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Note that this lift coefficient is for the total aircraft. It includes the lift of the wing, the fuselage lift and also the negative lift of the vertical stabilizer. There is also a small vertical component of thrust that is neglected.
If you know your airfoil lift curve slope, modify the result from the plot above by the ratio between the airfoil lift curve slope and $2\pi$. Now your lift coefficient will become: Now your lift coefficient will become:
The lift curve slope of a flat plate is the same as that of regular airfoils, complete with stall. They have a zero-lift angle of attack of 0° (obviously) and separated flow on the suction side. The maximum lift coefficient of 0.7 to 0.8 is reached at moderate angles of attack - details depend on the aspect ratio .
Above is the image of the question I want to ask. Please clarify on how the pressure coefficients of the upper and lower surfaces of an airfoil can be used, also if there is a set equation to calculate this lift coefficient.
$\begingroup$ @JamesDavis Changing the AoA changes the CL (Lift coefficient). For an AoA where CL is not zero, the total lift will be proportional to the square of the velocity. In your example chart CL goes to zero at an AoA of -5 degrees. If you maintained an AoA of -5 degrees, then, the wings would generate no lift, regardless of airspeed.
Eg: If I have a 2412NACA airfoil and would like to know the lift coefficient for an angle of attack $\alpha$ = 3°. What would be the corresponding lift coefficient? Thanks
I'm teaching an introductory course in fluid dynamics and I'd like to show some "real" data regarding drag and lift forces on an airfoil. It is easy to find data on the geometry of typical airfoils...
"The wing lift formula shows that lift of a wing is proportional to its area". That's absolutely true. However a wing generates both lift and drag. Drag is of two natures: Parasitic drag and lift-induced drag. The sum of all drag is the total drag: Source: Wikipedia. Induced drag decreases with speed, and has its origin in wing tip vortices.
Your intuition on lift/drag probably starts here -- but is more properly attacked to Normal/Axial force components. When angle of attack increases, the Normal force increases -- and while it still mostly points in the Lift direction, it rotates back and now also points in the drag direction. However, Lift is always perpendicular to velocity.
Note that the lift curve slope is per degree. $\frac{t}{c}$ is the relative thickness of the airfoil. In figure 21 of that chapter he uses the concept of the lift angle, the angle at which the lift coefficient reaches unity, to show how the lift curve slope varies over airfoil thickness.