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

EPPLER E662 AIRFOIL (e662-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E662 AIRFOIL (e662-il)
Reynolds number: 1,000,000
Max Cl/Cd: 138.7 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e662-il-1000000-n5.txt
Download as CSV file: xf-e662-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E662 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.750  -0.5139   0.04780   0.04444  -0.1561   0.8121   0.0075
 -14.250  -0.5511   0.03707   0.03338  -0.1625   0.8085   0.0075
 -14.000  -0.5681   0.03341   0.02955  -0.1628   0.8070   0.0076
 -13.750  -0.5642   0.03183   0.02789  -0.1625   0.8057   0.0077
 -13.500  -0.5686   0.02973   0.02565  -0.1615   0.8042   0.0078
 -13.250  -0.5662   0.02825   0.02405  -0.1604   0.8028   0.0079
 -13.000  -0.5693   0.02644   0.02210  -0.1586   0.8013   0.0080
 -12.750  -0.5610   0.02546   0.02102  -0.1574   0.8000   0.0082
 -12.500  -0.5582   0.02413   0.01956  -0.1555   0.7987   0.0082
 -12.250  -0.5515   0.02308   0.01838  -0.1538   0.7974   0.0083
 -12.000  -0.5444   0.02207   0.01725  -0.1519   0.7960   0.0084
 -11.500  -0.5250   0.02052   0.01551  -0.1481   0.7936   0.0085
 -11.250  -0.5167   0.01984   0.01474  -0.1458   0.7926   0.0086
 -11.000  -0.5073   0.01890   0.01370  -0.1437   0.7916   0.0088
 -10.750  -0.4918   0.01820   0.01292  -0.1425   0.7907   0.0090
 -10.500  -0.4730   0.01771   0.01238  -0.1416   0.7898   0.0092
 -10.250  -0.4520   0.01736   0.01200  -0.1410   0.7890   0.0094
 -10.000  -0.4314   0.01690   0.01149  -0.1403   0.7881   0.0095
  -9.750  -0.4098   0.01650   0.01105  -0.1398   0.7872   0.0097
  -9.500  -0.3884   0.01602   0.01049  -0.1392   0.7862   0.0100
  -9.250  -0.3664   0.01556   0.00998  -0.1387   0.7852   0.0102
  -9.000  -0.3441   0.01509   0.00944  -0.1382   0.7843   0.0104
  -8.750  -0.3213   0.01464   0.00893  -0.1378   0.7834   0.0106
  -8.500  -0.2981   0.01420   0.00842  -0.1374   0.7826   0.0109
  -8.250  -0.2742   0.01382   0.00796  -0.1371   0.7818   0.0111
  -8.000  -0.2497   0.01348   0.00757  -0.1369   0.7809   0.0113
  -7.750  -0.2251   0.01314   0.00717  -0.1367   0.7798   0.0115
  -7.500  -0.1999   0.01284   0.00683  -0.1366   0.7789   0.0116
  -7.250  -0.1774   0.01225   0.00620  -0.1361   0.7783   0.0122
  -7.000  -0.1523   0.01191   0.00585  -0.1360   0.7776   0.0125
  -6.750  -0.1268   0.01162   0.00554  -0.1359   0.7768   0.0129
  -6.500  -0.1010   0.01134   0.00524  -0.1358   0.7759   0.0132
  -6.250  -0.0751   0.01106   0.00494  -0.1358   0.7750   0.0135
  -6.000  -0.0489   0.01080   0.00466  -0.1358   0.7741   0.0139
  -5.750  -0.0225   0.01056   0.00439  -0.1358   0.7731   0.0142
  -5.500   0.0043   0.01034   0.00416  -0.1359   0.7721   0.0146
  -5.250   0.0313   0.01014   0.00393  -0.1360   0.7711   0.0149
  -5.000   0.0585   0.00996   0.00373  -0.1362   0.7701   0.0153
  -4.750   0.0853   0.00969   0.00345  -0.1363   0.7691   0.0162
  -4.500   0.1127   0.00951   0.00326  -0.1364   0.7681   0.0171
