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

EPPLER 1200 AIRFOIL (e1200-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 1200 AIRFOIL (e1200-il)
Reynolds number: 500,000
Max Cl/Cd: 107.43 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1200-il-500000.txt
Download as CSV file: xf-e1200-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1200 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
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.1157   0.09946   0.09671  -0.1035   0.8790   0.0369
 -13.000  -0.4498   0.05866   0.05569  -0.1195   0.9455   0.0247
 -12.750  -0.4582   0.04965   0.04636  -0.1317   0.9316   0.0240
 -12.500  -0.4711   0.04239   0.03860  -0.1399   0.9070   0.0234
 -12.250  -0.4936   0.03812   0.03390  -0.1400   0.8837   0.0230
 -12.000  -0.5097   0.03523   0.03067  -0.1382   0.8688   0.0228
 -11.750  -0.5271   0.03219   0.02723  -0.1355   0.8576   0.0224
 -11.500  -0.5292   0.03050   0.02533  -0.1333   0.8490   0.0224
 -11.250  -0.5285   0.02890   0.02348  -0.1309   0.8420   0.0225
 -11.000  -0.5249   0.02758   0.02201  -0.1286   0.8356   0.0225
 -10.750  -0.5171   0.02672   0.02103  -0.1264   0.8301   0.0229
 -10.500  -0.5051   0.02570   0.01985  -0.1248   0.8255   0.0232
 -10.250  -0.4915   0.02458   0.01860  -0.1234   0.8214   0.0235
 -10.000  -0.4752   0.02350   0.01737  -0.1223   0.8173   0.0238
  -9.750  -0.4569   0.02250   0.01623  -0.1213   0.8138   0.0240
  -9.500  -0.4371   0.02159   0.01517  -0.1205   0.8104   0.0243
  -9.250  -0.4169   0.02081   0.01428  -0.1197   0.8071   0.0246
  -9.000  -0.3961   0.02004   0.01344  -0.1190   0.8037   0.0249
  -8.750  -0.3749   0.01942   0.01273  -0.1182   0.8005   0.0251
  -8.500  -0.3535   0.01853   0.01176  -0.1175   0.7977   0.0255
  -8.250  -0.3334   0.01762   0.01082  -0.1167   0.7951   0.0260
  -8.000  -0.3130   0.01702   0.01019  -0.1159   0.7923   0.0267
  -7.750  -0.2928   0.01649   0.00966  -0.1150   0.7897   0.0273
  -7.500  -0.2720   0.01601   0.00916  -0.1142   0.7868   0.0280
  -7.250  -0.2501   0.01559   0.00871  -0.1135   0.7839   0.0288
  -7.000  -0.2275   0.01521   0.00828  -0.1129   0.7815   0.0298
  -6.750  -0.2054   0.01475   0.00778  -0.1123   0.7792   0.0305
  -6.500  -0.1844   0.01420   0.00720  -0.1115   0.7769   0.0319
  -6.250  -0.1609   0.01387   0.00685  -0.1111   0.7746   0.0332
  -6.000  -0.1371   0.01354   0.00651  -0.1107   0.7722   0.0348
  -5.750  -0.1131   0.01318   0.00614  -0.1102   0.7696   0.0368
  -5.500  -0.0884   0.01285   0.00582  -0.1100   0.7673   0.0404
  -5.250  -0.0633   0.01251   0.00550  -0.1097   0.7651   0.0475
  -5.000  -0.0396   0.01196   0.00514  -0.1094   0.7630   0.0811
  -4.750  -0.0149   0.01151   0.00488  -0.1094   0.7609   0.1280
  -4.500   0.0110   0.01121   0.00472  -0.1095   0.7587   0.1697
  -4.250   0.0358   0.01080   0.00454  -0.1094   0.7567   0.2211
  -4.000   0.0607   0.01031   0.00434  -0.1095   0.7545   0.2943
  -3.750   0.0856   0.00969   0.00411  -0.1097   0.7522   0.3957
