AIAA UNDERGRADUATE TEAM DESIGN COMPETITION PROPOSAL PDF Free Download

TADPOLE AIAA UNDERGRADUATE TEAM DESIGN COMPETITION PROPOSAL 2017 Conceptual Design of TADPOLE Multi-Mission Amphibian MIDDLE EAST TECHNICAL UNIVERSITY 5-10-2017

Team Member AIAA Number Contact Details Signature Koyuncuoğlu, Heyecan Utke 510055 heyecanutke@gmail.com Mert, Ümmehan 820119 mert.ummehan@gmail.com Tikenoğulları, Alp 820024 alp.tiken.ogullari@gmail.com Yutük, Kaan 820120 kaanyutuk@gmail.com Team Advisor AIAA Number Contact Detail Signature Prof. Dr. Serkan Özgen 327810 sozgen@ae.metu.edu.tr 1

Table of Contents List of Figures... 6 List of Tables... 7 List of Abbreviations... 8 List of Symbols... 9 1. CHAPTER 1... 12 1.1. Introduction... 12 1.2. Missions and Requirements... 12 1.3. Competitor Study... 15 1.4. Conclusions... 16 2. CHAPTER 2... 17 INITIAL SIZING... 17 2.1. Introduction... 17 2.2. Mission Profile... 17 2.3. First Estimation of Take-off Gross Weight... 17 2.3.1. Passenger Mission... 17 2.3.2. Low Cost Maritime Surveillance Mission... 19 2.4. Trade Study... 20 2.5. Conclusions... 21 3. CHAPTER 3... 23 Airfoil and Wing Planform Selection... 23 3.1. Introduction... 23 3.2. Design Lift Coefficient Calculation... 23 3.2.1. Passenger Mission... 23 3.2.2. Low Cost Maritime Surveillance Mission... 24 3.3. Selection of Airfoil and Aspect Ratio... 24 3.4. Estimation of Wing Sweep, Taper Ratio and Twist Value... 25 4. CHAPTER 4... 26 DESIGN PARAMETERS... 26 4.1. Introduction... 26 4.2. Calculation of Thrust-to-Weight Ratio... 26 4.3. Calculation of Wing Loading... 26 4.3.1. Take-off... 26 2

4.3.2. Cruise... 27 4.3.3. Loiter:... 28 4.3.4. Landing:... 29 4.4. Conclusions... 29 5. CHAPTER 5... 30 Refined Sizing... 30 5.1. Empty Weight Fraction Calculation... 30 5.1.1. Passenger Mission... 30 5.1.2. Low Cost Maritime Surveillance Mission... 32 6. CHAPTER 6... 35 Geometry Sizing and Configuration... 35 6.1. Calculation of Fuel Volume... 35 6.2. Fuselage Length... 35 6.2.1. Historical Trend... 35 6.2.2. Competitor Data... 35 6.3. Sizing and Planform of the Wing, Aerodynamic Parameters... 36 6.4. Tail Configuration and Size, Aerodynamic Parameters... 37 6.5. Estimation of Weight and Dimensions of the Engine... 38 7. CHAPTER 7... 39 CG Estimation, Landing Gear Sizing and Placement... 39 7.1. Decision for Location of Fuel... 39 7.2. Sizing of Fuselage... 39 7.3. Estimation of Weights of Major Components... 39 7.4. Estimation of Center of Gravity... 41 7.4.1. Longitudinal Center of Gravity... 41 7.4.2. Vertical Center of Gravity... 42 7.5. Sizing and Location of Propeller... 43 7.6. Calculation of Static Tail-Down Angle, Tip-Back Angle and Overturn Angle... 44 8. CHAPTER 8... 45 Aerodynamics... 45 8.1. Lift Curve Slope... 45 8.2. Maximum Lift Coefficient Calculation... 46 8.2.1. Maximum Lift Coefficient at Clean Configuration... 46 3

8.2.2. Maximum Lift Coefficient with High Lift Devices... 47 8.3. Drag Divergence Mach Number... 48 8.4. Total Parasite Drag and Wave Drag... 48 8.5. Calculation of Induced Drag Factor... 50 9. CHAPTER 9... 51 THRUST CURVES... 51 9.1. Introduction... 51 9.2. Uninstalled Thrust... 51 9.3. Installed Thrust Corrections... 53 9.3.1. Engine Related Losses... 53 9.3.2. Propeller Related Losses... 54 9.4. Installed Thrust and Propeller Efficiency... 55 9.5. Conclusion... 56 10. CHAPTER 10... 57 WEIGHTS AND LOADS... 57 10.1. Update of Take-off Gross Weight... 57 10.1.1. Passenger Mission... 57 10.1.2. Low Cost Maritime Surveillance Mission... 58 10.2. V-N Diagram... 59 10.2.1. Stall Line... 59 10.2.2. Corner Speed... 59 10.2.3. 1-G Stall Speed... 60 10.2.4. Dive Speed... 60 10.3. Gust Loads... 60 11. CHAPTER 11... 62 STABILITY AND CONTROL... 62 11.1. Longitudinal Stability Analysis... 62 11.2. Variation of Neutral Point with Mach Number... 62 11.3. Variation of Static Margin with Mach Number... 63 11.4. Variation of CmaPoint with Mach Number... 65 12. CHAPTER 12... 67 SIZING MATRIX AND CARPET PLOT... 67 13. CHAPTER 13... 71 4

TRIM ANALYSIS... 71 14. CHAPTER 14... 72 COST ANALYSİS... 72 14.1. Introduction... 72 14.2. Analysis... 72 14.2.1. Research, Development, Test and Evaluation... 74 14.2.2. Flyaway... 74 14.2.3. Operations and Maintenance cost... 76 15. CHAPTER 15... 77 DESIGN SUMMARY AND DRAWINGS... 77 15.1. Special Considerations... 77 15.2. Design Summary... 77 15.3. Drawings... 78 16.... 81 17.... 81 18.... 82 19. REFERENCES... 83 5

List of Figures Figure 1: Simple Cruise Mission Profile... 17 Figure 2: Fuel Burn per Passenger with Increasing Number of Passenger for Passenger Mission... 20 Figure 3: Endurance and Range Change with Increasing Speed Maritime Surveillance Mission... 20 Figure 4: Fuel Burn per Hour with Increasing Speed for Maritime Surveillance Mission... 21 Figure 5: Drawing of the Fuel Tanks in the Wing... 39 Figure 6: Lift Curve Slope for Subsonic Flight... 46 Figure 7: CLmaxValues for the Mach Numbers between 0.2 and 1.0 by adjusting... 47 Figure 8: Uninstalled Thrust vs. Mach Number Chart... 52 Figure 9: Propeller Efficiency vs. Mach Number by Uninstalled Thrust... 52 Figure 10: Installed Thrust vs. Mach Number... 55 Figure 11: Propeller Efficiency vs. Mach Number by Installed Thrust... 56 Figure 12: V-n Diagram... 61 Figure 13: Variation of Neutral Point with Mach Number... 63 Figure 14: Variation of Static Margin for Most Forward and Most Backward CG Cases... 64 Figure 15: Variation of Pitch Stiffness Derivative for Most Forward and Most Backward CG Cases... 66 Figure 16: Wing Area vs. Gross Weight at T=4135 hp... 68 Figure 17: Wing area vs. Gross Weight at T=2750 hp... 68 Figure 18: Wing area vs. Gross Weight at T=2100 hp... 69 Figure 19: Carpet Plot... 69 Figure 20: 3D View... 78 Figure 21Fuselage Structure... 79 Figure 22: Wing Structure... 79 Figure 23: Tail Structure... 80 Figure 24 Horizantal Tail Structure... 80 Figure 25: Front View... 81 6

List of Tables Table 1: Comparison of the Characteristics According to Design Requirements... 15 Table 2: Important Characteristics of the Competitor Aircrafts... 15 Table 3: Important Performance and Design Characteristics of the Competitor Aircrafts... 16 Table 4: Composite Material Trade for Passenger Mission... 21 Table 5: Composite Material Trade for Maritime Surveillance Mission... 21 Table 6: Aerodynamics Parameters of Similar Airfoil... 25 Table 7: Variables and Wing Loading Values for Both Missions... 27 Table 8: Parameters [2]... 30 Table 9: Wing Geometric Characteristics... 36 Table 10: Wing Geometric Characteristics for Horizontal and Vertical Tail... 38 Table 11: Planform and Wetted Areas for Components... 40 Table 12: Weights of the Major Components... 40 Table 13: Longitudinal Center of Gravity Location Calculation for Passenger Mission... 41 Table 14: Longitudinal Center of Gravity Location Calculation for Maritime Surveillance Mission... 41 Table 15: Vertical Center of Gravity Location Calculation for Passenger Mission... 42 Table 16: Vertical Center of Gravity Location Calculation for Maritime Surveillance Mission... 43 Table 17: Tail-down Angle Calculation... 44 Table 18: Tip-back Angle Calculation... 44 Table 19: Overturn Angle Calculation... 44 Table 20: CLmax Values for Mach Numbers between 0.2 and 0.6... 47 Table 21: CLmax Values for Mach Numbers between 0.2 and 0.6... 47 Table 22: Air Properties at Different Altitudes... 49 Table 23: Form Factor Calculation for Each Component... 49 Table 24: Some Parameters for the Components... 50 Table 25: Passenger Mission Statistical Method Weight Calculation... 57 Table 26: Maritime Surveillance Mission Statistical Method Weight Calculation... 58 Table 27: Load Factor Corresponding to Mach Number... 61 Table 28: Cost Analysis Corresponding to Material and Quantity... 72 Table 29: RDT & E Cost Analysis... 74 Table 30: Flyaway Cost Analysis... 74 Table 31: Production Cost & Selling Price Corresponding to Quantity... 75 Table 32: Operations & Maintenance Cost... 76 Table 33: Design Summary... 77 7

List of Abbreviations Abbreviation AIAA CFR EASA FAA FAR IFR ISA LE MIL-SPEC MSL PW SFC STOL TE TOP TOW VFR Description American Institute of Aeronautics and Astronautics Council on Foreign Relations European Aviation Safety Agency Federal Aviation Administration Federal Aviation Regulations Instrument Flight Rules International Standard Atmosphere Leading Edge Military Specification Mean Sea Level Pratt & Whitney Specific Fuel Consumption Short Takeoff and Landing Trailing Edge Take-off Parameter Take-off Weight Visual Flight Rules 8

List of Symbols Symbol Description Unit W i Weight lb R Range ft V Free Stream Velocity ft/s L Lift lb D Drag lb C Specific Fuel Consumption lb/(hp*h) C cruise Specific Fuel Consumption of Cruise lb/(hp*h) C loiter Specific Fuel Consumption of Loiter lb/(hp*h) C bhp Specific Fuel Consumption for Propeller Engines lb/(hp*h) η p Propeller Efficiency - K LD Lift-To-Drag Parameter for Retractable Prop - E Endurance h ρ Density slug/ft 3 Re Reynolds Number - c Wing Chord ft μ Dynamic Viscosity lbs/ft 2 b Wing Span ft AR Aspect Ratio - C L Lift Coefficient - C M0 Moment Coefficient - σ Density Ratio - C LTO Takeoff Lift Coefficient - e Oswald Efficiency Factor - C D0 Zero Lift Drag Coefficient - S wet Wet Area ft 2 S Area ft 2 9

