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

NACA 65(1)-412 (naca651412-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-412 (naca651412-il)
Reynolds number: 500,000
Max Cl/Cd: 98.55 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651412-il-500000.txt
Download as CSV file: xf-naca651412-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-412                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4399   0.08395   0.08170  -0.0529   1.0000   0.0228
  -9.750  -0.4472   0.07908   0.07686  -0.0553   1.0000   0.0233
  -9.500  -0.4634   0.07253   0.07038  -0.0595   1.0000   0.0232
  -9.250  -0.4935   0.06734   0.06522  -0.0615   0.9994   0.0231
  -9.000  -0.4972   0.05930   0.05701  -0.0727   0.9883   0.0232
  -8.750  -0.4902   0.05286   0.05034  -0.0795   0.9802   0.0241
  -8.500  -0.4745   0.04787   0.04473  -0.0842   0.9724   0.0264
  -8.250  -0.4602   0.04384   0.04034  -0.0864   0.9644   0.0265
  -8.000  -0.4553   0.03527   0.03144  -0.0890   0.9567   0.0277
  -7.750  -0.4356   0.03331   0.02942  -0.0894   0.9476   0.0284
  -7.500  -0.4142   0.03128   0.02722  -0.0899   0.9402   0.0295
  -7.250  -0.4011   0.02630   0.02154  -0.0874   0.9290   0.0256
  -7.000  -0.3854   0.02237   0.01718  -0.0862   0.9200   0.0240
  -6.750  -0.3635   0.02010   0.01458  -0.0853   0.9127   0.0240
  -6.500  -0.3405   0.01852   0.01276  -0.0846   0.9048   0.0244
  -6.250  -0.3155   0.01763   0.01166  -0.0841   0.8978   0.0255
  -6.000  -0.2904   0.01700   0.01087  -0.0836   0.8904   0.0262
  -5.750  -0.2656   0.01583   0.00954  -0.0831   0.8836   0.0268
  -5.500  -0.2423   0.01419   0.00783  -0.0824   0.8771   0.0277
  -5.250  -0.2181   0.01340   0.00700  -0.0820   0.8701   0.0288
  -5.000  -0.1927   0.01292   0.00647  -0.0817   0.8638   0.0302
  -4.750  -0.1675   0.01245   0.00597  -0.0813   0.8567   0.0319
  -4.500  -0.1420   0.01196   0.00540  -0.0810   0.8508   0.0331
  -4.250  -0.1163   0.01158   0.00496  -0.0807   0.8439   0.0342
  -4.000  -0.0917   0.01089   0.00418  -0.0804   0.8380   0.0362
  -3.750  -0.0657   0.01048   0.00376  -0.0803   0.8318   0.0396
  -3.500  -0.0389   0.01020   0.00343  -0.0803   0.8255   0.0429
  -3.250  -0.0118   0.00988   0.00303  -0.0802   0.8201   0.0481
  -3.000   0.0151   0.00957   0.00274  -0.0802   0.8139   0.0613
  -2.750   0.0374   0.00774   0.00219  -0.0808   0.8084   0.4039
  -2.500   0.0629   0.00713   0.00217  -0.0809   0.8028   0.5917
  -2.250   0.0906   0.00708   0.00217  -0.0809   0.7970   0.6300
  -2.000   0.1187   0.00709   0.00218  -0.0809   0.7922   0.6557
  -1.750   0.1462   0.00713   0.00225  -0.0807   0.7866   0.6836
  -1.500   0.1736   0.00722   0.00236  -0.0805   0.7813   0.7104
  -1.250   0.2014   0.00732   0.00242  -0.0803   0.7767   0.7256
  -1.000   0.2289   0.00734   0.00248  -0.0802   0.7711   0.7356
  -0.750   0.2574   0.00735   0.00244  -0.0804   0.7661   0.7414
  -0.500   0.2858   0.00737   0.00243  -0.0806   0.7617   0.7460
  -0.250   0.3141   0.00738   0.00245  -0.0808   0.7564   0.7515
   0.000   0.3426   0.00739   0.00244  -0.0810   0.7515   0.7567
   0.250   0.3710   0.00743   0.00245  -0.0811   0.7473   0.7615
