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

FX 63-100 AIRFOIL (fx63100-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: FX 63-100 AIRFOIL (fx63100-il)
Reynolds number: 200,000
Max Cl/Cd: 92.86 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx63100-il-200000.txt
Download as CSV file: xf-fx63100-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 63-100 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3713   0.09455   0.09118  -0.0419   1.0000   0.0500
  -8.250  -0.3848   0.09258   0.08931  -0.0405   1.0000   0.0501
  -8.000  -0.4038   0.09108   0.08791  -0.0381   1.0000   0.0502
  -7.750  -0.3883   0.08657   0.08340  -0.0341   1.0000   0.0516
  -7.500  -0.3999   0.08556   0.08247  -0.0302   1.0000   0.0520
  -7.250  -0.4144   0.08466   0.08164  -0.0265   1.0000   0.0525
  -7.000  -0.3979   0.08133   0.07832  -0.0299   0.9967   0.0542
  -6.750  -0.3707   0.07564   0.07261  -0.0400   0.9900   0.0579
  -6.500  -0.3284   0.05855   0.05526  -0.0757   0.9783   0.0629
  -6.250  -0.3070   0.05922   0.05606  -0.0707   0.9753   0.0657
  -6.000  -0.2557   0.04728   0.04330  -0.0961   0.9668   0.0755
  -5.750  -0.2089   0.03340   0.02845  -0.1077   0.9642   0.0516
  -5.500  -0.1651   0.02850   0.02308  -0.1129   0.9620   0.0478
  -5.250  -0.1181   0.02498   0.01886  -0.1180   0.9603   0.0493
  -5.000  -0.0798   0.02211   0.01546  -0.1209   0.9558   0.0507
  -4.750  -0.0431   0.02060   0.01392  -0.1234   0.9515   0.0553
  -4.500  -0.0016   0.01920   0.01224  -0.1261   0.9487   0.0592
  -4.250   0.0406   0.01775   0.01062  -0.1291   0.9467   0.0643
  -4.000   0.0733   0.01697   0.00980  -0.1301   0.9400   0.0688
  -3.750   0.1132   0.01633   0.00903  -0.1323   0.9362   0.0748
  -3.500   0.1554   0.01517   0.00796  -0.1353   0.9340   0.0824
  -3.250   0.2000   0.01430   0.00710  -0.1387   0.9322   0.0933
  -3.000   0.2348   0.01360   0.00646  -0.1402   0.9247   0.1084
  -2.750   0.2805   0.01251   0.00558  -0.1440   0.9210   0.1596
  -2.500   0.3281   0.01121   0.00523  -0.1487   0.9180   0.4288
  -2.250   0.3603   0.01099   0.00521  -0.1493   0.9091   0.4983
  -2.000   0.3986   0.01085   0.00511  -0.1509   0.9039   0.5538
  -1.750   0.4287   0.01087   0.00517  -0.1509   0.8949   0.5928
  -1.500   0.4632   0.01081   0.00513  -0.1517   0.8884   0.6264
  -1.250   0.4913   0.01085   0.00516  -0.1513   0.8784   0.6504
  -1.000   0.5227   0.01083   0.00513  -0.1515   0.8710   0.6732
  -0.750   0.5499   0.01088   0.00517  -0.1510   0.8607   0.6948
  -0.500   0.5770   0.01092   0.00523  -0.1504   0.8516   0.7143
  -0.250   0.6049   0.01094   0.00524  -0.1498   0.8427   0.7338
   0.000   0.6307   0.01099   0.00531  -0.1491   0.8325   0.7514
   0.250   0.6587   0.01100   0.00529  -0.1487   0.8240   0.7665
   0.500   0.6846   0.01103   0.00532  -0.1480   0.8136   0.7806
   0.750   0.7108   0.01105   0.00536  -0.1473   0.8037   0.7946
   1.000   0.7381   0.01105   0.00534  -0.1468   0.7948   0.8091
   1.250   0.7627   0.01106   0.00540  -0.1458   0.7837   0.8240
   1.500   0.7879   0.01107   0.00543  -0.1449   0.7733   0.8394
   1.750   0.8130   0.01104   0.00539  -0.1439   0.7625   0.8560
   2.000   0.8364   0.01097   0.00533  -0.1425   0.7488   0.8756
   2.250   0.8567   0.01086   0.00525  -0.1404   0.7341   0.9003
   2.500   0.8761   0.01068   0.00512  -0.1382   0.7204   0.9541
   2.750   0.9089   0.01076   0.00519  -0.1393   0.7066   1.0000
   3.000   0.9399   0.01089   0.00528  -0.1400   0.6925   1.0000