C fe Skin Friction Coefficient - s landing Landing Distance ft s a Obstacle-Clearance Distance ft M Mach Number - V stall Stall Speed ft/s V LO Lift-Off Velocity ft/s l F Fuselage Length ft d max Maximum Diameter ft c r Root Chord ft c t Tip Chord ft c Mean Chord ft λ Taper Ratio - LE LE Sweep Angle w Width ft P engine Power of Engine shp l engine Length of Engine ft x cg Longitudinal Center of Gravity ft z cg Vertical Center of Gravity ft C Lα Lift Curve Slope - β Beta - F Fuselage Lift Factor - Λ HL Hinged Line Angle C fc Flat Plate Skin Friction Coefficient - FF c Form Factor - Q c Interference Factor - C D,misc Drag of Flaps, Landing Gears, Upswept Aft Fuselage, Base Area - C D,L&P Drag of Leakages and Protuberances - 10

a Speed of Sound ft/s k Skin Roughness Value - t Thickness ft K Induced Drag Factor - C ram Ram recovery Correction Factor - C bleed Bleed Correction Factor - n Load Factor - V e Equivalent Speed ft/s U de Standard Gust Speed ft/s ε α Downwash derivative - h n Neutral Point of Aircraft - h nwb Neutral Point of Wing Body - C mα Pitch Stiffness Derivative - C lα Lift Curve Slope - 11

1. CHAPTER 1 1.1. Introduction There are some amphibious aircrafts that are used to transport passengers, cargo or act as military aircraft. However, these aircrafts are limited in size to less than 20 passengers or not useful for short runways. Therefore, the design of a new amphibious aircraft capable of multiple missions is established in this report. 1.2. Missions and Requirements There are three main missions that are Passenger Mission, Cargo Mission and Low cost Maritime Surveillance Mission. There are specific requirements for each mission that will be explained in detail in the following part. In this competition, there are two type of requirements: Mandatory Requirement (M) and Tradable Requirement (T). General Requirements: (M) Capable of taking off and landing from runways (dirt, grass, metal mat, gravel, asphalt & concrete) (M) Capable of taking off and landing from fresh and salt water (M) Minimum cruise speed of 200 knots on o (T) Target cruise speed: 250 knots or greater (M) Capable of VFR and IFR flight (M) Capable of flight in known icing conditions (M) Meets applicable certification rules in FAA 14 CFR Part 25 12

o All missions below assume reserves and equipment required to meet applicable FARs (M) Fuel burn per passenger at least 20% better than an existing aircraft on a similar mission length o Economic mission of 500 nmi in a single class configuration (M) Engine/propulsion system assumptions documented and the use of an engine that will be in service by 2027. o Assumptions on at least specific fuel consumption/efficiency, thrust/power and weight should be specified. Passenger Mission Requirements: (M) Crew: 2 flight crew, 1 cabin crew member (M) Minimum passenger capacity: 20 at a 28 or greater seat pitch o (T) maximum capacity of between 20 and 49 single-class seats Passenger and baggage weight assumptions o o Passenger weight of 193.6 lb. (88 kg) Baggage weight per passenger 37.4 lb. (17 kg) and volume of at least 4 cubic feet per passenger (M) 250 nmi Short Takeoff and Landing (STOL) runway mission with 20 passengers o Maximum takeoff and landing field lengths of 1,500 over a 50 obstacle to a runway with dry pavement (sea level ISA + 18 F day). o Takeoff, and landing performance should also be shown at 5,000 above mean sea level (ISA + 18 F) as well as for dirt, grass, metal mat, gravel, asphalt & concrete fields at sea level (ISA+18 F). (M) 250 nmi Short Takeoff and Landing (STOL) water mission with 20 passengers o o Maximum takeoff distance of 1,900 (sea level ISA + 18 F day) over a 50 obstacle. Takeoff, and landing performance should also be shown at ISA + 18 F at 5,000 MSL (ISA+18 F). This is to show performance into and out of a mountain lake o (M) Ability to take-off and land in Sea State 3 conditions 13

(T) Ability to take-off and land in Sea State 4 conditions (M) 1000 nmi design range for maximum density passenger mission o Show takeoff and landing field lengths over a 50 obstacle to a runway with dry pavement (sea level ISA + 18 F day). o Show takeoff and landing field lengths from water (sea level ISA + 18 F day). Cargo Mission: (M) 5,000 lb. payload (M) 500 nmi range o Show takeoff and landing field lengths over a 50 obstacle to a runway with dry pavement (sea level ISA + 18 F day). o Show takeoff and landing field lengths from water (sea level ISA + 18 F day). (M) Ability to unload, refuel and load cargo in no more than 60 minutes (T) Ability to carry odd shaped cargo or a standard container Low Cost Maritime Surveillance Mission: (M) 3,000 lb. or greater payload (M) 10-hours endurance at 500 above the water (ISA + 18 F day) o Show takeoff and landing field lengths over a 50 obstacle to a runway with dry pavement (sea level ISA + 18 F day). o o Show takeoff and landing field lengths from water (sea level ISA + 18 F day). (T) 12 hour or greater endurance (M) 150 knot cruise o (T) 200 knot or greater cruise speed 14

1.3. Competitor Study Competitor aircrafts are chosen by looking their similarity to cruise speed, passenger capacity and range for our requirements. Most similar aircrafts are Bombardier 415 and Canadair CL-215T in terms of the size. [1] Table 1: Comparison of the Characteristics According to Design Requirements General Requirements Twin Otter400 Bombardier 415 Canadair CL-215T Harbin SH-5 ShinMaywa US-2 Cruise Speed (200 knots) 170 180 180 243 259 SFC (lb/(hp h)) - - - - - Empty Weight (lb) 6881 28400 27100 58400 56504 Max TOW (lb) (Land/Water) 12500 / - 41000/37850 43500/37700 99208 105160 Power (bhp) 2*750 2*2380 4*3150 4*4592 Engine Dry Weight (lb) 2*270 2*929 4*972 4*1727 Passenger Missions Column1 Column2 Column3 Column4 Column5 Crew(2+1) 2 2 2 8 11 Passenger Capacity (20-49 ) 19 8 26-20 Range (>1000) (nmi) 980 1310 1200 2568 2538 Take-off Distance (Land/Water) (ft) 1200 / - 2700/2670 5000/- -/1798 2150/1820 Landing Distance (Land/Water) (ft) 1050 / - 2200/2180 - -/790 2655/720 Rate-of Climb (ft/s) 26.Tem 26,7 16,7 30 - Cargo Mission Column5 Column3 Column4 Column2 Column1 Payload (>5,000 lb) 4280 6400 12000 22000 - Low Cost Maritime Surveillance Column1 Column2 Column3 Column4 Column5 Endurance ( @500, 10h) (h) 9-8 15 - Table 2: Important Characteristics of the Competitor Aircrafts Geometric Characteristics Twin Otter400 Bombardier 415 Canadair CL-215T Harbin SH-5 ShinMaywa US-2 Length (ft) 51' 9" 65' 0½" 65' 0 ¼" 127' 7" 109' 1" Height (ft) Land/Water 19' 6" 29' 5½"/ 22' 10" 29' 5 ½" / 22' 7" 32' 2" 33' 0" / - Draught (ft) Wheels Up/Down - 3' 8" / 6' 8" 3'8" / 6'8" - - Wing Span (ft) 65' 0" 93' 11" 93' 10" 118' 1 ¼" 108' 9" Wing AR 10.May 8,2 8,2 9 8,1 Wing Area (sq ft) 420 1080 1079,9 1550 1462 Tailplane Span (ft) 20' 8" 36' 0" 36' 0" 34' 5 ½" 40' 8½" Tailplane Area (sq ft) 100 221,2 221,2-248 Propeller Diameter (ft) 8' 6" 13' 0¼" 13' 0¼" 12' 9 ½" 14' 6" Propeller Water Clearence (ft) - 3' 9¾" 4' 3¼" - - Propeller Ground Clearence (ft) - 9' 1" 9'1" - - 15

Table 3: Important Performance and Design Characteristics of the Competitor Aircrafts Performance and Design Characteristics Twin Otter400 Bombardier 415 Canadair CL-215T Harbin SH-5 ShinMaywa US-2 Max Internal Fuel Weight (lb) 2525 10250 10250 29579 38620 Max Payload Weight (lb) 4280 5500* 10560 22045 - Max TOW Land/Water (lb) 12500 / - 41000/37850 37700/37700 99208/- 99200/94800 Empty Weight (lb) 6881 30000 25990 58422 56505 Max Wing Loading Land/Water (lb/sq ft) 29.76/- 37.96/35.05-64.0/- 67,85 Max Power Loading Land/Water (lb/sph) 8.33/- 8.61/7.95-6.80/- (lb/ehp6,97 Empty Weight Fraction Land/Water 0.551/- 0.732/0.793 0.689/0.689 0.589/- 0.570/0.596 Fuel Weight Fraction Land/Water 0.202/- 0.25/0.271 0.272/0.272 0.298/- 0.389/0.407 Payload Weight Fraction Land/Water 0.342/- 0.134/0.145 0.280/0.280 0.222/- - 1.4. Conclusions Above aircrafts were all turboprop driven planes. Although similarities in design objectives and cruise speeds Harbin and ShinMaywa were too heavy with respect to our design, and Twin Otter was too light, Bombardier 415 and Canadair CL-215T were especially considered later. 16

2. CHAPTER 2 INITIAL SIZING 2.1. Introduction In this chapter, take-off gross weight has been established as a first estimation by using empty weight fraction and fuel weight fraction for both passenger and maritime surveillance missions, respectively. These calculations include some empirical equations and historical data. In addition, some trade studies have been conducted to see the effects of fuel burn with respect to range, speed and mission type; also for composite material effect. 2.2. Mission Profile The aircraft has to take-off and fly with minimum 20 passengers or cargo weight as described in the AIAA requirements and following figure shows a simple cruise mission profile graphically. Figure 1: Simple Cruise Mission Profile 2.3. First Estimation of Take-off Gross Weight 2.3.1. Passenger Mission Take-off gross weight is estimated by using following equation [2] and for calculation, number of passenger trade are made 17

