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

NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA 66-210 (naca66210-il)
Reynolds number: 1,000,000
Max Cl/Cd: 67.14 at α=1.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca66210-il-1000000-n5.txt
Download as CSV file: xf-naca66210-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-210                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.000  -0.5484   0.14744   0.14567  -0.0126   1.0000   0.0033
 -13.750  -0.5499   0.14152   0.13976  -0.0146   1.0000   0.0035
 -11.750  -0.9819   0.02722   0.02395  -0.0514   0.9806   0.0040
 -11.500  -0.9615   0.02463   0.02107  -0.0525   0.9705   0.0042
 -11.250  -0.9393   0.02296   0.01919  -0.0532   0.9609   0.0043
 -10.750  -0.8980   0.02104   0.01699  -0.0524   0.9421   0.0047
 -10.500  -0.8795   0.02001   0.01578  -0.0515   0.9341   0.0049
 -10.250  -0.8596   0.01922   0.01483  -0.0507   0.9265   0.0051
 -10.000  -0.8392   0.01843   0.01390  -0.0499   0.9199   0.0054
  -9.750  -0.8189   0.01759   0.01290  -0.0491   0.9131   0.0056
  -9.500  -0.7974   0.01691   0.01208  -0.0484   0.9072   0.0058
  -9.250  -0.7748   0.01635   0.01141  -0.0478   0.9011   0.0060
  -9.000  -0.7535   0.01557   0.01051  -0.0471   0.8955   0.0065
  -8.750  -0.7289   0.01533   0.01024  -0.0469   0.8907   0.0068
  -8.500  -0.7048   0.01494   0.00978  -0.0465   0.8855   0.0072
  -8.000  -0.6566   0.01411   0.00878  -0.0459   0.8764   0.0080
  -7.750  -0.6324   0.01363   0.00820  -0.0455   0.8716   0.0084
  -7.500  -0.6081   0.01322   0.00770  -0.0451   0.8673   0.0087
  -7.250  -0.5829   0.01288   0.00730  -0.0449   0.8630   0.0090
  -7.000  -0.5599   0.01221   0.00654  -0.0443   0.8583   0.0097
  -6.750  -0.5346   0.01194   0.00623  -0.0441   0.8540   0.0103
  -6.500  -0.5089   0.01168   0.00594  -0.0440   0.8503   0.0108
  -6.250  -0.4832   0.01139   0.00561  -0.0438   0.8464   0.0114
  -6.000  -0.4578   0.01106   0.00522  -0.0436   0.8425   0.0119
  -5.750  -0.4323   0.01077   0.00487  -0.0434   0.8389   0.0124
  -5.500  -0.4062   0.01053   0.00458  -0.0433   0.8356   0.0128
  -5.250  -0.3801   0.01027   0.00429  -0.0432   0.8320   0.0131
  -5.000  -0.3556   0.00977   0.00371  -0.0428   0.8284   0.0137
  -4.750  -0.3302   0.00942   0.00330  -0.0425   0.8249   0.0146
  -4.500  -0.3038   0.00919   0.00305  -0.0425   0.8217   0.0155
  -4.250  -0.2769   0.00900   0.00286  -0.0425   0.8184   0.0167
  -4.000  -0.2502   0.00881   0.00265  -0.0424   0.8152   0.0178
  -3.750  -0.2233   0.00864   0.00244  -0.0424   0.8122   0.0186
  -3.500  -0.1962   0.00850   0.00226  -0.0424   0.8095   0.0192
  -3.250  -0.1695   0.00825   0.00200  -0.0424   0.8068   0.0221
  -3.000  -0.1422   0.00810   0.00185  -0.0425   0.8039   0.0246
  -2.750  -0.1148   0.00797   0.00172  -0.0426   0.8010   0.0269
  -2.500  -0.0873   0.00787   0.00159  -0.0426   0.7983   0.0283
  -2.000  -0.0323   0.00763   0.00137  -0.0428   0.7931   0.0396
  -1.750  -0.0057   0.00735   0.00126  -0.0429   0.7904   0.0899
  -1.500   0.0188   0.00672   0.00110  -0.0427   0.7877   0.2386
  -1.250   0.0422   0.00592   0.00093  -0.0426   0.7851   0.4290
  -1.000   0.0641   0.00501   0.00081  -0.0420   0.7827   0.6664
  -0.750   0.0911   0.00489   0.00085  -0.0419   0.7805   0.7220
  -0.500   0.1189   0.00485   0.00088  -0.0420   0.7779   0.7452
  -0.250   0.1471   0.00485   0.00090  -0.0421   0.7746   0.7570
   0.000   0.1754   0.00486   0.00091  -0.0423   0.7705   0.7652
   0.250   0.2028   0.00486   0.00093  -0.0422   0.7612   0.7764
   0.500   0.2306   0.00489   0.00092  -0.0422   0.7505   0.7821
   0.750   0.2585   0.00491   0.00092  -0.0423   0.7407   0.7844
   1.250   0.3137   0.00498   0.00094  -0.0424   0.7120   0.7893
   1.500   0.3404   0.00507   0.00096  -0.0422   0.6841   0.7920
