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

WORTMANN FX 05-H-126 AIRFOIL (fx05h126-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: WORTMANN FX 05-H-126 AIRFOIL (fx05h126-il)
Reynolds number: 200,000
Max Cl/Cd: 70.81 at α=8.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx05h126-il-200000-n5.txt
Download as CSV file: xf-fx05h126-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX 05-H-126 AIRFOIL                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3896   0.09178   0.08729  -0.0022   0.6089   0.0312
  -8.000  -0.3969   0.08730   0.08286  -0.0073   0.6081   0.0328
  -7.750  -0.4072   0.08283   0.07841  -0.0126   0.6072   0.0331
  -7.500  -0.4079   0.07779   0.07329  -0.0176   0.6064   0.0333
  -7.000  -0.3981   0.06866   0.06386  -0.0226   0.6047   0.0335
  -6.750  -0.3929   0.06404   0.05924  -0.0231   0.6039   0.0342
  -6.500  -0.3792   0.06209   0.05732  -0.0229   0.6028   0.0351
  -6.250  -0.3712   0.05319   0.04801  -0.0251   0.6020   0.0255
  -6.000  -0.3573   0.04987   0.04458  -0.0256   0.6009   0.0249
  -5.750  -0.3420   0.04605   0.04054  -0.0260   0.5998   0.0244
  -5.500  -0.3252   0.04208   0.03628  -0.0260   0.5988   0.0239
  -5.250  -0.3071   0.03817   0.03202  -0.0256   0.5977   0.0236
  -5.000  -0.2875   0.03460   0.02805  -0.0250   0.5967   0.0235
  -4.750  -0.2659   0.03172   0.02476  -0.0243   0.5956   0.0243
  -4.500  -0.2431   0.02920   0.02180  -0.0236   0.5946   0.0249
  -4.250  -0.2193   0.02705   0.01927  -0.0229   0.5936   0.0249
  -4.000  -0.1944   0.02525   0.01710  -0.0223   0.5927   0.0249
  -3.750  -0.1686   0.02376   0.01527  -0.0218   0.5917   0.0252
  -3.500  -0.1421   0.02256   0.01378  -0.0214   0.5907   0.0254
  -3.250  -0.1152   0.02159   0.01256  -0.0211   0.5897   0.0258
  -3.000  -0.0881   0.02074   0.01154  -0.0208   0.5889   0.0261
  -2.750  -0.0615   0.01976   0.01051  -0.0207   0.5880   0.0274
  -2.500  -0.0340   0.01928   0.01005  -0.0209   0.5867   0.0288
  -2.250  -0.0066   0.01877   0.00952  -0.0209   0.5854   0.0296
  -2.000   0.0204   0.01831   0.00903  -0.0208   0.5842   0.0303
  -1.750   0.0472   0.01792   0.00862  -0.0206   0.5829   0.0312
  -1.500   0.0740   0.01761   0.00828  -0.0205   0.5816   0.0322
  -1.250   0.1009   0.01735   0.00798  -0.0204   0.5802   0.0333
  -1.000   0.1275   0.01708   0.00768  -0.0202   0.5790   0.0354
  -0.750   0.1546   0.01694   0.00753  -0.0201   0.5779   0.0395
  -0.500   0.1816   0.01681   0.00739  -0.0200   0.5769   0.0448
  -0.250   0.2085   0.01666   0.00727  -0.0198   0.5761   0.0603
   0.000   0.2343   0.01637   0.00718  -0.0196   0.5752   0.1213
   0.250   0.2492   0.01490   0.00714  -0.0177   0.5743   0.5407
   1.250   0.4578   0.01465   0.00785  -0.0382   0.5669   1.0000
   1.500   0.4837   0.01477   0.00791  -0.0380   0.5648   1.0000
   1.750   0.5095   0.01486   0.00794  -0.0377   0.5631   1.0000
   2.000   0.5353   0.01497   0.00799  -0.0374   0.5618   1.0000
   2.250   0.5611   0.01506   0.00803  -0.0370   0.5606   1.0000
   2.500   0.5869   0.01516   0.00807  -0.0367   0.5595   1.0000
   2.750   0.6128   0.01530   0.00817  -0.0364   0.5583   1.0000
   3.000   0.6395   0.01582   0.00880  -0.0370   0.5550   1.0000
   3.250   0.6656   0.01616   0.00918  -0.0372   0.5524   1.0000
   3.500   0.6916   0.01638   0.00941  -0.0371   0.5502   1.0000
   3.750   0.7175   0.01654   0.00957  -0.0370   0.5485   1.0000
   4.000   0.7434   0.01665   0.00970  -0.0367   0.5470   1.0000
   4.250   0.7694   0.01671   0.00975  -0.0364   0.5457   1.0000
   4.500   0.7954   0.01674   0.00977  -0.0360   0.5445   1.0000
