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MH 23 8% (mh23-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: MH 23 8% (mh23-il)
Reynolds number: 200,000
Max Cl/Cd: 66.86 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh23-il-200000.txt
Download as CSV file: xf-mh23-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 23  8%                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4851   0.08844   0.08499  -0.0282   1.0000   0.0352
  -8.500  -0.4870   0.08494   0.08153  -0.0297   1.0000   0.0359
  -8.250  -0.4928   0.08120   0.07784  -0.0315   1.0000   0.0367
  -8.000  -0.5046   0.07774   0.07445  -0.0330   1.0000   0.0369
  -7.750  -0.5207   0.07498   0.07173  -0.0318   1.0000   0.0370
  -7.500  -0.5326   0.07192   0.06867  -0.0308   1.0000   0.0374
  -7.250  -0.5434   0.06878   0.06550  -0.0295   1.0000   0.0381
  -7.000  -0.5508   0.06568   0.06234  -0.0280   1.0000   0.0389
  -6.750  -0.5560   0.06237   0.05892  -0.0265   1.0000   0.0403
  -6.500  -0.5586   0.06061   0.05673  -0.0243   1.0000   0.0426
  -6.250  -0.5584   0.05867   0.05451  -0.0214   1.0000   0.0429
  -6.000  -0.5641   0.05141   0.04729  -0.0199   1.0000   0.0443
  -5.750  -0.5576   0.04844   0.04436  -0.0180   1.0000   0.0457
  -5.500  -0.5501   0.04593   0.04177  -0.0160   1.0000   0.0478
  -5.250  -0.5412   0.04338   0.03904  -0.0139   1.0000   0.0509
  -5.000  -0.5346   0.04061   0.03563  -0.0107   1.0000   0.0572
  -4.750  -0.5222   0.03754   0.03265  -0.0093   1.0000   0.0595
  -4.500  -0.5085   0.03550   0.03046  -0.0074   1.0000   0.0643
  -4.250  -0.4786   0.02067   0.01575  -0.0066   1.0000   0.0728
  -4.000  -0.4654   0.01876   0.01351  -0.0046   1.0000   0.0840
  -3.750  -0.4519   0.02568   0.01921   0.0006   1.0000   0.0394
  -3.500  -0.4267   0.02231   0.01526   0.0031   1.0000   0.0266
  -3.250  -0.4022   0.01988   0.01235   0.0050   1.0000   0.0227
  -3.000  -0.3784   0.01870   0.01089   0.0066   1.0000   0.0208
  -2.750  -0.3561   0.01787   0.00995   0.0079   1.0000   0.0202
  -2.500  -0.3343   0.01682   0.00876   0.0093   1.0000   0.0199
  -2.250  -0.3139   0.01570   0.00759   0.0109   1.0000   0.0201
  -2.000  -0.2922   0.01470   0.00650   0.0121   0.9995   0.0214
  -1.750  -0.1208   0.01133   0.00660  -0.0151   1.0000   1.0000
  -1.500  -0.1095   0.01133   0.00647  -0.0123   1.0000   1.0000
  -1.250  -0.0974   0.01137   0.00638  -0.0098   1.0000   1.0000
  -1.000  -0.0835   0.01144   0.00630  -0.0075   0.9998   1.0000
  -0.750  -0.0365   0.01166   0.00638  -0.0120   0.9946   1.0000
  -0.500   0.0073   0.01179   0.00640  -0.0158   0.9884   1.0000
  -0.250   0.0554   0.01199   0.00648  -0.0204   0.9829   1.0000
   0.000   0.0986   0.01207   0.00650  -0.0239   0.9757   1.0000
   0.250   0.1484   0.01217   0.00655  -0.0287   0.9702   1.0000
   0.500   0.1906   0.01215   0.00653  -0.0319   0.9620   1.0000
