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WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT NASA) (fx6617ai-il)
Reynolds number: 50,000
Max Cl/Cd: 4.72 at α=9.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx6617ai-il-50000.txt
Download as CSV file: xf-fx6617ai-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT N
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.1878   0.11617   0.11074  -0.0613   0.9508   0.3202
 -10.000  -0.1656   0.11105   0.10553  -0.0661   0.9413   0.3223
  -9.750  -0.2607   0.09002   0.08461  -0.0885   0.9351   0.1573
  -9.500  -0.3809   0.07448   0.06894  -0.1041   0.9204   0.1322
  -9.250  -0.3791   0.06958   0.06391  -0.1071   0.9110   0.1297
  -9.000  -0.3961   0.06606   0.06020  -0.1076   0.8998   0.1273
  -8.750  -0.4129   0.06233   0.05613  -0.1081   0.8909   0.1247
  -8.500  -0.4264   0.05930   0.05269  -0.1073   0.8818   0.1226
  -8.250  -0.4324   0.05674   0.04968  -0.1061   0.8739   0.1214
  -8.000  -0.4285   0.05435   0.04691  -0.1052   0.8665   0.1215
  -7.750  -0.4133   0.05167   0.04379  -0.1054   0.8610   0.1219
  -7.500  -0.4141   0.05050   0.04235  -0.1027   0.8541   0.1219
  -7.250  -0.3928   0.04828   0.03976  -0.1028   0.8492   0.1229
  -7.000  -0.3756   0.04678   0.03799  -0.1020   0.8442   0.1248
  -6.750  -0.3700   0.04615   0.03717  -0.0996   0.8384   0.1276
  -6.500  -0.3462   0.04493   0.03545  -0.0995   0.8338   0.1332
  -6.250  -0.3110   0.04327   0.03393  -0.1003   0.8301   0.1421
  -6.000  -0.3149   0.04356   0.03413  -0.0963   0.8261   0.1457
  -5.750  -0.3058   0.04327   0.03394  -0.0937   0.8225   0.1530
  -5.500  -0.2883   0.04272   0.03347  -0.0922   0.8190   0.1682
  -5.250  -0.2630   0.04159   0.03255  -0.0920   0.8157   0.1975
  -5.000  -0.2657   0.04113   0.03245  -0.0888   0.8133   0.2296
  -4.750  -0.2759   0.03965   0.03270  -0.0848   0.8124   0.3697
  -4.500  -0.2879   0.04333   0.03694  -0.0743   0.8119   0.5509
  -4.250  -0.2911   0.04628   0.03981  -0.0664   0.8114   0.5920
  -4.000  -0.2920   0.04855   0.04199  -0.0596   0.8111   0.6223
  -3.750  -0.2913   0.05041   0.04374  -0.0534   0.8117   0.6501
  -3.500  -0.2921   0.05251   0.04581  -0.0455   0.8136   0.6830
  -3.250  -0.4713   0.04285   0.03408  -0.0420   0.9998   0.2130
  -3.000  -0.4619   0.04324   0.03712  -0.0359   0.9998   0.5840
  -2.750  -0.4615   0.04559   0.03940  -0.0278   0.9998   0.6364
  -2.500  -0.4580   0.04716   0.04085  -0.0213   0.9998   0.6721
  -2.250  -0.4543   0.04829   0.04188  -0.0153   0.9998   0.7031
  -2.000  -0.4557   0.04928   0.04282  -0.0074   0.9998   0.7420
  -1.750  -0.4618   0.04976   0.04330   0.0019   0.9998   0.7806
  -1.500  -0.4639   0.04985   0.04330   0.0094   0.9998   0.8185
  -1.250  -0.4606   0.04972   0.04303   0.0147   0.9998   0.8488
  -1.000  -0.4499   0.04956   0.04269   0.0173   0.9998   0.8692
  -0.750  -0.4336   0.04956   0.04250   0.0179   0.9998   0.8861
  -0.500  -0.4157   0.04967   0.04241   0.0180   0.9998   0.9029
  -0.250  -0.3906   0.05019   0.04271   0.0165   0.9977   0.9200
   0.000  -0.3509   0.05158   0.04386   0.0117   0.9903   0.9358
   0.250  -0.3028   0.05363   0.04566   0.0048   0.9822   0.9492
   0.500  -0.2530   0.05564   0.04743  -0.0027   0.9708   0.9606
   0.750  -0.2069   0.05735   0.04894  -0.0094   0.9580   0.9710
   1.000  -0.1613   0.05911   0.05052  -0.0161   0.9455   0.9803
   1.250  -0.1173   0.06099   0.05223  -0.0227   0.9343   0.9911
   1.500  -0.0778   0.06311   0.05418  -0.0284   0.9245   1.0002
   2.000  -0.0484   0.06369   0.05454  -0.0297   0.8968   1.0002
   2.250  -0.0304   0.06442   0.05514  -0.0311   0.8853   1.0002
   2.500   0.0104   0.06724   0.05780  -0.0367   0.8784   1.0002
   2.750   0.0284   0.06771   0.05818  -0.0379   0.8653   1.0002
   3.000   0.0502   0.06891   0.05928  -0.0399   0.8544   1.0002
   3.250   0.0922   0.07189   0.06212  -0.0455   0.8474   1.0002
   3.500   0.1077   0.07253   0.06270  -0.0463   0.8354   1.0002
   3.750   0.1364   0.07471   0.06477  -0.0494   0.8278   1.0002
   4.000   0.1669   0.07668   0.06667  -0.0527   0.8172   1.0002
   4.250   0.1835   0.07792   0.06785  -0.0537   0.8071   1.0002
   4.500   0.2223   0.08098   0.07081  -0.0582   0.8002   1.0002
   4.750   0.2338   0.08180   0.07160  -0.0583   0.7889   1.0002
   5.000   0.2764   0.08560   0.07531  -0.0633   0.7834   1.0002
   5.250   0.2819   0.08587   0.07556  -0.0624   0.7715   1.0002
   5.500   0.3140   0.08908   0.07871  -0.0657   0.7663   1.0002
   5.750   0.3268   0.09002   0.07962  -0.0658   0.7547   1.0002
   6.000   0.3426   0.09191   0.08148  -0.0666   0.7470   1.0002
   6.250   0.3679   0.09418   0.08372  -0.0684   0.7384   1.0002
   6.500   0.3801   0.09592   0.08544  -0.0687   0.7308   1.0002
   6.750   0.4057   0.09834   0.08784  -0.0705   0.7223   1.0002
   7.000   0.4152   0.09996   0.08947  -0.0704   0.7144   1.0002
   7.250   0.4395   0.10244   0.09194  -0.0720   0.7067   1.0002
   7.500   0.4501   0.10434   0.09384  -0.0720   0.6999   1.0002
   7.750   0.4709   0.10657   0.09608  -0.0731   0.6912   1.0002
   8.000   0.4851   0.10895   0.09848  -0.0736   0.6854   1.0002
   8.250   0.4988   0.11066   0.10022  -0.0739   0.6763   1.0002
   8.500   0.5350   0.11522   0.10481  -0.0769   0.6724   1.0002
   8.750   0.5256   0.11488   0.10450  -0.0746   0.6615   1.0002
   9.000   0.5583   0.11900   0.10868  -0.0771   0.6570   1.0002
   9.250   0.5508   0.11924   0.10896  -0.0753   0.6476   1.0002
   9.500   0.5790   0.12278   0.11256  -0.0771   0.6419   1.0002
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