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N-9 (n9-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: N-9 (n9-il)
Reynolds number: 50,000
Max Cl/Cd: 36.76 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n9-il-50000.txt
Download as CSV file: xf-n9-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-9                                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4715   0.11058   0.10359  -0.0053   1.0000   0.2159
  -9.750  -0.4607   0.10620   0.09922  -0.0050   1.0000   0.2233
  -9.500  -0.4762   0.10516   0.09834  -0.0074   1.0000   0.2320
  -9.250  -0.4565   0.10011   0.09325  -0.0057   1.0000   0.2437
  -9.000  -0.4515   0.09658   0.08978  -0.0056   1.0000   0.2539
  -8.500  -0.4481   0.09056   0.08391  -0.0048   1.0000   0.2827
  -8.250  -0.4438   0.08758   0.08099  -0.0036   1.0000   0.3011
  -8.000  -0.4467   0.08511   0.07862  -0.0033   1.0000   0.3201
  -7.750  -0.4468   0.08214   0.07575  -0.0028   1.0000   0.3393
  -7.500  -0.4324   0.07860   0.07223   0.0000   1.0000   0.3615
  -7.250  -0.4329   0.07589   0.06961   0.0012   1.0000   0.3854
  -7.000  -0.4322   0.07344   0.06723   0.0033   1.0000   0.4138
  -6.750  -0.4129   0.06998   0.06377   0.0082   1.0000   0.4508
  -6.500  -0.4047   0.06725   0.06110   0.0126   1.0000   0.4929
  -6.000  -0.3607   0.04497   0.03722  -0.0403   1.0000   0.1758
  -5.750  -0.3348   0.03995   0.03135  -0.0423   1.0000   0.1585
  -5.500  -0.3128   0.03661   0.02771  -0.0420   1.0000   0.1556
  -5.250  -0.2888   0.03379   0.02429  -0.0420   1.0000   0.1580
  -5.000  -0.2641   0.03121   0.02123  -0.0416   1.0000   0.1600
  -4.750  -0.2402   0.02909   0.01893  -0.0410   1.0000   0.1636
  -4.500  -0.2149   0.02753   0.01690  -0.0404   1.0000   0.1735
  -4.250  -0.1907   0.02596   0.01527  -0.0397   1.0000   0.1833
  -4.000  -0.1651   0.02455   0.01369  -0.0391   1.0000   0.1960
  -3.750  -0.1400   0.02340   0.01245  -0.0384   1.0000   0.2197
  -3.500  -0.1139   0.02213   0.01126  -0.0378   1.0000   0.2527
  -3.250  -0.0868   0.02082   0.01024  -0.0376   1.0000   0.3201
  -3.000  -0.0590   0.01701   0.00923  -0.0346   1.0000   1.0000
  -2.750  -0.0362   0.01723   0.00876  -0.0341   1.0000   1.0000
  -2.500  -0.0153   0.01749   0.00863  -0.0334   1.0000   1.0000
  -2.250   0.0053   0.01780   0.00863  -0.0329   1.0000   1.0000
  -2.000   0.0256   0.01817   0.00872  -0.0324   1.0000   1.0000
  -1.750   0.0459   0.01857   0.00890  -0.0321   1.0000   1.0000
  -1.500   0.0661   0.01902   0.00916  -0.0318   1.0000   1.0000
  -1.250   0.0863   0.01951   0.00948  -0.0315   1.0000   1.0000
  -1.000   0.1063   0.02005   0.00986  -0.0314   1.0000   1.0000
  -0.750   0.1263   0.02063   0.01032  -0.0313   1.0000   1.0000
  -0.500   0.1461   0.02126   0.01085  -0.0313   1.0000   1.0000
  -0.250   0.1657   0.02194   0.01146  -0.0313   1.0000   1.0000
   0.000   0.1850   0.02269   0.01214  -0.0315   1.0000   1.0000
   0.250   0.2443   0.02373   0.01312  -0.0390   0.9832   1.0000
   0.500   0.3005   0.02454   0.01394  -0.0456   0.9625   1.0000
   0.750   0.3570   0.02524   0.01466  -0.0519   0.9415   1.0000
   1.000   0.4135   0.02575   0.01526  -0.0577   0.9192   1.0000
   1.250   0.4645   0.02617   0.01578  -0.0622   0.8953   1.0000
   1.500   0.5253   0.02625   0.01602  -0.0677   0.8726   1.0000
   1.750   0.5691   0.02637   0.01630  -0.0699   0.8466   1.0000
   2.000   0.6087   0.02642   0.01649  -0.0710   0.8200   1.0000
   2.250   0.6472   0.02633   0.01654  -0.0715   0.7942   1.0000
   2.500   0.6857   0.02591   0.01627  -0.0712   0.7683   1.0000
   2.750   0.7215   0.02525   0.01569  -0.0698   0.7413   1.0000
   3.000   0.7521   0.02480   0.01527  -0.0677   0.7127   1.0000
   3.250   0.7785   0.02466   0.01518  -0.0654   0.6830   1.0000
   3.500   0.8032   0.02466   0.01519  -0.0630   0.6530   1.0000
   3.750   0.8259   0.02468   0.01523  -0.0605   0.6209   1.0000
   4.000   0.8488   0.02478   0.01541  -0.0583   0.5903   1.0000
   4.250   0.8712   0.02490   0.01559  -0.0561   0.5582   1.0000
   4.500   0.8935   0.02478   0.01543  -0.0536   0.5207   1.0000
   4.750   0.9130   0.02484   0.01537  -0.0506   0.4726   1.0000
   5.000   0.9307   0.02557   0.01575  -0.0475   0.4078   1.0000
   5.250   0.9489   0.02718   0.01684  -0.0450   0.3457   1.0000
   5.500   0.9699   0.02891   0.01840  -0.0436   0.3070   1.0000
   5.750   0.9930   0.03052   0.01993  -0.0426   0.2815   1.0000
   6.000   1.0147   0.03202   0.02153  -0.0415   0.2599   1.0000
   6.250   1.0380   0.03370   0.02332  -0.0407   0.2434   1.0000
   6.500   1.0605   0.03527   0.02489  -0.0397   0.2270   1.0000
   6.750   1.0824   0.03717   0.02691  -0.0388   0.2125   1.0000
   7.000   1.1012   0.03914   0.02908  -0.0375   0.1971   1.0000
   7.250   1.1184   0.04145   0.03162  -0.0360   0.1822   1.0000
   7.500   1.1325   0.04389   0.03435  -0.0343   0.1665   1.0000
   7.750   1.1451   0.04671   0.03740  -0.0326   0.1521   1.0000
   8.000   1.1565   0.05005   0.04093  -0.0309   0.1401   1.0000
   8.250   1.1730   0.05392   0.04475  -0.0299   0.1295   1.0000
   8.500   1.1635   0.05838   0.05003  -0.0268   0.1263   1.0000
   8.750   1.1539   0.06302   0.05520  -0.0244   0.1232   1.0000
   9.000   1.1439   0.06762   0.06014  -0.0225   0.1207   1.0000
   9.250   1.1216   0.07295   0.06582  -0.0207   0.1212   1.0000
   9.500   1.0879   0.07870   0.07182  -0.0194   0.1242   1.0000
   9.750   1.0591   0.08490   0.07814  -0.0200   0.1266   1.0000
  10.000   1.0357   0.09177   0.08507  -0.0222   0.1283   1.0000
  10.250   1.0177   0.09911   0.09244  -0.0249   0.1294   1.0000
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