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

FX 61-140 AIRFOIL (fx61140-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: FX 61-140 AIRFOIL (fx61140-il)
Reynolds number: 1,000,000
Max Cl/Cd: 115.25 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx61140-il-1000000-n5.txt
Download as CSV file: xf-fx61140-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 61-140 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5406   0.07636   0.07482  -0.0463   1.0000   0.0022
 -11.000  -0.7244   0.04042   0.03808  -0.0638   0.9992   0.0021
 -10.750  -0.7439   0.03329   0.03048  -0.0678   0.9951   0.0020
 -10.500  -0.7457   0.02869   0.02548  -0.0698   0.9896   0.0020
 -10.250  -0.7273   0.02519   0.02162  -0.0731   0.9860   0.0020
 -10.000  -0.7104   0.02271   0.01885  -0.0740   0.9818   0.0021
  -9.750  -0.6900   0.02085   0.01677  -0.0747   0.9767   0.0021
  -9.250  -0.6444   0.01817   0.01376  -0.0756   0.9652   0.0021
  -9.000  -0.6169   0.01706   0.01251  -0.0766   0.9598   0.0022
  -8.750  -0.5914   0.01605   0.01137  -0.0771   0.9507   0.0022
  -8.500  -0.5567   0.01512   0.01034  -0.0794   0.9434   0.0023
  -8.250  -0.5133   0.01423   0.00933  -0.0834   0.9345   0.0023
  -8.000  -0.4598   0.01292   0.00785  -0.0901   0.9224   0.0027
  -7.750  -0.3943   0.01226   0.00708  -0.0989   0.9023   0.0031
  -7.500  -0.3515   0.01187   0.00647  -0.1026   0.8597   0.0034
  -7.250  -0.3262   0.01164   0.00605  -0.1025   0.8222   0.0037
  -7.000  -0.3027   0.01144   0.00568  -0.1020   0.7903   0.0040
  -6.750  -0.2785   0.01123   0.00531  -0.1016   0.7627   0.0042
  -6.500  -0.2536   0.01091   0.00486  -0.1015   0.7385   0.0047
  -6.250  -0.2279   0.01071   0.00456  -0.1014   0.7170   0.0054
  -6.000  -0.2014   0.01051   0.00427  -0.1014   0.6972   0.0059
  -5.750  -0.1744   0.01035   0.00400  -0.1015   0.6788   0.0065
  -5.500  -0.1471   0.01016   0.00373  -0.1017   0.6614   0.0076
  -5.250  -0.1194   0.01000   0.00350  -0.1020   0.6459   0.0092
  -4.750  -0.0628   0.00967   0.00307  -0.1027   0.6182   0.0141
  -4.500  -0.0340   0.00953   0.00286  -0.1031   0.6062   0.0172
  -4.250  -0.0051   0.00939   0.00267  -0.1036   0.5943   0.0220
  -4.000   0.0242   0.00921   0.00250  -0.1041   0.5836   0.0354
  -3.750   0.0541   0.00898   0.00232  -0.1049   0.5737   0.0639
  -3.500   0.0863   0.00848   0.00207  -0.1066   0.5639   0.1468
  -3.250   0.1209   0.00781   0.00179  -0.1090   0.5542   0.2745
  -3.000   0.1564   0.00722   0.00156  -0.1115   0.5445   0.4016
  -2.750   0.1903   0.00687   0.00144  -0.1133   0.5357   0.4885
  -2.500   0.2224   0.00670   0.00141  -0.1146   0.5271   0.5544
  -2.250   0.2525   0.00669   0.00141  -0.1151   0.5186   0.5796
  -2.000   0.2821   0.00672   0.00141  -0.1155   0.5111   0.5936
  -1.750   0.3113   0.00677   0.00142  -0.1159   0.5035   0.6023
  -1.500   0.3406   0.00682   0.00142  -0.1162   0.4971   0.6097
  -1.250   0.3698   0.00687   0.00144  -0.1165   0.4908   0.6159
  -1.000   0.3990   0.00693   0.00146  -0.1168   0.4854   0.6215
  -0.750   0.4282   0.00698   0.00147  -0.1171   0.4795   0.6261
  -0.500   0.4571   0.00705   0.00151  -0.1174   0.4733   0.6304
  -0.250   0.4862   0.00711   0.00154  -0.1177   0.4677   0.6348
   0.000   0.5152   0.00717   0.00157  -0.1180   0.4619   0.6387
   0.250   0.5440   0.00724   0.00161  -0.1182   0.4566   0.6417
   0.500   0.5729   0.00731   0.00166  -0.1185   0.4503   0.6450
   1.000   0.6305   0.00747   0.00177  -0.1190   0.4385   0.6522
   1.250   0.6592   0.00755   0.00183  -0.1192   0.4326   0.6550
   1.500   0.6878   0.00763   0.00190  -0.1195   0.4274   0.6574
   1.750   0.7166   0.00770   0.00198  -0.1197   0.4222   0.6599
