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

MH 117 9.8% (mh117-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MH 117 9.8% (mh117-il)
Reynolds number: 50,000
Max Cl/Cd: 36.13 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh117-il-50000.txt
Download as CSV file: xf-mh117-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 117  9.8%                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4047   0.09994   0.09337  -0.0201   1.0000   0.2516
  -8.000  -0.3995   0.09653   0.09003  -0.0192   1.0000   0.2656
  -7.750  -0.3980   0.09352   0.08710  -0.0183   1.0000   0.2799
  -7.500  -0.3864   0.08951   0.08312  -0.0171   1.0000   0.2924
  -7.250  -0.3817   0.08595   0.07963  -0.0161   1.0000   0.3021
  -7.000  -0.3833   0.08327   0.07705  -0.0145   1.0000   0.3180
  -6.750  -0.4541   0.06793   0.06183  -0.0427   1.0000   0.1399
  -6.500  -0.4587   0.06065   0.05438  -0.0463   1.0000   0.1236
  -6.250  -0.4557   0.05622   0.04981  -0.0464   1.0000   0.1224
  -6.000  -0.4524   0.05145   0.04480  -0.0470   1.0000   0.1220
  -5.750  -0.4461   0.04656   0.03946  -0.0477   1.0000   0.1229
  -5.500  -0.4353   0.04156   0.03376  -0.0483   1.0000   0.1244
  -5.250  -0.4212   0.03918   0.03139  -0.0469   1.0000   0.1352
  -5.000  -0.4051   0.03621   0.02809  -0.0461   1.0000   0.1472
  -4.750  -0.3869   0.03361   0.02499  -0.0455   1.0000   0.1654
  -4.500  -0.3684   0.03160   0.02286  -0.0445   1.0000   0.1845
  -4.250  -0.3484   0.02987   0.02089  -0.0436   1.0000   0.2052
  -4.000  -0.3286   0.02851   0.01939  -0.0426   1.0000   0.2288
  -3.750  -0.3065   0.02710   0.01767  -0.0419   1.0000   0.2513
  -3.500  -0.2863   0.02612   0.01666  -0.0409   1.0000   0.2769
  -3.250  -0.2647   0.02511   0.01558  -0.0401   1.0000   0.3034
  -3.000  -0.2432   0.02431   0.01470  -0.0394   1.0000   0.3381
  -2.750  -0.2233   0.02356   0.01408  -0.0382   1.0000   0.3748
  -2.500  -0.2032   0.02292   0.01352  -0.0370   1.0000   0.4231
  -2.250  -0.1830   0.02231   0.01314  -0.0358   1.0000   0.4839
  -2.000  -0.1649   0.02170   0.01286  -0.0339   1.0000   0.5622
  -1.750  -0.1513   0.02103   0.01273  -0.0306   1.0000   0.6679
  -1.500  -0.1455   0.02013   0.01265  -0.0244   1.0000   0.8536
  -1.250  -0.1024   0.01992   0.01184  -0.0300   1.0000   1.0000
  -1.000  -0.0803   0.02030   0.01179  -0.0306   1.0000   1.0000
  -0.750  -0.0596   0.02073   0.01187  -0.0308   1.0000   1.0000
  -0.500  -0.0395   0.02121   0.01207  -0.0309   1.0000   1.0000
  -0.250  -0.0198   0.02172   0.01235  -0.0309   1.0000   1.0000
   0.000  -0.0003   0.02228   0.01270  -0.0309   1.0000   1.0000
   0.250   0.0189   0.02288   0.01309  -0.0308   1.0000   1.0000
   0.500   0.0576   0.02393   0.01394  -0.0345   0.9912   1.0000
   0.750   0.1015   0.02508   0.01491  -0.0390   0.9790   1.0000
   1.000   0.1443   0.02618   0.01586  -0.0432   0.9665   1.0000
   1.250   0.1862   0.02723   0.01680  -0.0470   0.9536   1.0000
