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

FX 63-100 AIRFOIL (fx63100-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: FX 63-100 AIRFOIL (fx63100-il)
Reynolds number: 100,000
Max Cl/Cd: 67.04 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx63100-il-100000.txt
Download as CSV file: xf-fx63100-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 63-100 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3679   0.09733   0.09271  -0.0366   1.0000   0.1030
  -7.750  -0.3950   0.09674   0.09229  -0.0360   1.0000   0.1034
  -7.500  -0.4165   0.09471   0.09039  -0.0401   1.0000   0.1037
  -7.250  -0.3739   0.08921   0.08480  -0.0294   1.0000   0.1080
  -7.000  -0.3836   0.08783   0.08350  -0.0263   1.0000   0.1099
  -6.750  -0.3952   0.08629   0.08205  -0.0242   1.0000   0.1124
  -6.500  -0.4087   0.08445   0.08030  -0.0240   1.0000   0.1153
  -6.250  -0.4272   0.07832   0.07420  -0.0429   1.0000   0.1196
  -6.000  -0.4256   0.07701   0.07300  -0.0331   1.0000   0.1211
  -5.750  -0.4233   0.07568   0.07172  -0.0276   1.0000   0.1238
  -5.500  -0.4104   0.06816   0.06403  -0.0454   1.0000   0.1357
  -5.250  -0.4100   0.06748   0.06350  -0.0373   1.0000   0.1381
  -5.000  -0.3902   0.06292   0.05885  -0.0443   0.9990   0.1527
  -4.750  -0.3556   0.05909   0.05496  -0.0498   0.9938   0.1692
  -4.500  -0.2314   0.03713   0.03048  -0.0865   0.9956   0.0891
  -4.250  -0.1827   0.03320   0.02602  -0.0920   0.9926   0.0881
  -4.000  -0.1336   0.03014   0.02229  -0.0971   0.9890   0.0888
  -3.750  -0.0896   0.02737   0.01938  -0.1014   0.9858   0.0943
  -3.500  -0.0457   0.02595   0.01757  -0.1047   0.9810   0.1003
  -3.250  -0.0030   0.02450   0.01583  -0.1078   0.9761   0.1064
  -3.000   0.0406   0.02361   0.01483  -0.1109   0.9716   0.1170
  -2.750   0.0779   0.02253   0.01384  -0.1130   0.9650   0.1270
  -2.500   0.1233   0.02168   0.01307  -0.1165   0.9604   0.1487
  -2.250   0.1636   0.02064   0.01235  -0.1194   0.9535   0.1931
  -2.000   0.2108   0.01948   0.01272  -0.1233   0.9487   0.5403
  -1.750   0.2431   0.01984   0.01312  -0.1236   0.9396   0.6136
  -1.500   0.2813   0.02007   0.01336  -0.1247   0.9329   0.6640
  -1.250   0.3063   0.02026   0.01361  -0.1234   0.9229   0.7004
  -1.000   0.3424   0.02028   0.01363  -0.1241   0.9167   0.7394
  -0.750   0.3676   0.02030   0.01366  -0.1231   0.9068   0.7703
  -0.500   0.3960   0.02019   0.01357  -0.1224   0.8990   0.8027
  -0.250   0.4230   0.01998   0.01338  -0.1215   0.8908   0.8345
   0.000   0.4471   0.01982   0.01323  -0.1202   0.8818   0.8645
   0.250   0.4775   0.01942   0.01283  -0.1199   0.8748   0.8959
   0.500   0.4993   0.01917   0.01261  -0.1184   0.8649   0.9394
   0.750   0.5534   0.01878   0.01217  -0.1234   0.8591   1.0000
   1.000   0.5972   0.01883   0.01215  -0.1270   0.8501   1.0000
   1.250   0.6467   0.01862   0.01188  -0.1311   0.8436   1.0000
   1.500   0.6857   0.01869   0.01192  -0.1334   0.8335   1.0000
   1.750   0.7323   0.01841   0.01160  -0.1365   0.8267   1.0000
   2.000   0.7658   0.01854   0.01172  -0.1376   0.8148   1.0000
   2.250   0.8021   0.01853   0.01170  -0.1388   0.8043   1.0000
   2.500   0.8423   0.01826   0.01141  -0.1404   0.7952   1.0000