   1.750   0.3571   0.00592   0.00112  -0.0403   0.5137   0.7950
   2.000   0.3732   0.00698   0.00148  -0.0385   0.3307   0.7981
   2.250   0.3927   0.00777   0.00176  -0.0374   0.1962   0.8012
   2.500   0.4143   0.00839   0.00201  -0.0365   0.0951   0.8044
   2.750   0.4381   0.00883   0.00222  -0.0360   0.0392   0.8075
   3.000   0.4645   0.00901   0.00238  -0.0359   0.0315   0.8103
   3.250   0.4907   0.00918   0.00254  -0.0358   0.0277   0.8131
   3.500   0.5168   0.00936   0.00274  -0.0356   0.0245   0.8162
   3.750   0.5433   0.00950   0.00290  -0.0355   0.0236   0.8196
   4.000   0.5696   0.00967   0.00309  -0.0354   0.0216   0.8231
   4.250   0.5952   0.00988   0.00330  -0.0352   0.0192   0.8264
   4.500   0.6202   0.01015   0.00359  -0.0348   0.0164   0.8300
   4.750   0.6458   0.01036   0.00382  -0.0345   0.0158   0.8337
   5.000   0.6712   0.01060   0.00409  -0.0343   0.0148   0.8374
   5.250   0.6961   0.01086   0.00439  -0.0339   0.0139   0.8409
   5.500   0.7208   0.01112   0.00467  -0.0335   0.0131   0.8448
   5.750   0.7451   0.01144   0.00501  -0.0330   0.0122   0.8490
   6.250   0.7897   0.01244   0.00618  -0.0313   0.0111   0.8577
   6.500   0.8129   0.01283   0.00662  -0.0307   0.0110   0.8626
   7.000   0.8582   0.01368   0.00761  -0.0292   0.0107   0.8723
   7.250   0.8815   0.01402   0.00800  -0.0287   0.0104   0.8776
   7.500   0.9040   0.01444   0.00849  -0.0280   0.0100   0.8826
   7.750   0.9258   0.01489   0.00903  -0.0271   0.0097   0.8882
   8.000   0.9478   0.01535   0.00955  -0.0263   0.0093   0.8941
   8.250   0.9694   0.01578   0.01005  -0.0255   0.0090   0.8999
   8.500   0.9913   0.01616   0.01048  -0.0247   0.0087   0.9065
   8.750   1.0131   0.01648   0.01083  -0.0239   0.0082   0.9131
   9.000   1.0322   0.01712   0.01154  -0.0227   0.0078   0.9210
   9.250   1.0487   0.01811   0.01270  -0.0211   0.0074   0.9292
   9.500   1.0681   0.01865   0.01334  -0.0199   0.0073   0.9383
   9.750   1.0871   0.01908   0.01387  -0.0186   0.0071   0.9494
  10.250   1.1327   0.02036   0.01539  -0.0181   0.0066   1.0000
  10.500   1.1518   0.02098   0.01609  -0.0171   0.0063   1.0000
  10.750   1.1697   0.02157   0.01674  -0.0159   0.0060   1.0000
  11.000   1.1878   0.02196   0.01717  -0.0147   0.0057   1.0000
  11.250   1.2050   0.02244   0.01770  -0.0134   0.0055   1.0000
  11.500   1.2227   0.02281   0.01810  -0.0122   0.0053   1.0000
  11.750   1.2372   0.02359   0.01896  -0.0107   0.0051   1.0000
  12.000   1.2490   0.02466   0.02014  -0.0088   0.0048   1.0000
  12.250   1.2619   0.02557   0.02116  -0.0072   0.0047   1.0000
  12.500   1.2731   0.02664   0.02236  -0.0055   0.0046   1.0000
  12.750   1.2817   0.02795   0.02381  -0.0035   0.0045   1.0000
  13.000   1.2903   0.02921   0.02521  -0.0017   0.0044   1.0000
  13.250   1.2984   0.03048   0.02661   0.0001   0.0042   1.0000
  13.500   1.3051   0.03188   0.02814   0.0019   0.0041   1.0000
  13.750   1.3097   0.03347   0.02987   0.0037   0.0040   1.0000
  14.000   1.3111   0.03539   0.03194   0.0056   0.0039   1.0000
  14.250   1.3157   0.03701   0.03368   0.0070   0.0037   1.0000
  14.500   1.3159   0.03908   0.03590   0.0085   0.0036   1.0000
  14.750   1.3171   0.04112   0.03805   0.0097   0.0036   1.0000
  15.000   1.3151   0.04353   0.04058   0.0108   0.0035   1.0000
  15.250   1.3011   0.04740   0.04467   0.0118   0.0035   1.0000
  15.500   1.2942   0.05064   0.04804   0.0122   0.0034   1.0000
  15.750   1.2857   0.05430   0.05182   0.0120   0.0033   1.0000
  16.000   1.2698   0.05917   0.05686   0.0111   0.0033   1.0000
  16.250   1.2345   0.06766   0.06561   0.0078   0.0033   1.0000
<< Back to NACA 66-210 (naca66210-il)

Polar data table (+)

Polar graphs


<< Back to NACA 66-210 (naca66210-il)