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

NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 63-212 AIRFOIL (n63212-il)
Reynolds number: 100,000
Max Cl/Cd: 49.56 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63212-il-100000.txt
Download as CSV file: xf-n63212-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-212 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5387   0.09680   0.09190  -0.0211   1.0000   0.1584
  -9.500  -0.5196   0.09404   0.08911  -0.0195   1.0000   0.1630
  -9.250  -0.5495   0.08850   0.08372  -0.0253   1.0000   0.1728
  -9.000  -0.5253   0.08601   0.08119  -0.0227   1.0000   0.1769
  -8.750  -0.5541   0.08061   0.07593  -0.0278   1.0000   0.1880
  -8.000  -0.6706   0.04786   0.04169  -0.0429   1.0000   0.0815
  -7.750  -0.6700   0.04296   0.03595  -0.0403   1.0000   0.0704
  -7.500  -0.6599   0.03939   0.03218  -0.0388   1.0000   0.0684
  -7.250  -0.6502   0.03641   0.02889  -0.0367   1.0000   0.0664
  -7.000  -0.6399   0.03379   0.02591  -0.0343   1.0000   0.0649
  -6.750  -0.6283   0.03154   0.02331  -0.0319   1.0000   0.0642
  -6.500  -0.6150   0.02965   0.02113  -0.0297   1.0000   0.0641
  -6.250  -0.6001   0.02800   0.01926  -0.0278   1.0000   0.0647
  -6.000  -0.5846   0.02671   0.01781  -0.0259   1.0000   0.0672
  -5.750  -0.5681   0.02560   0.01648  -0.0242   1.0000   0.0700
  -5.500  -0.5505   0.02422   0.01499  -0.0225   1.0000   0.0725
  -5.250  -0.5334   0.02281   0.01370  -0.0211   1.0000   0.0762
  -5.000  -0.5161   0.02182   0.01271  -0.0195   1.0000   0.0819
  -4.750  -0.4996   0.02071   0.01169  -0.0181   1.0000   0.0903
  -4.500  -0.4828   0.01971   0.01081  -0.0169   1.0000   0.1074
  -4.250  -0.4574   0.01827   0.00971  -0.0177   0.9974   0.1537
  -4.000  -0.4307   0.01578   0.00955  -0.0189   0.9907   0.5604
  -3.750  -0.3928   0.01644   0.01017  -0.0201   0.9824   0.6256
  -3.500  -0.3582   0.01706   0.01076  -0.0206   0.9734   0.6621
  -3.250  -0.3272   0.01779   0.01148  -0.0200   0.9643   0.6962
  -3.000  -0.2927   0.01861   0.01227  -0.0196   0.9568   0.7273
  -2.750  -0.2669   0.01906   0.01266  -0.0183   0.9469   0.7507
  -2.500  -0.2338   0.01949   0.01304  -0.0181   0.9395   0.7694
  -2.250  -0.2034   0.01965   0.01310  -0.0182   0.9307   0.7846
  -2.000  -0.1688   0.01972   0.01307  -0.0194   0.9231   0.7980
  -1.750  -0.1339   0.01973   0.01298  -0.0208   0.9152   0.8110
  -1.500  -0.1008   0.01980   0.01297  -0.0214   0.9077   0.8213
  -1.250  -0.0677   0.01977   0.01288  -0.0223   0.8998   0.8321
  -1.000  -0.0382   0.01975   0.01280  -0.0227   0.8915   0.8436
  -0.750  -0.0047   0.01970   0.01269  -0.0236   0.8840   0.8547
  -0.500   0.0218   0.01971   0.01267  -0.0233   0.8752   0.8650
  -0.250   0.0557   0.01963   0.01255  -0.0242   0.8682   0.8756
   0.000   0.0792   0.01964   0.01255  -0.0236   0.8586   0.8880
   0.250   0.1162   0.01955   0.01243  -0.0249   0.8523   0.8981
   0.500   0.1435   0.01959   0.01248  -0.0250   0.8429   0.9096
   0.750   0.1828   0.01947   0.01234  -0.0269   0.8368   0.9200
