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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 100,000
Max Cl/Cd: 49.17 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa50-il-100000-n5.txt
Download as CSV file: xf-usa50-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 50 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3805   0.09677   0.09207  -0.0207   1.0000   0.0437
  -9.000  -0.3819   0.09314   0.08847  -0.0216   1.0000   0.0447
  -8.500  -0.4783   0.09791   0.09300  -0.0183   1.0000   0.0401
  -8.250  -0.4780   0.09449   0.08964  -0.0192   1.0000   0.0406
  -8.000  -0.4778   0.09144   0.08665  -0.0198   1.0000   0.0430
  -7.750  -0.4806   0.08815   0.08342  -0.0208   1.0000   0.0429
  -7.500  -0.4855   0.08487   0.08023  -0.0218   1.0000   0.0433
  -7.250  -0.4860   0.08114   0.07654  -0.0239   1.0000   0.0437
  -7.000  -0.4849   0.07715   0.07258  -0.0259   1.0000   0.0431
  -6.500  -0.4783   0.06630   0.06163  -0.0307   1.0000   0.0260
  -6.250  -0.4728   0.06193   0.05721  -0.0317   1.0000   0.0243
  -6.000  -0.4656   0.05703   0.05217  -0.0328   1.0000   0.0225
  -5.500  -0.4393   0.04686   0.04137  -0.0331   1.0000   0.0186
  -5.250  -0.4287   0.04291   0.03719  -0.0322   1.0000   0.0182
  -5.000  -0.4155   0.03915   0.03312  -0.0312   1.0000   0.0179
  -4.750  -0.4004   0.03555   0.02916  -0.0299   1.0000   0.0177
  -4.500  -0.3835   0.03195   0.02509  -0.0284   1.0000   0.0176
  -4.250  -0.3648   0.02881   0.02146  -0.0268   1.0000   0.0176
  -4.000  -0.3444   0.02600   0.01812  -0.0252   1.0000   0.0178
  -3.750  -0.3222   0.02363   0.01518  -0.0236   1.0000   0.0192
  -3.500  -0.3006   0.02157   0.01282  -0.0225   1.0000   0.0213
  -3.250  -0.2774   0.01988   0.01087  -0.0212   1.0000   0.0226
  -2.750  -0.2325   0.01742   0.00803  -0.0185   1.0000   0.0294
  -2.500  -0.2106   0.01663   0.00715  -0.0173   1.0000   0.0366
  -2.250  -0.1885   0.01593   0.00646  -0.0161   1.0000   0.0585
  -2.000  -0.1668   0.01532   0.00588  -0.0151   1.0000   0.0992
  -1.750  -0.1443   0.01477   0.00548  -0.0143   0.9997   0.1452
  -1.500  -0.1102   0.01346   0.00520  -0.0165   0.9951   0.4101
  -1.250  -0.0476   0.01188   0.00505  -0.0226   1.0000   1.0000
  -1.000  -0.0086   0.01200   0.00487  -0.0252   0.9935   1.0000
  -0.750   0.0318   0.01210   0.00470  -0.0282   0.9857   1.0000
  -0.500   0.0710   0.01216   0.00458  -0.0309   0.9768   1.0000
  -0.250   0.1095   0.01220   0.00447  -0.0333   0.9672   1.0000
   0.000   0.1465   0.01220   0.00434  -0.0354   0.9562   1.0000
   0.250   0.1823   0.01217   0.00424  -0.0372   0.9437   1.0000
   0.500   0.2186   0.01212   0.00414  -0.0390   0.9309   1.0000
   0.750   0.2559   0.01205   0.00404  -0.0409   0.9177   1.0000
   1.000   0.2974   0.01194   0.00393  -0.0436   0.9021   1.0000
   1.250   0.3381   0.01180   0.00381  -0.0460   0.8788   1.0000
   1.500   0.3792   0.01169   0.00369  -0.0485   0.8516   1.0000
   1.750   0.4184   0.01163   0.00364  -0.0506   0.8227   1.0000
   2.000   0.4570   0.01163   0.00366  -0.0525   0.7910   1.0000
   2.250   0.4925   0.01169   0.00370  -0.0537   0.7580   1.0000
   2.500   0.5244   0.01184   0.00382  -0.0543   0.7259   1.0000
   2.750   0.5536   0.01204   0.00406  -0.0543   0.6965   1.0000
   3.000   0.5807   0.01227   0.00432  -0.0540   0.6692   1.0000
   3.250   0.6063   0.01253   0.00463  -0.0533   0.6422   1.0000
   3.500   0.6304   0.01282   0.00495  -0.0522   0.6111   1.0000
   3.750   0.6484   0.01332   0.00517  -0.0497   0.5414   1.0000
   4.000   0.6654   0.01385   0.00546  -0.0472   0.4640   1.0000
   4.250   0.6782   0.01478   0.00570  -0.0442   0.3171   1.0000
   4.500   0.6844   0.01719   0.00664  -0.0411   0.0729   1.0000
   4.750   0.7010   0.01854   0.00768  -0.0390   0.0269   1.0000
   5.000   0.7210   0.01944   0.00877  -0.0372   0.0234   1.0000
   5.250   0.7390   0.02056   0.01013  -0.0353   0.0197   1.0000
   5.500   0.7545   0.02194   0.01172  -0.0330   0.0178   1.0000
   5.750   0.7714   0.02322   0.01319  -0.0308   0.0172   1.0000
   6.000   0.7887   0.02472   0.01485  -0.0288   0.0167   1.0000
   6.250   0.8082   0.02652   0.01680  -0.0270   0.0163   1.0000
   6.500   0.8299   0.02856   0.01905  -0.0256   0.0157   1.0000
   6.750   0.8514   0.03066   0.02145  -0.0241   0.0144   1.0000
   7.000   0.8714   0.03298   0.02410  -0.0226   0.0136   1.0000
   7.250   0.8892   0.03579   0.02733  -0.0207   0.0135   1.0000
   7.500   0.9038   0.03880   0.03089  -0.0184   0.0137   1.0000
   7.750   0.9147   0.04212   0.03468  -0.0159   0.0139   1.0000
   8.000   0.9219   0.04557   0.03858  -0.0132   0.0142   1.0000
   8.250   0.9256   0.04916   0.04258  -0.0105   0.0146   1.0000
   8.500   0.9257   0.05278   0.04657  -0.0078   0.0149   1.0000
   8.750   0.9215   0.05655   0.05066  -0.0051   0.0152   1.0000
   9.000   0.9143   0.06014   0.05451  -0.0027   0.0155   1.0000
   9.250   0.9024   0.06352   0.05810  -0.0001   0.0157   1.0000
   9.500   0.8856   0.06709   0.06184   0.0021   0.0158   1.0000
   9.750   0.8696   0.07084   0.06574   0.0026   0.0160   1.0000
  10.000   0.8515   0.07557   0.07059   0.0011   0.0160   1.0000
  10.250   0.8357   0.08097   0.07610  -0.0023   0.0161   1.0000
  10.500   0.8202   0.08780   0.08300  -0.0074   0.0162   1.0000
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