COORDINATES OF NACA AIRFOILS
Created by naca456, version 1.0 (4 January 2002) on 17:39  4Jan2002
Ralph L. Carmichael, Public Domain Aeronautical Software
NACA 16-015 (same as 0015-45)                                                   
Four-digit-modified profile, t/c= 0.1500
   l.e. index= 4.0000   x of max. thickness= 0.5000
No camber, symmetrical section
SYMMETRICAL AIRFOIL DEFINITION
        x         y        dy/dx
    1  0.000000  0.000000**********
    2  0.005000  0.010321  1.014356
    3  0.007500  0.012591  0.821678
    4  0.012500  0.016155  0.628377
    5  0.025000  0.022584  0.433595
    6  0.050000  0.031400  0.295207
    7  0.075000  0.037927  0.233098
    8  0.100000  0.043250  0.195285
    9  0.150000  0.051736  0.148295
   10  0.200000  0.058339  0.117480
   11  0.250000  0.063597  0.093597
   12  0.300000  0.067754  0.073082
   13  0.350000  0.070932  0.054227
   14  0.400000  0.073189  0.036100
   15  0.450000  0.074545  0.018144
   16  0.500000  0.075000  0.000000
   17  0.550000  0.074514  0.019986
   18  0.600000  0.072946  0.043295
   19  0.650000  0.070129  0.069926
   20  0.700000  0.065898  0.099879
   21  0.750000  0.060086  0.133156
   22  0.800000  0.052527  0.169755
   23  0.850000  0.043055  0.209676
   24  0.900000  0.031504  0.252921
   25  0.950000  0.017708  0.299487
   26  1.000000  0.001500  0.349377
  1.500000  2.000000  1.500000  2.000000
  1.512500  2.025801  1.512500  1.974199
  1.518750  2.031478  1.518750  1.968522
  1.531250  2.040388  1.531250  1.959612
  1.562500  2.056460  1.562500  1.943540
  1.625000  2.078499  1.625000  1.921501
  1.687500  2.094818  1.687500  1.905182
  1.750000  2.108125  1.750000  1.891875
  1.875000  2.129340  1.875000  1.870660
  2.000000  2.145848  2.000000  1.854152
  2.125000  2.158992  2.125000  1.841008
  2.250000  2.169385  2.250000  1.830614
  2.375000  2.177331  2.375000  1.822669
  2.500000  2.182972  2.500000  1.817028
  2.625000  2.186363  2.625000  1.813637
  2.750000  2.187500  2.750000  1.812500
  2.875000  2.186285  2.875000  1.813715
  3.000000  2.182365  3.000000  1.817635
  3.125000  2.175323  3.125000  1.824677
  3.250000  2.164745  3.250000  1.835255
  3.375000  2.150215  3.375000  1.849785
  3.500000  2.131318  3.500000  1.868682
  3.625000  2.107638  3.625000  1.892362
  3.750000  2.078760  3.750000  1.921240
  3.875000  2.044269  3.875000  1.955731
  4.000000  2.003750  4.000000  1.996250
End of output for symmetrical airfoil