W 0 = W crew + W payload 1 W fuel W 0 W empty W 0 Crew weight was estimated as 700 lb for 2-flight crew and 1 cabin crew member with their baggage and passenger payload is calculated as 6468 lb for 28 passengers. Empty weight fraction is calculated by using below equation with the coefficients taken from Table 3.1 [2] assigned for flying boat that is the closest aircraft type to amphibians. W e W 0 = 1.09 W 0 0.05 (1.0) Then, fuel weight fraction is calculated by using following equation with the mission segment weight fraction: Where the mission segment weight fraction is W 6 W f W 0 = 1.06 (1 W 6 W 0 ) = W 6 W 5 W 4 W 3 W 2 W 1 W 0 W 5 W 4 W 3 W 2 W 1 W 0 For the mission segment weight fraction, weight fractions were estimated as 0.97, 0.985, 1 and 0.995 for engine start, climb, descent and landing from historical trends, respectively. Then, Brequet Range Equation was used to calculate weight fraction of climb. Brequet Range Equation: R = V C L D ln (W 2 W 3 ) or W 3 W 2 = e RC V ( L D ) Specific fuel consumption can be calculated by C = C bhp for propeller engines and η 550η p = 0.8 that is p propeller efficiency, C bhp = 0.5 [s/h/ft] for cruise and V = 337.562 ft s. Substituting the values V yields C cruise = 0.384. To calculate L D ) cruise, L D ) max can be calculated from L D ) max = K LD AR wet with values AR wet 1.6 and K LD = 11 (for retractable prop) and L ) 13.92. For propeller airplanes, D max the relations are L D ) cruise = L D ) max. Also, L D ) max is assumed as 14 from Figure 3.5 [2] so that it is decided 18

to take L D ) cruise as 14. Then, numerical values are substituted into range equation and it is found that W 3 W 2 = 0.854. Loiter segment fraction is calculated as 0.981 by using fallowing equation: W 5 W 4 = e EC ( L D ) loiter where E = 0.5 h, specific fuel consumption is C loiter = 0.460 and L D ) loiter = 12.133 since L D ) loiter = 0.866 L ) for propeller engines. D max So, mission segment weight fraction is calculated as W 6 W 0 = 0.797and using this value fuel weight fraction is found as W f W 0 = 0.215. Then, take-off gross weight is calculated as 40 807 lb by using this fractions and iterations and this take-off gross weight is used to calculate the empty weight fraction W e W 0 = 0.609. 2.3.2. Low Cost Maritime Surveillance Mission Same calculations are made for this mission with different values and by keeping empty weight as constant, endurance and speed trades are made for the calculation. Payload weight is taken as 3000lb that is the minimum mandatory requirement and crew weight is taken as same. Different values are V = 270.049 ft s, C cruise = 0.297 and C loiter = 0.357 so these values are substituted into weight fraction equations of cruise and loiter, and it is found that W 5 W 4 = 0.985 and W 6 W 0 = 0.717. Then, Fuel weight fraction is calculated as W f W 0 = 0.3 and take-off gross weight is found as 40 729 lb by using this fraction and iterations. In addition, empty weight fraction is calculated by using take-off gross weight as W e W 0 = 0.609. 19

2.4. Trade Study There is a requirement that needs fuel burn per passenger at least 20% better than an existing aircraft on a similar mission length. For that purpose, Canadair CL-215T is chosen as a reference aircraft for similar mission length with fuel burn 319.23lb per passenger and 80% of 319.23lb is target fuel burn per passenger for our design. As seen from Figure 2, the horizontal line represents the target fuel burn per passenger. These three curves are taken for different speeds and it can be seen that number of passenger are seen in the intersections as 28, 34 and 42. By considering the weight, lightest configuration is chosen with 200 knots and 28 number of passengers. Figure 2: Fuel Burn per Passenger with Increasing Number of Passenger for Passenger Mission Figure 3: Endurance and Range Change with Increasing Speed Maritime Surveillance Mission 20

Figure 4: Fuel Burn per Hour with Increasing Speed for Maritime Surveillance Mission It is seen that it is more efficient due to fuel burn with decreasing speed from Figure 3 and 4. Therefore, the speed is chosen as 155 knots that applies for the requirement because endurance should be greater than 12h for low cost surveillance maritime patrol mission with 12.6 h endurance. In addition, change in range due to speed is negligible. In addition, composite material trade is made for the design and When composite material is selected, endurance is chosen 12.6h and when metal is selected, endurance is chosen 11.8h for equality of weights. Table 4: Composite Material Trade for Passenger Mission Table 5: Composite Material Trade for Maritime Surveillance Mission 2.5. Conclusions In passenger mission, cruise velocity was taken as 337.562 ft/s and also was taken as 270.049 ft/s in maritime surveillance mission. To satisfy fuel burn per passenger requirement, trade studies are done 21

and 28 passenger should be transport. By choosing composite material, 8000 lb reduction was achieved in order to become operational cost lower. For passenger mission, take-off gross weight is found as 40807 lb in first weight calculation method and for maritime surveillance mission is found 40729 lb. Since they are structurally same aircraft, weights are calculated in both mission by keeping their empty weights same. 22

3. CHAPTER 3 Airfoil and Wing Planform Selection 3.1. Introduction This chapter establishes the selection of airfoil and wing planform by calculating the design lift coefficients for the missions. For this purpose, design lift coefficients and Reynolds numbers were calculated and similar airfoils were found to compare by looking these data and then airfoil type was chosen as best for both missions. Furthermore, some aerodynamic characteristics of the wing were calculated by estimating wing sweep, taper ratio and twist value. 3.2. Design Lift Coefficient Calculation 3.2.1. Passenger Mission Design lift coefficient is selected for the cruise condition by assuming W = L = 1 2 ρ V 2 c L S with the values ρ = 1.066 10 3 slug/ft 3 and V = 200 knots = 338 ft/s [3] for h = 25 000ft as cruise height. Firstly, wing loading is calculated as 35 lb/ft 2 by taking average of the closest competitors that have wing loadings as 29.76 and 37.96, respectively. Then, wing loading for cruise condition is calculated as 42.04 lb/ft 2 after 2-step iteration when the convergence is satisfied at later studies by using the equation W S ) cruise = W 0 S W 1 W 0 W 2 W 1 where W 1 W 0 = 0.97 and W 2 W 1 = 0.985. Then, substituting these values into Lift equation above, lift coefficient is found as C L = 0.69. 23

3.2.2. Low Cost Maritime Surveillance Mission For this mission, altitude is 500 ft and ρ = 2.3441 10 3 slug/ft 3[3] for this altitude. In addition, V = 155 knots = 262 ft/s with the same value of the wing loading for the cruise condition. So, lift coefficient is calculated as C L = 0.52 by using these values. 3.3. Selection of Airfoil and Aspect Ratio For selection of airfoil, Reynolds number is calculated by using equation Re = ρ V stall c. To calculate the μ wing chord, firstly wing area is calculated by substituting the values of wing loading and take-off gross weight, converged value after iterations through later studies, into S = W 0 42 190 = = 1003.56 ft2 W/S cruise 42.04 and then wing span is calculated by using the equation b = S AR for aspect ratio determined as 8 from Table 4.1 [2] and it is found as 89.60 ft. Then, wing chord is found as 11.20 ft from the equation c = b AR. After wing chord is found, V stall is assumed as 70 knots due to competitors that are Bombardier 415 and Canadair CL-215T having 68 knots and 66 knots, respectively. Substituting these values for both missions yield: Re = 4 389 766 for Passenger Mission Re = 8 137 512 for Low Cost Maritime Surveillance Mission C L value for low cost maritime surveillance mission is selected because if C L for passenger mission were selected, lift would be bigger than weight in low cost maritime surveillance mission. On the other hand, if the design C L value were chosen as maritime mission cruise C L value that is 0.52, at passenger mission designated cruise C L value can be accomplished with elevator trim. Hence, lower C L value is decided as a better choice for both missions and 2D and 3D airfoil analyses are made for maritime surveillance mission in XLFR5 that are listed in the following table. Eppler1210 was selected as airfoil at first 24

calculation of chapter 3 due to high C Lmax and C M0. After iterations, since take-off gross weight increased wing was decided to have 2-degree incidence angle. Table 6: Aerodynamics Parameters of Similar Airfoil 3.4. Estimation of Wing Sweep, Taper Ratio and Twist Value Wing sweep is selected as 0 from Figure 4.20 [2] because there is no need sweep angle at this speed regime. Moreover, sweep angle causes decrease in the normal direction of the freestream velocity to the LE that will result in lower lift. Then, taper ratio is estimated as 0.8 beginning at 2/3 of half span from root where leading edge has an angle of 0.75 to make sweep angle of quarter chord as 0. Twist angle is used for increasing stall angle. Since there is not a requirement for high climb angle, there is no need for twist angle in our design. 25

4. CHAPTER 4 DESIGN PARAMETERS 4.1. Introduction Thrust-to-weight ratio and wing loading were calculated in this chapter because these design parameters are the most important performance parameters since they have the effect on engine selection and takeoff gross weight. 4.2. Calculation of Thrust-to-Weight Ratio Since our aircraft is an turbo-prop driven amphibian, Power-to-Weight ratio (P/W) is calculated instead of Thrust-to-Weight ratio (T/W). As final iteration, since engine was decided at first calculation of 6 th chapter, actual horse power is divided by final take-off gross weight value. Therefore, hp W 0 = 0.130 shp/lb is first considered power to weight ratio. After engine was chosen, hp W 0 = 0.154 shp/lb is used as power to weight ratio. 4.3. Calculation of Wing Loading 4.3.1. Take-off Wing loading for take-off condition is calculated for only passenger mission since it is determinant mission in terms of payload. It is assumed that it 250 nmi Short Take-off runway mission with 20 passengers because it is a requirement of the competition. Therefore, calculations are done for maximum takeoff distance of 1500 ft for land and 1900 ft for water. Because of that, weight values are changed due to decreasing number of passenger and the fuel corresponding to this payload and range. Since empty weight is not change according to flight condition, it is taken as 30 228 lb. Then, payload weight is calculated as 5320 lb with 20 passengers and fuel weight is calculated as 4303 lb for 250 nmi. 26

Since fuel weight fraction is known and it is equal to the 0.108, take-off weight is calculated as 39 851 lb by using this fraction as below: W 0 = W e + W p 30228 + 5320 = = 39 851 lb 1 W f /W 0 1 0.108 After that, wing loading for take-off condition is calculated by using equation 5.8 [2] : W S ) = (TOP)σC LTO ( hp take off W ) where σ is density ratio and is equal to 0.862 by taking ratio of densities at mean sea level and 5000 ft, TOP is take-off parameter that is determined for corresponding take-off distance value by using Fig. 5.4 [2] and it is equal to 160, C LTO is lift coefficient for take-off and it is equal to 2.26 where flaps are in takeoff position. Then, wing loading for design is calculated as W 0 S = 42.23 lb/ft2. 4.3.2. Cruise It is calculated for both missions. Density and dynamic viscosity are chosen for mission altitudes, mean chord, aspect ratio and wetted, reference areas are known from previous study. Wing loading for cruise is calculated by using Equation 5.14 [2] : W S ) = q πarec D0 cruise where e is Oswald efficiency factor and it can be calculated from e = 1.78(1 0.045AR 0.68 ) 0.64 for Straight-Wing Aircraft from equation 12.49 [2] and C D0 can be calculated as C D0 = S wet C S fe from equation 2.37 [2] and C fe is the skin friction coefficient that can be found from Fig 2.55 [2] for a corresponding Reynolds number. All variables for both missions and wing loading values can be seen in Table 7: Table 7: Variables and Wing Loading Values for Both Missions Parameters Passenger Mission Maritime Surveillance Mission ρ (slug/ft 3 ) 1.07E-03 2.34E-03 μ, Dynamic Viscosity (lb s/ft 2 ) 3.22E-07 3.82E-07 c, Chord (ft) 10.52 11.20 AR 8 8 27

S wet /S 5 5 C fe, Skin Friction Coefficient 0.0042 0.0038 e, Oswald Efficiency Factor 0.811 0.811 V cruise,(ft/s) 338 261.95 Re 11,783,665 16,929,040 C D0, Zero Lift Drag Coefficient 0.021 0.019 q 60.89 80.42 W S cruise (lb/ft2 ) 39.83 50.04 After wing loading for cruise is calculated, it is used to check for design wing loading by using mission segment weight fractions for cruise as follows: W S ) cruise = W 0 W 1 W 2 S W 0 W 1 Wing loading was calculated as 41.96 lb/ft 2 for passenger mission and 52.37 lb/ft 2 for maritime surveillance mission. 4.3.3. Loiter: Wing loading for loiter condition for both missions is calculated by using equation 5.16 [2] that is given for propeller-driven aircrafts and the same variables in Table 7. W S ) = q 3πAReC D0 loiter Wing loading was calculated for loiter as 68.98 lb/ft 2 for passenger mission and 86.66 lb/ft 2 for maritime surveillance mission. Then wing loading is calculated as 75.31 lb/ft 2 for passenger mission and 94.62 lb/ft 2 for maritime surveillance mission by using mission segment weight fractions for both missions: W S ) = W 0 W 1 W 2 W 3 W 4 loiter S W 0 W 1 W 2 W 3 28

