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

FX 63-145 AIRFOIL (fx63145-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: FX 63-145 AIRFOIL (fx63145-il)
Reynolds number: 100,000
Max Cl/Cd: 56.6 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx63145-il-100000-n5.txt
Download as CSV file: xf-fx63145-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 63-145 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2822   0.10951   0.10488  -0.0399   1.0000   0.0485
  -9.750  -0.2818   0.10594   0.10138  -0.0410   1.0000   0.0482
  -9.500  -0.2989   0.09670   0.09222  -0.0468   1.0000   0.0323
  -9.000  -0.2959   0.08959   0.08522  -0.0466   1.0000   0.0261
  -8.750  -0.2968   0.08662   0.08234  -0.0466   1.0000   0.0255
  -8.500  -0.3020   0.08305   0.07888  -0.0471   1.0000   0.0248
  -8.250  -0.3094   0.07844   0.07437  -0.0490   0.9990   0.0239
  -8.000  -0.3315   0.06370   0.05944  -0.0663   0.9657   0.0213
  -7.750  -0.3226   0.05802   0.05364  -0.0744   0.9372   0.0210
  -7.500  -0.3076   0.05237   0.04772  -0.0825   0.9129   0.0205
  -7.250  -0.2841   0.04676   0.04169  -0.0902   0.8921   0.0199
  -7.000  -0.2543   0.04152   0.03590  -0.0968   0.8717   0.0191
  -6.750  -0.2219   0.03697   0.03071  -0.1018   0.8491   0.0184
  -6.500  -0.1859   0.03303   0.02600  -0.1058   0.8269   0.0176
  -6.250  -0.1507   0.03038   0.02271  -0.1082   0.8044   0.0172
  -6.000  -0.1179   0.02838   0.02023  -0.1097   0.7820   0.0171
  -5.750  -0.0887   0.02670   0.01815  -0.1101   0.7590   0.0170
  -5.500  -0.0605   0.02525   0.01640  -0.1103   0.7375   0.0171
  -5.250  -0.0334   0.02400   0.01490  -0.1101   0.7180   0.0172
  -5.000  -0.0080   0.02291   0.01365  -0.1097   0.6991   0.0175
  -4.500   0.0439   0.02118   0.01163  -0.1093   0.6680   0.0183
  -4.250   0.0707   0.02047   0.01075  -0.1094   0.6553   0.0190
  -4.000   0.0986   0.01986   0.00996  -0.1097   0.6446   0.0203
  -3.750   0.1269   0.01933   0.00929  -0.1102   0.6344   0.0230
  -3.500   0.1555   0.01893   0.00877  -0.1105   0.6249   0.0265
  -3.250   0.1848   0.01854   0.00820  -0.1109   0.6166   0.0312
  -3.000   0.2146   0.01800   0.00750  -0.1114   0.6081   0.0435
  -2.750   0.2464   0.01683   0.00683  -0.1133   0.6008   0.1877
  -2.500   0.2776   0.01583   0.00677  -0.1152   0.5935   0.4288
  -2.250   0.3044   0.01587   0.00702  -0.1146   0.5857   0.5111
  -2.000   0.3317   0.01606   0.00715  -0.1142   0.5779   0.5530
  -1.750   0.3593   0.01624   0.00722  -0.1139   0.5691   0.5825
  -1.500   0.3862   0.01636   0.00722  -0.1134   0.5614   0.5960
  -1.250   0.4139   0.01644   0.00716  -0.1133   0.5537   0.6056
  -1.000   0.4427   0.01653   0.00704  -0.1134   0.5477   0.6157
  -0.750   0.4699   0.01661   0.00707  -0.1132   0.5414   0.6237
  -0.500   0.4981   0.01672   0.00705  -0.1132   0.5355   0.6346
  -0.250   0.5247   0.01685   0.00710  -0.1128   0.5301   0.6457
   0.000   0.5511   0.01696   0.00723  -0.1124   0.5236   0.6573
   0.250   0.5780   0.01710   0.00733  -0.1120   0.5179   0.6702
   0.500   0.6053   0.01724   0.00741  -0.1119   0.5126   0.6839
   0.750   0.6303   0.01734   0.00756  -0.1111   0.5066   0.6945
   1.000   0.6568   0.01744   0.00764  -0.1108   0.5009   0.7043
   1.250   0.6830   0.01755   0.00774  -0.1104   0.4953   0.7140
   1.500   0.7088   0.01765   0.00784  -0.1099   0.4893   0.7227
