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

NACA 63-412 AIRFOIL (n63412-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 63-412 AIRFOIL (n63412-il)
Reynolds number: 500,000
Max Cl/Cd: 99.76 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63412-il-500000.txt
Download as CSV file: xf-n63412-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-412 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4722   0.08792   0.08560  -0.0423   1.0000   0.0281
 -10.250  -0.6475   0.04785   0.04503  -0.0681   1.0000   0.0183
 -10.000  -0.6706   0.04488   0.04194  -0.0661   1.0000   0.0182
  -9.750  -0.6890   0.04136   0.03822  -0.0638   1.0000   0.0181
  -9.500  -0.7076   0.03848   0.03516  -0.0597   1.0000   0.0180
  -9.250  -0.6960   0.03218   0.02826  -0.0632   0.9947   0.0180
  -9.000  -0.6719   0.02807   0.02359  -0.0659   0.9888   0.0185
  -8.750  -0.6445   0.02423   0.01927  -0.0687   0.9845   0.0193
  -8.500  -0.6128   0.02304   0.01801  -0.0708   0.9802   0.0201
  -8.250  -0.5800   0.02191   0.01679  -0.0727   0.9756   0.0210
  -8.000  -0.5458   0.02029   0.01496  -0.0748   0.9719   0.0218
  -7.750  -0.5139   0.01894   0.01341  -0.0761   0.9653   0.0228
  -7.500  -0.4805   0.01799   0.01227  -0.0776   0.9586   0.0236
  -7.250  -0.4521   0.01622   0.01039  -0.0784   0.9500   0.0249
  -7.000  -0.4217   0.01553   0.00968  -0.0793   0.9418   0.0262
  -6.750  -0.3950   0.01489   0.00897  -0.0793   0.9310   0.0274
  -6.500  -0.3686   0.01428   0.00826  -0.0792   0.9212   0.0286
  -6.250  -0.3424   0.01380   0.00768  -0.0789   0.9116   0.0297
  -6.000  -0.3192   0.01283   0.00664  -0.0784   0.9012   0.0316
  -5.750  -0.2936   0.01235   0.00611  -0.0781   0.8923   0.0335
  -5.500  -0.2679   0.01194   0.00564  -0.0778   0.8828   0.0354
  -5.250  -0.2415   0.01160   0.00521  -0.0776   0.8741   0.0374
  -5.000  -0.2159   0.01099   0.00454  -0.0774   0.8658   0.0408
  -4.750  -0.1892   0.01067   0.00418  -0.0773   0.8571   0.0449
  -4.500  -0.1622   0.01032   0.00375  -0.0772   0.8494   0.0498
  -4.250  -0.1351   0.00998   0.00341  -0.0772   0.8409   0.0582
  -4.000  -0.1080   0.00952   0.00302  -0.0773   0.8337   0.0865
  -3.750  -0.0819   0.00864   0.00264  -0.0777   0.8253   0.2049
  -3.500  -0.0559   0.00774   0.00232  -0.0781   0.8182   0.3748
  -3.250  -0.0285   0.00737   0.00224  -0.0783   0.8099   0.4647
  -3.000  -0.0006   0.00726   0.00217  -0.0783   0.8030   0.5087
  -2.750   0.0276   0.00718   0.00214  -0.0784   0.7949   0.5394
  -2.500   0.0557   0.00716   0.00209  -0.0783   0.7880   0.5625
  -2.250   0.0841   0.00714   0.00206  -0.0784   0.7800   0.5820
  -2.000   0.1120   0.00712   0.00205  -0.0783   0.7730   0.6057
  -1.750   0.1401   0.00710   0.00206  -0.0783   0.7652   0.6261
  -1.500   0.1683   0.00713   0.00206  -0.0782   0.7581   0.6436
  -1.250   0.1965   0.00714   0.00208  -0.0782   0.7504   0.6578
  -1.000   0.2246   0.00716   0.00208  -0.0781   0.7433   0.6678
  -0.750   0.2530   0.00718   0.00209  -0.0782   0.7357   0.6779
  -0.500   0.2814   0.00722   0.00208  -0.0782   0.7284   0.6878
  -0.250   0.3096   0.00722   0.00210  -0.0782   0.7208   0.6962
   0.000   0.3381   0.00727   0.00210  -0.0783   0.7135   0.7055
   0.250   0.3661   0.00728   0.00214  -0.0783   0.7059   0.7136
   0.500   0.3944   0.00731   0.00215  -0.0783   0.6984   0.7224
   0.750   0.4225   0.00734   0.00220  -0.0783   0.6907   0.7309
   1.000   0.4506   0.00738   0.00223  -0.0783   0.6830   0.7393
   1.250   0.4787   0.00741   0.00228  -0.0783   0.6750   0.7482
   1.500   0.5065   0.00746   0.00233  -0.0782   0.6672   0.7564
   1.750   0.5347   0.00751   0.00240  -0.0783   0.6588   0.7657
   2.000   0.5622   0.00757   0.00246  -0.0781   0.6510   0.7737
