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

MH 121 8.76% (mh121-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: MH 121 8.76% (mh121-il)
Reynolds number: 1,000,000
Max Cl/Cd: 117.71 at α=1.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-mh121-il-1000000-n5.txt
Download as CSV file: xf-mh121-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 121  8.76%                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3646   0.08955   0.08787  -0.0567   0.9918   0.0016
  -9.000  -0.3591   0.08482   0.08315  -0.0602   0.9896   0.0016
  -8.750  -0.3522   0.08001   0.07835  -0.0643   0.9870   0.0015
  -8.500  -0.3448   0.07462   0.07297  -0.0695   0.9842   0.0015
  -8.000  -0.3368   0.06378   0.06217  -0.0800   0.9710   0.0014
  -7.750  -0.3295   0.05391   0.05226  -0.0958   0.9575   0.0014
  -7.500  -0.3133   0.04443   0.04254  -0.1086   0.9483   0.0013
  -7.250  -0.2878   0.03604   0.03379  -0.1179   0.9421   0.0012
  -7.000  -0.2660   0.02796   0.02520  -0.1226   0.9322   0.0011
  -6.750  -0.2504   0.01939   0.01585  -0.1233   0.9204   0.0010
  -6.500  -0.2289   0.01561   0.01153  -0.1230   0.9125   0.0011
  -6.250  -0.2044   0.01358   0.00916  -0.1229   0.9060   0.0012
  -6.000  -0.1791   0.01239   0.00776  -0.1228   0.9004   0.0014
  -5.750  -0.1532   0.01156   0.00678  -0.1227   0.8952   0.0017
  -5.500  -0.1265   0.01097   0.00607  -0.1227   0.8910   0.0019
  -5.250  -0.0992   0.01095   0.00598  -0.1229   0.8866   0.0022
  -5.000  -0.0725   0.01049   0.00544  -0.1229   0.8822   0.0023
  -4.750  -0.0473   0.00927   0.00401  -0.1227   0.8782   0.0027
  -4.500  -0.0207   0.00875   0.00342  -0.1228   0.8744   0.0034
  -4.250   0.0065   0.00845   0.00308  -0.1229   0.8707   0.0041
  -4.000   0.0341   0.00801   0.00252  -0.1229   0.8672   0.0041
  -3.750   0.0619   0.00767   0.00206  -0.1230   0.8641   0.0042
  -3.500   0.0898   0.00744   0.00173  -0.1231   0.8609   0.0044
  -3.250   0.1179   0.00726   0.00148  -0.1232   0.8576   0.0058
  -3.000   0.1453   0.00675   0.00123  -0.1235   0.8545   0.0882
  -2.750   0.1724   0.00605   0.00102  -0.1241   0.8517   0.2447
  -2.500   0.2000   0.00574   0.00092  -0.1244   0.8488   0.3222
  -2.250   0.2277   0.00552   0.00086  -0.1246   0.8457   0.3855
  -2.000   0.2557   0.00543   0.00082  -0.1248   0.8428   0.4185
  -1.750   0.2838   0.00536   0.00079  -0.1250   0.8402   0.4503
  -1.500   0.3118   0.00528   0.00077  -0.1252   0.8375   0.4805
  -1.250   0.3398   0.00524   0.00077  -0.1253   0.8345   0.5014
  -1.000   0.3678   0.00521   0.00077  -0.1255   0.8313   0.5163
  -0.750   0.3958   0.00521   0.00076  -0.1256   0.8277   0.5269
  -0.500   0.4236   0.00520   0.00077  -0.1256   0.8232   0.5352
   0.000   0.4786   0.00520   0.00079  -0.1256   0.8110   0.5514
   0.250   0.5059   0.00520   0.00081  -0.1256   0.8034   0.5596
   0.500   0.5330   0.00521   0.00083  -0.1255   0.7952   0.5686
   0.750   0.5599   0.00523   0.00085  -0.1253   0.7854   0.5776
   1.000   0.5865   0.00526   0.00090  -0.1251   0.7729   0.5868
   1.500   0.6368   0.00541   0.00097  -0.1239   0.7240   0.6071
   1.750   0.6578   0.00568   0.00105  -0.1225   0.6676   0.6176
   2.000   0.6744   0.00628   0.00127  -0.1202   0.5716   0.6285
   2.250   0.6783   0.00793   0.00185  -0.1158   0.3247   0.6393
   2.500   0.6875   0.00947   0.00241  -0.1127   0.0986   0.6510
   2.750   0.7126   0.00968   0.00264  -0.1123   0.0794   0.6648
   3.000   0.7343   0.01020   0.00295  -0.1113   0.0240   0.6797
   3.250   0.7582   0.01053   0.00322  -0.1106   0.0041   0.6959
   3.500   0.7833   0.01074   0.00347  -0.1102   0.0015   0.7140
   3.750   0.8085   0.01092   0.00375  -0.1097   0.0012   0.7337
   4.250   0.8577   0.01138   0.00444  -0.1085   0.0011   0.7816
   4.500   0.8812   0.01168   0.00494  -0.1076   0.0011   0.8104
   4.750   0.9034   0.01202   0.00546  -0.1065   0.0011   0.8470
   5.000   0.9210   0.01235   0.00605  -0.1041   0.0011   0.9104
   5.250   0.9412   0.01289   0.00677  -0.1025   0.0011   1.0000
   5.500   0.9615   0.01377   0.00779  -0.1009   0.0011   1.0000
   5.750   0.9808   0.01490   0.00906  -0.0991   0.0012   1.0000
   6.000   1.0004   0.01631   0.01064  -0.0974   0.0013   1.0000
   6.250   1.0217   0.01817   0.01269  -0.0959   0.0014   1.0000
   6.500   1.0447   0.02073   0.01551  -0.0946   0.0015   1.0000
   6.750   1.0662   0.02375   0.01886  -0.0930   0.0017   1.0000
   7.750   1.1214   0.03963   0.03619  -0.0811   0.0032   1.0000
   8.000   1.1277   0.04315   0.03996  -0.0779   0.0031   1.0000
   8.250   1.1306   0.04694   0.04401  -0.0745   0.0030   1.0000
   8.500   1.1305   0.05074   0.04804  -0.0710   0.0030   1.0000
   8.750   1.1273   0.05446   0.05197  -0.0675   0.0029   1.0000
   9.000   1.1207   0.05808   0.05577  -0.0639   0.0028   1.0000
   9.250   1.1073   0.06140   0.05926  -0.0594   0.0028   1.0000
   9.500   1.0897   0.06473   0.06273  -0.0550   0.0028   1.0000
   9.750   1.0717   0.06821   0.06633  -0.0517   0.0028   1.0000
  10.000   1.0506   0.07273   0.07099  -0.0496   0.0028   1.0000
  10.250   1.0282   0.07800   0.07639  -0.0489   0.0028   1.0000
  10.500   1.0067   0.08376   0.08226  -0.0500   0.0028   1.0000
  10.750   0.9824   0.09153   0.09014  -0.0534   0.0029   1.0000
<< Back to MH 121 8.76% (mh121-il)

Polar data table (+)

Polar graphs


<< Back to MH 121 8.76% (mh121-il)