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=500,000 Ncrit=9


Details Polar file
Airfoil: FX 61-140 AIRFOIL (fx61140-il)
Reynolds number: 500,000
Max Cl/Cd: 96.64 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx61140-il-500000.txt
Download as CSV file: xf-fx61140-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.3294   0.08495   0.08295  -0.0420   1.0000   0.0171
 -10.000  -0.3381   0.07927   0.07730  -0.0436   1.0000   0.0175
  -9.750  -0.3503   0.07264   0.07070  -0.0458   1.0000   0.0175
  -9.500  -0.3758   0.06201   0.06007  -0.0514   1.0000   0.0169
  -9.250  -0.4088   0.05375   0.05174  -0.0552   1.0000   0.0166
  -9.000  -0.4445   0.04778   0.04568  -0.0562   1.0000   0.0161
  -8.750  -0.4757   0.04435   0.04220  -0.0544   1.0000   0.0160
  -8.500  -0.4899   0.03979   0.03749  -0.0578   0.9959   0.0160
  -8.250  -0.4844   0.03457   0.03205  -0.0642   0.9891   0.0163
  -8.000  -0.4684   0.02983   0.02708  -0.0690   0.9850   0.0172
  -7.750  -0.4545   0.02584   0.02282  -0.0713   0.9775   0.0183
  -7.500  -0.4192   0.02532   0.02190  -0.0741   0.9736   0.0207
  -7.250  -0.4140   0.01740   0.01340  -0.0765   0.9647   0.0216
  -7.000  -0.4294   0.02437   0.01964  -0.0753   0.9709   0.0129
  -6.750  -0.3999   0.02140   0.01634  -0.0763   0.9650   0.0125
  -6.500  -0.3628   0.02022   0.01496  -0.0785   0.9615   0.0136
  -6.250  -0.3339   0.01844   0.01301  -0.0788   0.9537   0.0137
  -6.000  -0.3014   0.01658   0.01100  -0.0797   0.9484   0.0136
  -5.750  -0.2722   0.01527   0.00960  -0.0800   0.9390   0.0138
  -5.500  -0.2342   0.01423   0.00847  -0.0822   0.9322   0.0142
  -5.250  -0.1949   0.01239   0.00655  -0.0855   0.9216   0.0161
  -5.000  -0.1454   0.01157   0.00567  -0.0906   0.9101   0.0189
  -4.750  -0.0889   0.01065   0.00465  -0.0973   0.8942   0.0256
  -4.500  -0.0398   0.01001   0.00389  -0.1022   0.8686   0.0405
  -4.250   0.0057   0.00791   0.00285  -0.1088   0.8382   0.3280
  -4.000   0.0438   0.00705   0.00254  -0.1121   0.8080   0.5219
  -3.750   0.0729   0.00708   0.00253  -0.1124   0.7802   0.5698
  -3.500   0.1008   0.00723   0.00256  -0.1122   0.7556   0.5973
  -3.250   0.1280   0.00744   0.00265  -0.1119   0.7330   0.6184
  -3.000   0.1553   0.00774   0.00281  -0.1115   0.7124   0.6389
  -2.750   0.1823   0.00805   0.00300  -0.1111   0.6938   0.6523
  -2.500   0.2082   0.00825   0.00312  -0.1104   0.6770   0.6578
  -2.250   0.2360   0.00842   0.00315  -0.1104   0.6613   0.6641
  -2.000   0.2636   0.00855   0.00317  -0.1103   0.6469   0.6687
  -1.750   0.2906   0.00870   0.00325  -0.1099   0.6336   0.6726
  -1.500   0.3187   0.00882   0.00327  -0.1100   0.6210   0.6772
  -1.250   0.3475   0.00892   0.00326  -0.1103   0.6095   0.6817
  -1.000   0.3744   0.00904   0.00333  -0.1100   0.5990   0.6844
  -0.750   0.4020   0.00916   0.00339  -0.1099   0.5888   0.6877
  -0.500   0.4305   0.00926   0.00342  -0.1101   0.5792   0.6916
  -0.250   0.4599   0.00936   0.00342  -0.1105   0.5703   0.6956
   0.000   0.4871   0.00944   0.00349  -0.1103   0.5618   0.6979
   0.250   0.5146   0.00955   0.00357  -0.1103   0.5540   0.7006
   0.500   0.5425   0.00966   0.00364  -0.1103   0.5460   0.7038
   0.750   0.5716   0.00975   0.00368  -0.1107   0.5383   0.7076
   1.000   0.6002   0.00984   0.00372  -0.1109   0.5310   0.7106
   1.250   0.6277   0.00992   0.00381  -0.1109   0.5239   0.7127
   1.500   0.6551   0.01003   0.00390  -0.1108   0.5169   0.7151
   1.750   0.6833   0.01012   0.00399  -0.1110   0.5101   0.7177
   2.000   0.7119   0.01022   0.00406  -0.1112   0.5033   0.7208
   2.250   0.7413   0.01034   0.00414  -0.1117   0.4970   0.7240
   2.500   0.7686   0.01040   0.00422  -0.1117   0.4904   0.7259
   2.750   0.7960   0.01053   0.00436  -0.1116   0.4844   0.7279
   3.000   0.8238   0.01061   0.00448  -0.1117   0.4782   0.7303
   3.250   0.8516   0.01076   0.00460  -0.1118   0.4723   0.7330
