aerotools package¶
Submodules¶
aerotools.vpm module¶
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class
aerotools.vpm.
Airfoil
(NACA_Name, Chord_Length=1, NUM_SAMPLES=100, Angle_Of_Attack=0)[source]¶ Bases:
object
Airfoil handles calculations made on 4-digit series NACA airfoil.
The Airfoil class calculates the x-y coordinates of all boundary points on an NACA 4-digit series airfoil.
Parameters: - NACA_ID – The NACA 4-digit series airfoil name i.e. “NACA0012”
- chord – The chord length of the airfoil
- NUM_SAMPLES – The number of samples / panels considered
- angle_of_attack – The angle of attack of the airfoil
- max_camber – The maximum camber of the airfoil i.e “0/100, 1/100”
- position_max_camber – The position of the max. camber i.e. “4/100”
- thickness – The maximum thickness of the airfoil i.e. “12/100”, “08/100”
- x_boundary_points – The x-locations of each boundary point on the airfoil
- y_boundary_points – The y-locations of each boundary point on the airfoil
- full_coefficient_lift – The coefficient of lift of the airfoil
- pressure_coefficient – The pressure coefficient at each (x,y) point
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get_airfoil_coordinates
()[source]¶ Returns the coordinates of the airfoil.
Returns: An array of x-coordinates and y-coordinates of boundary points. [X,Y] Return type: tuple
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get_coefficient_lift
()[source]¶ Returns the coefficient of lift.
Returns: The airfoil’s coefficient of lift (per meter span), Cl. Return type: float
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get_panel_coordinates
()[source]¶ Returns the coordinates of the midpoints of the panels.
Returns: An array of x-coordinates and y-coordiantes of the panel mid- points. [X, Y] Return type: tuple
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get_pressure_coefficients
()[source]¶ Returns the pressure coefficient at the midpoint of each panel.
Returns: Pressure coefficient at each boundary point, Cp. Return type: float[]
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set_airfoil
(NACA_ID)[source]¶ Sets the airfoil type used.
Parameters: NACA_ID – The 4-digit series airfoil name. Example: ‘NACA0012’
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set_angle_of_attack
(angle)[source]¶ Sets the angle of attack to use for next calculations.
Parameters: angle – The new angle of attack