P09123: MAV 1 & 3
/public/

# Aerodynamic Surfaces and Control Surfaces

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

## 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

## 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
• 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
• Total Drag Coefficient = 0.45
• Total Drag = 1.28 lbf
• This is a conservative estimate and will be tested once built