Hi I am currently analyzing different wing designs using ANSYS Fluent. One of my designs has a straight leading edge but a swept trailing edge thus it is tapered with wing root chord at 50mm and wing tip chord at length 200mm. Wing span is 250mm. (See Picture) The airfoil shape used is the S1223 which has a CL max around 2.3. Therefore using the Lift formula which (i know is approximate), I calculate the wing can produce 58.5N of lift at 36.452m/s.
However using ANSYS Fluent I find the wing can only produce a maximum of 39N of lift which is a CL of 1.5.
I am fairly certain the simulations I'm running are accurate but need advice on what could be causing this difference. Assuming the results I've obtained are correct. What is the reason for the large difference in coefficient of lift ? Let me know if I need to provide any additional information! Any help greatly appreciated! EDIT : Forgot to add that the simulation used assumes the wind tunnel boundary is at wing tip and wing root. Thus the wing span takes up the full wind tunnel.