   1.500   0.2271   0.02823   0.01772  -0.0506   0.9403   1.0000
   1.750   0.2669   0.02919   0.01862  -0.0539   0.9263   1.0000
   2.000   0.3050   0.03010   0.01950  -0.0567   0.9121   1.0000
   2.250   0.3408   0.03096   0.02035  -0.0590   0.8971   1.0000
   2.500   0.3751   0.03179   0.02121  -0.0610   0.8814   1.0000
   2.750   0.4091   0.03260   0.02205  -0.0627   0.8651   1.0000
   3.000   0.4445   0.03338   0.02288  -0.0645   0.8484   1.0000
   3.250   0.4830   0.03408   0.02368  -0.0665   0.8315   1.0000
   3.500   0.5271   0.03460   0.02431  -0.0690   0.8144   1.0000
   3.750   0.5578   0.03524   0.02504  -0.0695   0.7955   1.0000
   4.000   0.5940   0.03566   0.02562  -0.0705   0.7759   1.0000
   4.250   0.6479   0.03542   0.02559  -0.0730   0.7576   1.0000
   4.500   0.6767   0.03576   0.02607  -0.0724   0.7361   1.0000
   4.750   0.7243   0.03509   0.02565  -0.0730   0.7164   1.0000
   5.000   0.7574   0.03486   0.02561  -0.0720   0.6941   1.0000
   5.250   0.8054   0.03349   0.02446  -0.0715   0.6729   1.0000
   5.500   0.8368   0.03289   0.02409  -0.0696   0.6476   1.0000
   5.750   0.8725   0.03183   0.02320  -0.0676   0.6210   1.0000
   6.000   0.9080   0.03060   0.02210  -0.0653   0.5916   1.0000
   6.250   0.9443   0.02918   0.02077  -0.0630   0.5590   1.0000
   6.500   0.9689   0.02867   0.02030  -0.0601   0.5205   1.0000
   6.750   0.9939   0.02821   0.01976  -0.0573   0.4786   1.0000
   7.000   1.0165   0.02817   0.01952  -0.0546   0.4338   1.0000
   7.250   1.0367   0.02869   0.01976  -0.0520   0.3886   1.0000
   7.500   1.0540   0.02980   0.02069  -0.0496   0.3450   1.0000
   7.750   1.0721   0.03123   0.02186  -0.0476   0.3056   1.0000
   8.000   1.0894   0.03303   0.02360  -0.0457   0.2721   1.0000
   8.250   1.1080   0.03497   0.02546  -0.0441   0.2436   1.0000
   8.500   1.1284   0.03697   0.02733  -0.0428   0.2192   1.0000
   8.750   1.1455   0.03945   0.03004  -0.0412   0.2007   1.0000
   9.000   1.1626   0.04165   0.03230  -0.0398   0.1829   1.0000
   9.250   1.1786   0.04458   0.03549  -0.0383   0.1702   1.0000
   9.500   1.1904   0.04761   0.03887  -0.0365   0.1589   1.0000
   9.750   1.2015   0.05052   0.04197  -0.0348   0.1479   1.0000
  10.000   1.2158   0.05393   0.04556  -0.0334   0.1396   1.0000
  10.250   1.2102   0.05803   0.05018  -0.0307   0.1351   1.0000
  10.500   1.2010   0.06185   0.05444  -0.0280   0.1306   1.0000
  10.750   1.2100   0.06547   0.05808  -0.0268   0.1233   1.0000
  11.000   1.1915   0.07004   0.06305  -0.0243   0.1229   1.0000
  11.250   1.1691   0.07453   0.06781  -0.0220   0.1230   1.0000
  11.500   1.1445   0.07919   0.07265  -0.0205   0.1234   1.0000
  11.750   1.1203   0.08450   0.07811  -0.0203   0.1240   1.0000
  12.000   1.0984   0.09054   0.08425  -0.0214   0.1247   1.0000
  12.250   0.6879   0.14971   0.14312  -0.0641   0.3328   1.0000
<< Back to MH 117 9.8% (mh117-il)

Polar data table (+)

Polar graphs


<< Back to MH 117 9.8% (mh117-il)