   1.000   0.2136   0.01953   0.01242  -0.0279   0.8274   0.9323
   1.250   0.2616   0.01938   0.01229  -0.0314   0.8214   0.9409
   1.500   0.3035   0.01942   0.01238  -0.0346   0.8124   0.9508
   1.750   0.3508   0.01925   0.01223  -0.0381   0.8058   0.9599
   2.000   0.4002   0.01925   0.01232  -0.0428   0.7964   0.9680
   2.250   0.4507   0.01898   0.01210  -0.0469   0.7895   0.9759
   2.500   0.4955   0.01898   0.01222  -0.0509   0.7785   0.9860
   2.750   0.5426   0.01880   0.01212  -0.0548   0.7691   0.9957
   3.000   0.5716   0.01868   0.01206  -0.0552   0.7590   1.0000
   3.250   0.5826   0.01888   0.01234  -0.0529   0.7465   1.0000
   3.500   0.5945   0.01896   0.01247  -0.0504   0.7344   1.0000
   3.750   0.6083   0.01886   0.01240  -0.0476   0.7222   1.0000
   4.000   0.6243   0.01860   0.01216  -0.0449   0.7093   1.0000
   4.250   0.6410   0.01830   0.01191  -0.0422   0.6947   1.0000
   4.500   0.6587   0.01803   0.01167  -0.0398   0.6768   1.0000
   4.750   0.6812   0.01748   0.01113  -0.0376   0.6554   1.0000
   5.000   0.7057   0.01686   0.01049  -0.0357   0.6311   1.0000
   5.250   0.7303   0.01635   0.00997  -0.0340   0.6038   1.0000
   5.500   0.7531   0.01608   0.00974  -0.0324   0.5710   1.0000
   5.750   0.7750   0.01588   0.00954  -0.0306   0.5267   1.0000
   6.000   0.7920   0.01598   0.00941  -0.0281   0.4372   1.0000
   6.250   0.7862   0.01861   0.01035  -0.0236   0.1971   1.0000
   6.500   0.7919   0.02099   0.01204  -0.0210   0.1313   1.0000
   6.750   0.8066   0.02256   0.01341  -0.0193   0.1090   1.0000
   7.000   0.8246   0.02410   0.01482  -0.0180   0.0966   1.0000
   7.250   0.8460   0.02580   0.01630  -0.0173   0.0869   1.0000
   7.500   0.8711   0.02718   0.01781  -0.0167   0.0800   1.0000
   7.750   0.9003   0.02949   0.01993  -0.0170   0.0750   1.0000
   8.000   0.9264   0.03122   0.02191  -0.0166   0.0712   1.0000
   8.250   0.9516   0.03304   0.02389  -0.0162   0.0674   1.0000
   8.500   0.9770   0.03535   0.02639  -0.0159   0.0655   1.0000
   8.750   1.0006   0.03803   0.02928  -0.0155   0.0644   1.0000
   9.000   1.0216   0.04111   0.03264  -0.0148   0.0637   1.0000
   9.250   1.0388   0.04509   0.03689  -0.0142   0.0627   1.0000
   9.500   1.0489   0.04872   0.04092  -0.0128   0.0620   1.0000
   9.750   1.0579   0.05226   0.04485  -0.0114   0.0619   1.0000
  10.000   1.0688   0.05580   0.04875  -0.0101   0.0628   1.0000
  10.250   1.0400   0.06004   0.05401  -0.0054   0.0675   1.0000
  10.500   1.0253   0.06500   0.05933  -0.0034   0.0702   1.0000
  10.750   1.0170   0.06982   0.06431  -0.0022   0.0725   1.0000
  11.000   0.8970   0.06592   0.06098   0.0041   0.0712   1.0000
  11.250   0.8727   0.07169   0.06688   0.0036   0.0723   1.0000
  11.500   0.8677   0.07743   0.07273   0.0030   0.0758   1.0000
  11.750   0.8097   0.08586   0.08139  -0.0014   0.0763   1.0000
  12.000   0.7379   0.10296   0.09865  -0.0113   0.0886   1.0000
<< Back to NACA 63-212 AIRFOIL (n63212-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63-212 AIRFOIL (n63212-il)