4.3.4. Landing: Wing loading for landing condition is also calculated by taking safety margin as 1 and using the equation 5.11 [2] : W S ) = landing S landing S a 1 0.66 80 ( σc ) Lmax where S landing is landing distance, S a is obstacle-clearance distance that is equal to 450 for STOL and σ is density ratio. In addition, the reason of multiplying ground portion of the landing by 0.66 is that it may give an advantage for aircrafts equipped with reversible-pitch propellers. So, wing loading is found as 38.34 lb/ft 2. By using this wing loading, wing area is calculated as 951.46 ft 2 and wing loading for design is found as 44. 34 lb/ft 2. Among above wing loading values, the lowest one satisfies all wing loadings. Therefore, cruise segment wing loading was chosen as design wing loading. 4.4. Conclusions According Raymer s book [2], minimum value of wing loading should be selected to ensure the wing is large enough for all flight condition. Since wing loading value is minimum in take-off condition that equation of calculation includes hp/w value as our selection, 0.154, our power-to-weight ratio is suitable for the wing loading. Since the lowest wing loading satisfies the all requirements, and so take-off wing loading 42.23 lb/ft^2 is our design wing loading. The geometry satisfies the lowest wing loading will satisfy the other segments loading requirements. In addition, airfoil was chosen as Eppler1210 in before study. However, for higher lift coefficient, airfoil was changed to NACA 4415. 29

5. CHAPTER 5 Refined Sizing 5.1. Empty Weight Fraction Calculation 5.1.1. Passenger Mission To calculate the take-off gross weight, firstly empty weight fraction and fuel weight fraction should be calculated for the Equation 6.1 [2]. Since amphibian aircrafts use propeller-driven engine, Table 6.2 [2] should be used to calculate the empty weight fraction with the values for flying boat that is the closest aircraft type to amphibians. W e C = a + b W 1 W 0 A C 2 (hp/w 0 ) C 3 (W 0 /S) C C 4 V 5 max 0 Table 8: Parameters [2] a b C 1 C 2 C 3 C 4 C 5 hp W 0 (sph/lb) V max (knots) W 0 S (lb ft 2 ) 0 0.42-0.01 0.1 0.05-0.12 0.18 0.13 220 41.69 Remind that, minimum value of wing loading is chosen as 41.69 lb/ft 2 which is the wing loading of cruise mission. Then, by substituting the values into the empty weight fraction equation: W e W 0 = 0.781W 0 0.01 So, fuel weight fraction can be calculated by using Equation 6.2 [2] as follows: W f W 0 = 1.06 (1 W 6 W 0 ) Finally, take-off gross weight can be obtained with iterations by substituting the values of empty weight fraction and fuel weight fractions: W 0 = W crew + W payload 1 ( W f W ) ( W e 0 W ) 0 30

Where W crew = 700lb and W payload = 5313 lb for 23 passengers. 5.1.1.1. Engine Start, Take-off It is chosen as minimum, 0.97 for passenger mission since historical trend for this mission segment is W 1 W 0 = 0.97 0.99. 5.1.1.2. Climb Weight fraction of this segment can be calculated by using equation 6.9 [2] : W 2 W 1 = 1.0065 0.0325 M Firstly, Mach number at the beginning of climb should be calculated that corresponds to lift-off speed that is V LO = 1.1V stall. Since C Lmax ) take off is known as 2.26. Now, lift-off speed for climb is calculated as 122.62 ft/s by using weight fraction of engine start segment, wing loading, density of air at sea level and maximum take-of lift coefficient. Then, Mach number corresponding to lift-off speed is found as 0.11 by dividing lift-off speed by speed of sound at sea level and Mach number corresponding 25 000ft is calculated as 0.303 as same as. By substituting Mach number values into climb weight fraction equation, weight fraction is calculated as 0.994. 5.1.1.3. Cruise Weight fraction of cruise mission is calculated as 0.925 from Equation 6.12 [2] for propeller-driven aircrafts by taking range as 1253.5 ft that is the value with the addition of 253.5 ft that is the required value for FAR regulations for 45 minutes extra. W 3 W 2 = e R C bhp 550η p ( L D ) 31

5.1.1.4. Descent It is chosen as minimum for worst case, 0.99 since historical trend for this mission is W 4 W 3 = 0.990 0.995. 5.1.1.5. Loiter Weight fraction for this segment is calculated as 0.992 from the equation for propeller-driven is written from Equation 6.15 [2] : W 5 W 4 = e EV C bhp 550η p ( L D ) 5.1.1.6. Landing and Taxi-Back It is chosen as minimum for worst case, 0.992 since historical trend for this mission is 0.992-0.997. Finally, final weight fraction can be calculated as 0.867 by multiplying all segment fractions. And fuel weight fraction is calculated as 0.141 by using equation 6.2 [2] : By using these fractions, take-off gross weight is found as 40038.6657 lb for passenger mission and note that empty weight fraction is multiplied by 0.95 for composite trade. 5.1.2. Low Cost Maritime Surveillance Mission Same equations for fuel weight fraction, empty weight fraction and take-off gross weight are used as for the passenger mission. While crew payload weight is same as 700 lb, payload weight is changed to 3000 lb as required for this mission. 5.1.2.1. Engine Start it is chosen as minimum, 0.97, for maritime surveillance mission since historical trend for this mission segment is between 0.97 and 0.99. 32

5.1.2.2. Climb Same calculation has been made for this mission with the passenger mission with a different value of M 1 since cruise speed is different for that altitude and it is found as: M 1 = V cruise a @500ft = 261.95 1114.53 = 0.258 Then weight fraction for the climb section is found as 0.994. 5.1.2.3. Cruise Weight fraction of climb is found as 0.866 by using same Equation 6.12 [2] by taking q as 80.28 and calculating range by multiplying cruise speed and endurance which is 12.75 hour with added value of extra fuel for 45 minutes extra. 5.1.2.4. Descent It is chosen as minimum for worst case, 0.99 because historical trend is between 0.99 and 0.995 for this mission. 5.1.2.5. Loiter Weight fraction of this segment is calculated as 0.993 by using same equation in the passenger mission with the different value of lift-to-drag ratio that is 18.628 for loiter. 5.1.2.6. Landing and Taxi-Back It is chosen as minimum for worst case, 0.992 since historical trend is between 0.992 and 0.997 for this mission. Finally, weight fraction for this mission is calculated as 0.812 by using same equation that multiplies all fractions. Then, using Equation 6.2, fuel weight fraction is found as 0.199 and take-off gross weight is 33

found as 40046.6153 lb for this mission. Note that, empty weight fraction is multiplied with 0.95 for composite trade. 34

6. CHAPTER 6 Geometry Sizing and Configuration 6.1. Calculation of Fuel Volume Previously, take-off gross weight is found as 40 523 lb and fuel weight fraction is calculated as 0.199. To calculate the amount of fuel in terms of gallons, since the weight is known only density should be found and it is found as 6.7 lb/gal for JP-8 as MIL-SPEC density from Table 10.5 [2]. Since 85% of the wing volume can be used for fuel storage, it is divided by 0.85. Additionally, in order to consider fuel expansion due to temperature change add extra volume by 5% of total amount of fuel. Finally, it is calculated as 238.43 ft 3 that will be stored in wing since fuel occupies approximately 1/3 of total wing volume and it has a weight of 8055.38 pound for passenger mission and fuel volume is found as 168.35 ft 3 and fuel weight is calculated as 5687.66 lb for maritime surveillance mission 6.2. Fuselage Length 6.2.1. Historical Trend Fuselage length can be calculated by using the Table 6.3 [2] c l F = aw 0 Where a = 1.05 and c = 0.4 for a flying boat that is the closest type to amphibians. By substituting the value of take-off gross weight as an approximate value of 40523 lb, l F = 1.05(40523) 0.4 l F = 73.16 ft 6.2.2. Competitor Data Since ShinMaywa US-2 and Harbin SH-5, which are two of competitors, are too big that they have the lengths of 109 ft and 127 ft, respectively, they are not taken as consideration. In addition, Twin otter 400 has very small length as 51 ft so it is not taken into account. By taking the average of the lengths of Bombardier 415 and Canadair CL-215T which they both have fuselage lengths of 65 ft so: 35

l F = 65 ft By comparing these values, fuselage length is chosen as 73.16 ft. Then, maximum diameter of fuselage is calculated by using the fineness ratio [2] : Fineness Ratio = For l F = 73.16 ft, maximum diameter: Fuselage Length Maximum Diameter = 5 d max = 14.63 ft 6.3. Sizing and Planform of the Wing, Aerodynamic Parameters Firstly, area of the wing is calculated as 864.37 ft 2 by dividing the take-off gross weight of 35802 lb by wing loading value of 41.42. Wing span is calculated as 83.16 ft by using the aspect ratio equation. Mean aerodynamic chord is found as 10.39 ft from XFLR5 since the wing has a taper at 2/3 of chord. To calculate the root chord and tip chord, an equation is derived for the wing because taper ratio is at 2/3 of chord. Then, the wings are shown in the below and areas are calculated as equal: c* 3L c r c r c t 2L L Where c r = 6c 5+λ, c t = λc r and λ = 0.8 2c r L + c r + c t L = 3Lc 2 Table 9: Wing Geometric Characteristics Wing Characteristics Area (ft 2 ) 864.37 Span (ft) 83.16 36

Mean Chord (ft) 10.39 Root Chord (ft) 10.75 Tip Chord (ft) 8.60 LE Sweep Angle ( ) 0.74 Total Wing Volume (ft 3 ) 1076.60 6.4. Tail Configuration and Size, Aerodynamic Parameters T-tail configuration is chosen for the aircraft due to clearance from the water concerns. For an aircraft with two engines on the wings, tail arm should be about 50-55% of the fuselage length. It is chosen as 50% and vertical tail and horizontal tail areas are calculated by using equation 6.26 and 6.27 [2] : S VT = c VTb w S w L VT S HT = c HTc ws w L HT Vertical tail span can be calculated since area and aspect ratio are known now: b VT = S VT AR VT Then, root chords and tip chords of vertical and horizontal tail are calculated by using following equations: c r = 2S VT b VT + (b VT λ VT ) c t = λ VT c rvt So, mean aerodynamic chord of vertical and horizontal tail can be calculated by using the above root chord lengths: c = 2 3 c 1 + λ + λ 2 r 1 + λ b HT = Aspect ratio of the horizontal can be calculated now: 2 S HT c rht + (c rht λ HT ) 37

