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NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il)
Reynolds number: 50,000
Max Cl/Cd: 25.91 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ames01-il-50000.txt
Download as CSV file: xf-ames01-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES A-01 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5496   0.12557   0.11877   0.0195   1.0000   0.2527
  -9.750  -0.5711   0.12446   0.11777   0.0168   1.0000   0.2630
  -9.500  -0.5604   0.12046   0.11377   0.0173   1.0000   0.2780
  -9.250  -0.5484   0.11638   0.10971   0.0178   1.0000   0.2928
  -9.000  -0.5396   0.11260   0.10595   0.0181   1.0000   0.3077
  -8.750  -0.5357   0.10926   0.10264   0.0182   1.0000   0.3233
  -8.500  -0.5390   0.10646   0.09991   0.0180   1.0000   0.3396
  -8.250  -0.5081   0.10153   0.09492   0.0201   1.0000   0.3620
  -8.000  -0.5161   0.09947   0.09294   0.0206   1.0000   0.3853
  -7.750  -0.4903   0.09505   0.08849   0.0221   1.0000   0.4090
  -7.500  -0.4969   0.09278   0.08632   0.0232   1.0000   0.4336
  -7.250  -0.4709   0.08862   0.08212   0.0243   1.0000   0.4573
  -7.000  -0.4593   0.08539   0.07891   0.0256   1.0000   0.4847
  -6.500  -0.4326   0.07911   0.07266   0.0286   1.0000   0.5451
  -6.250  -0.5729   0.05611   0.04858  -0.0188   1.0000   0.2021
  -6.000  -0.5562   0.05071   0.04300  -0.0186   1.0000   0.1809
  -5.750  -0.5456   0.04633   0.03804  -0.0177   1.0000   0.1657
  -5.500  -0.5348   0.04299   0.03380  -0.0157   1.0000   0.1560
  -5.250  -0.5197   0.04029   0.03074  -0.0139   1.0000   0.1559
  -5.000  -0.5031   0.03780   0.02783  -0.0122   1.0000   0.1559
  -4.750  -0.4842   0.03538   0.02507  -0.0106   1.0000   0.1553
  -4.500  -0.4639   0.03329   0.02260  -0.0090   1.0000   0.1558
  -4.250  -0.4432   0.03126   0.02047  -0.0078   1.0000   0.1606
  -4.000  -0.4212   0.02967   0.01870  -0.0066   1.0000   0.1662
  -3.750  -0.3973   0.02821   0.01688  -0.0054   1.0000   0.1702
  -3.500  -0.3730   0.02674   0.01542  -0.0045   1.0000   0.1787
  -3.250  -0.3476   0.02553   0.01415  -0.0037   1.0000   0.1899
  -3.000  -0.3202   0.02447   0.01303  -0.0030   1.0000   0.2023
  -2.750  -0.2938   0.02333   0.01207  -0.0026   1.0000   0.2246
  -2.500  -0.0896   0.01992   0.01140  -0.0253   1.0000   1.0000
  -2.250  -0.0893   0.01974   0.01104  -0.0218   1.0000   1.0000
  -2.000  -0.0879   0.01961   0.01074  -0.0182   1.0000   1.0000
  -1.750  -0.0846   0.01954   0.01049  -0.0148   1.0000   1.0000
  -1.500  -0.0792   0.01951   0.01029  -0.0117   1.0000   1.0000
  -1.250  -0.0718   0.01953   0.01013  -0.0089   1.0000   1.0000
  -1.000  -0.0625   0.01960   0.01004  -0.0063   1.0000   1.0000
  -0.750  -0.0520   0.01970   0.00999  -0.0039   1.0000   1.0000
  -0.500  -0.0402   0.01985   0.01000  -0.0018   1.0000   1.0000
  -0.250  -0.0270   0.02004   0.01006   0.0001   1.0000   1.0000
   0.000  -0.0127   0.02027   0.01019   0.0017   1.0000   1.0000
   0.250   0.0025   0.02055   0.01037   0.0031   1.0000   1.0000
   0.500   0.0183   0.02088   0.01063   0.0043   1.0000   1.0000
   0.750   0.0346   0.02126   0.01094   0.0053   1.0000   1.0000
   1.000   0.0511   0.02171   0.01135   0.0061   1.0000   1.0000
   1.250   0.0675   0.02223   0.01184   0.0068   1.0000   1.0000
   1.500   0.1257   0.02315   0.01278  -0.0004   0.9849   1.0000
   1.750   0.2191   0.02397   0.01372  -0.0133   0.9499   1.0000
   2.000   0.3219   0.02405   0.01404  -0.0265   0.9081   1.0000
   2.250   0.4051   0.02314   0.01340  -0.0335   0.8583   1.0000
   2.500   0.4433   0.02188   0.01226  -0.0311   0.7977   1.0000
   2.750   0.4660   0.02067   0.01097  -0.0258   0.7238   1.0000
   3.000   0.4858   0.02010   0.01001  -0.0209   0.6496   1.0000
   3.250   0.5063   0.02035   0.00980  -0.0177   0.5924   1.0000
   3.500   0.5288   0.02096   0.01004  -0.0158   0.5528   1.0000
   3.750   0.5527   0.02170   0.01052  -0.0146   0.5227   1.0000
   4.000   0.5772   0.02250   0.01120  -0.0138   0.4983   1.0000
   4.250   0.6021   0.02333   0.01191  -0.0130   0.4795   1.0000
   4.500   0.6271   0.02422   0.01270  -0.0124   0.4628   1.0000
   4.750   0.6522   0.02517   0.01363  -0.0119   0.4479   1.0000
   5.000   0.6772   0.02621   0.01477  -0.0116   0.4346   1.0000
   5.250   0.7019   0.02733   0.01598  -0.0112   0.4220   1.0000
   5.500   0.7263   0.02853   0.01726  -0.0109   0.4105   1.0000
   5.750   0.7511   0.02976   0.01849  -0.0104   0.4005   1.0000
   6.000   0.7745   0.03114   0.02009  -0.0101   0.3895   1.0000
   6.250   0.7970   0.03271   0.02187  -0.0098   0.3793   1.0000
   6.500   0.8210   0.03414   0.02330  -0.0093   0.3701   1.0000
   6.750   0.8411   0.03596   0.02545  -0.0090   0.3596   1.0000
   7.000   0.8609   0.03793   0.02767  -0.0085   0.3499   1.0000
   7.250   0.8846   0.03943   0.02914  -0.0077   0.3399   1.0000
   7.500   0.8981   0.04196   0.03213  -0.0073   0.3292   1.0000
   7.750   0.9135   0.04445   0.03486  -0.0066   0.3193   1.0000
   8.000   0.9341   0.04622   0.03670  -0.0056   0.3079   1.0000
   8.250   0.9420   0.04933   0.04020  -0.0048   0.2974   1.0000
   8.500   0.9467   0.05280   0.04397  -0.0039   0.2872   1.0000
   8.750   0.9608   0.05516   0.04644  -0.0027   0.2755   1.0000
   9.000   0.9792   0.05696   0.04828  -0.0013   0.2622   1.0000
   9.250   0.9489   0.06397   0.05576  -0.0010   0.2582   1.0000
   9.500   0.9113   0.07194   0.06393  -0.0020   0.2575   1.0000
   9.750   0.8678   0.08091   0.07290  -0.0045   0.2603   1.0000
  10.000   0.8342   0.09071   0.08265  -0.0095   0.2638   1.0000
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