I am building an conventional plane with AR=6 and I find the wing area,A.So I have selected an airfoil for building my plane that is sd7062. Using xflr5 I got a cl vs alpha graph and I took the cl value at 4 degrees . using this cl value I calculated my lift and equated to the weight of my plane which is 2kg*9.81=19.6N .now equating L=W I find the area ,A and using this area and aspect ratio 6 or 7 I calculate the span and chord .Now when I do an analysis for the wing dimensions I got and plot a graph between Fz(lift) and alpha graph and find that at 4 degrees the Fz value is 13.078N ....why is there a decrease in my lift.....if there's a decrease then are the calculations I have done wrong?..please help me out![enter image description here]1

  • $\begingroup$ Difference due to 3D effects (finite wing)? $\endgroup$ – jjack Nov 30 '18 at 10:57
  • $\begingroup$ Sir,can you elaborate please? $\endgroup$ – sai teja Jan 10 '19 at 8:32
  • $\begingroup$ Does your wing have any twist or taper? $\endgroup$ – Koyovis Jul 31 '19 at 5:27
  • $\begingroup$ no twist, no taper just a simple rectangular wing... $\endgroup$ – sai teja Aug 1 '19 at 6:03

When you have taken the Cl of the airfoil concerned, the Co-efficient refers to an infinite wing. You have to compute the CL of the wing, which refers to a finite wing, taking into consideration the 3D nature of the flow.

  • $\begingroup$ So does this mean I have to build my wing with more area? $\endgroup$ – sai teja Sep 18 '19 at 4:50
  • $\begingroup$ Given you have tested this for 10 m/s velocity, I would suggest you to either increase your velocity or increase the wing area. Both of which will work. $\endgroup$ – Gaurav Sep 19 '19 at 5:49

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