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

RAE 100 AIRFOIL (rae100-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAE 100 AIRFOIL (rae100-il)
Reynolds number: 50,000
Max Cl/Cd: 26.08 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rae100-il-50000-n5.txt
Download as CSV file: xf-rae100-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 100 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.7759   0.08501   0.07745  -0.0146   1.0000   0.0610
 -11.000  -0.7996   0.07737   0.06975  -0.0199   1.0000   0.0606
 -10.750  -0.8237   0.07108   0.06335  -0.0236   1.0000   0.0602
 -10.500  -0.8482   0.06597   0.05808  -0.0250   1.0000   0.0601
 -10.250  -0.8691   0.06161   0.05350  -0.0245   1.0000   0.0603
 -10.000  -0.8835   0.05722   0.04874  -0.0237   1.0000   0.0608
  -9.750  -0.8870   0.05335   0.04451  -0.0228   1.0000   0.0616
  -9.500  -0.8783   0.05065   0.04171  -0.0219   1.0000   0.0632
  -9.250  -0.8693   0.04823   0.03912  -0.0209   1.0000   0.0659
  -9.000  -0.8610   0.04542   0.03597  -0.0198   1.0000   0.0690
  -8.750  -0.8515   0.04233   0.03232  -0.0185   1.0000   0.0720
  -8.500  -0.8354   0.03993   0.02978  -0.0176   1.0000   0.0750
  -8.250  -0.8184   0.03814   0.02787  -0.0166   1.0000   0.0797
  -8.000  -0.8008   0.03594   0.02515  -0.0155   1.0000   0.0853
  -7.750  -0.7811   0.03426   0.02357  -0.0146   1.0000   0.0903
  -7.500  -0.7612   0.03260   0.02157  -0.0136   1.0000   0.0982
  -7.250  -0.7412   0.03114   0.02019  -0.0127   1.0000   0.1057
  -7.000  -0.7207   0.02971   0.01864  -0.0116   1.0000   0.1157
  -6.750  -0.7000   0.02840   0.01723  -0.0106   1.0000   0.1277
  -6.500  -0.6797   0.02719   0.01602  -0.0096   1.0000   0.1422
  -6.250  -0.6597   0.02603   0.01492  -0.0085   1.0000   0.1592
  -6.000  -0.6394   0.02496   0.01386  -0.0074   1.0000   0.1812
  -5.750  -0.6193   0.02394   0.01295  -0.0063   1.0000   0.2066
  -5.500  -0.5991   0.02301   0.01213  -0.0052   1.0000   0.2370
  -5.250  -0.5786   0.02219   0.01141  -0.0041   1.0000   0.2723
  -5.000  -0.5583   0.02145   0.01081  -0.0029   1.0000   0.3095
  -4.750  -0.5376   0.02082   0.01027  -0.0016   1.0000   0.3493
  -4.500  -0.5170   0.02026   0.00985  -0.0002   1.0000   0.3885
  -4.250  -0.4960   0.01977   0.00946   0.0012   1.0000   0.4286
  -4.000  -0.4753   0.01933   0.00912   0.0027   1.0000   0.4669
  -3.750  -0.4548   0.01895   0.00882   0.0044   1.0000   0.5050
  -3.500  -0.4343   0.01860   0.00858   0.0062   1.0000   0.5410
  -3.250  -0.4138   0.01829   0.00835   0.0080   1.0000   0.5770
  -3.000  -0.3932   0.01800   0.00812   0.0098   1.0000   0.6126
  -2.750  -0.3726   0.01774   0.00796   0.0117   1.0000   0.6467
  -2.500  -0.3518   0.01751   0.00782   0.0136   1.0000   0.6808
  -2.250  -0.3308   0.01731   0.00769   0.0155   1.0000   0.7149
  -2.000  -0.3092   0.01714   0.00760   0.0173   1.0000   0.7494
  -1.750  -0.2865   0.01701   0.00753   0.0189   1.0000   0.7849
  -1.500  -0.2597   0.01694   0.00753   0.0197   1.0000   0.8199
  -1.250  -0.2271   0.01693   0.00754   0.0193   1.0000   0.8557
  -1.000  -0.1876   0.01696   0.00757   0.0173   1.0000   0.8908
  -0.750  -0.1417   0.01702   0.00759   0.0138   1.0000   0.9236