  -4.250   0.1404   0.00935   0.00308  -0.1367   0.7670   0.0178
  -4.000   0.1683   0.00921   0.00293  -0.1370   0.7659   0.0187
  -3.750   0.1966   0.00911   0.00280  -0.1373   0.7647   0.0196
  -3.500   0.2245   0.00896   0.00267  -0.1376   0.7636   0.0223
  -3.250   0.2520   0.00878   0.00254  -0.1378   0.7625   0.0308
  -3.000   0.2793   0.00855   0.00243  -0.1380   0.7610   0.0561
  -2.750   0.3069   0.00834   0.00233  -0.1383   0.7593   0.0829
  -2.500   0.3347   0.00813   0.00223  -0.1386   0.7575   0.1159
  -2.250   0.3627   0.00793   0.00214  -0.1390   0.7557   0.1515
  -2.000   0.3908   0.00767   0.00206  -0.1395   0.7541   0.2071
  -1.750   0.4192   0.00739   0.00197  -0.1401   0.7523   0.2733
  -1.500   0.4481   0.00707   0.00189  -0.1408   0.7503   0.3603
  -1.250   0.4775   0.00648   0.00178  -0.1420   0.7482   0.5192
  -1.000   0.5072   0.00595   0.00181  -0.1430   0.7460   0.6901
  -0.750   0.5358   0.00595   0.00187  -0.1433   0.7436   0.7237
  -0.500   0.5645   0.00596   0.00189  -0.1436   0.7411   0.7358
  -0.250   0.5932   0.00598   0.00191  -0.1438   0.7385   0.7452
   0.000   0.6221   0.00602   0.00190  -0.1442   0.7358   0.7541
   0.250   0.6501   0.00605   0.00197  -0.1443   0.7326   0.7615
   0.500   0.6785   0.00610   0.00201  -0.1446   0.7286   0.7693
   0.750   0.7061   0.00613   0.00206  -0.1446   0.7245   0.7749
   1.000   0.7337   0.00618   0.00209  -0.1446   0.7204   0.7792
   1.250   0.7618   0.00620   0.00213  -0.1449   0.7151   0.7821
   1.500   0.7894   0.00622   0.00213  -0.1450   0.7088   0.7840
   1.750   0.8169   0.00625   0.00214  -0.1451   0.7014   0.7854
   2.000   0.8433   0.00629   0.00214  -0.1449   0.6918   0.7867
   2.250   0.8686   0.00635   0.00216  -0.1446   0.6794   0.7880
   2.750   0.9154   0.00660   0.00229  -0.1431   0.6463   0.7904
   3.000   0.9366   0.00680   0.00240  -0.1419   0.6256   0.7916
   3.250   0.9559   0.00705   0.00256  -0.1403   0.6033   0.7929
   3.500   0.9744   0.00732   0.00274  -0.1386   0.5794   0.7941
   3.750   0.9915   0.00762   0.00293  -0.1366   0.5560   0.7954
   4.000   1.0068   0.00794   0.00315  -0.1343   0.5322   0.7968
   4.250   1.0198   0.00824   0.00335  -0.1314   0.5100   0.7983
   4.500   1.0315   0.00860   0.00362  -0.1284   0.4861   0.7997
   4.750   1.0440   0.00900   0.00391  -0.1256   0.4621   0.8009
   5.000   1.0574   0.00939   0.00421  -0.1231   0.4411   0.8019
   5.250   1.0711   0.00978   0.00453  -0.1206   0.4215   0.8031
   5.500   1.0837   0.01022   0.00489  -0.1180   0.3993   0.8045
   5.750   1.0960   0.01069   0.00528  -0.1154   0.3785   0.8058
   6.000   1.1084   0.01119   0.00570  -0.1129   0.3575   0.8071
   6.250   1.1200   0.01174   0.00617  -0.1104   0.3354   0.8083
   6.500   1.1316   0.01233   0.00667  -0.1079   0.3147   0.8096
   6.750   1.1453   0.01288   0.00716  -0.1059   0.2972   0.8108
   7.000   1.1581   0.01349   0.00770  -0.1038   0.2779   0.8121
   7.250   1.1713   0.01413   0.00826  -0.1018   0.2600   0.8134
