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RG 14A-1.4/7.0 AIRFOIL (rg14a147-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: RG 14A-1.4/7.0 AIRFOIL (rg14a147-il)
Reynolds number: 200,000
Max Cl/Cd: 56.07 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rg14a147-il-200000-n5.txt
Download as CSV file: xf-rg14a147-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 14A-1.4/7.0 AIRFOIL                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5175   0.09638   0.09282  -0.0139   1.0000   0.0221
  -8.750  -0.5163   0.09225   0.08873  -0.0161   1.0000   0.0222
  -8.500  -0.5155   0.08824   0.08476  -0.0183   1.0000   0.0222
  -8.250  -0.4454   0.07048   0.06721  -0.0200   1.0000   0.0121
  -8.000  -0.4516   0.06575   0.06251  -0.0218   1.0000   0.0118
  -7.750  -0.5275   0.07293   0.06955  -0.0241   1.0000   0.0123
  -7.500  -0.5301   0.06794   0.06454  -0.0271   1.0000   0.0119
  -7.250  -0.5298   0.06303   0.05958  -0.0295   1.0000   0.0114
  -7.000  -0.5271   0.05806   0.05453  -0.0313   1.0000   0.0112
  -6.750  -0.5223   0.05327   0.04960  -0.0323   1.0000   0.0108
  -6.500  -0.5158   0.04836   0.04450  -0.0326   1.0000   0.0104
  -6.250  -0.5070   0.04361   0.03951  -0.0324   1.0000   0.0101
  -6.000  -0.4963   0.03881   0.03440  -0.0317   1.0000   0.0097
  -5.750  -0.4836   0.03421   0.02942  -0.0305   1.0000   0.0094
  -5.500  -0.4687   0.03001   0.02476  -0.0290   1.0000   0.0091
  -5.250  -0.4516   0.02631   0.02057  -0.0274   1.0000   0.0090
  -5.000  -0.4322   0.02334   0.01710  -0.0258   1.0000   0.0091
  -4.750  -0.4113   0.02101   0.01430  -0.0244   1.0000   0.0093
  -4.500  -0.3893   0.01911   0.01204  -0.0231   1.0000   0.0098
  -4.250  -0.3669   0.01770   0.01037  -0.0219   1.0000   0.0108
  -4.000  -0.3439   0.01729   0.00980  -0.0209   1.0000   0.0128
  -3.750  -0.3220   0.01528   0.00760  -0.0196   1.0000   0.0140
  -3.500  -0.2999   0.01415   0.00637  -0.0185   1.0000   0.0151
  -3.250  -0.2770   0.01337   0.00551  -0.0176   0.9999   0.0169
  -3.000  -0.2411   0.01271   0.00472  -0.0193   0.9956   0.0214
  -2.750  -0.2063   0.01206   0.00402  -0.0209   0.9908   0.0337
  -2.500  -0.1720   0.01126   0.00348  -0.0226   0.9855   0.0995
  -2.250  -0.1385   0.01044   0.00321  -0.0245   0.9800   0.2345
  -2.000  -0.1086   0.00941   0.00300  -0.0257   0.9730   0.4461
  -1.750  -0.0809   0.00853   0.00293  -0.0256   0.9658   0.6540
  -1.500  -0.0551   0.00816   0.00303  -0.0243   0.9579   0.8164
  -1.250  -0.0253   0.00807   0.00295  -0.0239   0.9492   0.8793
  -1.000   0.0128   0.00803   0.00288  -0.0255   0.9427   0.9199
  -0.750   0.0538   0.00800   0.00279  -0.0279   0.9340   0.9480
  -0.500   0.1002   0.00796   0.00266  -0.0316   0.9264   0.9681
  -0.250   0.1473   0.00791   0.00254  -0.0355   0.9166   0.9822
   0.000   0.1960   0.00784   0.00241  -0.0398   0.9034   0.9913
