Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

CURTISS CR-1 AIRFOIL (cr1-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: CURTISS CR-1 AIRFOIL (cr1-il)
Reynolds number: 500,000
Max Cl/Cd: 89.77 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-cr1-il-500000-n5.txt
Download as CSV file: xf-cr1-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CURTISS CR-1 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
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.7626   0.05565   0.05266  -0.0820   0.9997   0.0258
 -13.000  -0.8317   0.03294   0.02933  -0.1078   0.9706   0.0257
 -12.750  -0.8167   0.02994   0.02611  -0.1102   0.9614   0.0261
 -12.500  -0.7911   0.02803   0.02408  -0.1127   0.9556   0.0266
 -12.250  -0.7652   0.02657   0.02249  -0.1145   0.9478   0.0271
 -12.000  -0.7359   0.02523   0.02101  -0.1166   0.9408   0.0276
 -11.750  -0.7093   0.02403   0.01966  -0.1178   0.9322   0.0281
 -11.250  -0.6591   0.02202   0.01730  -0.1190   0.9146   0.0294
 -11.000  -0.6360   0.02117   0.01625  -0.1189   0.9062   0.0299
 -10.750  -0.6151   0.02043   0.01534  -0.1181   0.8969   0.0303
 -10.500  -0.5957   0.01956   0.01433  -0.1171   0.8884   0.0307
 -10.250  -0.5765   0.01881   0.01349  -0.1160   0.8791   0.0312
 -10.000  -0.5554   0.01821   0.01278  -0.1150   0.8712   0.0317
  -9.750  -0.5337   0.01770   0.01217  -0.1141   0.8629   0.0323
  -9.500  -0.5113   0.01721   0.01158  -0.1132   0.8556   0.0328
  -9.250  -0.4893   0.01671   0.01097  -0.1123   0.8473   0.0334
  -9.000  -0.4666   0.01622   0.01036  -0.1114   0.8396   0.0339
  -8.750  -0.4439   0.01576   0.00979  -0.1104   0.8310   0.0345
  -8.500  -0.4206   0.01534   0.00924  -0.1096   0.8233   0.0350
  -8.250  -0.3969   0.01494   0.00873  -0.1088   0.8154   0.0354
  -8.000  -0.3739   0.01447   0.00816  -0.1079   0.8083   0.0361
  -7.750  -0.3505   0.01404   0.00767  -0.1070   0.8006   0.0369
  -7.500  -0.3262   0.01371   0.00727  -0.1063   0.7934   0.0378
  -7.250  -0.3015   0.01340   0.00689  -0.1056   0.7868   0.0386
  -7.000  -0.2766   0.01309   0.00650  -0.1050   0.7805   0.0395
  -6.750  -0.2515   0.01282   0.00613  -0.1043   0.7749   0.0404
  -6.500  -0.2261   0.01256   0.00579  -0.1037   0.7690   0.0412
  -6.250  -0.2009   0.01226   0.00544  -0.1031   0.7637   0.0423
  -6.000  -0.1755   0.01200   0.00512  -0.1025   0.7591   0.0436
  -5.750  -0.1496   0.01176   0.00486  -0.1021   0.7543   0.0454
  -5.500  -0.1235   0.01156   0.00460  -0.1016   0.7495   0.0471
  -5.250  -0.0975   0.01137   0.00435  -0.1011   0.7453   0.0490
  -5.000  -0.0714   0.01115   0.00412  -0.1006   0.7413   0.0517
  -4.750  -0.0450   0.01098   0.00391  -0.1002   0.7371   0.0547
  -4.500  -0.0187   0.01080   0.00372  -0.0997   0.7329   0.0581
  -4.250   0.0078   0.01066   0.00356  -0.0993   0.7290   0.0626
  -4.000   0.0344   0.01053   0.00341  -0.0989   0.7254   0.0673
  -3.750   0.0610   0.01041   0.00330  -0.0986   0.7212   0.0727
