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

GOE 121 (MVA H.1) AIRFOIL (goe121-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 121 (MVA H.1) AIRFOIL (goe121-il)
Reynolds number: 50,000
Max Cl/Cd: 40.59 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe121-il-50000.txt
Download as CSV file: xf-goe121-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 121 (MVA H.1) AIRFOIL                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3686   0.11477   0.10755  -0.0160   1.0000   0.1334
  -8.500  -0.3715   0.11391   0.10678  -0.0180   1.0000   0.1367
  -8.250  -0.3834   0.11480   0.10783  -0.0206   1.0000   0.1378
  -8.000  -0.3557   0.10625   0.09921  -0.0180   1.0000   0.1435
  -7.750  -0.3534   0.10405   0.09708  -0.0185   1.0000   0.1491
  -7.500  -0.3627   0.10427   0.09745  -0.0212   1.0000   0.1522
  -7.250  -0.3525   0.09936   0.09259  -0.0199   1.0000   0.1555
  -7.000  -0.3438   0.09597   0.08923  -0.0191   1.0000   0.1609
  -6.750  -0.3449   0.09490   0.08827  -0.0221   1.0000   0.1664
  -6.500  -0.3415   0.09206   0.08550  -0.0236   1.0000   0.1692
  -6.250  -0.3322   0.08818   0.08166  -0.0209   1.0000   0.1747
  -6.000  -0.3296   0.08735   0.08089  -0.0258   1.0000   0.1821
  -5.500  -0.3174   0.08110   0.07476  -0.0240   1.0000   0.1956
  -5.250  -0.3121   0.07816   0.07189  -0.0239   1.0000   0.2017
  -5.000  -0.3015   0.07696   0.07067  -0.0290   1.0000   0.2135
  -4.750  -0.2986   0.07301   0.06683  -0.0238   1.0000   0.2190
  -4.250  -0.2732   0.06876   0.06250  -0.0296   1.0000   0.2436
  -4.000  -0.2693   0.06536   0.05920  -0.0256   1.0000   0.2511
  -3.750  -0.2569   0.06276   0.05660  -0.0263   1.0000   0.2650
  -3.500  -0.2366   0.06069   0.05446  -0.0304   1.0000   0.2878
  -3.250  -0.2281   0.05767   0.05149  -0.0284   1.0000   0.3048
  -3.000  -0.2158   0.05516   0.04901  -0.0279   1.0000   0.3337
  -2.750  -0.2066   0.05257   0.04646  -0.0259   1.0000   0.3659
  -1.750  -0.1792   0.04265   0.03690  -0.0122   1.0000   0.5584
  -1.500  -0.1647   0.04015   0.03445  -0.0103   1.0000   0.5921
  -1.250  -0.1330   0.03777   0.03201  -0.0136   1.0000   0.6129
  -1.000   0.0899   0.03547   0.02698  -0.0679   1.0000   0.2092
  -0.750   0.1280   0.03348   0.02453  -0.0708   1.0000   0.1920
  -0.500   0.1642   0.03191   0.02242  -0.0731   1.0000   0.1848
  -0.250   0.1914   0.03102   0.02138  -0.0741   1.0000   0.1894
   0.000   0.2200   0.03040   0.02046  -0.0753   1.0000   0.1995
   0.250   0.2464   0.02990   0.01982  -0.0762   1.0000   0.2102
   0.500   0.2724   0.02962   0.01937  -0.0771   1.0000   0.2292
   0.750   0.3364   0.02904   0.01868  -0.0844   0.9855   0.2990
   1.000   0.3978   0.02841   0.01807  -0.0911   0.9687   0.3541
   1.250   0.4561   0.02784   0.01774  -0.0971   0.9497   0.4087
   1.500   0.5090   0.02617   0.01734  -0.1013   0.9310   1.0000
   1.750   0.5586   0.02664   0.01727  -0.1049   0.9071   1.0000
   2.000   0.6152   0.02680   0.01722  -0.1097   0.8865   1.0000
