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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. - PowerPoint PPT Presentation
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DESIGN OF THE 1903 WRIGHT FLYER
REPLICA
MADRAS INSTITUE OF TECHNOLOGYCHENNAI - 44
WEIGHT ESTIMATION
Elements Quantity Weight (N) Structure 7.3 Engine 1 5 Propeller 2 1 Landing gear 3 2.25 Servo motors 3 0.58 Radio Controls All 0.212 Fuel 0.3 litres 2.36 Mounting + belt 3 Fuel tank 2 0.6 Misc. 2.5
TOTAL WEIGHT 24.802 N
AERODYNAMIC DESIGN
Lift Calculation
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 .
CL Vs Alpha curve for inviscid flow
-1.5
-1
-0.5
0
0.5
1
1.5
2
2.5
3
-15 -10 -5 0 5 10 15 20 25
Alpha
CL
Drag Polar
Induced Drag EstimationAR for a biplane = 4 b/cSpan = 5 feetChord length = 12 inchesAR = 20Gap = 9 inches
CDi = 1/(AR)*(1+)CL2
CDi = 0.11136 CL 2
profile
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 isCD = 0.1303 + 0.1277CL
2
Wing warp
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
POWER PLANT SELECTION
Power available
0
10
20
30
40
50
60
70
80
90
100
1 2 3 4 5 6 7 8 9
velocity
po
we
r
1500
2500
3000
3500
4000
4500
5000
1000
5500
6000
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.
STRUCTURAL DESIGN
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
WING REAR SPAR
The maximum stress on the rear spar σz = 40 MPa The maximum allowable bending stress for spruce wood = 41 MPa
ELEVATOR AND RUDDER SPARS
ELEVATOR FRONT SPAR REAR SPAR
RUDDER SPAR
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.
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 = 8 mm
MATERIALS TO BE USED
S.NO COMPONENT MATERIAL
1 WING SPARS SPRUCE
2 OTHER STRUCTURAL COMPONENTS
BALSA
3 SKIN REYNOLDS PLASTIC
4 FUEL TANK PLASTIC
PERFORMANCE CALCULATION
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
LEVEL FLIGHT PERFORMANCE
Cruising Velocity = 4.7 m/sStalling Velocity = 2.35 m/s (CLmax = 2.04)VminD = 2.64 m/sDmin = 2.423 m/sPmin = 6.09 WVminP = 2.06 m/s
Range = 1.616 km (for cruise condition)
Endurance = 5 minutes 54 seconds
CLIMB PERFORMANCE
R/C = Excess Power / WeightExcess Power = Power Available – Power RequiredMaximum rate of climb occurs at 6 m/s
VelocityPower
AvailablePower
RequiredExcess Power
R/C maxAngle of Climb
m/s W W W m/s degree
2 8 6.108897 1.891103 0.075644 2.167557
3 12 7.83886 4.16114 0.166446 3.180502
4 30 13.4841 16.5159 0.660636 9.50645
5 60 22.52976 37.47024 1.49881 17.44327
6 90 36.55183 53.44817 2.137927 20.87438
7 90 60.97091 29.02909 1.161164 9.548366
8 91 90.17925 0.820751 0.03283 0.235128
EXCESS POWER
0
10
20
30
40
50
60
70
80
90
100
2 4 6 8
VELOCITY m/s
PO
WE
R W
POWERAVAILABLE
POWERREQUIRED
Take – Off
The take-off is curved up into 3 phasesThey are ground run, transition and initial climb upto 2 m and the same is repeated for landingGround run
Vavg = 0.7 VLO (lift off velocity)
= 0.84 Vstall
r = 0.1 for grass landVLO = 2.82 m/sCLLO= 0.8 CLmax
Ground Run = 6.3 mGround Run in transition = 2.1 mGround Run in climb = 4.48 mTotal take off distance = 12.88 m
GroundRun
Transition Climb
Landing & Turning performance
Landing distance total = 17.11 mMinimum turn radius = 0.4 m Corresponding time taken = 1.15 secondsV-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 aircraftThe maximum load factor = 275/25 = 11
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
0
1
2
3
4
5
6
7
8
9
10
11
12
0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35
Velocity
n
stall limit
Structural limit
Max velocity
STABILITY ANALYSIS
LONGITUDINAL STATIC STABILITY
DIRECTIONAL STATIC STABILITY
CROSS COUPLING EFFECT
Increment in Rolling moment due to pitch rate(constant for different pitch rates)
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 1 2 3 4 5 6 7 8 9 10
(deg)
Incr
emen
tal R
olli
ng
mo
men
t co
-eff
icie
nt
CR
Change in yaw co-efficient for different pitch rates (in rad/s)At cruising velocity of 4 m/s
0
0.0005
0.001
0.0015
0.002
0.0025
0.003
0 10 20 30 40 50 60 70 80 90 100
Wing warp deflection angle (deg)
In
cre
men
tal
Ya
w c
o-e
ffic
ien
t C
N
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
1.1
1.2
1.3
1.4
1.5
1.5708
COST ESTIMATION
Item Cost 4 channel radio control (with transmitter, receiver, 4- servos, Connectors etc.) 15000 Engine (0.25 bhp) 4000 Balsa Wood 2500 propellers 700 Fabrication cost 1000 Skin, belt, pulley, wires, LG etc. 1500 Total 24 700
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.
POSITION OF SERVOS
POSITION OF RECEIVER
PROBLEMS
We are amateur designersBut we are confident that we can overcome this problem after taking part in this workshopSince the stability of the aircraft is at a high risk we feel that flying the aircraft safely would require a lot of training