|Quoting David L (Reply 3):
I'm not sure I follow. Are you saying the climb angle or the angle of attack (or something else) is 5.33o? If it's the angle at which the aircraft climbs and you have a given speed, wouldn't simple trigonometry give you the rate of climb?
No, you add the angle of attack to the angle of climb (5.33*) to get the attitude, but the point is that the angle of attack depends on the airspeed, which in turn affects the drag quantity which allowed me to calculate excess thrust in the first place.
The angle of climb is trigonometry it is implie that if the aircraft had excess thrust to match its weight it would climb straight up, but because its excess thrust is only a fraction of its weight, it climbs at a certain angle depending on that amount.
However, it doesn't tell me the lift produced by the airfoil.
The calculation of the climb angle was sin-1(10000/107456) = 5.339748767 degrees.
How do I go about determining the lift co-efficient? I used the NASA foil simulator to create the wing to those specifications, and it shows a lift calculation in pounds and a co-efficient. But I'm wondering how it calculates them.
[Edited 2006-03-22 14:42:16]