AR HT = b 2 HT = 25.962 S HT 182.46 = 3.69 Also, leading edge sweep angle for vertical and horizontal tail is calculated as follows: LE = 180 π tan 1 ( c r c t ) b Note that root chord of horizontal tail is taken as equal to the tip chord of vertical tail and calculations are made due to this equality. Table 10: Wing Geometric Characteristics for Horizontal and Vertical Tail Tail Wing Characteristics Vertical Tail Horizontal Tail Area (ft 2 ) 112 153.81 Span (ft) 11.59 23.88 Root Chord (ft) 10.73 8.59 Tip Chord (ft) 8.59 4.29 Aspect Ratio 1.2 3.71 LE Sweep Angle ( ) 10.49 19.78 6.5. Estimation of Weight and Dimensions of the Engine To choose the engine, firstly it is looked at required thrust and loss due to cooling air and drag due to water landing that they have the values of 10% and 5%, respectively. Total required thrust is found as 2732 shp so that the engines are searched to satisfy this value. Then, PW127M is chosen because it has 2750 shp and the specifications of the engine [4] that are W engine,dry = 1061.97 lb, h engine = 33 in, w engine = 2.228 ft, l engine = 6.988 ft and P engine = 2499.665 shp. 38

7. CHAPTER 7 CG Estimation, Landing Gear Sizing and Placement 7.1. Decision for Location of Fuel Fuel is distributed between two wings as placing the tanks that stores the volume of 238.43 ft 3 that is calculated previously. Drawing and the sizing of the fuel tanks can be seen in the following figure: 7.2. Sizing of Fuselage Figure 5: Drawing of the Fuel Tanks in the Wing From the OpenVSP program, the fuselage length has length of 73.16ft from Chapter 6 and has the maximum diameter as 14.63 ft from OpenVSP program. Width of the fuselage is calculated by considering the passenger seats configuration that has 2+1 seat configuration. Seats have width of 20 in and aisle has also 20 in as width. In addition, 1.6 in is added to total width as thickness of material. Then, width of fuselage is calculated 84 in. 7.3. Estimation of Weights of Major Components Weights of the components are calculated by using Table 15.2 [2] with the data that are taken from OpenVSP. Areas and lengths that are taken from OpenVSP are tabulated as follows: 39

Table 11: Planform and Wetted Areas for Components Planform and Wetted Areas S exposed planform of Wing (ft 2 ) 929.14 S exposed planform of Horizontal Tail (ft 2 ) 185.85 S exposed planform of Vertical Tail (ft 2 ) 140.634 S wetted area of Fuselage (ft 2 ) 1688.24 From Table 15.2 [2], the coefficients are taken for transports and bombers and multiplied by the areas to calculate the approximate weights of the components. Weights of the major components were calculated by using reference parameters shown in following table: Table 12: Weights of the Major Components Weight (lb) Component Passenger Mission Maritime Surveillance Mission Reference Parameter Wing 7897.69 7897.69 Exposed Planform Area Horizontal Tail 848.41 848.41 Exposed Planform Area Vertical Tail 641.99 641.99 Exposed Planform Area Fuselage 9496.33 9496.33 Wetted area Installed Engine 2761.2 2761.2 - Fuel 6016.81 8286.54 - Landing Gears 1743.19 1702.85 - Cockpit 448 448 - Passenger Compartment 736 736 - Lavatory 20 20 - Payload 6013 3700 - All-Else Empty 6050.34 5882.47 - Since take-off gross weight is not known yet, firstly it is taken as 42174.66 lb as an estimation and weights of cockpit, passenger compartment and lavatory have been extracted from above take-off gross weight then it is iterated by using this assumption. It is calculated as 42672.96 lb for passenger mission and 41685.48 lb for maritime surveillance mission. 40

7.4. Estimation of Center of Gravity 7.4.1. Longitudinal Center of Gravity 7.4.1.1. Passenger Mission Firstly, center of gravity is calculated for the components weights except landing gears since they shouldn t change the center of gravity. Then, landing gear weights are added and checked for the equality of the center of gravity. All center of gravity positions of the components is taken with respect to the datum that is equal to the nose of the fuselage. Table 13: Longitudinal Center of Gravity Location Calculation for Passenger Mission Components Weights (lb) CG Position w.r.t Datum Weight*CG Fuselage 9496,33 30.47 289334.28 Wing 7897.69 32.32 255286.20 Horizontal Tail 848.41 66.70 56587.86 Vertical Tail 641.99 65.67 42162.12 Installed Engine 2761.12 30.16 83275.44 Fuel 6016.81 32.7 196749.8415 Main LG 1481.71 35 51859.92 Nose LG 261.48 11.14 2912.51 Cockpit 448 5.83 2613.33 Pass. Comp. 736 29.5 21712 Lavatory 20 42.2 846.78 Pass + Crew 4646.4 29.5 137068.80 Cargo 860.2 15.104 15.104 Pilots 506.40 6.67 3376 All Else Empty 6050.34 30.47 184341.68 x cg = 31.42784 ft 7.4.1.2. Low Cost Maritime Surveillance Mission Same calculation is made for this mission with different value of fuel and passenger compartment since it differs from the passenger mission according to range and payload. Table 14: Longitudinal Center of Gravity Location Calculation for Maritime Surveillance Mission Components Weights (lb) CG Position w.r.t Datum Weight*CG Fuselage 9496,33 31 294424.31 41

Wing 7897.69 27.37 255286.20 Horizontal Tail 848.41 66.70 56587.86 Vertical Tail 641.99 65.67 42162.12 Installed Engine 2761.12 30.16 83275.44 Fuel 8286.54 32.7 270969.897 Main LG 1481.71 36 53341.63 Nose LG 261.48 7.25 1895.49 Cockpit 448 5.83 2613.33 Pass. Comp. 0 29.5 0 Lavatory 20.07 42.2 846.78 Pass + Crew 193.60 29.5 5711.20 Cargo 2500 28.208 70520 Pilots 506.40 6.67 3376 All Else Empty 6050.34 28.21 170667.92 x cg = 31.68788 ft 7.4.2. Vertical Center of Gravity 7.4.2.1. Passenger Mission All center of gravity positions of the components is taken with respect to the datum that is equal to the nose of the fuselage. Table 15: Vertical Center of Gravity Location Calculation for Passenger Mission Components Weights (lb) CG Position w.r.t Datum Weight*CG Fuselage 9496,33 0.97 9163.96 Wing 7897.69 5.35 42228.95 Horizontal Tail 848.41 16.93 14363.66 Vertical Tail 641.99 9.06 5815.83 Installed Engine 2761.12 7.58 20923.78 Fuel 6016.81 5.37 32310.3 Main LG 1481.71-6.81-10087.49 Nose LG 261.48-7.31-1910.36 Cockpit 448 0 0,00 Pass. Comp. 736 0 0,00 Lavatory 20 0 0,00 Pass + Crew 4646.4 0 0,00 Cargo 860.2 1.67 1433.67 42

Pilots 506.40 0 0,00 All Else Empty 6050.34 0 0,00 z cg = 2.67715836 ft 7.4.2.2. Low Cost Maritime Surveillance Mission All center of gravity positions of the components is taken with respect to the datum that is equal to the nose of the fuselage. Table 16: Vertical Center of Gravity Location Calculation for Maritime Surveillance Mission Components Weights (lb) CG Position w.r.t Datum Weight*CG Fuselage 9496,33 0.97 9163.96 Wing 7897.69 5.35 42228.95 Horizontal Tail 848.41 16.93 14363.66 Vertical Tail 641.99 9.06 5815.83 Installed Engine 2761.12 7.58 20923.78 Fuel Wing 8286.54 5.37 32310.3 Main LG 1481.71-5.81-8605.78 Nose LG 261.48-6.31-1648.88 Cockpit 448 0 0,00 Pass. Comp. 0 0 0,00 Lavatory 20.07 0 0,00 Pass + Crew 193.60 0 0,00 Cargo 2500 1.67 4166.67 Pilots 506.40 0 0,00 All Else Empty 6050.34 0 0,00 z cg = 3.16248451 ft 7.5. Sizing and Location of Propeller First estimation on diameter of propeller was calculated by using below equation which is given in Raymer's book [2] : 4 D = K p Power 43

The engine selected is Pratt & Whitney PW 127M. EASA Certification [4] document of Pratt & Whitney Engine states that PW 127M engines were certified only with 6-bladed propeller configuration. Therefore, value of K p is 1.5 which is used for 4 and more bladed propeller configurations in British Units and power of engine is 2750. Therefore, propeller diameter is: 4 D = 1.5 2750 = 10.862 ft 7.6. Calculation of Static Tail-Down Angle, Tip-Back Angle and Overturn Angle All lengths are taken from OpenVSP program. Tail-down angle is calculated for the position that landing gears fully extended and Tip-back angle is calculated by adding this deflection angle so it is calculated for non-extended landing gears. Table 17: Tail-down Angle Calculation Vertical Difference between rear end of Fuselage and Main LG (ft) 9,96 Horizontal Difference between rear end of Fuselage and Main LG (ft) 38,16 Tail-down Angle ( ) 14,62 Table 18: Tip-back Angle Calculation Horizontal Difference between CG and Main LG (x-axis) (ft) 2,42 Vertical Difference between CG-Main LG (z-axis) (ft) 8,42 Tip-back Angle ( ) 12,44 Table 19: Overturn Angle Calculation Difference Between CG and Wheels (y-axis) (ft) 6 Difference Between CG and Wheels (z-axis) (ft) 9,49 Overturn Angle ( ) 57,68 44

8. CHAPTER 8 Aerodynamics 8.1. Lift Curve Slope Lift curve slope for the subsonic flight can be calculated from the equation 12.6 [2] : C Lα = 2πAR S exposed F 2 + 4 + AR2 β 2 S η 2 (1 + tan2 Λ max,t β 2 ) Where AR = 8, β 2 = 1 M 2, η is airfoil efficiency and it is 0.95 for most airfoils and F is fuselage lift factor that can be calculated from F = 1.07(1 + d b) 2. Maximum thickness for NACA4415 airfoil occurs at x c = 0.211 in the wing but since there is no sweep angle in our wings except for tapered section at the tips it can be assumed as 0 because sweep angle at the tapered section is around 0.625 according to drawing in OpenVSP and exposed area and total area of the wing are found as 2038.45 ft 2 and 1978.60 ft 2, respectively from the drawing. Fuselage lift factor is calculated as 1.26 by the drawing in OpenVSP. Substituting the all numerical values into C Lα equation for the Mach numbers between 0.2M and 1.0M following values are plotted: 45

14.00 12.00 10.00 Lift Curve Slope 8.00 6.00 4.00 2.00 8.2. Maximum Lift Coefficient Calculation Figure 6: Lift Curve Slope for Subsonic Flight 8.2.1. Maximum Lift Coefficient at Clean Configuration Firstly, it is checked whether the wing is a low AR wing or not by using equation 12.18 [2] and it is found that the wing violates the low wing equation since our AR is equal to 8 while RHS of the equation is 2.67. AR 3 (C 1 + 1)(cos Λ LE ) Where C 1 = 0.125 for λ = 0.8 from Fig.12.11 [2] and Λ LE = 0 Maximum lift coefficient for high AR wings can be calculated from the equation 12.16 and Fig 12.8-Fig 12.9 [2] : 0.00 0.00 0.20 0.40 0.60 0.80 1.00 1.20 Mach Number C Lmax = C lmax ( C L max C lmax ) + C Lmax From Fig 12.8, C Lmax C lmax is found as 0.9 for the corresponding sweep angle and since our leading-edge sweep angle is equal to 0. Also, C Lmax is found from Fig 12.9 for the zero-sweep angle and corresponding the Mach numbers between 0.2M and 0.6M. 46

