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

FX 66-17AII-182 AIRFOIL (fx6617a2-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: FX 66-17AII-182 AIRFOIL (fx6617a2-il)
Reynolds number: 1,000,000
Max Cl/Cd: 125.76 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx6617a2-il-1000000-n5.txt
Download as CSV file: xf-fx6617a2-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-17AII-182 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -18.000  -0.6957   0.11323   0.11099  -0.0485   0.9996   0.0051
 -17.750  -0.7651   0.09691   0.09445  -0.0562   0.9999   0.0050
 -17.500  -0.8006   0.08549   0.08284  -0.0633   0.9995   0.0051
 -17.250  -0.8209   0.07657   0.07375  -0.0698   0.9984   0.0051
 -17.000  -0.8511   0.06758   0.06458  -0.0752   0.9959   0.0050
 -16.750  -0.8659   0.06350   0.06043  -0.0751   0.9623   0.0051
 -16.500  -0.7951   0.05341   0.04993  -0.1017   0.9189   0.0051
 -16.250  -0.7645   0.04854   0.04436  -0.1129   0.7953   0.0052
 -16.000  -0.7727   0.04527   0.04085  -0.1137   0.7603   0.0051
 -15.750  -0.7804   0.04219   0.03758  -0.1143   0.7357   0.0051
 -15.500  -0.7783   0.04021   0.03545  -0.1144   0.7151   0.0052
 -15.250  -0.7814   0.03785   0.03291  -0.1144   0.6955   0.0052
 -15.000  -0.7832   0.03566   0.03059  -0.1143   0.6821   0.0053
 -14.750  -0.7833   0.03368   0.02849  -0.1141   0.6703   0.0054
 -14.500  -0.7806   0.03199   0.02668  -0.1136   0.6588   0.0055
 -14.250  -0.7808   0.03010   0.02468  -0.1130   0.6498   0.0056
 -14.000  -0.7751   0.02871   0.02319  -0.1124   0.6413   0.0056
 -13.750  -0.7693   0.02736   0.02175  -0.1117   0.6336   0.0057
 -13.500  -0.7617   0.02616   0.02047  -0.1109   0.6263   0.0058
 -13.250  -0.7539   0.02499   0.01922  -0.1101   0.6196   0.0059
 -13.000  -0.7461   0.02386   0.01800  -0.1091   0.6129   0.0059
 -12.750  -0.7372   0.02280   0.01687  -0.1082   0.6074   0.0061
 -12.500  -0.7277   0.02182   0.01582  -0.1072   0.6024   0.0062
 -12.250  -0.7178   0.02091   0.01484  -0.1061   0.5975   0.0063
 -12.000  -0.7090   0.01999   0.01384  -0.1049   0.5937   0.0064
 -11.750  -0.7003   0.01918   0.01297  -0.1033   0.5898   0.0065
 -11.500  -0.6929   0.01847   0.01219  -0.1011   0.5861   0.0066
 -11.250  -0.6846   0.01796   0.01161  -0.0986   0.5824   0.0067
 -11.000  -0.6714   0.01747   0.01105  -0.0968   0.5794   0.0069
 -10.750  -0.6551   0.01697   0.01049  -0.0955   0.5765   0.0070
 -10.500  -0.6373   0.01652   0.00998  -0.0943   0.5730   0.0072
 -10.250  -0.6189   0.01607   0.00948  -0.0932   0.5698   0.0076
 -10.000  -0.5991   0.01567   0.00904  -0.0923   0.5665   0.0079
  -9.750  -0.5780   0.01532   0.00866  -0.0915   0.5639   0.0084
  -9.500  -0.5561   0.01498   0.00829  -0.0909   0.5616   0.0090
  -9.250  -0.5338   0.01465   0.00793  -0.0903   0.5592   0.0095
  -9.000  -0.5099   0.01441   0.00770  -0.0899   0.5566   0.0105
  -8.750  -0.4858   0.01419   0.00744  -0.0895   0.5542   0.0114
  -8.500  -0.4617   0.01396   0.00716  -0.0891   0.5517   0.0122
  -8.000  -0.4132   0.01348   0.00663  -0.0883   0.5476   0.0135
  -7.750  -0.3880   0.01327   0.00640  -0.0881   0.5457   0.0142
  -7.500  -0.3628   0.01306   0.00616  -0.0878   0.5433   0.0150
  -7.250  -0.3377   0.01284   0.00590  -0.0875   0.5411   0.0155
  -7.000  -0.3124   0.01264   0.00565  -0.0873   0.5387   0.0158
