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NASA SC(2)-1006 AIRFOIL (sc21006-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-1006 AIRFOIL (sc21006-il)
Reynolds number: 200,000
Max Cl/Cd: 61.28 at α=-1°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc21006-il-200000-n5.txt
Download as CSV file: xf-sc21006-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.750  -0.5134   0.11331   0.10948  -0.0218   0.9999   0.0159
 -12.500  -0.5122   0.10977   0.10596  -0.0219   0.9999   0.0153
 -12.250  -0.5118   0.10615   0.10238  -0.0225   0.9999   0.0151
 -12.000  -0.5124   0.10245   0.09871  -0.0233   0.9999   0.0152
 -11.750  -0.5140   0.09864   0.09495  -0.0243   0.9999   0.0156
 -10.250  -0.5463   0.07821   0.07485  -0.0250   0.9999   0.0163
 -10.000  -0.5502   0.07342   0.07009  -0.0277   0.9999   0.0159
  -9.750  -0.5470   0.06701   0.06370  -0.0345   0.9999   0.0156
  -9.500  -0.5348   0.05803   0.05460  -0.0455   0.9999   0.0153
  -9.250  -0.5148   0.04892   0.04521  -0.0556   0.9999   0.0149
  -9.000  -0.4825   0.03971   0.03545  -0.0658   0.9999   0.0146
  -8.750  -0.4400   0.03180   0.02676  -0.0748   0.9999   0.0147
  -8.500  -0.3975   0.02663   0.02081  -0.0808   0.9999   0.0152
  -8.250  -0.3597   0.02360   0.01714  -0.0842   0.9999   0.0158
  -8.000  -0.3231   0.02111   0.01416  -0.0873   0.9999   0.0176
  -7.750  -0.2889   0.01937   0.01216  -0.0894   0.9999   0.0186
  -7.500  -0.2543   0.01794   0.01050  -0.0913   0.9999   0.0192
  -7.250  -0.2199   0.01680   0.00920  -0.0930   0.9999   0.0203
  -7.000  -0.1855   0.01587   0.00811  -0.0948   0.9999   0.0219
  -6.750  -0.1511   0.01511   0.00723  -0.0966   0.9999   0.0243
  -6.500  -0.1161   0.01444   0.00646  -0.0986   0.9999   0.0291
  -6.250  -0.0837   0.01403   0.00597  -0.0999   0.9999   0.0366
  -6.000  -0.0489   0.01351   0.00548  -0.1018   0.9999   0.0544
  -5.750  -0.0132   0.01297   0.00514  -0.1041   0.9999   0.1135
  -5.500   0.0244   0.01235   0.00495  -0.1070   0.9999   0.2278
  -5.250   0.0620   0.01181   0.00495  -0.1099   0.9999   0.3828
  -5.000   0.0943   0.01160   0.00501  -0.1112   0.9999   0.4680
  -4.750   0.1238   0.01154   0.00506  -0.1117   0.9999   0.5226
  -4.500   0.1524   0.01153   0.00516  -0.1120   0.9999   0.5659
  -4.250   0.1803   0.01157   0.00528  -0.1121   0.9999   0.6025
  -4.000   0.2076   0.01162   0.00544  -0.1120   0.9999   0.6394
  -3.750   0.2343   0.01172   0.00561  -0.1118   0.9999   0.6718
  -3.500   0.2606   0.01184   0.00580  -0.1115   0.9999   0.6970
  -3.250   0.2891   0.01195   0.00597  -0.1117   0.9989   0.7190
  -3.000   0.3205   0.01203   0.00612  -0.1125   0.9969   0.7379
  -2.750   0.3529   0.01213   0.00628  -0.1136   0.9953   0.7548
  -2.500   0.3842   0.01222   0.00644  -0.1144   0.9933   0.7692
  -2.250   0.4161   0.01228   0.00658  -0.1153   0.9895   0.7829
  -2.000   0.4570   0.01224   0.00664  -0.1180   0.9846   0.7964
  -1.750   0.4970   0.01203   0.00658  -0.1203   0.9751   0.8088
  -1.500   0.5455   0.01149   0.00619  -0.1241   0.9605   0.8201
  -1.000   0.6465   0.01055   0.00376  -0.1290   0.5205   0.8373
  -0.750   0.6593   0.01225   0.00434  -0.1262   0.2720   0.8467
  -0.500   0.6768   0.01373   0.00492  -0.1246   0.0797   0.8569
  -0.250   0.6995   0.01441   0.00543  -0.1235   0.0435   0.8691
   0.000   0.7227   0.01487   0.00593  -0.1223   0.0322   0.8827
   0.250   0.7453   0.01544   0.00663  -0.1210   0.0273   0.8965
   0.500   0.7679   0.01601   0.00733  -0.1196   0.0245   0.9100
   0.750   0.7897   0.01669   0.00809  -0.1181   0.0225   0.9247
   1.000   0.8096   0.01743   0.00891  -0.1163   0.0208   0.9524
   1.250   0.8339   0.01877   0.01031  -0.1155   0.0184   1.0001
   1.500   0.8613   0.01982   0.01147  -0.1153   0.0172   1.0001
   1.750   0.8886   0.02128   0.01312  -0.1149   0.0163   1.0001
   2.000   0.9162   0.02300   0.01505  -0.1145   0.0155   1.0001
   2.250   0.9434   0.02503   0.01736  -0.1140   0.0150   1.0001
   2.500   0.9696   0.02744   0.02013  -0.1131   0.0147   1.0001
   2.750   0.9940   0.03036   0.02351  -0.1119   0.0146   1.0001
   3.000   1.0161   0.03370   0.02733  -0.1102   0.0145   1.0001
   3.250   1.0371   0.03571   0.02961  -0.1092   0.0133   1.0001
   3.500   1.0567   0.03728   0.03139  -0.1083   0.0123   1.0001
   3.750   1.0705   0.04088   0.03536  -0.1064   0.0117   1.0001
   4.000   1.0793   0.04597   0.04096  -0.1037   0.0115   1.0001
   4.250   1.0846   0.05140   0.04686  -0.1007   0.0114   1.0001
   4.500   1.0885   0.05660   0.05246  -0.0979   0.0114   1.0001
   4.750   1.0919   0.06146   0.05764  -0.0954   0.0115   1.0001
   5.000   1.0943   0.06619   0.06266  -0.0930   0.0116   1.0001
   5.250   1.0961   0.07115   0.06788  -0.0909   0.0118   1.0001
   5.500   1.0774   0.08273   0.07989  -0.0890   0.0135   1.0001
   5.750   1.0608   0.08887   0.08618  -0.0882   0.0141   1.0001
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