   1.750   0.7356   0.01777   0.00791  -0.1098   0.4839   0.7318
   2.000   0.7615   0.01787   0.00803  -0.1094   0.4783   0.7395
   2.250   0.7883   0.01799   0.00817  -0.1092   0.4723   0.7478
   2.500   0.8141   0.01811   0.00827  -0.1089   0.4668   0.7553
   2.750   0.8401   0.01824   0.00845  -0.1086   0.4603   0.7631
   3.000   0.8663   0.01837   0.00863  -0.1084   0.4531   0.7706
   3.250   0.8914   0.01849   0.00883  -0.1079   0.4461   0.7781
   3.500   0.9178   0.01864   0.00903  -0.1078   0.4380   0.7865
   3.750   0.9422   0.01877   0.00920  -0.1072   0.4307   0.7941
   4.000   0.9682   0.01893   0.00936  -0.1070   0.4229   0.8027
   4.250   0.9915   0.01906   0.00954  -0.1062   0.4156   0.8103
   4.500   1.0163   0.01925   0.00974  -0.1058   0.4065   0.8194
   4.750   1.0386   0.01942   0.00997  -0.1049   0.3974   0.8286
   5.000   1.0617   0.01962   0.01026  -0.1042   0.3864   0.8396
   5.250   1.0845   0.01983   0.01060  -0.1034   0.3737   0.8522
   5.500   1.1043   0.02001   0.01082  -0.1020   0.3632   0.8654
   5.750   1.1235   0.02022   0.01104  -0.1005   0.3548   0.8818
   6.250   1.1599   0.02058   0.01151  -0.0973   0.3401   0.9923
   6.500   1.1888   0.02101   0.01206  -0.0982   0.3303   1.0000
   6.750   1.2153   0.02147   0.01263  -0.0986   0.3169   1.0000
   7.000   1.2379   0.02203   0.01316  -0.0983   0.3018   1.0000
   7.250   1.2583   0.02268   0.01374  -0.0977   0.2914   1.0000
   7.500   1.2767   0.02341   0.01440  -0.0968   0.2821   1.0000
   7.750   1.2982   0.02400   0.01513  -0.0963   0.2706   1.0000
   8.000   1.3200   0.02456   0.01584  -0.0958   0.2578   1.0000
   8.250   1.3358   0.02535   0.01665  -0.0945   0.2452   1.0000
   8.500   1.3476   0.02625   0.01752  -0.0927   0.2374   1.0000
   8.750   1.3578   0.02722   0.01848  -0.0907   0.2314   1.0000
   9.000   1.3672   0.02826   0.01956  -0.0887   0.2264   1.0000
   9.250   1.3808   0.02914   0.02060  -0.0873   0.2203   1.0000
   9.500   1.3905   0.03025   0.02182  -0.0856   0.2142   1.0000
   9.750   1.4051   0.03114   0.02295  -0.0845   0.2036   1.0000
  10.000   1.4077   0.03276   0.02458  -0.0825   0.1922   1.0000
  10.250   1.4051   0.03486   0.02665  -0.0806   0.1851   1.0000
  10.500   1.4023   0.03720   0.02896  -0.0791   0.1798   1.0000
  10.750   1.4015   0.03960   0.03140  -0.0780   0.1736   1.0000
  11.000   1.4005   0.04219   0.03406  -0.0773   0.1669   1.0000
  11.250   1.4095   0.04403   0.03615  -0.0770   0.1565   1.0000
  11.500   1.4128   0.04646   0.03871  -0.0768   0.1461   1.0000
  11.750   1.4092   0.04966   0.04193  -0.0767   0.1414   1.0000
  12.000   1.4068   0.05285   0.04517  -0.0767   0.1386   1.0000
  12.250   1.4060   0.05594   0.04833  -0.0768   0.1364   1.0000
  12.750   1.4047   0.06228   0.05485  -0.0772   0.1313   1.0000
  13.000   1.4033   0.06562   0.05830  -0.0776   0.1284   1.0000
  13.250   1.3997   0.06926   0.06199  -0.0780   0.1253   1.0000
  13.500   1.4015   0.07245   0.06539  -0.0786   0.1204   1.0000
  13.750   1.4060   0.07545   0.06870  -0.0793   0.1116   1.0000
  14.000   1.4016   0.07970   0.07307  -0.0805   0.0998   1.0000
  14.250   1.3884   0.08531   0.07860  -0.0823   0.0952   1.0000
  14.500   1.3757   0.09100   0.08422  -0.0841   0.0916   1.0000
<< Back to FX 63-145 AIRFOIL (fx63145-il)

Polar data table (+)

Polar graphs


<< Back to FX 63-145 AIRFOIL (fx63145-il)