   0.250   0.2428   0.00777   0.00228  -0.0438   0.8857   0.9997
   0.500   0.2753   0.00771   0.00215  -0.0446   0.8613   1.0000
   0.750   0.3030   0.00771   0.00206  -0.0444   0.8342   1.0000
   1.000   0.3279   0.00775   0.00200  -0.0436   0.8047   1.0000
   1.250   0.3516   0.00783   0.00197  -0.0426   0.7737   1.0000
   1.500   0.3748   0.00796   0.00199  -0.0414   0.7449   1.0000
   1.750   0.3975   0.00811   0.00204  -0.0403   0.7158   1.0000
   2.000   0.4199   0.00826   0.00210  -0.0390   0.6812   1.0000
   2.500   0.4643   0.00868   0.00225  -0.0365   0.6083   1.0000
   2.750   0.4867   0.00894   0.00241  -0.0353   0.5687   1.0000
   3.000   0.5091   0.00923   0.00256  -0.0342   0.5271   1.0000
   3.250   0.5315   0.00956   0.00273  -0.0331   0.4831   1.0000
   3.500   0.5542   0.00991   0.00294  -0.0321   0.4376   1.0000
   3.750   0.5770   0.01029   0.00318  -0.0312   0.3916   1.0000
   4.000   0.5998   0.01072   0.00350  -0.0304   0.3439   1.0000
   4.250   0.6218   0.01127   0.00382  -0.0295   0.2821   1.0000
   4.500   0.6437   0.01190   0.00420  -0.0287   0.2180   1.0000
   4.750   0.6664   0.01250   0.00463  -0.0279   0.1694   1.0000
   5.000   0.6890   0.01317   0.00511  -0.0272   0.1224   1.0000
   5.250   0.7113   0.01389   0.00568  -0.0265   0.0823   1.0000
   5.500   0.7336   0.01467   0.00640  -0.0257   0.0533   1.0000
   5.750   0.7558   0.01549   0.00715  -0.0249   0.0333   1.0000
   6.000   0.7777   0.01639   0.00808  -0.0240   0.0214   1.0000
   6.250   0.7992   0.01742   0.00928  -0.0229   0.0162   1.0000
   6.500   0.8197   0.01855   0.01047  -0.0218   0.0120   1.0000
   6.750   0.8412   0.01956   0.01165  -0.0208   0.0102   1.0000
   7.000   0.8614   0.02086   0.01311  -0.0196   0.0092   1.0000
   7.250   0.8813   0.02228   0.01470  -0.0184   0.0084   1.0000
   7.500   0.9008   0.02388   0.01648  -0.0172   0.0080   1.0000
   7.750   0.9190   0.02587   0.01869  -0.0159   0.0076   1.0000
   8.000   0.9356   0.02839   0.02159  -0.0145   0.0073   1.0000
   8.250   0.9491   0.03154   0.02514  -0.0128   0.0071   1.0000
   8.500   0.9596   0.03491   0.02896  -0.0110   0.0070   1.0000
   8.750   0.9658   0.03859   0.03309  -0.0090   0.0070   1.0000
   9.000   0.9676   0.04240   0.03734  -0.0069   0.0070   1.0000
   9.250   0.9643   0.04632   0.04164  -0.0047   0.0070   1.0000
   9.500   0.9566   0.05007   0.04573  -0.0026   0.0069   1.0000
   9.750   0.9435   0.05325   0.04913  -0.0002   0.0069   1.0000
  10.000   0.9283   0.05684   0.05293   0.0010   0.0070   1.0000
  10.250   0.9083   0.06158   0.05783  -0.0002   0.0069   1.0000
  10.500   0.8927   0.06693   0.06338  -0.0037   0.0069   1.0000
  10.750   0.8817   0.07284   0.06943  -0.0082   0.0070   1.0000
  11.000   0.8701   0.08087   0.07760  -0.0148   0.0072   1.0000
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