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(1)-212 (naca661212-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 66(1)-212 (naca661212-il)
Reynolds number: 50,000
Max Cl/Cd: 22.25 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca661212-il-50000-n5.txt
Download as CSV file: xf-naca661212-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66(1)-212                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.4959   0.11247   0.10507  -0.0456   1.0000   0.0634
 -11.250  -0.4933   0.10832   0.10094  -0.0464   1.0000   0.0617
 -11.000  -0.4953   0.10362   0.09630  -0.0482   1.0000   0.0600
 -10.750  -0.5007   0.09840   0.09114  -0.0508   1.0000   0.0584
 -10.500  -0.5124   0.09192   0.08474  -0.0549   1.0000   0.0566
 -10.250  -0.5384   0.08489   0.07775  -0.0595   1.0000   0.0546
  -9.750  -0.6114   0.07793   0.07068  -0.0574   1.0000   0.0523
  -9.500  -0.6288   0.07570   0.06847  -0.0538   1.0000   0.0521
  -9.250  -0.6448   0.07346   0.06623  -0.0501   1.0000   0.0518
  -9.000  -0.6627   0.07132   0.06404  -0.0460   1.0000   0.0516
  -8.750  -0.6790   0.06910   0.06174  -0.0418   1.0000   0.0514
  -8.500  -0.6928   0.06678   0.05931  -0.0379   1.0000   0.0511
  -8.250  -0.7047   0.06429   0.05665  -0.0340   1.0000   0.0509
  -8.000  -0.7140   0.06167   0.05383  -0.0303   1.0000   0.0508
  -7.750  -0.7205   0.05904   0.05093  -0.0268   1.0000   0.0508
  -7.500  -0.7250   0.05639   0.04793  -0.0234   1.0000   0.0512
  -7.250  -0.7259   0.05383   0.04494  -0.0201   1.0000   0.0519
  -7.000  -0.7235   0.05147   0.04208  -0.0170   1.0000   0.0525
  -6.750  -0.7166   0.04852   0.03895  -0.0150   1.0000   0.0534
  -6.500  -0.7031   0.04581   0.03591  -0.0136   0.9988   0.0539
  -6.250  -0.6817   0.04309   0.03276  -0.0134   0.9959   0.0544
  -6.000  -0.6584   0.04069   0.02999  -0.0133   0.9931   0.0552
  -5.750  -0.6339   0.03863   0.02762  -0.0132   0.9902   0.0565
  -5.500  -0.6078   0.03696   0.02564  -0.0134   0.9874   0.0597
  -5.250  -0.5793   0.03541   0.02368  -0.0136   0.9850   0.0626
  -5.000  -0.5516   0.03401   0.02195  -0.0133   0.9826   0.0643
  -4.750  -0.5228   0.03261   0.02049  -0.0132   0.9808   0.0665
  -4.500  -0.4952   0.03165   0.01949  -0.0132   0.9787   0.0715
  -4.250  -0.4665   0.03092   0.01852  -0.0132   0.9765   0.0771
  -4.000  -0.4414   0.03002   0.01754  -0.0129   0.9740   0.0816
  -3.750  -0.4205   0.02926   0.01670  -0.0120   0.9707   0.0883
  -3.500  -0.3993   0.02850   0.01591  -0.0114   0.9676   0.0999
  -3.250  -0.3774   0.02771   0.01514  -0.0110   0.9650   0.1160
  -3.000  -0.3600   0.02631   0.01427  -0.0102   0.9622   0.1843
  -2.750  -0.3695   0.02484   0.01573  -0.0006   0.9587   0.7271
  -2.500  -0.2451   0.02978   0.02017  -0.0059   0.9673   0.9273
  -2.250  -0.2071   0.02975   0.01979  -0.0087   0.9652   0.9383
  -2.000  -0.1641   0.02965   0.01940  -0.0127   0.9635   0.9444
  -1.750  -0.1375   0.02954   0.01907  -0.0135   0.9603   0.9522
  -1.500  -0.1022   0.02943   0.01877  -0.0161   0.9577   0.9576
  -1.250  -0.0727   0.02941   0.01860  -0.0177   0.9546   0.9640
  -1.000  -0.0350   0.02938   0.01843  -0.0208   0.9520   0.9683
  -0.750  -0.0027   0.02941   0.01833  -0.0229   0.9493   0.9735
  -0.500   0.0256   0.02938   0.01822  -0.0243   0.9455   0.9783
  -0.250   0.0565   0.02942   0.01819  -0.0261   0.9418   0.9827