  -3.500   0.0876   0.01030   0.00318  -0.0981   0.7165   0.0775
  -3.250   0.1140   0.01022   0.00307  -0.0977   0.7114   0.0825
  -3.000   0.1407   0.01015   0.00297  -0.0972   0.7062   0.0863
  -2.750   0.1670   0.01004   0.00287  -0.0968   0.7008   0.0909
  -2.500   0.1935   0.00998   0.00279  -0.0963   0.6959   0.0957
  -2.250   0.2201   0.00993   0.00270  -0.0959   0.6910   0.0993
  -2.000   0.2464   0.00982   0.00262  -0.0955   0.6857   0.1039
  -1.750   0.2727   0.00976   0.00255  -0.0950   0.6803   0.1087
  -1.500   0.2992   0.00973   0.00248  -0.0945   0.6756   0.1134
  -1.250   0.3255   0.00963   0.00243  -0.0941   0.6704   0.1200
  -1.000   0.3518   0.00957   0.00238  -0.0936   0.6648   0.1254
  -0.500   0.4040   0.00946   0.00230  -0.0926   0.6535   0.1382
  -0.250   0.4299   0.00942   0.00227  -0.0920   0.6468   0.1464
   0.000   0.4557   0.00938   0.00225  -0.0914   0.6397   0.1558
   0.250   0.4806   0.00936   0.00223  -0.0907   0.6289   0.1667
   0.500   0.5054   0.00935   0.00221  -0.0899   0.6155   0.1772
   0.750   0.5295   0.00936   0.00220  -0.0890   0.5999   0.1872
   1.000   0.5536   0.00939   0.00220  -0.0880   0.5848   0.1956
   1.250   0.5772   0.00944   0.00221  -0.0870   0.5688   0.2053
   1.500   0.5998   0.00952   0.00224  -0.0859   0.5500   0.2142
   2.000   0.6443   0.00975   0.00236  -0.0834   0.5123   0.2326
   2.250   0.6659   0.00989   0.00244  -0.0820   0.4922   0.2424
   2.500   0.6870   0.01004   0.00254  -0.0806   0.4723   0.2548
   2.750   0.7076   0.01020   0.00266  -0.0791   0.4530   0.2727
   3.000   0.7277   0.01033   0.00280  -0.0776   0.4351   0.2988
   3.250   0.7470   0.01043   0.00295  -0.0759   0.4172   0.3469
   3.750   0.9130   0.01017   0.00390  -0.1009   0.3604   1.0000
   4.000   0.9323   0.01042   0.00407  -0.0991   0.3474   1.0000
   4.250   0.9512   0.01067   0.00425  -0.0973   0.3348   1.0000
   4.500   0.9700   0.01092   0.00443  -0.0955   0.3226   1.0000
   4.750   0.9889   0.01115   0.00462  -0.0937   0.3116   1.0000
   5.000   1.0070   0.01141   0.00482  -0.0918   0.3011   1.0000
   5.250   1.0247   0.01168   0.00503  -0.0898   0.2905   1.0000
   5.500   1.0426   0.01193   0.00525  -0.0878   0.2811   1.0000
   5.750   1.0591   0.01221   0.00548  -0.0856   0.2720   1.0000
   6.000   1.0763   0.01246   0.00571  -0.0836   0.2633   1.0000
   6.250   1.0914   0.01275   0.00596  -0.0811   0.2554   1.0000
   6.500   1.1052   0.01299   0.00618  -0.0783   0.2481   1.0000
   6.750   1.1172   0.01329   0.00644  -0.0753   0.2411   1.0000
   7.000   1.1314   0.01355   0.00670  -0.0727   0.2337   1.0000
   7.250   1.1438   0.01391   0.00701  -0.0699   0.2255   1.0000
   7.500   1.1589   0.01422   0.00731  -0.0676   0.2186   1.0000
   7.750   1.1720   0.01462   0.00768  -0.0650   0.2112   1.0000
   8.000   1.1885   0.01494   0.00801  -0.0631   0.2071   1.0000
   8.250   1.2043   0.01529   0.00837  -0.0611   0.2029   1.0000