   4.750   0.8207   0.01731   0.01047  -0.0364   0.5410   1.0000
   5.000   0.8459   0.01778   0.01104  -0.0366   0.5373   1.0000
   5.250   0.8715   0.01795   0.01126  -0.0365   0.5347   1.0000
   5.500   0.8975   0.01799   0.01133  -0.0362   0.5328   1.0000
   5.750   0.9238   0.01793   0.01132  -0.0358   0.5312   1.0000
   6.000   0.9506   0.01771   0.01110  -0.0353   0.5297   1.0000
   6.250   0.9742   0.01824   0.01182  -0.0357   0.5215   1.0000
   6.500   1.0023   0.01753   0.01112  -0.0350   0.5173   1.0000
   6.750   1.0276   0.01756   0.01127  -0.0349   0.5099   1.0000
   7.000   1.0545   0.01725   0.01101  -0.0346   0.5039   1.0000
   7.250   1.0797   0.01729   0.01118  -0.0346   0.4944   1.0000
   7.500   1.1065   0.01698   0.01094  -0.0343   0.4862   1.0000
   7.750   1.1312   0.01714   0.01122  -0.0343   0.4752   1.0000
   8.000   1.1558   0.01724   0.01141  -0.0342   0.4611   1.0000
   8.250   1.1801   0.01733   0.01155  -0.0340   0.4415   1.0000
   8.500   1.2053   0.01707   0.01122  -0.0334   0.4176   1.0000
   8.750   1.2257   0.01731   0.01127  -0.0326   0.3862   1.0000
   9.000   1.2384   0.01837   0.01214  -0.0317   0.3512   1.0000
   9.250   1.2416   0.02031   0.01393  -0.0311   0.3131   1.0000
   9.500   1.2294   0.02316   0.01662  -0.0295   0.2813   1.0000
   9.750   1.2148   0.02592   0.01921  -0.0272   0.2536   1.0000
  10.000   1.2031   0.02871   0.02183  -0.0254   0.2251   1.0000
  10.250   1.1924   0.03152   0.02447  -0.0239   0.1973   1.0000
  10.500   1.1844   0.03419   0.02701  -0.0225   0.1741   1.0000
  10.750   1.1807   0.03657   0.02929  -0.0214   0.1543   1.0000
  11.000   1.1780   0.03891   0.03155  -0.0204   0.1353   1.0000
  11.250   1.1770   0.04118   0.03373  -0.0195   0.1179   1.0000
  11.500   1.1769   0.04346   0.03592  -0.0188   0.1025   1.0000
  11.750   1.1786   0.04568   0.03809  -0.0182   0.0927   1.0000
  12.000   1.1818   0.04780   0.04019  -0.0176   0.0845   1.0000
  12.250   1.1860   0.04987   0.04227  -0.0172   0.0779   1.0000
  12.500   1.1908   0.05192   0.04434  -0.0168   0.0724   1.0000
  12.750   1.1960   0.05397   0.04643  -0.0165   0.0669   1.0000
  13.000   1.2032   0.05584   0.04838  -0.0162   0.0623   1.0000
  13.250   1.2071   0.05809   0.05063  -0.0160   0.0577   1.0000
  13.500   1.2156   0.05988   0.05252  -0.0158   0.0535   1.0000
  13.750   1.2218   0.06194   0.05463  -0.0156   0.0488   1.0000
  14.000   1.2270   0.06413   0.05686  -0.0156   0.0444   1.0000
  14.250   1.2334   0.06624   0.05903  -0.0155   0.0391   1.0000
  14.500   1.2378   0.06860   0.06144  -0.0155   0.0338   1.0000
  14.750   1.2410   0.07113   0.06399  -0.0155   0.0282   1.0000
  15.000   1.2431   0.07382   0.06671  -0.0156   0.0241   1.0000
  15.250   1.2441   0.07671   0.06964  -0.0157   0.0204   1.0000
  15.500   1.2438   0.07982   0.07277  -0.0160   0.0182   1.0000
  15.750   1.2443   0.08286   0.07591  -0.0163   0.0164   1.0000
  16.000   1.2431   0.08619   0.07930  -0.0167   0.0150   1.0000
  16.250   1.2416   0.08963   0.08281  -0.0172   0.0140   1.0000
  16.500   1.2409   0.09304   0.08635  -0.0178   0.0133   1.0000
  16.750   1.2398   0.09653   0.08996  -0.0185   0.0126   1.0000
  17.000   1.2380   0.10017   0.09372  -0.0193   0.0121   1.0000
  17.250   1.2351   0.10407   0.09775  -0.0203   0.0117   1.0000
  17.500   1.2315   0.10816   0.10195  -0.0214   0.0113   1.0000
  17.750   1.2267   0.11248   0.10639  -0.0228   0.0110   1.0000
  18.000   1.2204   0.11710   0.11111  -0.0243   0.0107   1.0000
<< Back to WORTMANN FX 05-H-126 AIRFOIL (fx05h126-il)

Polar data table (+)

Polar graphs


<< Back to WORTMANN FX 05-H-126 AIRFOIL (fx05h126-il)