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

NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il)
Reynolds number: 500,000
Max Cl/Cd: 76.47 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ames01-il-500000.txt
Download as CSV file: xf-ames01-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.6393   0.13696   0.13480   0.0264   1.0000   0.0213
 -11.750  -1.0581   0.03936   0.03598  -0.0208   1.0000   0.0134
 -11.500  -1.0513   0.03731   0.03380  -0.0197   1.0000   0.0138
 -11.250  -1.0279   0.03737   0.03392  -0.0195   1.0000   0.0141
 -11.000  -1.0041   0.03740   0.03398  -0.0194   1.0000   0.0144
 -10.750  -0.9865   0.03624   0.03271  -0.0190   1.0000   0.0149
 -10.500  -0.9777   0.03323   0.02936  -0.0181   1.0000   0.0156
 -10.250  -0.9664   0.03021   0.02593  -0.0171   1.0000   0.0163
 -10.000  -0.9489   0.02809   0.02346  -0.0166   1.0000   0.0168
  -9.750  -0.9333   0.02547   0.02054  -0.0159   1.0000   0.0175
  -9.500  -0.9079   0.02496   0.01999  -0.0161   1.0000   0.0180
  -9.250  -0.8820   0.02446   0.01946  -0.0163   1.0000   0.0185
  -9.000  -0.8568   0.02368   0.01857  -0.0164   1.0000   0.0192
  -8.750  -0.8322   0.02269   0.01742  -0.0164   1.0000   0.0199
  -8.500  -0.8079   0.02162   0.01617  -0.0163   1.0000   0.0207
  -8.250  -0.7787   0.02094   0.01529  -0.0169   0.9931   0.0213
  -8.000  -0.7519   0.01934   0.01354  -0.0175   0.9880   0.0224
  -7.750  -0.7213   0.01910   0.01331  -0.0184   0.9840   0.0233
  -7.500  -0.6936   0.01883   0.01300  -0.0186   0.9787   0.0242
  -7.250  -0.6672   0.01840   0.01248  -0.0184   0.9736   0.0253
  -7.000  -0.6419   0.01793   0.01190  -0.0179   0.9686   0.0263
  -6.750  -0.6175   0.01751   0.01135  -0.0172   0.9634   0.0269
  -6.500  -0.5967   0.01619   0.00993  -0.0160   0.9588   0.0283
  -6.250  -0.5714   0.01580   0.00953  -0.0155   0.9542   0.0293
  -6.000  -0.5470   0.01556   0.00926  -0.0148   0.9497   0.0305
  -5.750  -0.5232   0.01538   0.00903  -0.0138   0.9457   0.0321
  -5.500  -0.4957   0.01531   0.00892  -0.0138   0.9416   0.0337
  -5.250  -0.4721   0.01446   0.00801  -0.0130   0.9374   0.0358
  -5.000  -0.4482   0.01426   0.00783  -0.0121   0.9334   0.0378
  -4.750  -0.4222   0.01404   0.00759  -0.0116   0.9289   0.0398
  -4.500  -0.3972   0.01376   0.00727  -0.0109   0.9233   0.0417
  -4.250  -0.3746   0.01363   0.00708  -0.0094   0.9181   0.0431
  -4.000  -0.3507   0.01270   0.00613  -0.0086   0.9117   0.0462
  -3.750  -0.3273   0.01238   0.00579  -0.0074   0.9058   0.0486
  -3.500  -0.3022   0.01211   0.00550  -0.0066   0.9000   0.0514
  -3.250  -0.2760   0.01188   0.00524  -0.0060   0.8938   0.0540
  -3.000  -0.2529   0.01142   0.00472  -0.0047   0.8885   0.0570
  -2.750  -0.2261   0.01101   0.00434  -0.0044   0.8815   0.0613
  -2.500  -0.2009   0.01076   0.00405  -0.0035   0.8751   0.0653
  -2.250  -0.1739   0.01049   0.00376  -0.0031   0.8683   0.0699
  -2.000  -0.1479   0.01016   0.00344  -0.0025   0.8611   0.0777
  -1.750  -0.1210   0.00987   0.00318  -0.0021   0.8537   0.0892
  -1.500  -0.0956   0.00923   0.00288  -0.0016   0.8458   0.1719
  -1.250  -0.0750   0.00769   0.00258  -0.0008   0.8386   0.5063
  -1.000  -0.0526   0.00692   0.00247   0.0004   0.8305   0.6816
  -0.750  -0.0296   0.00644   0.00243   0.0020   0.8225   0.7968
  -0.500  -0.0062   0.00618   0.00241   0.0037   0.8137   0.8806
  -0.250   0.0238   0.00615   0.00247   0.0040   0.8030   0.9348
   0.000   0.0606   0.00625   0.00253   0.0027   0.7897   0.9665
   0.250   0.1042   0.00639   0.00259  -0.0002   0.7724   0.9806
   0.500   0.1519   0.00655   0.00264  -0.0042   0.7486   0.9890
   0.750   0.2008   0.00678   0.00268  -0.0084   0.7051   0.9971
   1.000   0.2389   0.00705   0.00261  -0.0106   0.6312   1.0000
   1.250   0.2687   0.00752   0.00256  -0.0114   0.5163   1.0000
   1.500   0.2987   0.00817   0.00263  -0.0125   0.3895   1.0000
   1.750   0.3264   0.00854   0.00268  -0.0129   0.3318   1.0000
   2.000   0.3530   0.00875   0.00273  -0.0128   0.3052   1.0000