Table 20: C Lmax Values for Mach Numbers between 0.2 and 0.6 Mach Number 0.2 0.3 0.4 0.5 0.6 C Lmax 0.00-0.24-0.42-0.55-0.64 Then, C Lmax is found as 2.05 from XLFR5 for NACA 4415 airfoil for the Reynolds Number that is calculated as 14,576,979 for the M=0.2. Then, C Lmax is calculated by using these values for the Mach numbers between 0.2 and 0.6 that can be seen from below table: Table 21: C Lmax Values for Mach Numbers between 0.2 and 0.6 Mach Number 0.2 0.3 0.4 0.5 0.6 C Lmax 1.845 1.602 1.430 1.295 1.204 Then, C Lmax values are corrected by using adjustment factor for high Mach numbers in the Fig 12.14 [2] and it is plotted: Lift Coefficient 2.000 1.800 1.600 1.400 1.200 1.000 0.800 0.600 0.400 0.200 0.000 0.00 0.20 0.40 0.60 0.80 1.00 1.20 Mach Number Figure 7: C Lmax Values for the Mach Numbers between 0.2 and 1.0 by adjusting 8.2.2. Maximum Lift Coefficient with High Lift Devices Double slotted flap is used for trailing-edge and it has C Lmax = 1.6 as lift contribution with c c = 1 for trailing-edge and slat is used for leading-edge and it has C Lmax = 0.42 as lift contribution with c c = 1.05 by assuming chord is increasing about 5% with devices for leading-edge edge according to Table 12.2 of [2]. Then, equation 12.21 [2] is used to calculate the maximum lift coefficient with high lift devices: 47

So, substituting the numerical values yield at: C Lmax = C lmax ( S flapped ) cos Λ S HL C Lmax ) flapped = 0.89 C Lmax ) slatted = 0.26 Then, maximum lift coefficient can be calculated by using lift coefficient of clean configuration and contributions of the high lift devices: C Lmax = C Lmax ) clean + C lmax ) high lift C Lmax ) landing = 2.99 This value corresponds to lift coefficient of the landing and lift coefficient for the take-off is 60-80% of landing value and by taking 80% of this value: C Lmax ) take off = 0.8 C Lmax ) landing = 2.39 8.3. Drag Divergence Mach Number Drag divergence Mach number is not applicable for our aircraft because our sweep angle is equal to zero and aircraft cannot reach to 0.6M in reality when it is looked at Fig 12.26 [2] to calculate. 8.4. Total Parasite Drag and Wave Drag Total parasite drag is calculated for only subsonic region because aircraft cannot be in supersonic region and it is calculated by using the equation 12.24 [2] : C D0 ) subsonic = C f c FF c Q c S wet,c S Where C fc : Flat plate skin friction coefficient FF c : Form Factor Q c : Interference Factor + C D,misc + C D,L&P C D,misc : drag of flaps, landing gears, upswept aft fuselage, base area 48

C D,L&P : Drag of leakages and protuberances It is assumed as fully turbulent flow and skin friction coefficient can be calculated by using equation 12.27 [2] : C fc = 0.455 (logre) 2.58 (1 + 0.144M 2 ) 0.65 Since it is asked to calculate for the altitudes as sea level, 15 000ft and 24 000ft, all air parameters are found for these altitudes [3] Table 22: Air Properties at Different Altitudes Properties Sea level 15 000ft 24 000ft ρ (slug/ft 3 ) 0.002377 0.001496 0.0011034 a (ft/s) 1116.45 1057.31 1020.19 μ (lbs/ft 2 ) 3.74E-07 3.43E-07 3.27E-07 Also, cut-off Reynold number is calculated by using following equation 12.28 [2] for subsonic flow from assuming if the surface is rough and smaller one of the Reynolds numbers is taken account to calculated the skin friction coefficient. Where l: Length of component Re cut off = 38.21(l k) 1.053 k = 2.08 10 5 skin roughness value for smooth paint from Table 12.4 [2] Component form factor is calculated by using equation 12.30, 12.31 and 12.33 [2] for wing, tails, fuselage and nacelle. Table 23: Form Factor Calculation for Each Component Components FF Equation Wing, HT and VT FF = [1 + 0.6 ( t 4 (x c) m c ) + 100 (t c ) ] [1.34M 0.18 (cos Λ max ) 0.28 ] Fuselage FF = (1 + 60 f 3 + f 400 ) Nacelle FF = 1 + ( 0.35 f ) 49

Then, wetted areas, reference areas, thickness ratios of airfoils, interference factors and maximumthickness angles are found and tabulated in below table for the components from the drawing in OpenVSP and tables [2] : Note that, airfoil is decided as NACA0012 for horizontal and vertical tail. Table 24: Some Parameters for the Components Parameters Wing Horizontal Tail Vertical Tail Fuselage Nacelle S wet (ft 2 ) 1978.60 372.07 264.81 1692.659 50.052 S(ft 2 ) 2038.45 373.68 291.02 1767.694 50.852 t/c 0.159 0.12 0.12 - - x/c 0.211 0.3 0.3 - - Q c 1 1.04 1.04 1 1.3 Λ max,t ( ) 0.625 7.174 7.386 - - Supersonic wave drag and transonic parasite drag are not calculated for this aircraft since it cannot fly as supersonic. 8.5. Calculation of Induced Drag Factor Induced drag factor can be found for subsonic flights from the equation 12.48 [2] : K = 1 π AR e Where e = 1.78(1 0.045AR 0.68 ) 0.64 from the equation 12.49 [2] for straight-wing aircrafts AR = 8 So, by substituting the numerical values, K = 1 π 8 0.8105392 K = 0.0491 50

9. CHAPTER 9 THRUST CURVES 9.1. Introduction Uninstalled thrust is a value obtained by manufacturer after some tests in test facilities. Since these tests are conducted only with engine, without installing it on an airplane, the real thrust values during flight operations are not the same as uninstalled thrust. In order to calculate installed thrust, below study were done. 9.2. Uninstalled Thrust First of all, an internet search was done to find uninstalled thrust or power values of PW 127M engine with respect to Mach number and altitude but no such specifications could be found. Since the lack of information about uninstalled thrust values, Appendix A.4 of Raymer s Aircraft Design book was used. This Appendix gives typical turboprop engine characteristics. Appendix A.4 gives the sea level static power of that engine as 6500 hp. Since one of our engine has 2750 hp power, static power values were scaled and this scale factor used for scaling uninstalled thrust chart. S. C = 2 2750 6500 = 0.8461 Scaled uninstalled values were obtained and plotted below in one chart. Note that, these values are not so accurate, they are just approximated values obtained by scaling a sample engine s data. 51

30000.000 25000.000 Uninstalled Thrust 20000.000 15000.000 10000.000 5000.000 Sea Level 5000 ft 10 000 ft 15 000 ft 20 000 ft 25 000 ft 30 000 ft 0.000 0.000 0.100 0.200 0.300 0.400 0.500 0.600 0.700 Mach Number Figure 8: Uninstalled Thrust vs. Mach Number Chart Since the engine is turboprop, losses on propeller efficiency should also be calculated. First, obtain the propeller efficiency without any effects or losses. Propeller efficiency is obtained by using below equations. 1.000 0.900 0.800 Propeller Efficieny 0.700 0.600 0.500 0.400 0.300 0.200 0.100 Sea Level 5 000 ft 10 000 ft 15 000 ft 20 000 ft 25 000 ft 30 000 ft 0.000 0.000 0.100 0.200 0.300 0.400 0.500 0.600 Mach Number Figure 9: Propeller Efficiency vs. Mach Number by Uninstalled Thrust 52

Then the one should calculates the losses related to both engine and propeller, and should take into account losses in order to determine installed thrust and propeller efficiency. 9.3. Installed Thrust Corrections 9.3.1. Engine Related Losses These types of losses are resulted from core engine which is the same as core engine of a turbofan or turbojet. There are three types of engine related losses: 9.3.1.1. Inlet Recovery %Thrust loss = C ram [( P 1 ( P 1 ] x 100 P 0 P )ref 0 )actual Loss is multiplied by 2 since aircraft uses 2 engines. Therefore: %Thrust loss = 8.10 9.3.1.2. Bleed Air bleed air mass flow %Thrust loss = C bleed ( ) x 100 engine mass flow According to Appendix A.4, bleed air mass flow rate is 0.8 lb/s and engine mass flow rate is 42.3 lb/s. Since both mass flow rates scaled by using scale factor, equation can be written as follows, where C bleed = 2.0 %Thrust loss = 7.6 9.3.1.3. Power Extraction According to Appendix A.4, extracted power from engine is 54 kw for the 6500-hp-powered engine which is approximately 73.6 hp. To implement our engine, use scale factor. P ext = (P ext ) 6500hp 2 2750 6500 = 62.277 hp 53

9.3.2. Propeller Related Losses 9.3.2.1. Blockage Effect The blockage effect is a result of being a body just behind the propeller. The plane has wing-mounted two engines. Since there is a nacelle mounts the engine to the wing just behind each of the propellers, nacelles affect advance ratio in negative way, existence of these parts decreases advance ratio. J corrected = J (1 0.329 ( S C D 2 )) 9.3.2.2. Compressibility Effects Compressibility effect is related to local Mach number of propeller blades and more realted to tip Mach number of blades. Effect of compressibility is negative on propeller efficiency. 0.16 η pcorrected = η p (M tip 0.89) ( ) for tip M > 0.89 0.48 3t/c M tip = V + (πnd) 2 a For the tip Mach number less than 0.89, compressibility effects does not account. 9.3.2.3. Scrubbing Drag η peffective = η p [1 1558 D 2 ρ ρ SL (C fc S wet ) washed ] Above equation is calculated for wetted skins which are in the wake of propellers, since they are exposed to scrubbing drag effects. For this aircraft, two engines are mounted on the wing so wetted areas affected by scrubbing drag are nacelles of engines and a part of wing which is behind the propeller disk. 9.3.2.4. Cooling and Miscellaneous Drag D q cooling Where T is temperature in Rankine and ρ ρ SL = σ = ( 4.9 x 10 7 bhp T2 ) σ V 54

D q = (2 x 10 4 )bhp misc 9.4. Installed Thrust and Propeller Efficiency By using above two sections, calculate the installed thrust and propeller efficiency at each altitude. T installed = T uninstalled (thrust loss T uninstalled ) (thrus loss bleed T uninstalled ) D cooling D misc ( P ext V inf ) 25000 Installed Thrust 20000 15000 10000 5000 0 0.000 0.200 0.400 0.600 0.800 Mach Number Sea Level 5 000 ft 10 000 ft 15 000 ft 20 000 ft 25 000 ft 30 000 ft Figure 10: Installed Thrust vs. Mach Number Below, propeller efficiencies plotted with respect to Mach number at each altitude. 55

Propeller Efficiency 1.000 0.900 0.800 0.700 0.600 0.500 0.400 0.300 0.200 0.100 0.000 0.000 0.100 0.200 0.300 0.400 0.500 0.600 Mach Number Sea Level 5 000 ft 10 000 ft 15 000 ft 20 000 ft 25 000 ft 30 000 ft Figure 11: Propeller Efficiency vs. Mach Number by Installed Thrust 9.5. Conclusion From calculations, above installed thrust and propeller efficiency values were obtained. There are significant decreases in thrust values due to engine related losses; inlet recovery, bleed air and power extension. Installed thrust values at flight altitudes seems like to be sufficient enough to meet the requirements but to judge correctly that if this installed thrust values are sufficient or not, the one should do performance analysis in a detailed manner with more accurate engine data. 56