   3.500   0.8803   0.01084   0.00473  -0.1121   0.4659   0.7358
   3.750   0.9089   0.01097   0.00484  -0.1124   0.4596   0.7385
   4.000   0.9360   0.01106   0.00498  -0.1123   0.4534   0.7404
   4.250   0.9632   0.01116   0.00512  -0.1123   0.4465   0.7424
   4.500   0.9904   0.01132   0.00531  -0.1123   0.4403   0.7447
   4.750   1.0181   0.01142   0.00547  -0.1123   0.4335   0.7472
   5.000   1.0458   0.01160   0.00563  -0.1125   0.4270   0.7497
   5.250   1.0741   0.01170   0.00578  -0.1127   0.4186   0.7522
   5.500   1.1004   0.01180   0.00594  -0.1126   0.4087   0.7542
   5.750   1.1262   0.01191   0.00607  -0.1123   0.3945   0.7561
   6.000   1.1522   0.01206   0.00624  -0.1120   0.3804   0.7581
   6.250   1.1782   0.01224   0.00643  -0.1119   0.3662   0.7604
   6.500   1.2041   0.01246   0.00664  -0.1117   0.3486   0.7630
   6.750   1.2293   0.01278   0.00690  -0.1115   0.3259   0.7654
   7.000   1.2533   0.01316   0.00720  -0.1110   0.2996   0.7674
   7.250   1.2759   0.01361   0.00758  -0.1103   0.2708   0.7693
   7.500   1.2964   0.01427   0.00808  -0.1094   0.2320   0.7714
   7.750   1.3119   0.01542   0.00889  -0.1078   0.1705   0.7736
   8.000   1.3240   0.01688   0.00995  -0.1058   0.1087   0.7759
   8.250   1.3375   0.01815   0.01097  -0.1039   0.0667   0.7784
   8.500   1.3519   0.01931   0.01196  -0.1022   0.0393   0.7807
   8.750   1.3662   0.02027   0.01287  -0.1003   0.0243   0.7826
   9.000   1.3765   0.02147   0.01403  -0.0978   0.0124   0.7845
   9.250   1.3885   0.02242   0.01503  -0.0955   0.0091   0.7868
   9.500   1.3968   0.02339   0.01611  -0.0926   0.0078   0.7893
   9.750   1.4070   0.02431   0.01712  -0.0901   0.0071   0.7919
  10.000   1.4158   0.02535   0.01825  -0.0876   0.0066   0.7942
  10.250   1.4213   0.02656   0.01957  -0.0849   0.0061   0.7963
  10.500   1.4223   0.02811   0.02125  -0.0819   0.0058   0.7984
  10.750   1.4177   0.03020   0.02348  -0.0787   0.0055   0.8009
  11.000   1.4125   0.03257   0.02601  -0.0761   0.0054   0.8036
  11.250   1.4126   0.03478   0.02837  -0.0745   0.0053   0.8062
  11.500   1.4114   0.03736   0.03108  -0.0734   0.0052   0.8086
  11.750   1.4094   0.04019   0.03405  -0.0726   0.0051   0.8107
  12.000   1.4062   0.04335   0.03735  -0.0721   0.0050   0.8128
  12.250   1.4026   0.04677   0.04092  -0.0719   0.0050   0.8152
  12.500   1.3990   0.05037   0.04466  -0.0720   0.0049   0.8177
  12.750   1.3948   0.05419   0.04861  -0.0724   0.0049   0.8201
  13.000   1.3911   0.05809   0.05263  -0.0728   0.0048   0.8223
  13.250   1.3871   0.06205   0.05673  -0.0734   0.0048   0.8243
  13.500   1.3832   0.06609   0.06092  -0.0740   0.0048   0.8261
  13.750   1.3802   0.07012   0.06509  -0.0747   0.0048   0.8281
  14.000   1.3768   0.07431   0.06941  -0.0755   0.0047   0.8301
  14.250   1.3742   0.07853   0.07378  -0.0764   0.0047   0.8322
  14.500   1.3718   0.08281   0.07821  -0.0774   0.0047   0.8343
  14.750   1.3684   0.08736   0.08291  -0.0785   0.0047   0.8363
  15.000   1.3655   0.09193   0.08764  -0.0799   0.0048   0.8383
  15.250   1.3610   0.09685   0.09274  -0.0814   0.0048   0.8403
  15.500   1.3555   0.10208   0.09815  -0.0831   0.0048   0.8423
  15.750   1.3492   0.10758   0.10384  -0.0853   0.0048   0.8444
  16.000   1.3423   0.11338   0.10984  -0.0879   0.0049   0.8467
  16.250   1.3311   0.12014   0.11681  -0.0911   0.0049   0.8484
  16.500   1.3210   0.12686   0.12371  -0.0946   0.0050   0.8505
  16.750   1.3103   0.13386   0.13090  -0.0985   0.0050   0.8524
  17.000   1.2983   0.14131   0.13854  -0.1029   0.0051   0.8545
  17.250   1.2865   0.14894   0.14635  -0.1077   0.0051   0.8568
  17.500   1.2723   0.15739   0.15499  -0.1132   0.0052   0.8591
  17.750   1.2570   0.16647   0.16423  -0.1193   0.0053   0.8615
<< Back to FX 61-140 AIRFOIL (fx61140-il)

Polar data table (+)

Polar graphs


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