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

EPPLER 393 AIRFOIL (e393-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 393 AIRFOIL (e393-il)
Reynolds number: 500,000
Max Cl/Cd: 124.74 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e393-il-500000.txt
Download as CSV file: xf-e393-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 393 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3248   0.09804   0.09590  -0.0422   1.0000   0.0177
  -9.250  -0.3266   0.09531   0.09319  -0.0420   1.0000   0.0182
  -9.000  -0.3305   0.09264   0.09056  -0.0415   1.0000   0.0187
  -8.750  -0.3276   0.08880   0.08675  -0.0438   0.9990   0.0193
  -8.500  -0.3131   0.08291   0.08087  -0.0509   0.9958   0.0205
  -8.250  -0.2954   0.07613   0.07410  -0.0603   0.9909   0.0219
  -6.750  -0.2061   0.02185   0.01763  -0.1199   0.9201   0.0145
  -6.500  -0.1796   0.01858   0.01384  -0.1209   0.9063   0.0141
  -6.250  -0.1523   0.01657   0.01145  -0.1212   0.8915   0.0144
  -6.000  -0.1252   0.01532   0.00992  -0.1212   0.8762   0.0150
  -5.750  -0.0983   0.01454   0.00891  -0.1211   0.8610   0.0155
  -5.500  -0.0737   0.01287   0.00700  -0.1208   0.8463   0.0164
  -5.250  -0.0477   0.01213   0.00614  -0.1206   0.8323   0.0178
  -5.000  -0.0209   0.01179   0.00569  -0.1205   0.8190   0.0199
  -4.750   0.0061   0.01138   0.00516  -0.1203   0.8066   0.0218
  -4.500   0.0329   0.01052   0.00417  -0.1203   0.7953   0.0262
  -4.250   0.0604   0.01022   0.00376  -0.1203   0.7847   0.0315
  -4.000   0.0880   0.00986   0.00331  -0.1203   0.7745   0.0403
  -3.750   0.1159   0.00949   0.00296  -0.1204   0.7648   0.0582
  -3.500   0.1436   0.00910   0.00268  -0.1207   0.7560   0.1029
  -3.250   0.1714   0.00875   0.00249  -0.1209   0.7473   0.1619
  -3.000   0.1994   0.00840   0.00233  -0.1213   0.7392   0.2331
  -2.750   0.2273   0.00807   0.00219  -0.1216   0.7316   0.3190
  -2.500   0.2553   0.00781   0.00215  -0.1219   0.7239   0.4024
  -2.250   0.2833   0.00770   0.00215  -0.1220   0.7171   0.4714
  -2.000   0.3115   0.00765   0.00214  -0.1221   0.7100   0.5113
  -1.750   0.3396   0.00765   0.00212  -0.1221   0.7034   0.5369
  -1.500   0.3678   0.00765   0.00211  -0.1221   0.6968   0.5580
  -1.250   0.3959   0.00766   0.00211  -0.1221   0.6906   0.5812
  -1.000   0.4240   0.00767   0.00212  -0.1222   0.6846   0.6016
  -0.750   0.4521   0.00767   0.00213  -0.1222   0.6782   0.6203
  -0.500   0.4801   0.00772   0.00215  -0.1222   0.6728   0.6396
  -0.250   0.5081   0.00771   0.00218  -0.1222   0.6667   0.6588
   0.000   0.5360   0.00773   0.00221  -0.1222   0.6611   0.6777
   0.250   0.5639   0.00775   0.00226  -0.1222   0.6555   0.6984
   0.500   0.5915   0.00775   0.00231  -0.1221   0.6498   0.7200
   1.000   0.6463   0.00777   0.00243  -0.1219   0.6388   0.7686
   1.250   0.6730   0.00778   0.00249  -0.1216   0.6335   0.7956
   1.500   0.6992   0.00780   0.00258  -0.1211   0.6282   0.8254
   1.750   0.7240   0.00777   0.00265  -0.1204   0.6225   0.8599
   2.000   0.7460   0.00775   0.00269  -0.1189   0.6173   0.9025
   2.250   0.7713   0.00764   0.00267  -0.1181   0.6113   0.9830
   2.500   0.8005   0.00772   0.00273  -0.1185   0.6048   1.0000
   2.750   0.8289   0.00783   0.00281  -0.1187   0.5980   1.0000
   3.000   0.8572   0.00792   0.00288  -0.1189   0.5913   1.0000
   3.250   0.8852   0.00803   0.00297  -0.1190   0.5846   1.0000
   3.500   0.9130   0.00812   0.00307  -0.1191   0.5769   1.0000
   3.750   0.9406   0.00823   0.00317  -0.1191   0.5697   1.0000
   4.000   0.9681   0.00834   0.00328  -0.1191   0.5621   1.0000
   4.250   0.9954   0.00844   0.00339  -0.1191   0.5535   1.0000
   4.500   1.0223   0.00859   0.00352  -0.1190   0.5450   1.0000
