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USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 41 AIRFOIL (usa41-il)
Reynolds number: 200,000
Max Cl/Cd: 79.31 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa41-il-200000.txt
Download as CSV file: xf-usa41-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 41 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3394   0.09752   0.09398  -0.0258   1.0000   0.0345
  -8.000  -0.3444   0.09609   0.09263  -0.0274   1.0000   0.0350
  -7.750  -0.3506   0.09461   0.09123  -0.0280   1.0000   0.0353
  -7.500  -0.3498   0.09217   0.08885  -0.0299   1.0000   0.0355
  -7.250  -0.3466   0.08946   0.08619  -0.0317   1.0000   0.0356
  -7.000  -0.3497   0.08478   0.08159  -0.0318   1.0000   0.0360
  -6.750  -0.3498   0.08147   0.07832  -0.0275   1.0000   0.0369
  -6.500  -0.3490   0.07923   0.07611  -0.0256   1.0000   0.0376
  -6.250  -0.3491   0.07715   0.07406  -0.0240   1.0000   0.0386
  -6.000  -0.3491   0.07497   0.07192  -0.0234   1.0000   0.0395
  -5.750  -0.3472   0.07260   0.06957  -0.0236   1.0000   0.0406
  -5.500  -0.3425   0.06990   0.06688  -0.0249   1.0000   0.0425
  -5.250  -0.3301   0.06652   0.06348  -0.0292   1.0000   0.0451
  -5.000  -0.2863   0.05931   0.05614  -0.0423   0.9959   0.0488
  -4.750  -0.2587   0.05683   0.05365  -0.0443   0.9916   0.0531
  -4.500  -0.2003   0.04935   0.04585  -0.0593   0.9863   0.0611
  -4.250  -0.1446   0.04671   0.04275  -0.0683   0.9816   0.0726
  -4.000  -0.1246   0.04208   0.03830  -0.0696   0.9761   0.0755
  -3.750  -0.0799   0.03855   0.03453  -0.0758   0.9719   0.0885
  -2.750   0.0834   0.02036   0.01434  -0.0900   0.9508   0.0662
  -2.500   0.1255   0.01790   0.01106  -0.0919   0.9462   0.0601
  -2.250   0.1690   0.01600   0.00890  -0.0949   0.9431   0.0615
  -2.000   0.2045   0.01494   0.00766  -0.0959   0.9350   0.0610
  -1.750   0.2459   0.01381   0.00645  -0.0982   0.9300   0.0618
  -1.500   0.2799   0.01305   0.00567  -0.0990   0.9208   0.0632
  -1.250   0.3191   0.01238   0.00501  -0.1008   0.9140   0.0667
  -1.000   0.3498   0.01195   0.00455  -0.1008   0.9019   0.0708
  -0.750   0.3802   0.01154   0.00411  -0.1008   0.8893   0.0744
  -0.500   0.4093   0.01118   0.00377  -0.1005   0.8755   0.0856
  -0.250   0.4370   0.01084   0.00374  -0.0999   0.8603   0.1781
   0.000   0.4645   0.01075   0.00363  -0.0993   0.8438   0.2183
   0.250   0.4913   0.01058   0.00344  -0.0987   0.8265   0.2438
   0.500   0.5161   0.01034   0.00330  -0.0977   0.8055   0.2781
   0.750   0.5387   0.00952   0.00329  -0.0966   0.7854   0.5601
   1.000   0.5855   0.00869   0.00310  -0.1001   0.7622   1.0000
   1.250   0.6113   0.00881   0.00301  -0.0992   0.7391   1.0000
   1.500   0.6366   0.00898   0.00300  -0.0983   0.7147   1.0000
   1.750   0.6619   0.00917   0.00301  -0.0975   0.6914   1.0000
   2.000   0.6871   0.00939   0.00306  -0.0967   0.6693   1.0000
   2.250   0.7122   0.00961   0.00314  -0.0959   0.6479   1.0000
   2.500   0.7373   0.00985   0.00326  -0.0951   0.6278   1.0000
   2.750   0.7625   0.01011   0.00339  -0.0944   0.6096   1.0000
   3.000   0.7876   0.01036   0.00355  -0.0937   0.5912   1.0000
   3.250   0.8127   0.01063   0.00374  -0.0930   0.5740   1.0000
   3.500   0.8377   0.01091   0.00396  -0.0923   0.5575   1.0000
   3.750   0.8628   0.01119   0.00419  -0.0917   0.5418   1.0000
   4.000   0.8878   0.01149   0.00444  -0.0910   0.5267   1.0000
   4.250   0.9129   0.01179   0.00471  -0.0904   0.5123   1.0000
   4.500   0.9378   0.01209   0.00502  -0.0897   0.4981   1.0000
   4.750   0.9626   0.01238   0.00531  -0.0890   0.4836   1.0000
   5.000   0.9873   0.01267   0.00562  -0.0884   0.4696   1.0000
   5.250   1.0115   0.01294   0.00590  -0.0876   0.4539   1.0000
   5.500   1.0339   0.01313   0.00609  -0.0864   0.4299   1.0000
   5.750   1.0560   0.01336   0.00627  -0.0852   0.4046   1.0000
   6.000   1.0786   0.01360   0.00654  -0.0842   0.3794   1.0000
   6.250   1.0999   0.01394   0.00682  -0.0829   0.3480   1.0000
   6.500   1.1202   0.01439   0.00717  -0.0816   0.3029   1.0000
   6.750   1.1290   0.01611   0.00805  -0.0789   0.1530   1.0000
   7.000   1.1335   0.01880   0.00993  -0.0756   0.0505   1.0000
   7.250   1.1505   0.01992   0.01114  -0.0738   0.0421   1.0000
   7.500   1.1629   0.02142   0.01273  -0.0713   0.0381   1.0000
   7.750   1.1781   0.02258   0.01402  -0.0692   0.0357   1.0000
   8.000   1.1925   0.02379   0.01529  -0.0672   0.0328   1.0000
   8.250   1.2038   0.02536   0.01692  -0.0648   0.0306   1.0000
   8.500   1.2144   0.02762   0.01920  -0.0624   0.0293   1.0000
   8.750   1.2322   0.02993   0.02158  -0.0609   0.0285   1.0000
   9.000   1.2528   0.03177   0.02356  -0.0597   0.0280   1.0000
   9.250   1.2733   0.03375   0.02572  -0.0585   0.0274   1.0000
   9.500   1.2907   0.03542   0.02761  -0.0570   0.0260   1.0000
   9.750   1.3085   0.03778   0.03023  -0.0556   0.0253   1.0000
  10.000   1.3244   0.04077   0.03353  -0.0540   0.0254   1.0000
  10.250   1.3358   0.04428   0.03740  -0.0520   0.0258   1.0000
  10.500   1.3419   0.04825   0.04175  -0.0496   0.0266   1.0000
  10.750   1.3431   0.05302   0.04686  -0.0473   0.0276   1.0000
  11.000   1.2900   0.04940   0.04376  -0.0393   0.0300   1.0000
  11.250   1.2403   0.05415   0.04924  -0.0319   0.0342   1.0000
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