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DESIGN OF THE 1903 WRIGHT FLYER REPLICA

DESIGN OF THE 1903 WRIGHT FLYER REPLICA. MADRAS INSTITUE OF TECHNOLOGY CHENNAI - 44. WEIGHT ESTIMATION. TOTAL WEIGHT 24.802 N. AERODYNAMIC DESIGN. Lift Calculation. CL Vs Alpha curve for inviscid flow. 3. 2.5. 2. 1.5. 1. C L. 0.5. 0. -15. -10. -5. 0. 5. 10. 15. 20.

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DESIGN OF THE 1903 WRIGHT FLYER REPLICA

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  1. DESIGN OF THE 1903 WRIGHT FLYER REPLICA MADRAS INSTITUE OF TECHNOLOGY CHENNAI - 44

  2. WEIGHT ESTIMATION TOTAL WEIGHT 24.802 N

  3. AERODYNAMIC DESIGN

  4. Lift Calculation CL Vs Alpha curve for inviscid flow 3 2.5 2 1.5 1 CL 0.5 0 -15 -10 -5 0 5 10 15 20 25 -0.5 -1 -1.5 Alpha • As the t/c ratio of the airfoil is less than 0.05 the classical theory of thin airfoils can be employed, by using this theory all the parameters other than drag is forecasted .

  5. Drag Polar • Induced Drag Estimation AR for a biplane = 4 b/c Span = 5 feet Chord length = 12 inches AR = 20 Gap = 9 inches • CDi = 1/(AR)*(1+)CL2 • CDi = 0.11136 CL 2profile • Profile Drag Calculation CD wet /Cf = 1+ 1.5 (t/c)3/2 +7 (t/c)3 CDp/Cf = 60 (t/c CL/5)4 • The drag polar of our model is CD = 0.1303 + 0.1277CL2

  6. Wing warp

  7. Rolling moment for Both wings = 0.56 (k/c) sin  (l+ k cos )2 Where c is the chord of the wing  is the angle of warp from the undisturbed configuration k is the length of wing warp

  8. POWER PLANT SELECTION

  9. Power available 100 90 80 1500 70 2500 3000 60 3500 power 4000 50 4500 5000 40 1000 5500 30 6000 20 10 0 1 2 3 4 5 6 7 8 9 velocity

  10. specifications • From drag calculations the power required 0.25 bHp • Diameter of the propeller ( 2-blade propeller) 10 inches The diameter is determined from the thrust to be produced. The ground clearance was also taken into account while determining the diameter of the propeller.

  11. STRUCTURAL DESIGN

  12. WING FRONT SPAR The bending moment about X axis (Mx) = 14.96 Nm The formula used, Mxc =(Mx-(My*Ixy/Iyy)) /( 1-Ixy²/ (Ixx*Iyy)) =36.65 Nm Myc =(My-(Mx*Ixy/Ixx)) / (1-Ixy²/ (Ixx*Iyy)) = -108.04 Nm The maximum stress on the front spar σz = 32 MPa The maximum allowable bending stress for spruce wood = 41 MPa

  13. WING REAR SPAR • The maximum stress on the rear spar σz = 40 MPa • The maximum allowable bending stress for spruce wood = 41 MPa

  14. ELEVATOR AND RUDDER SPARS ELEVATOR FRONT SPAR REAR SPAR RUDDER SPAR

  15. Design of truss members • Though the diameter of the truss members are different, for fabrication simplicity all the members are designed with diameter 5 mm.

  16. PROPELLER SHAFT DESIGN The formula used to calculate the diameter of the shaft Me = (M +√(M²+T²)) / 2 = 0.15306 Nm Te = √(M²+T²) = 0.7938 Nm Maximum bending strength of the balsa wood σb = 1.18934*10^7 N/m τ = 2482113 N/m² Dmoment =7.15 mm Dtorque =7.95 mm Therefore the required diameter for the propeller shafts = 8mm

  17. MATERIALS TO BE USED

  18. PERFORMANCE CALCULATION

  19. INTRODUCTION • The performance design covers the five major calculations which are listed below • Steady level flight performance • Climb performance • Range & Endurance • Take – Off & Landing • Turn Performance

  20. LEVEL FLIGHT PERFORMANCE Cruising Velocity = 4.7 m/s Stalling Velocity = 2.35 m/s (CLmax = 2.04) VminD = 2.64 m/s Dmin = 2.423 m/s Pmin = 6.09 W VminP = 2.06 m/s Range = 1.616 km (for cruise condition) Endurance = 5 minutes 54 seconds

  21. CLIMB PERFORMANCE • R/C = Excess Power / Weight • Excess Power = Power Available – Power Required • Maximum rate of climb occurs at 6 m/s

  22. Take – Off • The take-off is curved up into 3 phases • They are ground run, transition and initial climb upto 2 m and the same is repeated for landing • Ground run Vavg = 0.7 VLO (lift off velocity) = 0.84 Vstall r = 0.1 for grass land VLO = 2.82 m/s CLLO= 0.8 CLmax Ground Run = 6.3 m Ground Run in transition = 2.1 m Ground Run in climb = 4.48 m Total take off distance = 12.88 m Transition Climb Ground Run

  23. Landing & Turning performance • Landing distance total = 17.11 m • Minimum turn radius = 0.4 m • Corresponding time taken = 1.15 seconds • V-n diagram is a plot between the velocity and load factor ( n = L/W) • It gives the structural limit (max) of the aircraft and the highest and lowest possible velocity that can be reached by the aircraft • The maximum load factor = 275/25 = 11

  24. V-n DIAGRAM From the v-n diagram it is clear that n is maximum for the velocity of 8 m/s and the maximum velocity can be 35.75 m/s for the n value less than 11

  25. STABILITY ANALYSIS

  26. LONGITUDINAL STATIC STABILITY

  27. DIRECTIONAL STATIC STABILITY

  28. CROSS COUPLING EFFECT

  29. Change in yaw co-efficient for different pitch rates (in rad/s) At cruising velocity of 4 m/s 0.003 0.0025 0.1 0.2 0.3 N 0.4 0.002 0.5 0.6 0.7 Incremental Yaw co-efficient C 0.8 0.0015 0.9 1 1.1 1.2 0.001 1.3 1.4 1.5 1.5708 0.0005 0 0 10 20 30 40 50 60 70 80 90 100 Wing warp deflection angle (deg)

  30. COST ESTIMATION

  31. RADIO CONTROL COMPONENTS • Engine throttle is controlled by servo motor. • Four channel receiver set with 4 servo motors and connectors are used. • The R/C unit weighs about 0.75 N. • The R/C unit is placed just below the wing so that it reduces the bending moment caused by the lift.

  32. POSITION OF SERVOS

  33. POSITION OF RECEIVER

  34. PROBLEMS • We are amateur designers • But we are confident that we can overcome this problem after taking part in this workshop • Since the stability of the aircraft is at a high risk we feel that flying the aircraft safely would require a lot of training

  35. INTERACTION....

  36. THANK YOU

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