# Lift Calculations.

## Main Question or Discussion Point

This is a simple question hopefully.

The equation for determining lift produced by an aircraft is the following if I am correct:
L = (1/2) d v2 s CL

If I was designing an aircraft that would maintain flight at standard air pressure at 1000 feet .002308 ISA at 130 MPH with a wing area of 60 Sq Ft. and a Coefficient of lift of 1, then I would plug the numbers in accordingly..
(.5)*(.002308)*(130^2)*(60)*(1) = 1170
Would this mean that the maximum lift capacity for this type of plane be 1170 lbs, or am I way off base here?

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What you calculated is the lift force on the wing at those conditions. You don't know that it's the maximum achieveable by the profile of the wing. The lift on the wing will change with angle of attack and altitude, and it isn't neccessarily max at 0 degrees.

The equation relating these is complex, you'd have to do experiments to find out. I've done it using a wind tunnel and CFD simulations.

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Also your velocity should be in ft/s rather than mph so your units cancel appropriately.

yes.. keep the units straight...

the above formula is good for approximations....

but keep in mind.. the fuselage can create lift.. as well as the tail.. (Lift downward actually)...
so the sum of all lifting forces need to be considered...
as well as a particular wings Cl max...

Think of:

1. The actual 3D wing will have a CL max, depending of the shape of a 2D section or profile and on the shape tec. of the actual wing.
2. The Lift generated depends on the angle of attack and thus the CL. So what angle of attack do you consider?
3.