Table of Contents

NOTE: For detailed calculations please see logbooks
Airfoil Selection
 Defined preliminary flight envelope
 Minimum velocity at 10 mph
 Cruise velocity at 25 mph
 Maximum velocity at 35 mph
 Calculated operating parameters at each extreme of
flight envelope assuming a three foot span and six inch
chord and a weight of 1.5 kg (3.31 lbf)
 Minimum Reynolds number = 50000
 Minimum Clmin = 8.62
 Cruise Reynolds number = 150000
 Cruise Clmin = 2.1564
 Maximum Reynolds number = 180000
 Maximum Clmin = 0.7041
 Benchmarked from P08121
 Used a list generated by last years team to start
 Added more airfoils to analysis from the UIUC
airfoil database
 Looked for low Reynolds Number airfoils
 Analyzed list of airfoils in XFoil v. LR5
Cl vs. Angle sf Attack for various airfoils
Other Comparison Plots
 Chose Selig S1210 airfoil due to high lift coefficiend and high operating envelope
S1210 profile
S1210 Cl vs. Angle of Attack
Finite Wing Calculations
 Calculated minimum CL for the wing using finite wing
length and lift curve slope corrections
 Original dimensions of 6in chord and 3ft span gave a static angle of attack at cruise of 7 degrees
 Resized wing through iteration (can be seen in
logbook) to get new dimensions
 c = 7in
 span = 3.5ft
 static AoA at cruise = 4.5 degrees
Horizontal Tail
Van put your stuff here
Vertical Tail
 Using S9033 airfoil (same as P08121)
 From references (see benchmarking section) calculated
vertical tail dimensions
 chord of vertical tail = 4in
 span of vertical tail = 7 in
Control Surfaces
 From references, the control surfaces were sized
accordingly and are listed below
 Aileron
 V aileron = 0.4 (from reference)
 span aileron = 3.15ft
 chord aileron = 0.84 in
 Elevator
 Vbar = 0.3
 span elevator = 0.9*span wing = 1.004 ft
 chord elevator = 2.66 in
 Rudder
 Vbar = 0.11
 span rudder = 6.3 in
 chord rudder = 2.6 in
 Aileron
Drag Calculations
 Break the total drag coefficient up into four parts
 Wing Drag
 Add 2D drag (from Xfoil) to calculated
induced drag
 2D drag wing = 0.0269
 Induced drag wing = 0.0627
 Total wing drag = 0.0896
 Add 2D drag (from Xfoil) to calculated
induced drag
 Horizontal Stabilizer Drag
 2D drag HS = 0.0308
 Induced Drag HS = 0.0773
 Total HS drag = 0.1081
 Vertical Tail Drag
 2D drag VT = 0.00838
 No induced drag due to a 0 degree angle of attack and symmertic airfoil
 Fuselage Drag
 From Reference = 0.25
 Wing Drag
 Total Drag Coefficient = 0.45
 Total Drag = 1.28 lbf
 This is a conservative estimate and will be tested once built