  -3.500   0.1110   0.00912   0.00401  -0.1098   0.7499   0.5219
  -3.250   0.1392   0.00905   0.00407  -0.1099   0.7477   0.5788
  -3.000   0.1684   0.00911   0.00414  -0.1101   0.7459   0.6072
  -2.750   0.1979   0.00923   0.00423  -0.1104   0.7440   0.6268
  -2.500   0.2268   0.00939   0.00437  -0.1105   0.7419   0.6414
  -2.250   0.2548   0.00949   0.00448  -0.1105   0.7397   0.6539
  -2.000   0.2832   0.00961   0.00459  -0.1106   0.7373   0.6640
  -1.750   0.3118   0.00967   0.00464  -0.1107   0.7348   0.6705
  -1.500   0.3407   0.00975   0.00470  -0.1108   0.7324   0.6768
  -1.250   0.3705   0.00983   0.00471  -0.1112   0.7301   0.6837
  -1.000   0.3998   0.00990   0.00477  -0.1114   0.7278   0.6892
  -0.750   0.4281   0.01004   0.00490  -0.1114   0.7249   0.6965
  -0.500   0.4551   0.01008   0.00496  -0.1112   0.7212   0.7035
  -0.250   0.4828   0.01010   0.00500  -0.1110   0.7173   0.7089
   0.000   0.5123   0.01009   0.00495  -0.1113   0.7136   0.7141
   0.250   0.5426   0.01012   0.00488  -0.1118   0.7095   0.7183
   0.500   0.5685   0.01003   0.00485  -0.1113   0.7046   0.7212
   0.750   0.5966   0.00997   0.00479  -0.1114   0.7000   0.7240
   1.000   0.6260   0.00994   0.00472  -0.1117   0.6962   0.7269
   1.250   0.6551   0.00996   0.00471  -0.1120   0.6926   0.7300
   1.500   0.6823   0.00994   0.00471  -0.1120   0.6883   0.7330
   1.750   0.7106   0.00990   0.00466  -0.1121   0.6840   0.7355
   2.000   0.7392   0.00984   0.00460  -0.1123   0.6800   0.7378
   2.250   0.7671   0.00985   0.00461  -0.1123   0.6758   0.7404
   2.500   0.7935   0.00982   0.00463  -0.1121   0.6706   0.7430
   2.750   0.8214   0.00979   0.00459  -0.1122   0.6656   0.7457
   3.000   0.8501   0.00979   0.00455  -0.1124   0.6608   0.7486
   3.250   0.8760   0.00978   0.00459  -0.1122   0.6543   0.7514
   3.500   0.9027   0.00972   0.00454  -0.1120   0.6482   0.7539
   3.750   0.9285   0.00971   0.00457  -0.1116   0.6411   0.7563
   4.000   0.9542   0.00969   0.00457  -0.1112   0.6329   0.7590
   4.250   0.9795   0.00971   0.00461  -0.1108   0.6241   0.7619
   4.500   1.0046   0.00974   0.00461  -0.1103   0.6139   0.7651
   4.750   1.0285   0.00979   0.00467  -0.1097   0.6009   0.7681
   5.000   1.0508   0.00984   0.00472  -0.1086   0.5858   0.7707
   5.250   1.0711   0.00997   0.00480  -0.1072   0.5668   0.7735
   5.500   1.0890   0.01017   0.00495  -0.1054   0.5422   0.7766
   5.750   1.1032   0.01049   0.00515  -0.1029   0.5138   0.7802
   6.000   1.1126   0.01093   0.00543  -0.0996   0.4825   0.7840
   6.250   1.1158   0.01139   0.00576  -0.0951   0.4533   0.7876
   6.500   1.1202   0.01196   0.00622  -0.0910   0.4260   0.7913
   6.750   1.1272   0.01255   0.00671  -0.0875   0.4014   0.7954
   7.000   1.1337   0.01321   0.00726  -0.0841   0.3784   0.7998
   7.250   1.1404   0.01386   0.00784  -0.0809   0.3576   0.8041
   7.500   1.1485   0.01453   0.00846  -0.0780   0.3386   0.8088
   7.750   1.1575   0.01523   0.00911  -0.0754   0.3218   0.8139