  -6.750  -0.2873   0.01241   0.00538  -0.0870   0.5367   0.0163
  -6.500  -0.2631   0.01211   0.00504  -0.0866   0.5347   0.0169
  -6.250  -0.2379   0.01187   0.00478  -0.0863   0.5329   0.0175
  -6.000  -0.2120   0.01166   0.00454  -0.0861   0.5315   0.0181
  -5.750  -0.1859   0.01146   0.00432  -0.0860   0.5299   0.0187
  -5.500  -0.1597   0.01128   0.00412  -0.0858   0.5282   0.0194
  -5.250  -0.1332   0.01112   0.00393  -0.0857   0.5262   0.0200
  -5.000  -0.1071   0.01093   0.00371  -0.0855   0.5242   0.0211
  -4.750  -0.0808   0.01076   0.00350  -0.0854   0.5222   0.0218
  -4.500  -0.0544   0.01061   0.00333  -0.0852   0.5203   0.0225
  -4.250  -0.0278   0.01049   0.00318  -0.0851   0.5184   0.0231
  -4.000  -0.0008   0.01035   0.00303  -0.0851   0.5171   0.0240
  -3.750   0.0264   0.01023   0.00290  -0.0851   0.5156   0.0252
  -3.500   0.0534   0.01009   0.00277  -0.0851   0.5138   0.0293
  -3.250   0.0796   0.00988   0.00262  -0.0849   0.5120   0.0484
  -3.000   0.1053   0.00963   0.00248  -0.0847   0.5103   0.0801
  -2.750   0.1306   0.00934   0.00233  -0.0845   0.5086   0.1249
  -2.500   0.1557   0.00906   0.00220  -0.0842   0.5067   0.1748
  -2.250   0.1799   0.00870   0.00205  -0.0838   0.5049   0.2505
  -2.000   0.2039   0.00834   0.00190  -0.0834   0.5029   0.3194
  -1.750   0.2253   0.00770   0.00173  -0.0826   0.5018   0.4578
  -1.500   0.2522   0.00759   0.00174  -0.0825   0.5005   0.5012
  -1.250   0.2802   0.00757   0.00176  -0.0826   0.4991   0.5202
  -1.000   0.3085   0.00759   0.00178  -0.0828   0.4975   0.5340
  -0.750   0.3366   0.00760   0.00179  -0.0829   0.4959   0.5412
  -0.500   0.3650   0.00763   0.00180  -0.0831   0.4941   0.5464
  -0.250   0.3932   0.00767   0.00181  -0.0833   0.4922   0.5508
   0.000   0.4211   0.00770   0.00185  -0.0833   0.4904   0.5584
   0.250   0.4491   0.00776   0.00187  -0.0835   0.4886   0.5621
   0.500   0.4774   0.00781   0.00190  -0.0837   0.4872   0.5650
   0.750   0.5056   0.00782   0.00193  -0.0839   0.4860   0.5679
   1.000   0.5339   0.00785   0.00196  -0.0840   0.4846   0.5703
   1.250   0.5621   0.00788   0.00200  -0.0842   0.4830   0.5725
   1.500   0.5903   0.00791   0.00203  -0.0844   0.4813   0.5748
   1.750   0.6183   0.00796   0.00207  -0.0846   0.4794   0.5773
   2.000   0.6462   0.00801   0.00211  -0.0847   0.4775   0.5799
   2.250   0.6738   0.00807   0.00216  -0.0848   0.4756   0.5819
   2.500   0.7010   0.00812   0.00221  -0.0848   0.4734   0.5841
   2.750   0.7288   0.00816   0.00228  -0.0850   0.4717   0.5862
   3.000   0.7567   0.00820   0.00234  -0.0851   0.4699   0.5882
   3.250   0.7845   0.00824   0.00240  -0.0853   0.4681   0.5901
   3.500   0.8122   0.00830   0.00247  -0.0854   0.4664   0.5920
   3.750   0.8397   0.00836   0.00254  -0.0855   0.4646   0.5940
   4.000   0.8670   0.00842   0.00262  -0.0856   0.4628   0.5960
   4.500   0.9206   0.00857   0.00279  -0.0856   0.4591   0.5995
   4.750   0.9475   0.00864   0.00289  -0.0856   0.4572   0.6014
   5.000   0.9748   0.00867   0.00297  -0.0857   0.4533   0.6034
   5.250   1.0009   0.00875   0.00305  -0.0855   0.4476   0.6054
   5.500   1.0262   0.00886   0.00316  -0.0853   0.4433   0.6077
   5.750   1.0532   0.00892   0.00327  -0.0853   0.4411   0.6100
