CLARK V AIRFOIL 17. 17. 0.000000 0.000000 0.012500 0.019500 0.025000 0.028910 0.050000 0.042410 0.075000 0.052220 0.100000 0.060130 0.150000 0.071550 0.200000 0.079560 0.300000 0.088590 0.400000 0.091420 0.500000 0.088650 0.600000 0.080980 0.700000 0.068810 0.800000 0.051940 0.900000 0.029970 0.950000 0.016380 1.000000 0.001400 0.000000 0.000000 0.012500 -0.013800 0.025000 -0.018290 0.050000 -0.023090 0.075000 -0.025580 0.100000 -0.027370 0.150000 -0.028950 0.200000 -0.029440 0.300000 -0.027810 0.400000 -0.024180 0.500000 -0.020150 0.600000 -0.016120 0.700000 -0.012090 0.800000 -0.008060 0.900000 -0.004030 0.950000 -0.002010 1.000000 0.000000