   2.250   0.6567   0.02707   0.01741  -0.1117   0.8621   1.0000
   2.500   0.7050   0.02695   0.01727  -0.1142   0.8419   1.0000
   2.750   0.7382   0.02713   0.01744  -0.1142   0.8174   1.0000
   3.000   0.7796   0.02681   0.01710  -0.1148   0.7971   1.0000
   3.250   0.8091   0.02687   0.01720  -0.1138   0.7730   1.0000
   3.500   0.8453   0.02643   0.01676  -0.1130   0.7526   1.0000
   3.750   0.8720   0.02652   0.01686  -0.1114   0.7283   1.0000
   4.000   0.9035   0.02622   0.01657  -0.1099   0.7071   1.0000
   4.250   0.9274   0.02653   0.01691  -0.1081   0.6830   1.0000
   4.500   0.9549   0.02666   0.01705  -0.1067   0.6627   1.0000
   4.750   0.9804   0.02710   0.01753  -0.1054   0.6432   1.0000
   5.000   1.0040   0.02774   0.01830  -0.1043   0.6237   1.0000
   5.250   1.0297   0.02820   0.01884  -0.1031   0.6056   1.0000
   5.500   1.0569   0.02848   0.01916  -0.1018   0.5876   1.0000
   5.750   1.0804   0.02910   0.01988  -0.1004   0.5674   1.0000
   6.000   1.1051   0.02957   0.02044  -0.0988   0.5464   1.0000
   6.250   1.1317   0.02998   0.02085  -0.0973   0.5259   1.0000
   6.500   1.1525   0.03103   0.02208  -0.0958   0.5043   1.0000
   6.750   1.1779   0.03146   0.02251  -0.0941   0.4818   1.0000
   7.000   1.1983   0.03156   0.02265  -0.0916   0.4521   1.0000
   7.250   1.2178   0.03160   0.02276  -0.0891   0.4230   1.0000
   7.500   1.2347   0.03111   0.02224  -0.0860   0.3886   1.0000
   7.750   1.2464   0.03071   0.02189  -0.0824   0.3488   1.0000
   8.000   1.2494   0.03080   0.02204  -0.0778   0.2921   1.0000
   8.250   1.2394   0.03251   0.02310  -0.0723   0.1965   1.0000
   8.500   1.2391   0.03586   0.02575  -0.0686   0.1476   1.0000
   8.750   1.2510   0.03840   0.02814  -0.0662   0.1265   1.0000
   9.000   1.2658   0.04054   0.03026  -0.0641   0.1133   1.0000
   9.250   1.2816   0.04268   0.03227  -0.0624   0.1037   1.0000
   9.500   1.3054   0.04528   0.03507  -0.0611   0.0971   1.0000
   9.750   1.3376   0.04871   0.03837  -0.0611   0.0924   1.0000
  10.000   1.3542   0.05216   0.04237  -0.0595   0.0899   1.0000
  10.250   1.3653   0.05575   0.04641  -0.0576   0.0872   1.0000
  10.500   1.3738   0.05947   0.05048  -0.0558   0.0851   1.0000
  10.750   1.3766   0.06357   0.05498  -0.0537   0.0843   1.0000
  11.000   1.3691   0.06797   0.05986  -0.0511   0.0847   1.0000
  11.250   1.3527   0.07239   0.06471  -0.0484   0.0853   1.0000
  11.500   1.3292   0.07680   0.06946  -0.0458   0.0861   1.0000
  11.750   1.3022   0.08179   0.07474  -0.0446   0.0869   1.0000
  12.000   1.2728   0.08763   0.08083  -0.0452   0.0878   1.0000
  12.250   1.2418   0.09453   0.08794  -0.0476   0.0889   1.0000
  12.500   1.2116   0.10253   0.09608  -0.0515   0.0902   1.0000
  12.750   1.1872   0.11097   0.10460  -0.0560   0.0915   1.0000
  13.000   1.1714   0.11894   0.11260  -0.0597   0.0925   1.0000
<< Back to GOE 121 (MVA H.1) AIRFOIL (goe121-il)

Polar data table (+)

Polar graphs


<< Back to GOE 121 (MVA H.1) AIRFOIL (goe121-il)