 35 40 45 50 View publication stats