Did Ashok publish an equation or a set of tabular data for the wind tunnel results? I hate trying to read data off of graphs. (See http://www.krnet.org/as504x/repeat_1m.gif )
Looking at the graph it appears the best CL to operate at is approx 0.72. The point where Cd starts to increase with increasing CL. That gives an angle of attack of 5 degrees. The max CL that is achievable is approximately 1.2 somewhere around 14 degrees AoA (see table 2 on http://www.krnet.org/as504x/design.html) As the light bulb goes on, that CL is going to be the best CL/Cd ratio so it can be used to calculate the best glide speed. Lift = Weight = (1/2) rho V^2 CL S solving for V V = Sqrt(2W/ (CL rho S)) solving for CL CL = 2W / (S rho V^2) Assuming the stall speed is computed at Sea level rho = 0.0023769 lb s^2/ft^4 For the KR2, CL = 1.67 based on gross = 900 lbs, stall = 52 mph, wing area 78 ft^2 (from the manual, the RR site says 80 ft^2) For the KR2S CL = 1.73 based on gross = 980 lbs, stall = 52 mph, wing area 82 ft^2 Since they both use the same airfoil, it makes sense that CL should be 1.7 and we only have 2 significant digits anyway. This seems very high compared to the new airfoil. It looks like the new airfoil is going to have a stall speed about 10 miles per hour faster than the standard KR2 airfoil. Hopefully the data below is readable I copied and pasted from excel. All data were computed using a wing area of 78 sq ft. The first row is gross weight of the airplane * load factor. The first set of columns are predicted best glide for the new airfoil. The second set are predicted stall speed for the new airfoil. The last column is stall speed for a standard KR2 with the original airfoil. The airspeeds are true airspeeds not indicated speeds and are based on the standard atmosphere. altitude rho lb s^2/ft^4 900 1200 1800 2400 900 1200 1800 2400 900 0 0.0023769 79.2 91.4 112.0 129.3 61.3 70.8 86.7 100.1 51.5 1000 0.0023081 80.3 92.8 113.6 131.2 62.2 71.9 88.0 101.6 52.3 2000 0.0022409 81.5 94.2 115.3 133.2 63.2 72.9 89.3 103.1 53.1 4000 0.0021110 84.0 97.0 118.8 137.2 65.1 75.1 92.0 106.3 54.7 8000 0.0018685 89.3 103.1 126.3 145.8 69.2 79.9 97.8 113.0 58.1
