1. The problem statement, all variables and given/known data "The lift gradient of a wing under actual flight conditions is 0.1179 per degree. Calculate the lift-drag ratio of the wing with an angle of attak of 3 degrees?" Given is: altitude=5000 m velocity=225 m/s wing area S=149 m2 wing span b=34.5 m span efficiency factor e1=0.82 Cd (profile drag coefficient) at 3 degrees=0.0062. ρ∞=0.736 kg/m3 p∞=5.41*10^4 Pa, T∞=255.7 K cp=1008 J/kg*K 2. Relevant equations I don't exactly know which equation to use, that's the whole point of my question. You might can use a=dCl/dalpha = a0/(1+(a0/pi*A*e1)). Note that everyting in the equation is in radians. Maybe you can calculate Cl with this equation, and as you know Cd (given) you can calculate the L/D ratio. 3. The attempt at a solution I've tried several things like the lift gradient equation a=dCl/dalpha=a0/(1+(a0/pi*A*e1), is I can calculate the aspect ratio A (S/b^2). A=35.4^/149≈0.231. I also know the span efficiency factor e1, as this is given (0.82). The fact is that I don't know if this equation is right. The other thing I know, is that the equation for lift drag ratio is L/D = Cl/Cd. Am I right to say that we know Cd? This is 0.0062. Then we only have to calculate the lift coefficient, but I don't know how to do that without the lift given or the mass of the aircraft. (as the equation is L=Cl*(0.5*ρ*V^2)*S. Can anyone help me with this problem?