   2.250   0.3792   0.00891   0.00279  -0.0127   0.2891   1.0000
   2.500   0.4051   0.00908   0.00286  -0.0124   0.2774   1.0000
   2.750   0.4309   0.00921   0.00293  -0.0121   0.2682   1.0000
   3.000   0.4567   0.00937   0.00303  -0.0118   0.2603   1.0000
   3.250   0.4824   0.00950   0.00312  -0.0115   0.2533   1.0000
   3.500   0.5080   0.00971   0.00326  -0.0111   0.2471   1.0000
   3.750   0.5337   0.00981   0.00338  -0.0107   0.2422   1.0000
   4.000   0.5593   0.00998   0.00351  -0.0103   0.2369   1.0000
   4.250   0.5847   0.01022   0.00370  -0.0100   0.2314   1.0000
   4.500   0.6105   0.01034   0.00384  -0.0096   0.2271   1.0000
   4.750   0.6362   0.01052   0.00400  -0.0092   0.2225   1.0000
   5.000   0.6613   0.01083   0.00425  -0.0088   0.2171   1.0000
   5.250   0.6873   0.01093   0.00441  -0.0085   0.2132   1.0000
   5.500   0.7131   0.01109   0.00457  -0.0081   0.2085   1.0000
   5.750   0.7385   0.01135   0.00480  -0.0077   0.2035   1.0000
   6.000   0.7642   0.01154   0.00502  -0.0074   0.1990   1.0000
   6.250   0.7902   0.01166   0.00516  -0.0071   0.1938   1.0000
   6.500   0.8157   0.01191   0.00538  -0.0068   0.1886   1.0000
   6.750   0.8415   0.01210   0.00561  -0.0064   0.1841   1.0000
   7.000   0.8678   0.01222   0.00577  -0.0062   0.1786   1.0000
   7.250   0.8933   0.01248   0.00598  -0.0059   0.1720   1.0000
   7.500   0.9200   0.01256   0.00614  -0.0058   0.1659   1.0000
   7.750   0.9457   0.01280   0.00635  -0.0056   0.1594   1.0000
   8.000   0.9720   0.01296   0.00657  -0.0054   0.1529   1.0000
   8.250   0.9975   0.01323   0.00681  -0.0052   0.1448   1.0000
   8.500   1.0237   0.01342   0.00705  -0.0050   0.1365   1.0000
   8.750   1.0491   0.01372   0.00734  -0.0048   0.1269   1.0000
   9.000   1.0741   0.01407   0.00766  -0.0045   0.1156   1.0000
   9.250   1.0986   0.01449   0.00806  -0.0043   0.1021   1.0000
   9.500   1.1220   0.01505   0.00854  -0.0039   0.0863   1.0000
   9.750   1.1441   0.01579   0.00918  -0.0034   0.0704   1.0000
  10.000   1.1659   0.01654   0.00989  -0.0028   0.0576   1.0000
  10.250   1.1868   0.01737   0.01068  -0.0021   0.0476   1.0000
  10.500   1.2075   0.01820   0.01151  -0.0014   0.0409   1.0000
  10.750   1.2282   0.01896   0.01233  -0.0007   0.0364   1.0000
  11.000   1.2474   0.01983   0.01322   0.0002   0.0327   1.0000
  11.250   1.2651   0.02082   0.01428   0.0012   0.0299   1.0000
  11.500   1.2843   0.02155   0.01510   0.0021   0.0282   1.0000
  11.750   1.3016   0.02242   0.01605   0.0032   0.0265   1.0000
  12.000   1.3142   0.02363   0.01732   0.0047   0.0249   1.0000
  12.250   1.3208   0.02516   0.01897   0.0070   0.0237   1.0000
  12.500   1.3323   0.02607   0.01998   0.0088   0.0231   1.0000
  12.750   1.3402   0.02717   0.02120   0.0108   0.0224   1.0000
  13.000   1.3472   0.02845   0.02258   0.0126   0.0218   1.0000
  13.250   1.3533   0.02988   0.02410   0.0142   0.0211   1.0000
  13.500   1.3586   0.03145   0.02577   0.0156   0.0205   1.0000
  13.750   1.3616   0.03330   0.02773   0.0169   0.0199   1.0000
  14.000   1.3603   0.03565   0.03017   0.0180   0.0194   1.0000
  14.250   1.3537   0.03870   0.03334   0.0187   0.0189   1.0000
  14.500   1.3415   0.04261   0.03740   0.0189   0.0185   1.0000
  14.750   1.3310   0.04669   0.04163   0.0185   0.0182   1.0000
  15.000   1.3288   0.05015   0.04523   0.0174   0.0181   1.0000
  15.250   1.3234   0.05426   0.04948   0.0159   0.0179   1.0000
  15.500   1.3156   0.05899   0.05436   0.0138   0.0178   1.0000
  15.750   1.3063   0.06422   0.05974   0.0112   0.0176   1.0000
  16.000   1.2948   0.06994   0.06561   0.0083   0.0175   1.0000
  16.250   1.2825   0.07597   0.07177   0.0052   0.0174   1.0000
  16.500   1.2692   0.08216   0.07810   0.0021   0.0173   1.0000
  16.750   1.2553   0.08845   0.08451  -0.0010   0.0171   1.0000
  17.000   1.2412   0.09480   0.09097  -0.0041   0.0171   1.0000
<< Back to NASA/AMES A-01 AIRFOIL (ames01-il)

Polar data table (+)

Polar graphs


<< Back to NASA/AMES A-01 AIRFOIL (ames01-il)