   3.250   0.9697   0.01104   0.00539  -0.1403   0.6779   1.0000
   3.500   0.9982   0.01117   0.00547  -0.1403   0.6611   1.0000
   3.750   1.0258   0.01132   0.00556  -0.1401   0.6425   1.0000
   4.000   1.0526   0.01149   0.00569  -0.1397   0.6226   1.0000
   4.250   1.0786   0.01168   0.00583  -0.1392   0.5997   1.0000
   4.500   1.1037   0.01189   0.00598  -0.1385   0.5731   1.0000
   4.750   1.1283   0.01215   0.00616  -0.1377   0.5450   1.0000
   5.000   1.1525   0.01247   0.00638  -0.1369   0.5162   1.0000
   5.250   1.1760   0.01285   0.00667  -0.1360   0.4846   1.0000
   5.500   1.1988   0.01331   0.00701  -0.1350   0.4519   1.0000
   5.750   1.2211   0.01383   0.00741  -0.1340   0.4185   1.0000
   6.000   1.2424   0.01442   0.00786  -0.1329   0.3825   1.0000
   6.250   1.2628   0.01508   0.00838  -0.1316   0.3382   1.0000
   6.500   1.2810   0.01594   0.00898  -0.1301   0.2846   1.0000
   6.750   1.2970   0.01705   0.00977  -0.1284   0.2304   1.0000
   7.000   1.3122   0.01830   0.01071  -0.1267   0.1775   1.0000
   7.250   1.3250   0.01982   0.01187  -0.1246   0.1239   1.0000
   7.500   1.3375   0.02135   0.01317  -0.1224   0.0962   1.0000
   7.750   1.3513   0.02270   0.01446  -0.1203   0.0840   1.0000
   8.000   1.3630   0.02419   0.01590  -0.1180   0.0770   1.0000
   8.250   1.3793   0.02525   0.01705  -0.1162   0.0721   1.0000
   8.500   1.3928   0.02652   0.01834  -0.1142   0.0679   1.0000
   8.750   1.4046   0.02809   0.01992  -0.1119   0.0646   1.0000
   9.000   1.4200   0.02915   0.02111  -0.1101   0.0617   1.0000
   9.250   1.4342   0.03030   0.02233  -0.1082   0.0586   1.0000
   9.500   1.4477   0.03180   0.02380  -0.1065   0.0557   1.0000
   9.750   1.4651   0.03361   0.02569  -0.1052   0.0532   1.0000
  10.000   1.4790   0.03480   0.02707  -0.1034   0.0507   1.0000
  10.250   1.4922   0.03613   0.02850  -0.1017   0.0482   1.0000
  10.500   1.5061   0.03772   0.03009  -0.1002   0.0459   1.0000
  10.750   1.5247   0.04019   0.03270  -0.0994   0.0434   1.0000
  11.000   1.5341   0.04175   0.03449  -0.0973   0.0415   1.0000
  11.250   1.5446   0.04359   0.03650  -0.0955   0.0397   1.0000
  11.500   1.5540   0.04543   0.03844  -0.0939   0.0380   1.0000
  11.750   1.5734   0.04935   0.04244  -0.0938   0.0356   1.0000
  12.000   1.5701   0.05126   0.04465  -0.0909   0.0347   1.0000
  12.250   1.5688   0.05381   0.04749  -0.0886   0.0337   1.0000
  12.500   1.5675   0.05686   0.05082  -0.0867   0.0328   1.0000
  12.750   1.5647   0.06007   0.05427  -0.0850   0.0319   1.0000
  13.000   1.5615   0.06336   0.05777  -0.0837   0.0311   1.0000
  13.250   1.5563   0.06681   0.06141  -0.0826   0.0304   1.0000
  13.500   1.5510   0.07044   0.06520  -0.0820   0.0298   1.0000
  13.750   1.5472   0.07454   0.06942  -0.0816   0.0290   1.0000
  14.000   1.5309   0.08072   0.07586  -0.0817   0.0286   1.0000
  14.250   1.5086   0.08747   0.08291  -0.0826   0.0284   1.0000
  14.500   1.4882   0.09372   0.08942  -0.0841   0.0284   1.0000
  15.000   1.4418   0.10781   0.10400  -0.0901   0.0283   1.0000
  15.250   1.1995   0.18317   0.18052  -0.1428   0.0409   1.0000
  15.500   1.1906   0.19367   0.19098  -0.1488   0.0423   1.0000
  15.750   1.1942   0.19871   0.19603  -0.1510   0.0435   1.0000
  16.000   0.9528   0.21175   0.20953  -0.1390   0.0653   1.0000
  16.250   0.9338   0.21885   0.21659  -0.1455   0.0639   1.0000
<< Back to FX 63-100 AIRFOIL (fx63100-il)

Polar data table (+)

Polar graphs


<< Back to FX 63-100 AIRFOIL (fx63100-il)