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

YS-900 (ys900-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: YS-900 (ys900-il)
Reynolds number: 100,000
Max Cl/Cd: 20.97 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ys900-il-100000.txt
Download as CSV file: xf-ys900-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: YS-900                                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6287   0.09637   0.09187  -0.0283   1.0000   0.0742
  -8.750  -0.6466   0.09316   0.08865  -0.0279   1.0000   0.0745
  -8.500  -0.6671   0.09066   0.08615  -0.0256   1.0000   0.0748
  -8.250  -0.6903   0.08859   0.08405  -0.0216   1.0000   0.0751
  -8.000  -0.7107   0.08646   0.08183  -0.0182   1.0000   0.0756
  -7.750  -0.7322   0.08493   0.08010  -0.0138   1.0000   0.0760
  -7.500  -0.7200   0.07816   0.07356  -0.0132   1.0000   0.0795
  -7.250  -0.7238   0.07508   0.07044  -0.0102   1.0000   0.0834
  -7.000  -0.6650   0.06174   0.05742  -0.0119   1.0000   0.0967
  -6.750  -0.7525   0.07053   0.06537  -0.0017   1.0000   0.0902
  -6.500  -0.7415   0.06588   0.06091  -0.0002   1.0000   0.0963
  -6.250  -0.7546   0.06410   0.05870   0.0048   1.0000   0.1034
  -6.000  -0.7454   0.06012   0.05485   0.0068   1.0000   0.1116
  -5.750  -0.7455   0.05696   0.05156   0.0102   1.0000   0.1212
  -5.500  -0.7442   0.05414   0.04862   0.0137   1.0000   0.1342
  -5.250  -0.7415   0.05162   0.04595   0.0172   1.0000   0.1507
  -5.000  -0.7396   0.04877   0.04304   0.0208   1.0000   0.1730
  -3.500  -0.6212   0.03163   0.02266   0.0388   1.0000   0.0726
  -3.250  -0.5892   0.02987   0.02017   0.0415   1.0000   0.0497
  -3.000  -0.5589   0.02680   0.01684   0.0421   1.0000   0.0421
  -2.750  -0.5316   0.02600   0.01564   0.0437   1.0000   0.0373
  -2.500  -0.5003   0.02368   0.01325   0.0437   1.0000   0.0348
  -2.250  -0.4741   0.02210   0.01162   0.0448   1.0000   0.0333
  -2.000  -0.4546   0.02100   0.01045   0.0468   1.0000   0.0324
  -1.750  -0.4371   0.02007   0.00907   0.0493   1.0000   0.0311
  -1.500  -0.4178   0.01939   0.00817   0.0514   1.0000   0.0328
  -1.250  -0.3968   0.01893   0.00748   0.0530   1.0000   0.0369
  -1.000  -0.0638   0.02056   0.01170  -0.0073   1.0000   1.0000
  -0.750  -0.0480   0.02050   0.01157  -0.0055   1.0000   1.0000
  -0.500  -0.0320   0.02045   0.01148  -0.0037   1.0000   1.0000
  -0.250  -0.0160   0.02042   0.01142  -0.0018   1.0000   1.0000
   0.000   0.0000   0.02041   0.01140   0.0000   1.0000   1.0000
   0.250   0.0161   0.02042   0.01142   0.0018   1.0000   1.0000
   0.500   0.0321   0.02045   0.01148   0.0036   1.0000   1.0000
   0.750   0.0480   0.02050   0.01157   0.0054   1.0000   1.0000
   1.000   0.0639   0.02057   0.01170   0.0072   1.0000   1.0000
   1.250   0.0797   0.02065   0.01188   0.0091   1.0000   1.0000
   1.500   0.0954   0.02076   0.01207   0.0109   1.0000   1.0000
   1.750   0.1109   0.02090   0.01230   0.0128   1.0000   1.0000
   2.000   0.1262   0.02105   0.01259   0.0147   1.0000   1.0000
   2.250   0.1415   0.02123   0.01296   0.0166   1.0000   1.0000
   2.500   0.1564   0.02145   0.01337   0.0185   1.0000   1.0000
   2.750   0.1712   0.02169   0.01389   0.0205   1.0000   1.0000
   3.000   0.5584   0.02663   0.01658  -0.0424   0.0411   1.0000
   3.250   0.5885   0.02920   0.01952  -0.0414   0.0487   1.0000
   3.500   0.6184   0.03207   0.02278  -0.0401   0.0627   1.0000
   3.750   0.7148   0.03606   0.03020  -0.0407   0.3370   1.0000
   4.000   0.7264   0.03866   0.03290  -0.0372   0.3155   1.0000
   4.250   0.7311   0.04106   0.03534  -0.0329   0.2794   1.0000
   4.500   0.7350   0.04363   0.03785  -0.0288   0.2414   1.0000
   5.000   0.7409   0.04862   0.04291  -0.0211   0.1750   1.0000
   5.250   0.7491   0.05195   0.04613  -0.0184   0.1563   1.0000
   5.500   0.7460   0.05401   0.04849  -0.0141   0.1365   1.0000
   5.750   0.7473   0.05681   0.05142  -0.0106   0.1230   1.0000
   6.000   0.7591   0.06284   0.05713  -0.0096   0.1154   1.0000
   6.250   0.7531   0.06304   0.05777  -0.0044   0.1044   1.0000
   6.500   0.7439   0.06575   0.06077  -0.0001   0.0977   1.0000
   6.750   0.7481   0.06901   0.06402   0.0025   0.0914   1.0000
   7.000   0.7437   0.07353   0.06857   0.0054   0.0885   1.0000
   7.250   0.7263   0.07492   0.07028   0.0099   0.0843   1.0000
   7.500   0.7199   0.07782   0.07324   0.0130   0.0806   1.0000
   7.750   0.7377   0.08519   0.08033   0.0132   0.0762   1.0000
   8.000   0.7158   0.08659   0.08194   0.0177   0.0758   1.0000
   8.250   0.6950   0.08859   0.08405   0.0213   0.0753   1.0000
   8.500   0.6720   0.09063   0.08613   0.0252   0.0750   1.0000
   8.750   0.6518   0.09303   0.08852   0.0279   0.0748   1.0000
   9.000   0.6319   0.09618   0.09167   0.0285   0.0744   1.0000
<< Back to YS-900 (ys900-il)

Polar data table (+)

Polar graphs


<< Back to YS-900 (ys900-il)