10. CHAPTER 10 WEIGHTS AND LOADS 10.1. Update of Take-off Gross Weight Below tables are taken by using statistical weight groups methods. This method is a bit complicated and much more accurate way to estimate the weight. In this method, lots of component of aircraft are considered separately one by one. 10.1.1. Passenger Mission Table 25: Passenger Mission Statistical Method Weight Calculation Passenger Weight [lbs] Weight [lbs] Equipment Structures 13688.17 5843.90 Wing 3476.43 Flight Controls 778.82 Horizontal Tail 257.98 Instruments 130.09 Vertical Tail 264.27 Hydraulics 151.97 Fuselage 7336.45 Electrical 578.58 Main LG 1393.94 Avionics 1840.27 Nose LG 232.33 Furnishing 507.83 Engine Mounts 726.77 Air Conditioning 570.27 Anti-icing 71.36 Handling Gear 10.70 Misc. 1204.00 We allowance 1129.60 Propulsion 3059.94763 Total We 23721.62 Engines - installed 2761.12 Starter (pneumatic) 54.80 Useful Load 11959.5551 Engine Controls 85.21 Crew+pass+luggage 6013.00 Fuel System 158.81 Fuel 5946.56 Take-off Gross Weight 35681.17 It can be observed that there is difference between iteration steps. Empty weight at first iteration is 22 lb higher than the last iteration. Also, it can be observed that there is difference between weights calculated by approximate methods and these statistical methods. Statistical empty weight is 22.6% 57

lower than the approximated one and statistical take-off gross weight is 6991 lb lower than approximated one. 10.1.2. Low Cost Maritime Surveillance Mission Table 26: Maritime Surveillance Mission Statistical Method Weight Calculation Maritime Weight [lbs] Weight [lbs] Equipment Structures 13456.16 5015.34 Wing 3414.87 Flight Controls 777.08 Horizontal Tail 252.74 Instruments 130.09 Vertical Tail 259.60 Hydraulics 151.97 Fuselage 7219.73 Electrical 578.58 Main LG 1354.87 Avionics 1840.27 Nose LG 227.57 Furnishing 419.60 Engine Mounts 726.77 Air Conditioning 570.27 Anti-icing 69.11 Handling Gear 10.37 Misc. 468.00 We allowance 1079.86 Propulsion 3125.73011 Total We 22677.09 Engines - 2761.12 installed Starter 54.80 Useful Load 11877.785 (pneumatic) Engine Controls 85.21 Crew+pass+luggage 3700.00 Fuel System 224.60 Fuel 8177.79 Take-off Gross Weight 34554.88 It can be observed that there is difference between iteration steps. Empty weight at first iteration is 22 lb higher than the last iteration. Also, it can be observed that there is difference between weights calculated by approximate methods and these statistical methods. Statistical empty weight is 23.6% lower than the approximated one and statistical take-off gross weight is 7131 lb lower than approximated one. At the beginning of conceptual design process, weight information was estimated very roughly based on historical data. These data did not represent truly. Then, aircraft's weight is estimated by considering big parts. Now, aircraft's weight is estimated by considering almost each part. Thus, these statistical 58

methods give us more accurate results while having more information about the aircraft and weights were compared at each iteration step during statistical method. Also, statistical method was compared with initial sizing. From before weight calculation to this method, major components of aircraft become smaller so it caused huge weight reduction. Results of above method will be verified in Size Matrix and Carpet Plots chapter. Cargo mission was not considered for this method since maritime and passenger missions' requirements also satisfy the weight consideration of cargo mission. 10.2. V-N Diagram 10.2.1. Stall Line n max = 1 2 ρ SL V E 2 C Lmax W S ρ SL = 0.002377 slug/ft 2 By using above equation, for positive and negative maximum lift coefficient values, stall lines are calculated. 10.2.2. Corner Speed By taking limit load factor as 3.5, corner speed at sea level was calculated as follow 3.5 = 1 2 0.002377 V 2 1.735 E,corner 41.28 Here C Lmax value was found as result of a set of iterations. C Lmax was found as 1.668 at corner speed. The iteration can be found in appendix 2. Below, corner speeds of positive and negative load cases are written V E,corner,positive load = 264.6958 ft/s V E,corner,negative load = 148.1993 ft/s 59

10.2.3. 1-G Stall Speed 1 = 1 2 0.002377 V 2 1.845 E,stall 41.28 V E,stall = 137.206 ft/s 10.2.4. Dive Speed Dive speed is, V dive = 1.5 V cruise or V dive = 1.2 V max Cruise speed is 223.28 ft/s which was determined in previous chapters, so V E,dive@sea level = 1.5 223.28 = 334.92 Since the drag divergence Mach number is not applicable to our aircraft, use the equation of T R = D to find the maximum speed. Maximum equivalent speed was found as 368.412 ft/s at sea level. Therefore, V E,dive@sea level = 1.2 368.412 = 442.094 As the largest value, V E,dive = 442.094 10.3. Gust Loads Calculate the mass ratio at sea level: μ = 2 W/S ρ g c C Lα Gust alleviation factor for subsonic flight: K = 0.88 μ 5.3 + μ U = U de K Where U de = 30 ft/s. U de is used as standard gust speed. 60

It is assumed that aircraft is in 1-g load when the gust load is experienced. Therefore, add the Δn to 1 n gust = 1 + Δn Table 27: Load Factor Corresponding to Mach Number Mach V e n(+) 0.1 111.64 1.344094 0.2 223.28 1.699191 0.3 334.92 2.073579 0.4 446.56 2.478447 0.5 558.2 2.930567 0.6 669.84 3.458271 Above table shows the gust load factors with respect to flight speed at sea level. Since at corner speed, cruise speed and dive speed the gust load factor does not exceed 3.5, which is limit load factor, gust load does not change V-n diagram boundaries determined in part 1 of this study. Use above V-n diagram and gust load calculations. Below, V-n diagram including gust load is plotted. Figure 12: V-n Diagram 61

11. CHAPTER 11 STABILITY AND CONTROL 11.1. Longitudinal Stability Analysis In order to determine the longitudinal stability, calculate neutral point first, and then static margin due to neutral point and center of gravity of aircraft. To calculate neutral point of aircraft following equation is used: h n = h n,wb + a HT a ac V HT (1 ε α ) where, h n,wb is neutral point of wing-body of aircraft; a ac is lift curve slope of entire aircraft; a HT is lift curve slope of horizontal tail; V HT is horizontal tail volume ratio and ε is downwash derivative which includes the downwash effect into the stability calculation. Using ε includes downwash effect and gives α an accurate result for neutral point. α Ratio of lift curve slopes includes the compressibility effect into the downwash derivative equation. Since the aircraft exceed subsonic conditions, only subsonic downwash derivative equation is used. 11.2. Variation of Neutral Point with Mach Number As stated section 1, neutral point is found as a function of neutral point of wing-body, lift curve slopes of horizontal tail and wing-body, horizontal tail volume ratio and downwash derivative. Here, excluding tail volume ratio, all other parameters are function of Mach number. Therefore, neutral point location of aircraft changes with Mach number. Below figure shows the relation. 62

Static Margin for Most Backward CG 0.300 0.280 0.260 0.240 0.220 0.200 0.180 0.160 0 0.2 0.4 0.6 0.8 1 Mach Number Figure 13: Variation of Neutral Point with Mach Number 11.3. Variation of Static Margin with Mach Number Static margin is a fundamental concept in order to analyze the stability of an aircraft. Static margin is denoted as K n. K n = h n h For a stable plane, static margin must be positive. In other words, cg location must be ahead of neutral point of aircraft. Below, the one can see the static margin change with respect to Mach number, similar to neutral point location variation. The plots are due to both most forward and most backward cg positions. 63

Static Margin for Most Forward CG 0.300 0.280 0.260 0.240 0.220 0.200 0.180 0 0.2 0.4 0.6 0.8 1 Mach NUmber Static Margin for Most Backward CG 0.300 0.280 0.260 0.240 0.220 0.200 0.180 0.160 0 0.2 0.4 0.6 0.8 1 Mach Number Figure 14: Variation of Static Margin for Most Forward and Most Backward CG Cases Static margin is positive in all flight regime, which means the aircraft stable. It was desired since the aircraft is used in civilian purposes. Static margin just slightly changes with Mach number between 0 and 0.4 Mach, which is actually includes the operational interval of the aircraft. It means that stability characteristics of the plane is almost not change due to Mach number. Moreover, by comparing above two plots, for most forward and most 64

backward cg positions, static margin changes from 0.24 to 0.22. This shows with cg change, stability characteristics of aircraft does almost not change. 11.4. Variation of C ma Point with Mach Number C ma curve is found by using neutral point and cg of aircraft. C ma = C Lα (h n h) For stability, C ma must be negative. Below two figures are pitch stiffness curves with respect to Mach number. Pitch Stiffness Derivative for most forward CG 0.000-0.500-1.000-1.500-2.000-2.500-3.000 0 0.2 0.4 0.6 0.8 1 Mach Number 65

Pitch Stiffness Derivative for most forward CG 0.000-0.500-1.000-1.500-2.000-2.500-3.000 0 0.2 0.4 0.6 0.8 1 Mach Number Figure 15: Variation of Pitch Stiffness Derivative for Most Forward and Most Backward CG Cases 66

12. CHAPTER 12 SIZING MATRIX AND CARPET PLOT The aircraft designed for mission requirements which played important roles in performance and hence dimensions of aircraft. Required wing loadings and powers were determined, weight was calculated for several times during design process. But there is last thing needed to be done in order to verify that if previous calculations are done properly and if they really gives the optimum design, or what is the optimum design. This last step is sizing matrix and carpet plot. For the design, wing area and power is selected as most important requirement driven parameters. Thus, these two parameters are manipulated by +-20% and new configurations are calculated. Then below matrix can be created. 1214.06 ft 2 1011.74 ft 2 809.39 ft 2 4152 hp W 0 = 39,540.45 W 0 = 38,640.05 W 0 = 37,708.94 2750 hp W 0 = 37,413.13 W 0 = 36,534.21 W 0 = 35,625.79 2100 hp W 0 = 36,997.34 W 0 = 36,122.67 W 0 = 35,218.75 By using above sizing matrix, one can calculate gross weight for each wing area at each thrust, and can expand these data. Below, wing area vs. gross weight plots for each thrust value can are given. 67

40000.00 39500.00 y = 4.5257x + 34051 Gross Weight (lb) 39000.00 38500.00 38000.00 37500.00 700.00 800.00 900.00 1000.001100.001200.001300.00 Wing Area (ft^2) Figure 16: Wing Area vs. Gross Weight at T=4135 hp Gross Weight (lb) 37600.00 37400.00 y = 4.4165x + 32056 37200.00 37000.00 36800.00 36600.00 36400.00 36200.00 36000.00 35800.00 35600.00 35400.00 700.00 800.00 900.00 1000.001100.001200.001300.00 Wing Are (ft^2) Figure 17: Wing area vs. Gross Weight at T=2750 hp 68

