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

FX 71-L-150/25 AIRFOIL (fx711525-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: FX 71-L-150/25 AIRFOIL (fx711525-il)
Reynolds number: 100,000
Max Cl/Cd: 38.74 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx711525-il-100000.txt
Download as CSV file: xf-fx711525-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 71-L-150/25 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
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.7440   0.10229   0.09686  -0.0172   1.0000   0.0669
 -13.000  -0.7533   0.09585   0.09041  -0.0202   1.0000   0.0655
 -12.750  -0.8147   0.08310   0.07744  -0.0286   1.0000   0.0633
 -12.500  -0.9324   0.07133   0.06495  -0.0347   1.0000   0.0608
 -12.250  -0.9277   0.06759   0.06117  -0.0349   1.0000   0.0602
 -12.000  -0.9370   0.06380   0.05723  -0.0348   1.0000   0.0597
 -11.750  -0.9488   0.06025   0.05350  -0.0341   1.0000   0.0591
 -11.500  -0.9625   0.05687   0.04989  -0.0326   1.0000   0.0585
 -11.250  -0.9743   0.05380   0.04656  -0.0303   1.0000   0.0579
 -11.000  -0.9788   0.05066   0.04310  -0.0283   1.0000   0.0574
 -10.750  -0.9773   0.04758   0.03968  -0.0267   1.0000   0.0569
 -10.500  -0.9700   0.04475   0.03652  -0.0253   1.0000   0.0567
 -10.250  -0.9568   0.04215   0.03365  -0.0243   1.0000   0.0568
 -10.000  -0.9386   0.03983   0.03110  -0.0236   1.0000   0.0571
  -9.750  -0.9171   0.03767   0.02875  -0.0232   1.0000   0.0578
  -9.500  -0.8930   0.03566   0.02657  -0.0230   1.0000   0.0589
  -9.250  -0.8671   0.03381   0.02453  -0.0229   1.0000   0.0603
  -9.000  -0.8372   0.03204   0.02284  -0.0233   1.0000   0.0624
  -8.750  -0.8123   0.03071   0.02153  -0.0232   1.0000   0.0658
  -8.500  -0.7878   0.02934   0.02026  -0.0230   1.0000   0.0698
  -8.250  -0.7665   0.02804   0.01901  -0.0224   1.0000   0.0757
  -8.000  -0.7496   0.02669   0.01780  -0.0213   1.0000   0.0837
  -7.750  -0.7407   0.02528   0.01663  -0.0195   1.0000   0.0946
  -7.500  -0.7418   0.02419   0.01572  -0.0162   1.0000   0.1086
  -7.250  -0.7569   0.02362   0.01525  -0.0104   1.0000   0.1187
  -7.000  -0.7664   0.02279   0.01453  -0.0057   0.9991   0.1374
  -6.750  -0.7453   0.02059   0.01301  -0.0074   0.9866   0.2085
  -6.500  -0.7252   0.01872   0.01193  -0.0086   0.9757   0.3236
  -6.250  -0.6936   0.01806   0.01185  -0.0102   0.9677   0.4339
  -6.000  -0.6607   0.01830   0.01221  -0.0111   0.9588   0.4969
  -5.750  -0.6178   0.01917   0.01308  -0.0129   0.9532   0.5432
  -5.500  -0.5869   0.02002   0.01389  -0.0126   0.9447   0.5719
  -5.250  -0.5420   0.02150   0.01538  -0.0135   0.9395   0.5933
  -5.000  -0.5057   0.02256   0.01636  -0.0136   0.9324   0.6100
  -4.750  -0.4714   0.02328   0.01697  -0.0139   0.9248   0.6262
  -4.500  -0.4226   0.02477   0.01842  -0.0147   0.9204   0.6349
  -4.250  -0.3986   0.02518   0.01874  -0.0136   0.9119   0.6484
  -4.000  -0.3597   0.02608   0.01958  -0.0137   0.9057   0.6576
  -3.750  -0.3298   0.02641   0.01981  -0.0133   0.8990   0.6692
  -3.500  -0.3044   0.02674   0.02008  -0.0123   0.8908   0.6793
  -3.250  -0.2766   0.02676   0.02001  -0.0120   0.8843   0.6885
  -3.000  -0.2506   0.02703   0.02022  -0.0112   0.8771   0.6976
  -2.750  -0.2304   0.02694   0.02007  -0.0101   0.8690   0.7076
  -2.500  -0.2007   0.02703   0.02010  -0.0097   0.8633   0.7150
  -2.250  -0.1883   0.02678   0.01979  -0.0082   0.8547   0.7251
  -2.000  -0.1596   0.02690   0.01988  -0.0077   0.8482   0.7317
  -1.750  -0.1433   0.02670   0.01961  -0.0064   0.8419   0.7424
  -1.500  -0.1189   0.02682   0.01973  -0.0057   0.8342   0.7485
  -1.250  -0.1039   0.02652   0.01935  -0.0045   0.8276   0.7585
  -1.000  -0.0781   0.02663   0.01946  -0.0039   0.8218   0.7639
  -0.750  -0.0668   0.02646   0.01926  -0.0026   0.8138   0.7739