   0.500   0.3992   0.00745   0.00249  -0.0813   0.7423   0.7672
   0.750   0.4275   0.00747   0.00252  -0.0815   0.7375   0.7724
   1.000   0.4558   0.00752   0.00256  -0.0816   0.7333   0.7777
   1.250   0.4841   0.00756   0.00263  -0.0818   0.7287   0.7837
   1.500   0.5121   0.00758   0.00270  -0.0819   0.7238   0.7889
   1.750   0.5403   0.00764   0.00276  -0.0821   0.7195   0.7947
   2.000   0.5685   0.00769   0.00284  -0.0822   0.7143   0.8011
   2.250   0.5957   0.00770   0.00289  -0.0821   0.7075   0.8063
   2.500   0.6230   0.00772   0.00294  -0.0820   0.6986   0.8129
   2.750   0.6496   0.00772   0.00293  -0.0817   0.6874   0.8190
   3.000   0.6756   0.00772   0.00294  -0.0812   0.6735   0.8256
   3.250   0.7011   0.00770   0.00295  -0.0807   0.6543   0.8323
   3.500   0.7263   0.00773   0.00300  -0.0801   0.6359   0.8388
   3.750   0.7518   0.00781   0.00306  -0.0796   0.6161   0.8464
   4.000   0.7761   0.00790   0.00316  -0.0788   0.5919   0.8533
   4.250   0.7992   0.00811   0.00329  -0.0779   0.5536   0.8614
   4.500   0.8139   0.00876   0.00354  -0.0755   0.4542   0.8695
   4.750   0.8174   0.01042   0.00433  -0.0716   0.2761   0.8795
   5.000   0.8222   0.01206   0.00519  -0.0681   0.1189   0.8908
   5.250   0.8331   0.01310   0.00587  -0.0653   0.0499   0.9021
   5.500   0.8504   0.01356   0.00636  -0.0634   0.0402   0.9134
   5.750   0.8646   0.01415   0.00697  -0.0609   0.0340   0.9270
   6.000   0.8806   0.01443   0.00734  -0.0586   0.0319   0.9434
   6.250   0.8972   0.01479   0.00778  -0.0566   0.0295   0.9699
   6.500   0.9194   0.01561   0.00863  -0.0563   0.0271   1.0000
   6.750   0.9357   0.01663   0.00971  -0.0548   0.0257   1.0000
   7.000   0.9554   0.01724   0.01037  -0.0538   0.0248   1.0000
   7.250   0.9734   0.01793   0.01109  -0.0525   0.0237   1.0000
   7.500   0.9911   0.01863   0.01183  -0.0512   0.0227   1.0000
   7.750   1.0081   0.01939   0.01261  -0.0498   0.0217   1.0000
   8.000   1.0240   0.02035   0.01359  -0.0483   0.0210   1.0000
   8.250   1.0402   0.02173   0.01500  -0.0470   0.0203   1.0000
   8.500   1.0663   0.02381   0.01714  -0.0472   0.0197   1.0000
   8.750   1.0889   0.02478   0.01821  -0.0467   0.0194   1.0000
   9.000   1.1130   0.02597   0.01950  -0.0464   0.0191   1.0000
   9.250   1.1372   0.02730   0.02097  -0.0462   0.0186   1.0000
   9.500   1.1587   0.02856   0.02235  -0.0457   0.0180   1.0000
   9.750   1.1817   0.03027   0.02423  -0.0454   0.0175   1.0000
  10.000   1.2028   0.03224   0.02640  -0.0449   0.0172   1.0000
  10.250   1.2211   0.03472   0.02913  -0.0440   0.0171   1.0000
  10.500   1.2344   0.03776   0.03248  -0.0425   0.0171   1.0000
  10.750   1.2380   0.04163   0.03674  -0.0398   0.0174   1.0000
  11.000   1.2316   0.04570   0.04117  -0.0360   0.0179   1.0000
  11.250   1.2208   0.04952   0.04528  -0.0322   0.0183   1.0000
  11.500   1.2072   0.05300   0.04902  -0.0286   0.0185   1.0000
  11.750   1.1906   0.05684   0.05308  -0.0256   0.0187   1.0000
<< Back to NACA 65(1)-412 (naca651412-il)

Polar data table (+)

Polar graphs


<< Back to NACA 65(1)-412 (naca651412-il)