Question - Thesis - Lift coefficient #3
Description
Thanks for sharing this project. I have the following question:
In your thesis:
In appendix B.3:
- C_L(alpha) is fitted to the value 0.1081
Right below formula 3.14:
- C_L(alpha) is defined with the value 0.008905
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Does this mean the value 0.1081 is the lift coefficient as used within formula 3.11 specific for the airfoil shape and the value 0.008905 is the 'overall' lift coefficient considering the profile, length and 'wingspan' of your used thrust vane ?