   2.000   0.7451   0.00779   0.00206  -0.1199   0.4171   0.6623
   2.250   0.7737   0.00788   0.00214  -0.1201   0.4126   0.6647
   2.500   0.8023   0.00796   0.00222  -0.1204   0.4070   0.6671
   3.000   0.8591   0.00815   0.00242  -0.1207   0.3962   0.6711
   3.250   0.8872   0.00826   0.00253  -0.1209   0.3897   0.6731
   3.500   0.9154   0.00837   0.00265  -0.1210   0.3830   0.6753
   3.750   0.9434   0.00849   0.00277  -0.1212   0.3758   0.6776
   4.000   0.9715   0.00862   0.00290  -0.1213   0.3696   0.6799
   4.250   0.9993   0.00876   0.00303  -0.1214   0.3612   0.6819
   4.500   1.0269   0.00891   0.00318  -0.1215   0.3510   0.6838
   4.750   1.0534   0.00916   0.00337  -0.1214   0.3316   0.6856
   5.000   1.0787   0.00955   0.00361  -0.1212   0.3003   0.6874
   5.250   1.1022   0.01012   0.00396  -0.1206   0.2547   0.6892
   5.500   1.1259   0.01067   0.00433  -0.1201   0.2183   0.6911
   5.750   1.1464   0.01157   0.00491  -0.1192   0.1585   0.6930
   6.000   1.1636   0.01280   0.00571  -0.1178   0.0864   0.6948
   6.250   1.1840   0.01367   0.00635  -0.1168   0.0450   0.6963
   6.500   1.2051   0.01443   0.00697  -0.1158   0.0185   0.6979
   6.750   1.2288   0.01490   0.00742  -0.1153   0.0107   0.6996
   7.000   1.2509   0.01550   0.00801  -0.1145   0.0029   0.7012
   7.250   1.2746   0.01593   0.00848  -0.1140   0.0022   0.7030
   7.500   1.2984   0.01634   0.00893  -0.1135   0.0020   0.7048
   7.750   1.3218   0.01676   0.00940  -0.1129   0.0018   0.7066
   8.000   1.3445   0.01722   0.00991  -0.1123   0.0017   0.7085
   8.250   1.3666   0.01772   0.01046  -0.1115   0.0015   0.7102
   8.500   1.3879   0.01827   0.01107  -0.1106   0.0014   0.7117
   8.750   1.4082   0.01885   0.01174  -0.1096   0.0013   0.7132
   9.000   1.4272   0.01952   0.01248  -0.1084   0.0012   0.7148
   9.250   1.4450   0.02021   0.01325  -0.1070   0.0012   0.7165
   9.500   1.4609   0.02100   0.01412  -0.1053   0.0011   0.7182
   9.750   1.4720   0.02196   0.01518  -0.1028   0.0011   0.7199
  10.000   1.4799   0.02288   0.01619  -0.0998   0.0010   0.7215
  10.250   1.4845   0.02402   0.01742  -0.0964   0.0010   0.7231
  10.500   1.4890   0.02522   0.01872  -0.0933   0.0010   0.7247
  11.000   1.4964   0.02794   0.02166  -0.0878   0.0010   0.7281
  11.250   1.4991   0.02953   0.02335  -0.0854   0.0010   0.7297
  11.500   1.5007   0.03135   0.02528  -0.0833   0.0010   0.7313
  11.750   1.5011   0.03344   0.02749  -0.0816   0.0010   0.7330
  12.000   1.5014   0.03575   0.02990  -0.0803   0.0010   0.7346
  12.250   1.5004   0.03840   0.03265  -0.0793   0.0010   0.7363
  12.500   1.4983   0.04138   0.03575  -0.0787   0.0010   0.7379
  12.750   1.4961   0.04459   0.03907  -0.0785   0.0010   0.7394
  13.000   1.4937   0.04799   0.04259  -0.0786   0.0010   0.7410
  13.250   1.4902   0.05169   0.04642  -0.0789   0.0010   0.7424
  13.500   1.4859   0.05564   0.05049  -0.0794   0.0010   0.7437
  13.750   1.4823   0.05959   0.05455  -0.0802   0.0009   0.7451
  14.000   1.4770   0.06391   0.05900  -0.0811   0.0009   0.7464
  14.250   1.4746   0.06795   0.06314  -0.0821   0.0009   0.7478
  14.500   1.4687   0.07257   0.06789  -0.0834   0.0009   0.7491
  14.750   1.4613   0.07762   0.07305  -0.0849   0.0009   0.7503
  15.000   1.4557   0.08253   0.07808  -0.0865   0.0009   0.7516
  15.250   1.4488   0.08774   0.08341  -0.0883   0.0009   0.7528
  15.750   1.4409   0.09760   0.09350  -0.0922   0.0009   0.7553
  16.000   1.4313   0.10370   0.09973  -0.0946   0.0009   0.7564
  16.250   1.4257   0.10915   0.10530  -0.0970   0.0009   0.7576
<< Back to FX 61-140 AIRFOIL (fx61140-il)

Polar data table (+)

Polar graphs


<< Back to FX 61-140 AIRFOIL (fx61140-il)