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N-12 (n12-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: N-12 (n12-il)
Reynolds number: 200,000
Max Cl/Cd: 46.45 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n12-il-200000.txt
Download as CSV file: xf-n12-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.8601   0.04001   0.03429  -0.0142   1.0000   0.0607
  -8.250  -0.8562   0.03472   0.02836  -0.0100   1.0000   0.0523
  -8.000  -0.8510   0.03122   0.02400  -0.0053   1.0000   0.0482
  -7.750  -0.8348   0.02956   0.02203  -0.0031   1.0000   0.0477
  -7.500  -0.8175   0.02712   0.01938  -0.0014   1.0000   0.0483
  -7.250  -0.7982   0.02519   0.01734   0.0000   1.0000   0.0495
  -7.000  -0.7777   0.02367   0.01565   0.0015   1.0000   0.0499
  -6.750  -0.7567   0.02230   0.01417   0.0029   1.0000   0.0505
  -6.500  -0.7359   0.02114   0.01293   0.0043   1.0000   0.0513
  -6.250  -0.7160   0.02014   0.01189   0.0058   1.0000   0.0528
  -6.000  -0.6965   0.01923   0.01093   0.0075   1.0000   0.0541
  -5.750  -0.6778   0.01838   0.01002   0.0093   1.0000   0.0549
  -5.500  -0.6600   0.01762   0.00923   0.0113   1.0000   0.0557
  -5.250  -0.6429   0.01696   0.00853   0.0134   1.0000   0.0566
  -5.000  -0.6266   0.01635   0.00788   0.0156   1.0000   0.0580
  -4.750  -0.6121   0.01566   0.00718   0.0181   1.0000   0.0601
  -4.500  -0.5962   0.01517   0.00667   0.0204   1.0000   0.0620
  -4.250  -0.5795   0.01478   0.00625   0.0225   1.0000   0.0644
  -4.000  -0.5623   0.01447   0.00590   0.0245   1.0000   0.0672
  -3.750  -0.5454   0.01414   0.00561   0.0266   1.0000   0.0734
  -3.500  -0.5290   0.01376   0.00542   0.0286   1.0000   0.1018
  -3.250  -0.5130   0.01336   0.00521   0.0306   1.0000   0.1427
  -3.000  -0.4977   0.01291   0.00501   0.0327   1.0000   0.1942
  -2.750  -0.4854   0.01225   0.00489   0.0352   1.0000   0.2993
  -2.500  -0.4708   0.01193   0.00487   0.0374   1.0000   0.3813
  -2.250  -0.4336   0.01169   0.00494   0.0350   0.9944   0.4527
  -2.000  -0.4027   0.01118   0.00501   0.0341   0.9873   0.5623
  -1.750  -0.3707   0.01081   0.00520   0.0334   0.9803   0.6882
  -1.500  -0.3287   0.01064   0.00537   0.0310   0.9732   0.7742
  -1.250  -0.2828   0.01053   0.00545   0.0280   0.9646   0.8429
  -1.000  -0.2227   0.01056   0.00555   0.0222   0.9596   0.8889
  -0.750  -0.1498   0.01090   0.00592   0.0140   0.9584   0.9226
  -0.500  -0.0897   0.01107   0.00606   0.0082   0.9512   0.9350
  -0.250  -0.0331   0.01111   0.00606   0.0029   0.9454   0.9433
   0.000   0.0212   0.01111   0.00605  -0.0020   0.9366   0.9482
   0.250   0.0803   0.01108   0.00601  -0.0080   0.9283   0.9543
   0.500   0.1384   0.01102   0.00596  -0.0139   0.9182   0.9596
   0.750   0.1998   0.01087   0.00581  -0.0206   0.9031   0.9614
   1.000   0.2575   0.01075   0.00567  -0.0266   0.8791   0.9631
   1.250   0.3060   0.01073   0.00560  -0.0307   0.8482   0.9669
   1.500   0.3464   0.01077   0.00557  -0.0330   0.8126   0.9727
   1.750   0.3888   0.01078   0.00545  -0.0357   0.7650   0.9784
   2.000   0.4246   0.01088   0.00530  -0.0369   0.6958   0.9852
   2.250   0.4607   0.01101   0.00511  -0.0385   0.6232   0.9916
   2.500   0.4961   0.01119   0.00496  -0.0402   0.5450   0.9978
   2.750   0.5189   0.01141   0.00491  -0.0394   0.4792   1.0000
   3.000   0.5334   0.01170   0.00494  -0.0370   0.4227   1.0000
   3.250   0.5482   0.01202   0.00501  -0.0348   0.3713   1.0000
   3.500   0.5644   0.01230   0.00514  -0.0327   0.3362   1.0000
   3.750   0.5814   0.01256   0.00530  -0.0308   0.3062   1.0000
   4.000   0.5978   0.01287   0.00548  -0.0288   0.2672   1.0000
   4.250   0.6103   0.01358   0.00566  -0.0262   0.1463   1.0000
   4.500   0.6197   0.01477   0.00646  -0.0230   0.0936   1.0000
   4.750   0.6352   0.01530   0.00697  -0.0207   0.0809   1.0000
   5.000   0.6517   0.01572   0.00739  -0.0185   0.0741   1.0000
   5.250   0.6662   0.01633   0.00798  -0.0160   0.0698   1.0000
   5.500   0.6823   0.01682   0.00851  -0.0138   0.0659   1.0000
   5.750   0.6981   0.01737   0.00906  -0.0115   0.0625   1.0000
   6.000   0.7119   0.01835   0.01000  -0.0090   0.0599   1.0000
   6.250   0.7288   0.01919   0.01088  -0.0068   0.0584   1.0000
   6.500   0.7469   0.01994   0.01170  -0.0049   0.0561   1.0000
   6.750   0.7651   0.02075   0.01257  -0.0031   0.0536   1.0000
   7.000   0.7838   0.02176   0.01362  -0.0014   0.0519   1.0000
   7.250   0.8031   0.02306   0.01494   0.0001   0.0503   1.0000
   7.500   0.8227   0.02530   0.01728   0.0014   0.0485   1.0000
   7.750   0.8403   0.02625   0.01847   0.0034   0.0471   1.0000
   8.000   0.8577   0.02788   0.02034   0.0053   0.0461   1.0000
   8.250   0.8733   0.02974   0.02249   0.0075   0.0452   1.0000
   8.500   0.8871   0.03151   0.02451   0.0099   0.0440   1.0000
   8.750   0.9000   0.03328   0.02647   0.0121   0.0427   1.0000
   9.000   0.9132   0.03511   0.02843   0.0142   0.0416   1.0000
   9.250   0.9146   0.03844   0.03228   0.0181   0.0422   1.0000
   9.500   0.9049   0.04328   0.03772   0.0231   0.0440   1.0000
   9.750   0.8951   0.04787   0.04269   0.0272   0.0460   1.0000
  10.000   0.8873   0.05200   0.04703   0.0304   0.0473   1.0000
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