   2.750   0.8742   0.01818   0.01132  -0.1405   0.7810   1.0000
   3.000   0.9063   0.01801   0.01116  -0.1405   0.7661   1.0000
   3.250   0.9372   0.01790   0.01105  -0.1403   0.7511   1.0000
   3.500   0.9671   0.01789   0.01105  -0.1400   0.7364   1.0000
   3.750   0.9966   0.01790   0.01109  -0.1397   0.7214   1.0000
   4.000   1.0256   0.01793   0.01114  -0.1393   0.7059   1.0000
   4.250   1.0543   0.01797   0.01119  -0.1388   0.6900   1.0000
   4.500   1.0831   0.01794   0.01117  -0.1383   0.6728   1.0000
   4.750   1.1088   0.01795   0.01122  -0.1372   0.6514   1.0000
   5.000   1.1352   0.01786   0.01111  -0.1360   0.6282   1.0000
   5.250   1.1597   0.01790   0.01116  -0.1347   0.6030   1.0000
   5.500   1.1845   0.01797   0.01121  -0.1335   0.5773   1.0000
   5.750   1.2073   0.01810   0.01132  -0.1319   0.5470   1.0000
   6.000   1.2281   0.01832   0.01150  -0.1302   0.5111   1.0000
   6.250   1.2480   0.01868   0.01172  -0.1283   0.4720   1.0000
   6.500   1.2664   0.01925   0.01217  -0.1264   0.4296   1.0000
   6.750   1.2831   0.01999   0.01276  -0.1243   0.3832   1.0000
   7.000   1.2973   0.02092   0.01350  -0.1220   0.3274   1.0000
   7.250   1.3064   0.02236   0.01451  -0.1191   0.2620   1.0000
   7.500   1.3113   0.02443   0.01606  -0.1158   0.2002   1.0000
   7.750   1.3170   0.02660   0.01781  -0.1128   0.1588   1.0000
   8.000   1.3277   0.02842   0.01946  -0.1103   0.1337   1.0000
   8.250   1.3413   0.03020   0.02112  -0.1083   0.1196   1.0000
   8.500   1.3593   0.03198   0.02287  -0.1069   0.1097   1.0000
   8.750   1.3803   0.03391   0.02470  -0.1060   0.1017   1.0000
   9.000   1.4030   0.03563   0.02652  -0.1053   0.0952   1.0000
   9.250   1.4321   0.03810   0.02889  -0.1057   0.0894   1.0000
   9.500   1.4525   0.03992   0.03100  -0.1046   0.0845   1.0000
   9.750   1.4776   0.04221   0.03329  -0.1045   0.0799   1.0000
  10.000   1.4997   0.04524   0.03657  -0.1039   0.0759   1.0000
  10.250   1.5143   0.04774   0.03948  -0.1022   0.0728   1.0000
  10.500   1.5296   0.05023   0.04219  -0.1008   0.0694   1.0000
  10.750   1.5514   0.05498   0.04690  -0.1013   0.0656   1.0000
  11.000   1.5501   0.05768   0.05016  -0.0978   0.0645   1.0000
  11.250   1.5455   0.06095   0.05394  -0.0943   0.0634   1.0000
  11.500   1.5356   0.06438   0.05780  -0.0906   0.0624   1.0000
  11.750   1.5207   0.06778   0.06156  -0.0867   0.0616   1.0000
  12.000   1.5032   0.07146   0.06557  -0.0834   0.0609   1.0000
  12.250   1.4833   0.07566   0.07009  -0.0809   0.0605   1.0000
  12.500   1.4605   0.08046   0.07520  -0.0794   0.0604   1.0000
  12.750   1.4341   0.08612   0.08116  -0.0789   0.0607   1.0000
  13.000   1.4049   0.09263   0.08794  -0.0798   0.0611   1.0000
  13.250   1.3743   0.10002   0.09560  -0.0821   0.0618   1.0000
  13.500   1.3434   0.10840   0.10419  -0.0859   0.0627   1.0000
  13.750   1.3142   0.11749   0.11346  -0.0908   0.0637   1.0000
  14.000   1.2880   0.12721   0.12330  -0.0965   0.0646   1.0000
  14.250   1.2689   0.13689   0.13304  -0.1018   0.0654   1.0000
<< Back to FX 63-100 AIRFOIL (fx63100-il)

Polar data table (+)

Polar graphs


<< Back to FX 63-100 AIRFOIL (fx63100-il)