  -0.500  -0.0374   0.02645   0.01924  -0.0021   0.8084   0.7788
  -0.250  -0.0236   0.02649   0.01928  -0.0008   0.8005   0.7879
   0.000   0.0000   0.02648   0.01928   0.0000   0.7940   0.7940
   0.250   0.0236   0.02649   0.01928   0.0008   0.7879   0.8005
   0.500   0.0374   0.02645   0.01925   0.0021   0.7788   0.8085
   0.750   0.0668   0.02646   0.01926   0.0026   0.7739   0.8137
   1.000   0.0781   0.02663   0.01947   0.0039   0.7639   0.8218
   1.250   0.1038   0.02652   0.01935   0.0045   0.7585   0.8276
   1.500   0.1189   0.02682   0.01973   0.0057   0.7485   0.8342
   1.750   0.1434   0.02669   0.01960   0.0064   0.7424   0.8419
   2.000   0.1595   0.02690   0.01988   0.0077   0.7317   0.8482
   2.250   0.1883   0.02678   0.01979   0.0082   0.7251   0.8547
   2.500   0.2008   0.02703   0.02010   0.0097   0.7150   0.8633
   2.750   0.2304   0.02693   0.02007   0.0101   0.7076   0.8690
   3.000   0.2506   0.02702   0.02022   0.0112   0.6976   0.8771
   3.250   0.2766   0.02675   0.02001   0.0120   0.6885   0.8842
   3.500   0.3045   0.02674   0.02007   0.0123   0.6794   0.8908
   3.750   0.3298   0.02640   0.01980   0.0133   0.6692   0.8990
   4.000   0.3597   0.02609   0.01959   0.0136   0.6577   0.9057
   4.250   0.3986   0.02519   0.01875   0.0136   0.6484   0.9119
   4.500   0.4227   0.02475   0.01840   0.0147   0.6349   0.9204
   4.750   0.4713   0.02328   0.01696   0.0139   0.6262   0.9248
   5.000   0.5057   0.02256   0.01635   0.0136   0.6100   0.9324
   5.250   0.5420   0.02151   0.01539   0.0135   0.5933   0.9395
   5.500   0.5869   0.02002   0.01389   0.0126   0.5719   0.9448
   5.750   0.6177   0.01916   0.01308   0.0129   0.5432   0.9533
   6.000   0.6608   0.01829   0.01219   0.0111   0.4965   0.9588
   6.250   0.6936   0.01806   0.01185   0.0102   0.4336   0.9678
   6.500   0.7252   0.01872   0.01193   0.0086   0.3231   0.9758
   6.750   0.7454   0.02059   0.01300   0.0074   0.2082   0.9867
   7.000   0.7665   0.02279   0.01454   0.0057   0.1369   0.9992
   7.250   0.7563   0.02363   0.01526   0.0105   0.1178   1.0000
   7.500   0.7418   0.02418   0.01572   0.0162   0.1084   1.0000
   7.750   0.7409   0.02528   0.01662   0.0194   0.0947   1.0000
   8.000   0.7494   0.02673   0.01782   0.0213   0.0834   1.0000
   8.250   0.7668   0.02805   0.01901   0.0224   0.0758   1.0000
   8.500   0.7883   0.02934   0.02026   0.0229   0.0700   1.0000
   8.750   0.8126   0.03072   0.02153   0.0231   0.0657   1.0000
   9.000   0.8380   0.03207   0.02287   0.0232   0.0625   1.0000
   9.250   0.8678   0.03381   0.02452   0.0228   0.0603   1.0000
   9.500   0.8933   0.03566   0.02658   0.0230   0.0589   1.0000
   9.750   0.9173   0.03767   0.02876   0.0232   0.0578   1.0000
  10.000   0.9389   0.03984   0.03111   0.0236   0.0572   1.0000
  10.250   0.9570   0.04219   0.03368   0.0242   0.0568   1.0000
  10.500   0.9702   0.04475   0.03653   0.0252   0.0568   1.0000
  10.750   0.9776   0.04760   0.03970   0.0266   0.0569   1.0000
  11.000   0.9792   0.05064   0.04309   0.0283   0.0574   1.0000
  11.250   0.9745   0.05383   0.04659   0.0302   0.0580   1.0000
  11.500   0.9631   0.05690   0.04993   0.0325   0.0586   1.0000
  11.750   0.9495   0.06026   0.05351   0.0340   0.0591   1.0000
  12.000   0.9385   0.06379   0.05722   0.0347   0.0597   1.0000
  12.250   0.9276   0.06768   0.06126   0.0348   0.0602   1.0000
  12.500   0.9333   0.07140   0.06503   0.0346   0.0608   1.0000
  12.750   0.8144   0.08326   0.07760   0.0283   0.0633   1.0000
  13.000   0.7547   0.09581   0.09037   0.0202   0.0656   1.0000
  13.250   0.7447   0.10242   0.09700   0.0171   0.0670   1.0000
<< Back to FX 71-L-150/25 AIRFOIL (fx711525-il)

Polar data table (+)

Polar graphs


<< Back to FX 71-L-150/25 AIRFOIL (fx711525-il)