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NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 500,000
Max Cl/Cd: 91.95 at α=-2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc21006-il-500000-n5.txt
Download as CSV file: xf-sc21006-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-1006 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.5394   0.09235   0.09006  -0.0229   0.9999   0.0066
 -11.500  -0.5455   0.08920   0.08695  -0.0225   0.9999   0.0067
 -11.250  -0.5540   0.08622   0.08401  -0.0217   0.9999   0.0067
 -11.000  -0.5503   0.08116   0.07896  -0.0260   0.9979   0.0068
 -10.750  -0.5408   0.07503   0.07284  -0.0331   0.9950   0.0070
 -10.500  -0.5284   0.06672   0.06454  -0.0443   0.9918   0.0071
 -10.250  -0.4955   0.03920   0.03649  -0.0840   0.9863   0.0070
 -10.000  -0.4605   0.03004   0.02669  -0.0945   0.9855   0.0075
  -9.750  -0.4251   0.02531   0.02144  -0.1000   0.9851   0.0082
  -9.500  -0.3872   0.02057   0.01602  -0.1051   0.9850   0.0086
  -9.250  -0.3521   0.01799   0.01301  -0.1081   0.9848   0.0090
  -9.000  -0.3185   0.01641   0.01114  -0.1101   0.9846   0.0095
  -8.750  -0.2858   0.01535   0.00984  -0.1118   0.9843   0.0099
  -8.500  -0.2527   0.01415   0.00844  -0.1136   0.9841   0.0102
  -8.250  -0.2188   0.01282   0.00694  -0.1157   0.9840   0.0110
  -8.000  -0.1864   0.01230   0.00636  -0.1171   0.9837   0.0121
  -7.750  -0.1533   0.01171   0.00569  -0.1187   0.9835   0.0128
  -7.500  -0.1201   0.01116   0.00502  -0.1203   0.9833   0.0135
  -7.250  -0.0893   0.01073   0.00452  -0.1212   0.9825   0.0148
  -7.000  -0.0595   0.01042   0.00415  -0.1219   0.9814   0.0162
  -6.750  -0.0281   0.01004   0.00371  -0.1229   0.9807   0.0203
  -6.500   0.0021   0.00983   0.00348  -0.1237   0.9798   0.0250
  -6.250   0.0329   0.00960   0.00326  -0.1246   0.9790   0.0365
  -6.000   0.0648   0.00928   0.00309  -0.1259   0.9785   0.0738
  -5.750   0.0965   0.00901   0.00297  -0.1271   0.9780   0.1178
  -5.500   0.1284   0.00875   0.00287  -0.1283   0.9774   0.1680
  -5.250   0.1617   0.00838   0.00279  -0.1300   0.9771   0.2613
  -5.000   0.1950   0.00804   0.00278  -0.1317   0.9769   0.3660
  -4.750   0.2267   0.00789   0.00278  -0.1328   0.9763   0.4159
  -4.500   0.2579   0.00780   0.00280  -0.1337   0.9756   0.4514
  -4.250   0.2892   0.00772   0.00283  -0.1347   0.9750   0.4944
  -4.000   0.3204   0.00766   0.00288  -0.1355   0.9744   0.5298
  -3.750   0.3520   0.00760   0.00293  -0.1364   0.9737   0.5622
  -3.500   0.3861   0.00749   0.00294  -0.1379   0.9727   0.5942
  -3.250   0.4263   0.00726   0.00281  -0.1405   0.9710   0.6198
  -3.000   0.4593   0.00705   0.00270  -0.1414   0.9650   0.6436
  -2.750   0.4986   0.00669   0.00244  -0.1436   0.9583   0.6668
  -2.500   0.5309   0.00633   0.00215  -0.1439   0.9388   0.6858
  -2.250   0.5609   0.00610   0.00188  -0.1435   0.8965   0.7014
  -2.000   0.5797   0.00685   0.00177  -0.1406   0.6901   0.7134
  -1.500   0.6166   0.00941   0.00258  -0.1372   0.2728   0.7334
  -1.250   0.6396   0.01031   0.00295  -0.1365   0.1298   0.7446
  -1.000   0.6644   0.01093   0.00325  -0.1360   0.0525   0.7557
  -0.750   0.6906   0.01128   0.00353  -0.1356   0.0306   0.7649
  -0.500   0.7172   0.01155   0.00382  -0.1352   0.0240   0.7745
  -0.250   0.7434   0.01195   0.00425  -0.1347   0.0187   0.7840
   0.000   0.7698   0.01224   0.00464  -0.1343   0.0167   0.7922
   0.250   0.7960   0.01256   0.00499  -0.1339   0.0142   0.8001
   0.500   0.8217   0.01301   0.00549  -0.1333   0.0126   0.8088
   0.750   0.8455   0.01384   0.00643  -0.1323   0.0114   0.8173
   1.000   0.8707   0.01432   0.00700  -0.1316   0.0109   0.8270
   1.250   0.8955   0.01495   0.00775  -0.1308   0.0104   0.8366
   1.500   0.9197   0.01570   0.00862  -0.1298   0.0098   0.8442
   1.750   0.9440   0.01663   0.00967  -0.1289   0.0094   0.8513
   2.000   0.9678   0.01772   0.01091  -0.1278   0.0091   0.8577
   2.500   1.0163   0.02012   0.01365  -0.1259   0.0084   0.8712
   2.750   1.0408   0.02081   0.01445  -0.1252   0.0079   0.8788
   3.000   1.0643   0.02167   0.01548  -0.1243   0.0075   0.8869
   3.250   1.0867   0.02316   0.01720  -0.1231   0.0072   0.8963
   3.500   1.1052   0.02626   0.02075  -0.1211   0.0069   0.9075
   3.750   1.1216   0.02930   0.02429  -0.1186   0.0066   0.9232
   4.000   1.1350   0.03246   0.02794  -0.1153   0.0064   1.0001
   4.250   1.1415   0.03992   0.03612  -0.1109   0.0061   1.0001
   4.500   1.1436   0.04813   0.04491  -0.1063   0.0058   1.0001
   4.750   1.1462   0.05501   0.05218  -0.1028   0.0054   1.0001
   5.000   1.1479   0.06094   0.05840  -0.1000   0.0050   1.0001
   5.250   1.1452   0.06707   0.06478  -0.0974   0.0049   1.0001
   5.500   1.1383   0.07326   0.07119  -0.0952   0.0049   1.0001
   5.750   1.1278   0.07930   0.07741  -0.0935   0.0048   1.0001
   6.000   1.1136   0.08505   0.08329  -0.0923   0.0048   1.0001
   6.250   1.0900   0.09057   0.08890  -0.0910   0.0049   1.0001
   6.500   1.0688   0.09707   0.09548  -0.0931   0.0049   1.0001
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