Calculating IAS from lift equation
 

Hi,
Given Cl = 0.2, Weight of aircraft = 170 tonnes and Smainwing = 250m^2 and Stailwing = 56m^2 is it possible to calculate the IAS (in km/hr) the aircraft should travel so that L = W ?

I am really struggling to work this out - the numbers I am coming up with just don't make sense, so if anyone out there could help me out it would be greatly appreciated!

Thanks in advance!

Given that L = 0.5*rho *V^2 *S *Cl

Rewriting to solve for V yields:

V = sqrt(L / (0.5*rho*S*Cl))

This gives the true airspeed, which is equal to IAS at sea level.
At Sea level, rho = 1.225 kg/m^3
Neglecting the contribution of the tailplane to the total lift (its Cl is usually different than that of the main wing and can be negative):

V = sqrt(170E3 /(0.5*1.225* 250*0.2)) = sqrt(5551) = 74.5 m/s

74.5 *3600/1000 = 268 km / hr

This is the true airspeed at sea level, equal to the indicated airspeed at sea level. For higher altitude, the true airspeed needs to go up. Neglecting compressibility effects the IAS will stay roughly the same for higher altitudes.

Best,

ATCast

Multiply TAS by the square root of relative density (sigma = rho/rho_0) to get CAS, then cross-correct using the calibration chart in the back of the POH to give IAS.

TAS is equal to EAS at sea level, at low speeds this is also equal to CAS, it is not equal to IAS.

G

Thanks so much! makes sense now! :)