   6.000   1.0798   0.00900   0.00338  -0.0853   0.4382   0.6120
   6.250   1.1051   0.00910   0.00350  -0.0850   0.4322   0.6140
   6.500   1.1297   0.00921   0.00363  -0.0846   0.4267   0.6163
   6.750   1.1553   0.00930   0.00377  -0.0844   0.4218   0.6193
   7.000   1.1794   0.00943   0.00391  -0.0840   0.4160   0.6217
   7.250   1.2035   0.00957   0.00408  -0.0835   0.4094   0.6248
   7.500   1.2247   0.00980   0.00427  -0.0825   0.3958   0.6278
   7.750   1.2434   0.01010   0.00452  -0.0811   0.3781   0.6306
   8.000   1.2540   0.01066   0.00494  -0.0783   0.3462   0.6334
   8.250   1.2454   0.01163   0.00565  -0.0719   0.2976   0.6360
   8.500   1.2219   0.01314   0.00682  -0.0632   0.2372   0.6388
   8.750   1.2052   0.01453   0.00800  -0.0563   0.1954   0.6416
   9.000   1.1807   0.01641   0.00966  -0.0491   0.1521   0.6441
   9.250   1.1644   0.01835   0.01149  -0.0441   0.1216   0.6466
   9.500   1.1508   0.02059   0.01365  -0.0403   0.0961   0.6496
   9.750   1.1482   0.02247   0.01552  -0.0380   0.0818   0.6529
  10.000   1.1360   0.02522   0.01820  -0.0355   0.0601   0.6559
  10.250   1.1373   0.02714   0.02012  -0.0341   0.0521   0.6592
  10.500   1.1296   0.02981   0.02276  -0.0323   0.0373   0.6624
  10.750   1.1332   0.03166   0.02464  -0.0313   0.0329   0.6665
  11.000   1.1353   0.03366   0.02666  -0.0303   0.0286   0.6709
  11.250   1.1314   0.03616   0.02916  -0.0291   0.0171   0.6753
  11.500   1.1311   0.03845   0.03147  -0.0281   0.0134   0.6799
  11.750   1.1371   0.04025   0.03333  -0.0276   0.0130   0.6858
  12.000   1.1424   0.04214   0.03527  -0.0270   0.0123   0.6918
  12.250   1.1467   0.04419   0.03739  -0.0266   0.0114   0.6985
  12.750   1.1601   0.04797   0.04132  -0.0259   0.0103   0.7161
  13.000   1.1680   0.04979   0.04323  -0.0258   0.0102   0.7290
  13.250   1.1746   0.05184   0.04539  -0.0257   0.0100   0.7480
  13.500   1.1848   0.05374   0.04748  -0.0263   0.0096   0.7910
  13.750   1.2829   0.05956   0.05382  -0.0496   0.0084   0.9440
  14.000   1.2874   0.06181   0.05611  -0.0489   0.0082   0.9998
  14.250   1.2959   0.06399   0.05834  -0.0492   0.0080   1.0000
  14.500   1.3015   0.06625   0.06064  -0.0492   0.0077   1.0000
  14.750   1.3078   0.06841   0.06284  -0.0491   0.0074   1.0000
  15.000   1.3122   0.07080   0.06528  -0.0490   0.0071   1.0000
  15.250   1.3158   0.07331   0.06783  -0.0489   0.0070   1.0000
  15.500   1.3211   0.07562   0.07019  -0.0489   0.0068   1.0000
  15.750   1.3242   0.07823   0.07284  -0.0490   0.0065   1.0000
  16.000   1.3274   0.08082   0.07546  -0.0490   0.0061   1.0000
  16.250   1.3318   0.08327   0.07797  -0.0491   0.0060   1.0000
  16.500   1.3351   0.08591   0.08066  -0.0492   0.0059   1.0000
  16.750   1.3408   0.08822   0.08302  -0.0494   0.0055   1.0000
  17.000   1.3445   0.09082   0.08567  -0.0496   0.0053   1.0000
  17.250   1.3471   0.09358   0.08849  -0.0498   0.0052   1.0000
  17.500   1.3513   0.09609   0.09104  -0.0500   0.0050   1.0000
  17.750   1.3547   0.09875   0.09374  -0.0504   0.0048   1.0000
  18.000   1.3566   0.10159   0.09662  -0.0507   0.0046   1.0000
  18.250   1.3588   0.10443   0.09950  -0.0511   0.0044   1.0000
<< Back to FX 66-17AII-182 AIRFOIL (fx6617a2-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-17AII-182 AIRFOIL (fx6617a2-il)