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

N-10 (n10-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: N-10 (n10-il)
Reynolds number: 500,000
Max Cl/Cd: 95.34 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n10-il-500000.txt
Download as CSV file: xf-n10-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-10                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.3058   0.08501   0.08288  -0.0394   0.9999   0.0338
 -11.000  -0.3072   0.08227   0.08017  -0.0388   0.9999   0.0345
 -10.750  -0.3128   0.07918   0.07711  -0.0383   0.9999   0.0352
 -10.500  -0.5830   0.03745   0.03426  -0.0716   0.9931   0.0265
 -10.250  -0.5592   0.03464   0.03134  -0.0740   0.9893   0.0273
 -10.000  -0.4453   0.07693   0.07487  -0.0365   0.9999   0.0342
  -9.750  -0.4721   0.04351   0.04087  -0.0732   0.9874   0.0303
  -9.500  -0.4705   0.02890   0.02503  -0.0803   0.9803   0.0305
  -9.250  -0.4442   0.02538   0.02096  -0.0819   0.9757   0.0315
  -9.000  -0.4149   0.02215   0.01723  -0.0838   0.9731   0.0328
  -8.750  -0.3805   0.02080   0.01580  -0.0859   0.9715   0.0339
  -8.500  -0.3509   0.01980   0.01470  -0.0865   0.9660   0.0348
  -8.250  -0.3186   0.01864   0.01338  -0.0876   0.9617   0.0357
  -8.000  -0.2844   0.01765   0.01222  -0.0890   0.9581   0.0371
  -7.750  -0.2563   0.01673   0.01114  -0.0890   0.9497   0.0379
  -7.500  -0.2243   0.01588   0.01013  -0.0897   0.9439   0.0386
  -7.250  -0.1962   0.01531   0.00942  -0.0895   0.9343   0.0391
  -7.000  -0.1685   0.01359   0.00759  -0.0896   0.9261   0.0406
  -6.750  -0.1407   0.01294   0.00689  -0.0894   0.9155   0.0416
  -6.500  -0.1132   0.01242   0.00630  -0.0892   0.9042   0.0425
  -6.250  -0.0856   0.01195   0.00576  -0.0889   0.8937   0.0435
  -5.750  -0.0311   0.01120   0.00485  -0.0882   0.8718   0.0460
  -5.500  -0.0039   0.01095   0.00452  -0.0878   0.8605   0.0471
  -5.250   0.0226   0.01048   0.00396  -0.0873   0.8494   0.0486
  -5.000   0.0493   0.01012   0.00354  -0.0869   0.8381   0.0508
  -4.750   0.0765   0.00990   0.00329  -0.0866   0.8272   0.0533
  -4.500   0.1040   0.00973   0.00305  -0.0863   0.8167   0.0563
  -4.250   0.1311   0.00949   0.00277  -0.0860   0.8057   0.0623
  -4.000   0.1584   0.00927   0.00257  -0.0857   0.7939   0.0749
  -3.750   0.1855   0.00898   0.00237  -0.0855   0.7826   0.1080
  -3.500   0.2122   0.00864   0.00227  -0.0853   0.7716   0.1834
  -3.250   0.2396   0.00850   0.00220  -0.0852   0.7596   0.2242
  -3.000   0.2670   0.00836   0.00212  -0.0850   0.7466   0.2555
  -2.500   0.3207   0.00789   0.00200  -0.0847   0.7202   0.3901
  -2.250   0.3464   0.00751   0.00199  -0.0844   0.7064   0.5224
  -2.000   0.3673   0.00681   0.00204  -0.0829   0.6929   0.7573
  -1.750   0.3894   0.00651   0.00213  -0.0807   0.6795   0.9402
  -1.500   0.4370   0.00659   0.00212  -0.0847   0.6644   0.9888
  -1.250   0.4785   0.00671   0.00210  -0.0877   0.6495   1.0001
  -1.000   0.5034   0.00681   0.00210  -0.0871   0.6356   1.0001
  -0.750   0.5282   0.00693   0.00211  -0.0864   0.6201   1.0001
  -0.500   0.5529   0.00707   0.00213  -0.0857   0.6027   1.0001
  -0.250   0.5777   0.00720   0.00217  -0.0851   0.5817   1.0001
   0.000   0.6022   0.00736   0.00220  -0.0844   0.5563   1.0001
   0.250   0.6264   0.00755   0.00224  -0.0837   0.5199   1.0001
   0.500   0.6489   0.00790   0.00232  -0.0827   0.4691   1.0001
   0.750   0.6718   0.00829   0.00247  -0.0819   0.4317   1.0001
   1.000   0.6961   0.00859   0.00263  -0.0813   0.4092   1.0001
   1.250   0.7212   0.00885   0.00279  -0.0808   0.3951   1.0001
   1.500   0.7467   0.00908   0.00294  -0.0805   0.3846   1.0001
   1.750   0.7723   0.00930   0.00310  -0.0802   0.3749   1.0001
   2.000   0.7984   0.00950   0.00326  -0.0799   0.3671   1.0001