   7.500   1.1835   0.01483   0.00888  -0.0997   0.2409   0.8146
   7.750   1.1964   0.01553   0.00951  -0.0978   0.2216   0.8158
   8.000   1.2087   0.01629   0.01017  -0.0959   0.2021   0.8171
   8.250   1.2203   0.01710   0.01089  -0.0940   0.1825   0.8184
   8.500   1.2322   0.01791   0.01162  -0.0921   0.1628   0.8197
   8.750   1.2422   0.01885   0.01245  -0.0900   0.1410   0.8209
   9.000   1.2556   0.01963   0.01318  -0.0884   0.1275   0.8221
   9.500   1.2825   0.02123   0.01469  -0.0854   0.1026   0.8246
   9.750   1.2943   0.02216   0.01556  -0.0837   0.0887   0.8259
  10.000   1.3077   0.02301   0.01637  -0.0824   0.0783   0.8271
  10.250   1.3203   0.02392   0.01725  -0.0809   0.0681   0.8284
  10.500   1.3339   0.02479   0.01810  -0.0796   0.0602   0.8297
  10.750   1.3458   0.02579   0.01906  -0.0782   0.0511   0.8309
  11.000   1.3572   0.02685   0.02008  -0.0767   0.0431   0.8321
  11.250   1.3693   0.02787   0.02108  -0.0754   0.0366   0.8333
  11.500   1.3819   0.02886   0.02208  -0.0741   0.0317   0.8348
  11.750   1.3928   0.02999   0.02322  -0.0727   0.0262   0.8363
  12.000   1.4048   0.03107   0.02431  -0.0715   0.0228   0.8378
  12.250   1.4164   0.03221   0.02547  -0.0703   0.0192   0.8393
  12.500   1.4264   0.03350   0.02675  -0.0690   0.0151   0.8408
  12.750   1.4363   0.03482   0.02809  -0.0678   0.0114   0.8423
  13.250   1.4541   0.03775   0.03106  -0.0654   0.0065   0.8452
  13.500   1.4637   0.03919   0.03255  -0.0643   0.0057   0.8466
  14.000   1.4833   0.04212   0.03560  -0.0624   0.0048   0.8496
  14.250   1.4919   0.04375   0.03728  -0.0614   0.0044   0.8513
  14.500   1.5001   0.04545   0.03905  -0.0605   0.0042   0.8530
  14.750   1.5077   0.04724   0.04092  -0.0597   0.0039   0.8548
  15.000   1.5160   0.04901   0.04276  -0.0589   0.0037   0.8567
  15.250   1.5235   0.05088   0.04471  -0.0583   0.0036   0.8586
  15.500   1.5306   0.05285   0.04676  -0.0577   0.0035   0.8605
  15.750   1.5370   0.05494   0.04893  -0.0571   0.0034   0.8623
  16.000   1.5424   0.05718   0.05126  -0.0566   0.0033   0.8645
  16.250   1.5472   0.05953   0.05369  -0.0562   0.0032   0.8668
  16.500   1.5511   0.06203   0.05628  -0.0558   0.0031   0.8693
  16.750   1.5540   0.06472   0.05906  -0.0556   0.0030   0.8718
  17.000   1.5554   0.06767   0.06210  -0.0555   0.0029   0.8741
  17.250   1.5567   0.07072   0.06524  -0.0555   0.0028   0.8763
  17.500   1.5567   0.07399   0.06863  -0.0556   0.0027   0.8787
  17.750   1.5541   0.07768   0.07244  -0.0559   0.0026   0.8812
  18.000   1.5512   0.08150   0.07638  -0.0564   0.0026   0.8838
  18.250   1.5474   0.08557   0.08056  -0.0571   0.0025   0.8865
  18.500   1.5461   0.08934   0.08444  -0.0579   0.0025   0.8895
  19.000   1.5391   0.09771   0.09304  -0.0601   0.0025   0.8958
  19.250   1.5348   0.10210   0.09756  -0.0615   0.0024   0.8995
<< Back to EPPLER E662 AIRFOIL (e662-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E662 AIRFOIL (e662-il)