- #1
marv
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Homework Statement
"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
Homework 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.
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?