Gross Weight (lb) 37200.00 37000.00 y = 4.3949x + 31666 36800.00 36600.00 36400.00 36200.00 36000.00 35800.00 35600.00 35400.00 35200.00 35000.00 700.00 800.00 900.00 1000.001100.001200.001300.00 Wing Area (ft^2) Figure 18: Wing area vs. Gross Weight at T=2100 hp Above plots are used to expand area vs gross weight values. Then, by using these new power and gross weight values, wing loading calculations (performed in chapter 4) are repeated. Result of these calculations below figure is plotted. 42000.00 41000.00 40000.00 Gross Weight (lb) 39000.00 38000.00 37000.00 36000.00 4152 hp 2750 hp 35000.00 2100 hp 2100 W/S 34000.00 >2565 W/S 33000.00 Baseline Optimum 800 900 1000 1100 1200 1300 Wing Area (ft^2) Figure 19: Carpet Plot 69

On above plot, three parallel lines having less slopes (light blue, blue and range colored lines) are power curves. That is, if there is a 2750 hp engine wing area to gross weight values should be on blue line. On the other hand, yellow and purple lines are wing loading curves. There are two wing loading curves on the graph, because above 2565 hp cruise wing loading becomes determinant while below 2565 hp takeoff wing loading is determinant. Only one of the candidate engines has power below 2565 hp. Therefore on the graph this wing loading curve is named as 2100 W/S At chapter 11 it was said that carpet plot will be used to verify if iterations doing in chapter 11 gave best weight results or not. Therefore on above graph two aircraft configurations shown as points. The first one is called as baseline and it is the configuration that before doing iterations at chapter 11. It can be seen that the point has nearly 970 ft^2 wing area and 36600 lb. gross weight. The second point is called as optimum and that is the final configuration after iterations at chapter 11. The optimum configuration is clearly seen that it is so close to intersection of 2750hp power curve and cruise wing loading curve. It means that, results of iterations at chapter 11 gave almost the best wing loading and power result. As it was stated first, carpet plot performed its task satisfactorily and verified the iterations done at chapter 11. 70

13. CHAPTER 13 TRIM ANALYSIS 71

14. CHAPTER 14 COST ANALYSİS 14.1. Introduction Cost analysis is based on the trade study which was conducted at the beginning of design process. From the trade study, material used decided to be composite. The idea behind using composite material is that getting a lighter aircraft at the end. Lighter aircraft provides more cost efficient flight operations, although production cost per aircraft at target production quantity, which is 300, is slightly higher for composite materials. Here, between the production cost and operational cost, a trade was done and it was determined that less operational costs compensate higher production cost in long term. 14.2. Analysis Below two figures make clear the comparison between metal and composite materials in terms of production cost and operational costs. Table 28: Cost Analysis Corresponding to Material and Quantity Production Costs Operations and Maintenance Costs Production Quantitiy Composite Material (Million $) Aluminum (Million $) Composite Material (Million $) Aluminum (Million $) 1 513.524 467.743 7.991 8.124 2 305.505 276.801 7.991 8.124 3 227.667 205.747 7.991 8.124 4 185.681 167.545 7.991 8.125 5 158.987 143.311 7.991 8.125 6 140.327 126.401 7.992 8.125 7 126.449 113.842 7.992 8.125 8 115.666 104.096 7.992 8.125 9 107.011 96.280 7.992 8.125 10 99.887 89.853 7.992 8.125 20 64.543 58.048 7.993 8.126 30 50.680 45.618 7.994 8.127 40 42.955 38.704 7.995 8.128 50 37.917 34.200 7.995 8.128 72

60 34.321 30.987 7.996 8.129 70 31.597 28.555 7.996 8.129 80 29.446 26.635 7.997 8.129 90 27.696 25.073 7.997 8.130 100 26.236 23.771 7.998 8.130 110 24.996 22.665 7.998 8.131 120 23.926 21.710 7.998 8.131 130 22.991 20.876 7.999 8.131 140 22.164 20.139 7.999 8.132 150 21.428 19.482 7.999 8.132 160 20.766 18.892 8.000 8.132 170 20.167 18.357 8.000 8.133 180 19.621 17.871 8.001 8.133 190 19.122 17.425 8.001 8.133 200 18.662 17.016 8.001 8.134 250 16.815 15.368 8.003 8.135 300 15.472 14.168 8.004 8.136 400 11.604 10.626 8.004 8.136 500 9.283 8.501 8.004 8.136 Above, it can be seen that production cost of a composite aircraft is 780 000 dollars higher then aluminum aircraft. In addition, operational costs of a composite airplane are less then aluminum plane at an amount of 130 000 dollars per year. It means that for a 500-plane production, advantage of composite aircrafts on operations costs compensates production cost disadvantage in 6 years and for above 6-year operational life span composite aircrafts make profit with respect to aluminum aircrafts. Life-span cost of an airplane is shown in Figure 18.1 on Raymer s Aircraft Design book. There four major groups for a civilian aircraft as can be seen. They are, RDT&E and Flyaway Cost which are calculated together and includes Special Construction cost group, and Operations and Maintenance cost group. RDT&E and Flyaway costs are related to manufacturer and they determine directly the price of an airplane. On the other hand, maintenance and operation costs are related to operator. 73

14.2.1. Research, Development, Test and Evaluation Research, development, test and engineering costs are in fact non-recurring costs. Test costs are compulsory outcomes so that pass from FAA/EASA Certifications. In addition, this subtitle covers production tooling and facilities costs. Below table shows the costs of these components per aircraft for 500 production quantity. Table 29: RDT & E Cost Analysis Passenger Maritime Engineering Costs (Million $/Aircraft) 1.283113 1.239019 Tooling Costs (Million $/Aircraft) 0.988161 0.954203 Quality Control Costs (Million $/Aircraft) 0.799475 0.770508 Development Costs (Million $/Aircraft) 0.119673 0.116328 Flight Test Costs (Million $/Aircraft) 0.001458 0.001437 14.2.2. Flyaway The actual material cost which is consumed during production process in order to create an aircraft. In other words, flyaway cost consists of airframe, engine and avionics costs. Below table shows flyaway costs. Table 30: Flyaway Cost Analysis Passenger Maritime Manufacturing Costs (Million $/Aircraft) 4.731759 4.560319 Manufacturing Materials Cost (Million $/Aircraft) 0.014025 0.013455 Engine Cost (Million $/Aircraft) 3.361381 3.361381 Avionics Cost (Million $/Aircraft) 0.015396 0.015396 74

Considering above two cost groups cost of an airplane is tabulated with respect to production quantity. Table 31: Production Cost & Selling Price Corresponding to Quantity Passenger Maritime Quantitiy Cost per A/C (million $) Selling Price (million $) Cost per A/C (million $) Selling Price (million $) 1 513.51442 590.5 496.46846 570.9 2 305.49908 351.3 295.24759 339.5 3 227.66226 261.8 219.97302 253.0 4 185.67716 213.5 179.37670 206.3 5 158.98350 182.8 153.56915 176.6 6 140.32370 161.4 135.53054 155.9 7 126.44600 145.4 122.11593 140.4 8 115.66348 133.0 111.69395 128.4 9 107.00882 123.1 103.32918 118.8 10 99.88515 114.9 96.44451 110.9 20 64.54212 74.2 62.29358 71.6 30 50.67887 58.3 48.90181 56.2 40 42.95416 49.4 41.44118 47.7 50 37.91643 43.6 36.57633 42.1 60 34.31983 39.5 33.10352 38.1 70 31.59608 36.3 30.47376 35.0 80 29.44587 33.9 28.39790 32.7 90 27.69528 31.8 26.70795 30.7 100 26.23570 30.2 25.29903 29.1 110 24.99552 28.7 24.10194 27.7 120 23.92540 27.5 23.06906 26.5 130 22.99015 26.4 22.16640 25.5 140 22.16393 25.5 21.36900 24.6 150 21.42727 24.6 20.65806 23.8 160 20.76524 23.9 20.01916 23.0 170 20.16612 23.2 19.44101 22.4 180 19.62063 22.6 18.91462 21.8 190 19.12127 22.0 18.43275 21.2 200 18.66193 21.5 17.98952 20.7 250 16.81510 19.3 16.20757 18.6 300 15.47117 17.8 14.91099 17.1 400 11.60338 13.3 11.18324 12.9 75

500 9.28270 10.7 8.94659 10.3 Above Table shows the estimated production costs and estimated selling price in order to making 15% profit from each aircraft. The target production quantity is 500 aircrafts which means, each aircraft is planned to be sold by 10.70 million dollars for passenger transport configuration and 10.30 million dollars for maritime surveillance configuration. Cargo plane configuration is more like to be maritime surveillance configuration since the plane does not involve passenger seats and passenger compartment. Therefore, an aircraft which is dedicated to cargo transportation mission has selling price of 10.30 million dollars. 14.2.3. Operations and Maintenance cost Operations and Maintenance cost can be divided by three main groups; Oil and fuel costs, Crew salaries and maintenance cost. Each cost group is respect to one-flight-hour term. Here, the one assumed that one airplane flies 2500 hours per year. Table 32: Operations & Maintenance Cost Passenger Maritime Oil and Fuel Costs ($/Flight Hour) 364.00 500.58 Crew Salaries ($/Flight Hour) 1232.65 781.44 Maintanence ($/Flight Hour) 1615.08 1468.67 Total Cost ($/Flight Hour) 3211.73 2750.68 76

15. CHAPTER 15 DESIGN SUMMARY AND DRAWINGS 15.1. Special Considerations Cargo will be loaded as a container. Cargo door is big enough to take whole container. Cargo can load and unload very quickly with this way. Also, cargo door and fuel cap will be located on different sides of aircraft since it can be loaded and refuel in same time. In order to take-off from runways such as dirt, grass, concrete and etc., aircraft has sufficient propeller clearance. It is achieved by putting engine above the wing and fuselage also has clearance. In addition, high-wing configuration was chosen. Propeller ground clearance is 9.09 ft. Fuselage hull was designed to take-off from water. There are sharp corners under fuselage to release from water easily. All calculations are done by considering fresh water case, since fresh water has a lower density. Position of the aircraft on the water is separately considered by calculating buoyancy and buoyancy center was found close to center of gravity with a little bit difference and it creates 1.9 incidence angle while aircraft positioning on the water. First, passenger number is considered as 28. Since our engine is new, 23 passenger is also satisfy the fuel burn per passenger requirement. To satisfy sea state level 3 requirement, propeller was calculated above water with enough clearance. Propeller water clearance is higher 3.17 ft. It is enough to satisfy. 15.2. Design Summary Table 33: Design Summary Geometric Characteristics Dimensions Length (ft) 73.16 Height (ft) Land/Water 11.8 Draught (ft) Wheels Down 3.6 Wing Span (ft) 83.13 Wing Root Chord (ft) 10.73 Wing Tip Chord (ft) 8.59 Wing AR 8 Wing Area (sq ft) 864.37 HT Span (ft) 23.88 HT Root Chord (ft) 8.59 HT Tip Chord (ft) 4.29 77

HT AR 3,71 HT Area (sq ft) 153.81 VT Span (ft) 11.59 VT Root Chord (ft) 10.73 VT Tip Chord (ft) 8.59 VT AR 1.2 VT Area (sq ft) 112.00 Propeller Diameter (ft) 10,86 Propeller Water Clearance (ft) 3,17 Propeller Ground Clearance (ft) 9,09 15.3. Drawings Figure 20: 3D View 78

Figure 21Fuselage Structure Figure 22: Wing Structure 79

Figure 23: Tail Structure Figure 24 Horizantal Tail Structure 80

16. Figure 25: Front View 17. 81