   8.500   1.2197   0.01568   0.00877  -0.0591   0.1990   1.0000
   8.750   1.2346   0.01613   0.00921  -0.0571   0.1949   1.0000
   9.000   1.2508   0.01653   0.00964  -0.0554   0.1897   1.0000
   9.250   1.2651   0.01705   0.01014  -0.0534   0.1839   1.0000
   9.500   1.2800   0.01756   0.01065  -0.0517   0.1793   1.0000
   9.750   1.2959   0.01804   0.01116  -0.0501   0.1745   1.0000
  10.000   1.3097   0.01865   0.01175  -0.0483   0.1689   1.0000
  10.250   1.3245   0.01922   0.01234  -0.0467   0.1642   1.0000
  10.500   1.3404   0.01976   0.01292  -0.0452   0.1608   1.0000
  10.750   1.3548   0.02040   0.01358  -0.0437   0.1564   1.0000
  11.000   1.3679   0.02113   0.01431  -0.0420   0.1513   1.0000
  11.250   1.3832   0.02176   0.01497  -0.0407   0.1463   1.0000
  11.500   1.3954   0.02259   0.01579  -0.0391   0.1397   1.0000
  11.750   1.4090   0.02336   0.01659  -0.0377   0.1325   1.0000
  12.000   1.4196   0.02434   0.01754  -0.0361   0.1218   1.0000
  12.250   1.4254   0.02569   0.01878  -0.0342   0.1023   1.0000
  12.500   1.4271   0.02738   0.02035  -0.0321   0.0858   1.0000
  12.750   1.4306   0.02900   0.02194  -0.0302   0.0744   1.0000
  13.000   1.4318   0.03085   0.02374  -0.0283   0.0608   1.0000
  13.250   1.4192   0.03390   0.02660  -0.0258   0.0330   1.0000
  13.500   1.4158   0.03631   0.02901  -0.0240   0.0240   1.0000
  13.750   1.4199   0.03815   0.03089  -0.0228   0.0214   1.0000
  14.000   1.4225   0.04018   0.03298  -0.0216   0.0189   1.0000
  14.250   1.4271   0.04209   0.03496  -0.0207   0.0178   1.0000
  14.500   1.4315   0.04404   0.03698  -0.0198   0.0169   1.0000
  14.750   1.4343   0.04621   0.03922  -0.0190   0.0161   1.0000
  15.000   1.4345   0.04873   0.04182  -0.0183   0.0152   1.0000
  15.250   1.4375   0.05102   0.04419  -0.0178   0.0147   1.0000
  15.500   1.4387   0.05354   0.04681  -0.0174   0.0142   1.0000
  15.750   1.4383   0.05634   0.04969  -0.0170   0.0137   1.0000
  16.000   1.4369   0.05933   0.05278  -0.0169   0.0133   1.0000
  16.250   1.4337   0.06264   0.05618  -0.0169   0.0129   1.0000
  16.500   1.4287   0.06626   0.05989  -0.0171   0.0126   1.0000
  16.750   1.4228   0.07009   0.06383  -0.0174   0.0123   1.0000
  17.000   1.4179   0.07390   0.06775  -0.0180   0.0121   1.0000
  17.250   1.4104   0.07818   0.07216  -0.0187   0.0120   1.0000
  17.500   1.4029   0.08252   0.07661  -0.0196   0.0116   1.0000
  17.750   1.3934   0.08724   0.08145  -0.0207   0.0115   1.0000
  18.000   1.3826   0.09226   0.08659  -0.0220   0.0113   1.0000
  18.250   1.3702   0.09759   0.09205  -0.0235   0.0111   1.0000
  18.500   1.3599   0.10263   0.09719  -0.0251   0.0110   1.0000
  18.750   1.3452   0.10851   0.10319  -0.0270   0.0109   1.0000
  19.000   1.3315   0.11430   0.10910  -0.0290   0.0108   1.0000
<< Back to CURTISS CR-1 AIRFOIL (cr1-il)

Polar data table (+)

Polar graphs


<< Back to CURTISS CR-1 AIRFOIL (cr1-il)