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E184 (8.33%) (e184-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: E184 (8.33%) (e184-il)
Reynolds number: 500,000
Max Cl/Cd: 73.85 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e184-il-500000.txt
Download as CSV file: xf-e184-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E184  (8.33%)                                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6757   0.08234   0.08052   0.0187   1.0000   0.0186
  -8.250  -0.6884   0.07662   0.07484   0.0135   1.0000   0.0185
  -8.000  -0.6955   0.07005   0.06822   0.0086   1.0000   0.0185
  -7.750  -0.6950   0.06435   0.06244   0.0053   1.0000   0.0192
  -7.500  -0.6922   0.05830   0.05625   0.0027   1.0000   0.0200
  -7.250  -0.6767   0.05281   0.05045   0.0012   1.0000   0.0228
  -7.000  -0.6638   0.04880   0.04608   0.0005   0.9600   0.0231
  -6.750  -0.6791   0.03927   0.03594   0.0025   0.9074   0.0241
  -6.500  -0.6679   0.03638   0.03288   0.0038   0.8843   0.0251
  -6.250  -0.6510   0.03459   0.03092   0.0048   0.8663   0.0263
  -6.000  -0.6330   0.03249   0.02856   0.0058   0.8511   0.0283
  -5.750  -0.6038   0.03436   0.03008   0.0073   0.8371   0.0330
  -5.500  -0.5966   0.02234   0.01705   0.0104   0.8280   0.0208
  -5.250  -0.5737   0.02007   0.01440   0.0116   0.8176   0.0208
  -5.000  -0.5496   0.01751   0.01145   0.0127   0.8076   0.0197
  -4.750  -0.5245   0.01574   0.00942   0.0136   0.7982   0.0194
  -4.500  -0.4992   0.01454   0.00803   0.0144   0.7899   0.0197
  -4.250  -0.4736   0.01359   0.00693   0.0151   0.7809   0.0202
  -4.000  -0.4481   0.01281   0.00603   0.0157   0.7730   0.0208
  -3.750  -0.4223   0.01218   0.00528   0.0163   0.7652   0.0215
  -3.500  -0.3961   0.01170   0.00472   0.0167   0.7577   0.0225
  -3.250  -0.3700   0.01120   0.00411   0.0172   0.7506   0.0230
  -3.000  -0.3439   0.01060   0.00337   0.0177   0.7436   0.0249
  -2.750  -0.3172   0.01022   0.00293   0.0182   0.7369   0.0296
  -2.500  -0.2913   0.00965   0.00254   0.0186   0.7304   0.0750
  -2.250  -0.2662   0.00904   0.00228   0.0189   0.7239   0.1683
  -2.000  -0.2414   0.00846   0.00208   0.0193   0.7179   0.2799
  -1.750  -0.2169   0.00787   0.00193   0.0197   0.7114   0.4011
  -1.500  -0.1951   0.00716   0.00180   0.0207   0.7061   0.5652
  -1.250  -0.1747   0.00647   0.00171   0.0224   0.6997   0.7215
  -1.000  -0.1294   0.00596   0.00184   0.0196   0.6942   0.9265
  -0.750  -0.0755   0.00619   0.00200   0.0144   0.6883   0.9680
  -0.500  -0.0332   0.00633   0.00205   0.0115   0.6821   0.9815
  -0.250   0.0140   0.00642   0.00203   0.0074   0.6765   0.9901
   0.000   0.0602   0.00643   0.00199   0.0034   0.6699   0.9972
   0.500   0.1212   0.00641   0.00186   0.0022   0.6572   1.0000
   0.750   0.1467   0.00642   0.00179   0.0027   0.6504   1.0000
   1.000   0.1727   0.00640   0.00176   0.0030   0.6423   1.0000
   1.250   0.1985   0.00641   0.00172   0.0035   0.6350   1.0000
   1.500   0.2246   0.00640   0.00170   0.0038   0.6260   1.0000
   1.750   0.2508   0.00642   0.00169   0.0042   0.6181   1.0000
   2.000   0.2771   0.00643   0.00168   0.0045   0.6096   1.0000
   2.250   0.3035   0.00644   0.00170   0.0049   0.5997   1.0000
   2.500   0.3298   0.00647   0.00172   0.0052   0.5903   1.0000
   2.750   0.3562   0.00650   0.00174   0.0055   0.5804   1.0000
   3.000   0.3828   0.00653   0.00178   0.0058   0.5685   1.0000
   3.250   0.4093   0.00657   0.00183   0.0061   0.5553   1.0000
   3.500   0.4359   0.00662   0.00190   0.0064   0.5411   1.0000
   3.750   0.4624   0.00670   0.00197   0.0067   0.5241   1.0000
   4.000   0.4891   0.00679   0.00205   0.0070   0.5000   1.0000
   4.250   0.5155   0.00698   0.00213   0.0072   0.4568   1.0000
   4.500   0.5417   0.00749   0.00232   0.0071   0.3675   1.0000
   4.750   0.5674   0.00839   0.00272   0.0068   0.2536   1.0000
   5.000   0.5926   0.00936   0.00321   0.0064   0.1508   1.0000
   5.250   0.6175   0.01019   0.00370   0.0063   0.0811   1.0000
   5.500   0.6416   0.01148   0.00462   0.0064   0.0163   1.0000
   5.750   0.6669   0.01199   0.00523   0.0068   0.0144   1.0000
   6.000   0.6916   0.01260   0.00593   0.0072   0.0133   1.0000
   6.250   0.7157   0.01332   0.00674   0.0077   0.0125   1.0000
   6.500   0.7382   0.01434   0.00784   0.0082   0.0114   1.0000
   6.750   0.7581   0.01587   0.00949   0.0091   0.0107   1.0000
   7.000   0.7801   0.01678   0.01051   0.0099   0.0105   1.0000
   7.250   0.8015   0.01784   0.01169   0.0108   0.0101   1.0000
   7.500   0.8219   0.01920   0.01318   0.0119   0.0099   1.0000
   7.750   0.8416   0.02085   0.01498   0.0131   0.0097   1.0000
   8.000   0.8603   0.02299   0.01733   0.0144   0.0098   1.0000
   8.250   0.8764   0.02601   0.02065   0.0158   0.0102   1.0000
   8.500   0.8945   0.02849   0.02327   0.0170   0.0112   1.0000
  12.500   0.5950   0.12240   0.12020  -0.0059   0.0221   1.0000
  12.750   0.5909   0.12763   0.12541  -0.0082   0.0214   1.0000
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