   0.750   0.2442   0.01211   0.00650  -0.0373   0.9568   1.0000
   1.000   0.2871   0.01200   0.00641  -0.0404   0.9480   1.0000
   1.250   0.3453   0.01178   0.00625  -0.0466   0.9435   1.0000
   1.500   0.3902   0.01152   0.00604  -0.0499   0.9341   1.0000
   1.750   0.4345   0.01118   0.00580  -0.0529   0.9248   1.0000
   2.000   0.4717   0.01079   0.00549  -0.0543   0.9128   1.0000
   2.250   0.4988   0.01044   0.00521  -0.0534   0.8968   1.0000
   2.500   0.5231   0.01007   0.00490  -0.0519   0.8784   1.0000
   2.750   0.5427   0.00977   0.00473  -0.0495   0.8537   1.0000
   3.000   0.5650   0.00946   0.00446  -0.0475   0.8231   1.0000
   3.250   0.5891   0.00918   0.00413  -0.0458   0.7776   1.0000
   3.500   0.6124   0.00916   0.00390  -0.0440   0.7122   1.0000
   3.750   0.6317   0.00949   0.00391  -0.0417   0.6358   1.0000
   4.000   0.6489   0.01000   0.00409  -0.0393   0.5574   1.0000
   4.250   0.6652   0.01059   0.00441  -0.0370   0.4800   1.0000
   4.500   0.6811   0.01124   0.00474  -0.0348   0.4037   1.0000
   4.750   0.6972   0.01191   0.00512  -0.0327   0.3320   1.0000
   5.000   0.7131   0.01266   0.00555  -0.0307   0.2607   1.0000
   5.250   0.7291   0.01346   0.00604  -0.0288   0.1928   1.0000
   5.500   0.7449   0.01437   0.00662  -0.0269   0.1303   1.0000
   5.750   0.7598   0.01544   0.00745  -0.0247   0.0848   1.0000
   6.000   0.7737   0.01669   0.00859  -0.0220   0.0595   1.0000
   6.250   0.7877   0.01802   0.00994  -0.0194   0.0485   1.0000
   6.500   0.8020   0.01946   0.01137  -0.0170   0.0390   1.0000
   6.750   0.8210   0.02004   0.01218  -0.0154   0.0298   1.0000
   7.000   0.8362   0.02114   0.01333  -0.0134   0.0183   1.0000
   7.250   0.8505   0.02361   0.01602  -0.0106   0.0152   1.0000
   7.500   0.8678   0.02580   0.01849  -0.0084   0.0138   1.0000
   7.750   0.8833   0.02840   0.02144  -0.0059   0.0129   1.0000
   8.000   0.8952   0.03141   0.02487  -0.0030   0.0127   1.0000
   8.250   0.9024   0.03477   0.02870   0.0004   0.0129   1.0000
   8.500   0.9044   0.03841   0.03280   0.0042   0.0132   1.0000
   8.750   0.9017   0.04228   0.03710   0.0082   0.0137   1.0000
   9.000   0.8953   0.04605   0.04123   0.0121   0.0142   1.0000
   9.250   0.8842   0.04985   0.04534   0.0160   0.0146   1.0000
   9.500   0.8697   0.05321   0.04892   0.0198   0.0149   1.0000
   9.750   0.8512   0.05634   0.05223   0.0235   0.0152   1.0000
  10.000   0.8325   0.05978   0.05583   0.0258   0.0154   1.0000
  10.250   0.8127   0.06382   0.06002   0.0264   0.0156   1.0000
  10.500   0.7953   0.06824   0.06457   0.0251   0.0157   1.0000
  10.750   0.7758   0.07426   0.07072   0.0213   0.0157   1.0000
  11.000   0.7573   0.08242   0.07897   0.0149   0.0157   1.0000
  11.250   0.7427   0.09214   0.08871   0.0080   0.0159   1.0000
  11.500   0.7315   0.10013   0.09667   0.0038   0.0164   1.0000
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