   0.000   0.0890   0.02952   0.01824  -0.0283   0.9384   0.9866
   0.250   0.1231   0.02963   0.01831  -0.0308   0.9350   0.9902
   0.500   0.1481   0.02971   0.01840  -0.0315   0.9301   0.9946
   0.750   0.1792   0.02986   0.01855  -0.0334   0.9258   0.9984
   1.000   0.2096   0.03007   0.01879  -0.0351   0.9218   1.0000
   1.250   0.2199   0.03022   0.01897  -0.0329   0.9148   1.0000
   1.500   0.2410   0.03043   0.01923  -0.0327   0.9091   1.0000
   1.750   0.2589   0.03064   0.01950  -0.0319   0.9028   1.0000
   2.000   0.2745   0.03084   0.01976  -0.0306   0.8956   1.0000
   2.250   0.2979   0.03110   0.02010  -0.0307   0.8896   1.0000
   2.500   0.3092   0.03129   0.02039  -0.0286   0.8810   1.0000
   2.750   0.3313   0.03154   0.02075  -0.0283   0.8742   1.0000
   3.000   0.3450   0.03175   0.02106  -0.0265   0.8652   1.0000
   3.250   0.3604   0.03197   0.02140  -0.0250   0.8563   1.0000
   3.500   0.3855   0.03219   0.02180  -0.0251   0.8482   1.0000
   3.750   0.3957   0.03236   0.02210  -0.0225   0.8370   1.0000
   4.000   0.4152   0.03253   0.02244  -0.0214   0.8266   1.0000
   4.250   0.4479   0.03263   0.02278  -0.0224   0.8173   1.0000
   4.500   0.4625   0.03269   0.02306  -0.0203   0.8041   1.0000
   5.000   0.5709   0.02728   0.01841  -0.0201   0.7086   1.0000
   5.250   0.5576   0.02659   0.01775  -0.0119   0.6640   1.0000
   5.500   0.5464   0.02641   0.01761  -0.0050   0.6105   1.0000
   5.750   0.5843   0.02626   0.01462  -0.0011   0.1669   1.0000
   6.000   0.5810   0.02762   0.01550   0.0032   0.1217   1.0000
   6.250   0.5874   0.02879   0.01652   0.0061   0.1041   1.0000
   6.500   0.5982   0.02993   0.01762   0.0083   0.0929   1.0000
   6.750   0.6120   0.03111   0.01877   0.0099   0.0834   1.0000
   7.000   0.6317   0.03218   0.01997   0.0111   0.0765   1.0000
   7.250   0.6622   0.03356   0.02130   0.0109   0.0711   1.0000
   7.500   0.7075   0.03487   0.02292   0.0088   0.0641   1.0000
   7.750   0.7622   0.03720   0.02530   0.0049   0.0593   1.0000
   8.000   0.8020   0.03955   0.02799   0.0034   0.0568   1.0000
   8.250   0.8284   0.04171   0.03054   0.0035   0.0539   1.0000
   8.500   0.8501   0.04393   0.03309   0.0041   0.0514   1.0000
   8.750   0.8699   0.04654   0.03602   0.0049   0.0504   1.0000
   9.000   0.8856   0.04931   0.03912   0.0061   0.0497   1.0000
   9.250   0.8971   0.05221   0.04236   0.0076   0.0492   1.0000
   9.500   0.9044   0.05523   0.04573   0.0096   0.0490   1.0000
   9.750   0.9096   0.05831   0.04909   0.0115   0.0485   1.0000
  10.000   0.9115   0.06150   0.05253   0.0134   0.0479   1.0000
  10.250   0.9110   0.06510   0.05632   0.0152   0.0473   1.0000
  10.500   0.9000   0.06802   0.05953   0.0183   0.0470   1.0000
  10.750   0.8847   0.07094   0.06272   0.0214   0.0468   1.0000
  11.000   0.8690   0.07421   0.06621   0.0236   0.0467   1.0000
  11.250   0.8522   0.07778   0.06999   0.0251   0.0467   1.0000
  11.500   0.8342   0.08177   0.07415   0.0257   0.0467   1.0000
  11.750   0.8170   0.08617   0.07869   0.0255   0.0469   1.0000
  12.000   0.7991   0.09110   0.08374   0.0243   0.0470   1.0000
  12.250   0.7832   0.09642   0.08914   0.0224   0.0472   1.0000
<< Back to NACA 66(1)-212 (naca661212-il)

Polar data table (+)

Polar graphs


<< Back to NACA 66(1)-212 (naca661212-il)