   2.250   0.5902   0.00761   0.00255  -0.0781   0.6418   0.7829
   2.500   0.6174   0.00767   0.00262  -0.0779   0.6326   0.7912
   2.750   0.6448   0.00773   0.00269  -0.0778   0.6219   0.7999
   3.000   0.6719   0.00778   0.00278  -0.0776   0.6100   0.8088
   3.250   0.6987   0.00784   0.00287  -0.0773   0.5969   0.8172
   3.500   0.7257   0.00792   0.00295  -0.0771   0.5831   0.8267
   3.750   0.7512   0.00801   0.00303  -0.0765   0.5633   0.8351
   4.000   0.7771   0.00813   0.00314  -0.0761   0.5434   0.8445
   4.250   0.8030   0.00825   0.00328  -0.0757   0.5276   0.8539
   4.500   0.8284   0.00838   0.00342  -0.0752   0.5082   0.8633
   4.750   0.8533   0.00857   0.00359  -0.0746   0.4857   0.8738
   5.000   0.8769   0.00879   0.00376  -0.0738   0.4524   0.8847
   5.250   0.8975   0.00921   0.00399  -0.0725   0.3978   0.8967
   5.500   0.9139   0.00995   0.00438  -0.0706   0.3167   0.9112
   5.750   0.9267   0.01089   0.00493  -0.0682   0.2274   0.9310
   6.000   0.9431   0.01183   0.00549  -0.0666   0.1482   0.9792
   6.250   0.9638   0.01286   0.00615  -0.0661   0.0872   1.0000
   6.500   0.9848   0.01374   0.00681  -0.0655   0.0553   1.0000
   6.750   1.0072   0.01442   0.00744  -0.0649   0.0438   1.0000
   7.000   1.0298   0.01503   0.00803  -0.0643   0.0375   1.0000
   7.250   1.0513   0.01569   0.00872  -0.0635   0.0333   1.0000
   7.500   1.0724   0.01633   0.00937  -0.0627   0.0301   1.0000
   7.750   1.0890   0.01731   0.01039  -0.0612   0.0276   1.0000
   8.000   1.1092   0.01794   0.01108  -0.0602   0.0261   1.0000
   8.250   1.1280   0.01863   0.01181  -0.0590   0.0246   1.0000
   8.500   1.1447   0.01942   0.01262  -0.0576   0.0233   1.0000
   8.750   1.1501   0.02091   0.01416  -0.0545   0.0219   1.0000
   9.000   1.1656   0.02160   0.01491  -0.0528   0.0215   1.0000
   9.250   1.1791   0.02245   0.01584  -0.0509   0.0209   1.0000
   9.500   1.1921   0.02338   0.01683  -0.0490   0.0203   1.0000
   9.750   1.2046   0.02440   0.01792  -0.0472   0.0197   1.0000
  10.000   1.2171   0.02547   0.01906  -0.0455   0.0192   1.0000
  10.250   1.2296   0.02659   0.02024  -0.0439   0.0187   1.0000
  10.500   1.2418   0.02781   0.02151  -0.0424   0.0183   1.0000
  10.750   1.2536   0.02915   0.02290  -0.0410   0.0178   1.0000
  11.000   1.2671   0.03130   0.02512  -0.0398   0.0173   1.0000
  11.250   1.2796   0.03293   0.02689  -0.0386   0.0168   1.0000
  11.500   1.2910   0.03433   0.02842  -0.0373   0.0166   1.0000
  11.750   1.3017   0.03595   0.03019  -0.0360   0.0163   1.0000
  12.000   1.3114   0.03780   0.03220  -0.0347   0.0160   1.0000
  12.250   1.3191   0.03979   0.03436  -0.0335   0.0157   1.0000
  12.500   1.3250   0.04203   0.03678  -0.0322   0.0155   1.0000
  12.750   1.3285   0.04448   0.03940  -0.0310   0.0153   1.0000
  13.000   1.3295   0.04713   0.04224  -0.0299   0.0151   1.0000
  13.250   1.3284   0.05001   0.04531  -0.0290   0.0150   1.0000
  13.500   1.3245   0.05322   0.04871  -0.0283   0.0148   1.0000
  13.750   1.3201   0.05648   0.05214  -0.0280   0.0147   1.0000
  14.000   1.3150   0.05988   0.05569  -0.0279   0.0145   1.0000
  14.250   1.3092   0.06348   0.05944  -0.0282   0.0144   1.0000
  14.500   1.3033   0.06714   0.06322  -0.0288   0.0142   1.0000
  14.750   1.2931   0.07160   0.06784  -0.0299   0.0141   1.0000
  15.000   1.2820   0.07645   0.07284  -0.0315   0.0140   1.0000
  15.250   1.2686   0.08194   0.07850  -0.0336   0.0139   1.0000
  15.500   1.2509   0.08854   0.08529  -0.0365   0.0139   1.0000
  15.750   1.2277   0.09666   0.09365  -0.0407   0.0139   1.0000
  16.000   1.1903   0.10844   0.10574  -0.0475   0.0141   1.0000
<< Back to NACA 63-412 AIRFOIL (n63412-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63-412 AIRFOIL (n63412-il)