   4.750   1.0494   0.00868   0.00366  -0.1189   0.5358   1.0000
   5.000   1.0762   0.00881   0.00380  -0.1188   0.5265   1.0000
   5.250   1.1023   0.00897   0.00395  -0.1185   0.5154   1.0000
   5.500   1.1284   0.00912   0.00412  -0.1183   0.5036   1.0000
   5.750   1.1545   0.00927   0.00430  -0.1180   0.4913   1.0000
   6.000   1.1800   0.00946   0.00450  -0.1177   0.4771   1.0000
   6.250   1.2046   0.00969   0.00471  -0.1172   0.4600   1.0000
   6.500   1.2281   0.00999   0.00497  -0.1165   0.4340   1.0000
   6.750   1.2491   0.01044   0.00528  -0.1154   0.3953   1.0000
   7.000   1.2671   0.01113   0.00573  -0.1139   0.3443   1.0000
   7.250   1.2826   0.01204   0.00634  -0.1121   0.2879   1.0000
   7.500   1.2966   0.01305   0.00705  -0.1101   0.2346   1.0000
   7.750   1.3106   0.01404   0.00777  -0.1081   0.1875   1.0000
   8.000   1.3244   0.01498   0.00852  -0.1061   0.1490   1.0000
   8.250   1.3382   0.01587   0.00924  -0.1041   0.1181   1.0000
   8.500   1.3517   0.01672   0.00996  -0.1020   0.0941   1.0000
   8.750   1.3620   0.01757   0.01070  -0.0993   0.0736   1.0000
   9.000   1.3709   0.01846   0.01149  -0.0965   0.0556   1.0000
   9.250   1.3780   0.01947   0.01241  -0.0935   0.0386   1.0000
   9.500   1.3786   0.02091   0.01370  -0.0897   0.0192   1.0000
   9.750   1.3814   0.02231   0.01509  -0.0865   0.0126   1.0000
  10.000   1.3875   0.02357   0.01643  -0.0838   0.0107   1.0000
  10.250   1.3954   0.02476   0.01773  -0.0816   0.0098   1.0000
  10.500   1.4015   0.02615   0.01919  -0.0794   0.0089   1.0000
  10.750   1.4042   0.02787   0.02101  -0.0770   0.0084   1.0000
  11.000   1.4035   0.02999   0.02326  -0.0747   0.0080   1.0000
  11.250   1.4051   0.03205   0.02543  -0.0728   0.0078   1.0000
  11.500   1.4089   0.03402   0.02751  -0.0713   0.0076   1.0000
  11.750   1.4111   0.03622   0.02982  -0.0699   0.0075   1.0000
  12.000   1.4129   0.03857   0.03228  -0.0687   0.0073   1.0000
  12.250   1.4137   0.04110   0.03493  -0.0676   0.0072   1.0000
  12.500   1.4142   0.04375   0.03770  -0.0667   0.0070   1.0000
  12.750   1.4152   0.04645   0.04051  -0.0659   0.0069   1.0000
  13.000   1.4155   0.04929   0.04346  -0.0652   0.0068   1.0000
  13.250   1.4168   0.05211   0.04640  -0.0647   0.0066   1.0000
  13.500   1.4186   0.05493   0.04935  -0.0644   0.0064   1.0000
  13.750   1.4195   0.05794   0.05246  -0.0643   0.0062   1.0000
  14.000   1.4203   0.06106   0.05568  -0.0642   0.0061   1.0000
  14.250   1.4202   0.06437   0.05911  -0.0643   0.0059   1.0000
  14.500   1.4193   0.06788   0.06275  -0.0645   0.0059   1.0000
  14.750   1.4180   0.07155   0.06654  -0.0648   0.0058   1.0000
  15.000   1.4161   0.07542   0.07054  -0.0652   0.0057   1.0000
  15.250   1.4139   0.07942   0.07468  -0.0658   0.0057   1.0000
  15.500   1.4104   0.08370   0.07910  -0.0667   0.0056   1.0000
  15.750   1.4065   0.08820   0.08375  -0.0677   0.0056   1.0000
  16.000   1.4013   0.09300   0.08870  -0.0690   0.0055   1.0000
  16.250   1.3951   0.09809   0.09396  -0.0706   0.0055   1.0000
  16.500   1.3878   0.10350   0.09954  -0.0726   0.0055   1.0000
  16.750   1.3796   0.10924   0.10545  -0.0749   0.0055   1.0000
  17.000   1.3701   0.11537   0.11175  -0.0776   0.0055   1.0000
  17.250   1.3603   0.12172   0.11828  -0.0807   0.0055   1.0000
  17.500   1.3496   0.12845   0.12518  -0.0842   0.0055   1.0000
  17.750   1.3384   0.13552   0.13243  -0.0882   0.0055   1.0000
  18.000   1.3263   0.14302   0.14011  -0.0927   0.0055   1.0000
  18.250   1.3134   0.15103   0.14830  -0.0978   0.0056   1.0000
<< Back to EPPLER 393 AIRFOIL (e393-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 393 AIRFOIL (e393-il)