   8.000   1.1670   0.01596   0.00979  -0.0731   0.3064   0.8188
   8.250   1.1760   0.01672   0.01052  -0.0707   0.2923   0.8240
   8.750   1.1986   0.01820   0.01199  -0.0668   0.2687   0.8364
   9.000   1.2098   0.01896   0.01275  -0.0649   0.2585   0.8436
   9.250   1.2202   0.01982   0.01359  -0.0631   0.2487   0.8515
   9.500   1.2333   0.02050   0.01433  -0.0616   0.2400   0.8604
   9.750   1.2427   0.02140   0.01523  -0.0596   0.2315   0.8713
  10.000   1.2562   0.02205   0.01596  -0.0582   0.2241   0.8858
  10.250   1.2630   0.02288   0.01685  -0.0557   0.2171   0.9138
  10.500   1.2813   0.02350   0.01757  -0.0554   0.2102   1.0000
  10.750   1.2944   0.02452   0.01855  -0.0545   0.2028   1.0000
  11.000   1.3098   0.02539   0.01944  -0.0538   0.1965   1.0000
  11.250   1.3236   0.02635   0.02039  -0.0529   0.1901   1.0000
  11.500   1.3356   0.02743   0.02146  -0.0519   0.1841   1.0000
  11.750   1.3507   0.02832   0.02238  -0.0512   0.1779   1.0000
  12.000   1.3609   0.02956   0.02358  -0.0501   0.1719   1.0000
  12.250   1.3757   0.03049   0.02456  -0.0494   0.1666   1.0000
  12.500   1.3881   0.03161   0.02569  -0.0485   0.1610   1.0000
  12.750   1.3979   0.03293   0.02699  -0.0475   0.1552   1.0000
  13.000   1.4122   0.03394   0.02803  -0.0469   0.1489   1.0000
  13.250   1.4207   0.03542   0.02949  -0.0460   0.1431   1.0000
  13.500   1.4342   0.03653   0.03065  -0.0454   0.1375   1.0000
  13.750   1.4440   0.03796   0.03207  -0.0446   0.1316   1.0000
  14.000   1.4536   0.03942   0.03355  -0.0438   0.1265   1.0000
  14.250   1.4654   0.04072   0.03489  -0.0433   0.1212   1.0000
  14.500   1.4723   0.04248   0.03665  -0.0425   0.1165   1.0000
  14.750   1.4832   0.04392   0.03814  -0.0420   0.1122   1.0000
  15.000   1.4926   0.04552   0.03977  -0.0415   0.1078   1.0000
  15.250   1.4978   0.04754   0.04178  -0.0408   0.1031   1.0000
  15.500   1.5076   0.04916   0.04347  -0.0405   0.0992   1.0000
  15.750   1.5158   0.05096   0.04531  -0.0401   0.0950   1.0000
  16.000   1.5186   0.05336   0.04771  -0.0396   0.0906   1.0000
  16.250   1.5278   0.05513   0.04956  -0.0394   0.0874   1.0000
  16.500   1.5341   0.05724   0.05172  -0.0392   0.0836   1.0000
  16.750   1.5364   0.05983   0.05433  -0.0390   0.0800   1.0000
  17.000   1.5415   0.06217   0.05674  -0.0389   0.0769   1.0000
  17.250   1.5471   0.06448   0.05911  -0.0390   0.0737   1.0000
  17.500   1.5482   0.06738   0.06204  -0.0391   0.0702   1.0000
  17.750   1.5479   0.07050   0.06522  -0.0393   0.0675   1.0000
  18.000   1.5529   0.07303   0.06784  -0.0396   0.0650   1.0000
  18.250   1.5542   0.07608   0.07096  -0.0400   0.0622   1.0000
  18.500   1.5512   0.07977   0.07469  -0.0407   0.0596   1.0000
  18.750   1.5500   0.08327   0.07828  -0.0414   0.0572   1.0000
  19.000   1.5506   0.08658   0.08168  -0.0422   0.0548   1.0000
  19.250   1.5474   0.09049   0.08565  -0.0433   0.0524   1.0000
<< Back to EPPLER 1200 AIRFOIL (e1200-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 1200 AIRFOIL (e1200-il)