  -0.500  -0.0900   0.01706   0.00759   0.0089   1.0000   0.9527
  -0.250  -0.0362   0.01704   0.00756   0.0031   1.0000   0.9798
   0.000   0.0000   0.01700   0.00752   0.0000   1.0000   1.0000
   0.250   0.0362   0.01704   0.00756  -0.0031   0.9798   1.0000
   0.500   0.0900   0.01706   0.00759  -0.0089   0.9527   1.0000
   0.750   0.1416   0.01702   0.00759  -0.0138   0.9236   1.0000
   1.000   0.1877   0.01696   0.00757  -0.0173   0.8908   1.0000
   1.250   0.2271   0.01692   0.00754  -0.0193   0.8557   1.0000
   1.500   0.2597   0.01694   0.00753  -0.0197   0.8200   1.0000
   1.750   0.2866   0.01701   0.00753  -0.0189   0.7850   1.0000
   2.000   0.3093   0.01714   0.00760  -0.0173   0.7496   1.0000
   2.250   0.3308   0.01731   0.00769  -0.0155   0.7149   1.0000
   2.500   0.3518   0.01751   0.00782  -0.0137   0.6808   1.0000
   2.750   0.3726   0.01774   0.00796  -0.0117   0.6467   1.0000
   3.000   0.3933   0.01800   0.00812  -0.0098   0.6126   1.0000
   3.250   0.4138   0.01829   0.00835  -0.0080   0.5769   1.0000
   3.500   0.4344   0.01860   0.00858  -0.0062   0.5410   1.0000
   3.750   0.4549   0.01895   0.00882  -0.0044   0.5050   1.0000
   4.000   0.4754   0.01933   0.00912  -0.0028   0.4669   1.0000
   4.250   0.4961   0.01977   0.00946  -0.0012   0.4285   1.0000
   4.500   0.5171   0.02026   0.00985   0.0002   0.3885   1.0000
   4.750   0.5377   0.02082   0.01027   0.0016   0.3494   1.0000
   5.000   0.5584   0.02145   0.01081   0.0029   0.3095   1.0000
   5.250   0.5787   0.02219   0.01141   0.0041   0.2723   1.0000
   5.500   0.5991   0.02301   0.01213   0.0052   0.2370   1.0000
   5.750   0.6193   0.02394   0.01295   0.0063   0.2066   1.0000
   6.000   0.6394   0.02495   0.01386   0.0074   0.1812   1.0000
   6.250   0.6596   0.02603   0.01492   0.0085   0.1592   1.0000
   6.500   0.6797   0.02719   0.01603   0.0096   0.1423   1.0000
   6.750   0.6999   0.02840   0.01723   0.0106   0.1277   1.0000
   7.000   0.7206   0.02971   0.01864   0.0117   0.1158   1.0000
   7.250   0.7411   0.03114   0.02019   0.0127   0.1057   1.0000
   7.500   0.7610   0.03260   0.02156   0.0136   0.0983   1.0000
   7.750   0.7810   0.03426   0.02356   0.0147   0.0903   1.0000
   8.000   0.8007   0.03594   0.02515   0.0155   0.0852   1.0000
   8.250   0.8182   0.03812   0.02785   0.0167   0.0797   1.0000
   8.500   0.8353   0.03991   0.02977   0.0176   0.0750   1.0000
   8.750   0.8514   0.04230   0.03229   0.0185   0.0720   1.0000
   9.000   0.8609   0.04540   0.03595   0.0198   0.0690   1.0000
   9.250   0.8691   0.04820   0.03910   0.0209   0.0659   1.0000
   9.500   0.8781   0.05064   0.04171   0.0219   0.0633   1.0000
   9.750   0.8869   0.05332   0.04449   0.0228   0.0616   1.0000
  10.000   0.8834   0.05720   0.04871   0.0238   0.0608   1.0000
  10.250   0.8689   0.06160   0.05349   0.0245   0.0603   1.0000
  10.500   0.8494   0.06587   0.05797   0.0251   0.0602   1.0000
  10.750   0.8243   0.07100   0.06327   0.0237   0.0603   1.0000
  11.000   0.7997   0.07737   0.06974   0.0199   0.0606   1.0000
  11.250   0.7767   0.08490   0.07733   0.0147   0.0610   1.0000
<< Back to RAE 100 AIRFOIL (rae100-il)

Polar data table (+)

Polar graphs


<< Back to RAE 100 AIRFOIL (rae100-il)