   2.250   0.8243   0.00971   0.00342  -0.0797   0.3600   1.0001
   2.500   0.8502   0.00991   0.00360  -0.0795   0.3542   1.0001
   2.750   0.8766   0.01008   0.00377  -0.0793   0.3485   1.0001
   3.000   0.9024   0.01030   0.00395  -0.0791   0.3433   1.0001
   3.250   0.9284   0.01051   0.00416  -0.0789   0.3389   1.0001
   3.500   0.9550   0.01066   0.00434  -0.0788   0.3342   1.0001
   3.750   0.9807   0.01087   0.00454  -0.0786   0.3288   1.0001
   4.000   1.0063   0.01111   0.00477  -0.0784   0.3239   1.0001
   4.250   1.0328   0.01124   0.00495  -0.0783   0.3189   1.0001
   4.500   1.0584   0.01145   0.00516  -0.0781   0.3132   1.0001
   4.750   1.0839   0.01166   0.00538  -0.0779   0.3067   1.0001
   5.000   1.1096   0.01183   0.00557  -0.0777   0.2980   1.0001
   5.250   1.1351   0.01202   0.00578  -0.0776   0.2894   1.0001
   5.500   1.1598   0.01225   0.00599  -0.0772   0.2795   1.0001
   5.750   1.1851   0.01243   0.00620  -0.0771   0.2661   1.0001
   6.000   1.2092   0.01270   0.00642  -0.0767   0.2438   1.0001
   6.250   1.2288   0.01332   0.00678  -0.0757   0.1957   1.0001
   6.500   1.2458   0.01417   0.00740  -0.0743   0.1626   1.0001
   6.750   1.2654   0.01479   0.00795  -0.0733   0.1486   1.0001
   7.000   1.2848   0.01540   0.00851  -0.0722   0.1376   1.0001
   7.250   1.3057   0.01587   0.00898  -0.0714   0.1285   1.0001
   7.500   1.3260   0.01636   0.00948  -0.0705   0.1207   1.0001
   7.750   1.3447   0.01695   0.01004  -0.0694   0.1101   1.0001
   8.000   1.3639   0.01749   0.01054  -0.0684   0.0952   1.0001
   8.250   1.3685   0.01891   0.01158  -0.0652   0.0467   1.0001
   8.500   1.3705   0.02034   0.01286  -0.0614   0.0261   1.0001
   8.750   1.3811   0.02128   0.01382  -0.0591   0.0223   1.0001
   9.000   1.3909   0.02228   0.01489  -0.0568   0.0204   1.0001
   9.250   1.4008   0.02331   0.01600  -0.0547   0.0193   1.0001
   9.500   1.4110   0.02433   0.01711  -0.0528   0.0184   1.0001
   9.750   1.4193   0.02552   0.01839  -0.0508   0.0177   1.0001
  10.000   1.4257   0.02689   0.01984  -0.0488   0.0170   1.0001
  10.250   1.4280   0.02864   0.02169  -0.0466   0.0164   1.0001
  10.500   1.4246   0.03096   0.02412  -0.0443   0.0158   1.0001
  10.750   1.4300   0.03268   0.02595  -0.0430   0.0155   1.0001
  11.000   1.4340   0.03461   0.02798  -0.0419   0.0151   1.0001
  11.250   1.4358   0.03681   0.03029  -0.0408   0.0148   1.0001
  11.500   1.4360   0.03929   0.03288  -0.0399   0.0144   1.0001
  11.750   1.4347   0.04201   0.03571  -0.0392   0.0142   1.0001
  12.000   1.4322   0.04501   0.03881  -0.0387   0.0139   1.0001
  12.250   1.4285   0.04827   0.04217  -0.0385   0.0137   1.0001
  12.500   1.4234   0.05179   0.04580  -0.0384   0.0135   1.0001
  12.750   1.4172   0.05559   0.04969  -0.0386   0.0133   1.0001
  13.000   1.4096   0.05963   0.05384  -0.0389   0.0131   1.0001
  13.250   1.4011   0.06388   0.05819  -0.0394   0.0130   1.0001
  13.500   1.3920   0.06828   0.06268  -0.0399   0.0128   1.0001
  13.750   1.3826   0.07266   0.06715  -0.0403   0.0126   1.0001
  14.000   1.3747   0.07675   0.07130  -0.0405   0.0125   1.0001
  14.250   1.3724   0.08059   0.07526  -0.0414   0.0124   1.0001
  14.500   1.3698   0.08450   0.07928  -0.0424   0.0123   1.0001
  14.750   1.3668   0.08846   0.08335  -0.0435   0.0122   1.0001
  15.000   1.3637   0.09251   0.08751  -0.0446   0.0120   1.0001
  15.250   1.3605   0.09664   0.09175  -0.0459   0.0119   1.0001
  15.500   1.3573   0.10074   0.09595  -0.0471   0.0117   1.0001
  15.750   1.3540   0.10492   0.10023  -0.0485   0.0115   1.0001
  16.000   1.3507   0.10907   0.10447  -0.0499   0.0113   1.0001
<< Back to N-10 (n10-il)

Polar data table (+)

Polar graphs


<< Back to N-10 (n10-il)