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

GOE 276 (DAIMLER VII) AIRFOIL (goe276-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 276 (DAIMLER VII) AIRFOIL (goe276-il)
Reynolds number: 200,000
Max Cl/Cd: 72.3 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe276-il-200000.txt
Download as CSV file: xf-goe276-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 276 (DAIMLER VII) AIRFOIL                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3959   0.11056   0.10692  -0.0179   1.0000   0.0285
  -8.750  -0.3971   0.10861   0.10502  -0.0209   1.0000   0.0287
  -8.500  -0.3990   0.10643   0.10291  -0.0230   1.0000   0.0288
  -8.250  -0.3971   0.10367   0.10019  -0.0261   1.0000   0.0289
  -8.000  -0.3940   0.09815   0.09472  -0.0249   1.0000   0.0293
  -7.750  -0.3866   0.09386   0.09044  -0.0215   1.0000   0.0298
  -7.500  -0.3813   0.09063   0.08723  -0.0210   1.0000   0.0303
  -7.250  -0.3767   0.08762   0.08425  -0.0214   1.0000   0.0308
  -7.000  -0.3717   0.08461   0.08129  -0.0225   1.0000   0.0314
  -6.750  -0.3661   0.08157   0.07826  -0.0239   1.0000   0.0322
  -6.500  -0.3596   0.07848   0.07520  -0.0256   1.0000   0.0330
  -6.250  -0.3523   0.07537   0.07211  -0.0275   1.0000   0.0340
  -6.000  -0.3436   0.07228   0.06903  -0.0298   1.0000   0.0353
  -5.750  -0.3158   0.06941   0.06598  -0.0390   1.0000   0.0374
  -5.500  -0.2957   0.06667   0.06302  -0.0428   1.0000   0.0377
  -5.250  -0.2966   0.06073   0.05721  -0.0414   1.0000   0.0386
  -5.000  -0.2905   0.05790   0.05442  -0.0399   1.0000   0.0393
  -4.750  -0.2792   0.05525   0.05176  -0.0396   1.0000   0.0404
  -4.500  -0.2627   0.05247   0.04892  -0.0405   1.0000   0.0420
  -4.250  -0.2305   0.04890   0.04518  -0.0446   0.9986   0.0450
  -4.000  -0.1711   0.04359   0.03929  -0.0539   0.9955   0.0503
  -3.750  -0.1407   0.04020   0.03597  -0.0570   0.9925   0.0526
  -3.500  -0.1008   0.03727   0.03285  -0.0612   0.9887   0.0572
  -3.250  -0.0504   0.03379   0.02882  -0.0666   0.9856   0.0642
  -3.000  -0.0137   0.03123   0.02627  -0.0701   0.9814   0.0694
  -2.750   0.0298   0.02873   0.02340  -0.0738   0.9766   0.0794
  -2.500   0.0733   0.02710   0.02139  -0.0773   0.9733   0.0922
  -2.250   0.1112   0.02541   0.01950  -0.0800   0.9687   0.1062
  -2.000   0.1477   0.02317   0.01726  -0.0828   0.9635   0.1225
  -1.750   0.1881   0.02146   0.01556  -0.0864   0.9599   0.1539
  -1.500   0.2257   0.02095   0.01486  -0.0886   0.9535   0.1942
  -1.250   0.2655   0.01868   0.01268  -0.0917   0.9486   0.2259
  -1.000   0.3226   0.01702   0.00991  -0.0923   0.9449   0.0927
  -0.750   0.3588   0.01574   0.00856  -0.0932   0.9353   0.0829
  -0.500   0.3979   0.01499   0.00762  -0.0944   0.9252   0.0747
  -0.250   0.4350   0.01386   0.00645  -0.0955   0.9117   0.0714
   0.000   0.4682   0.01311   0.00566  -0.0958   0.8937   0.0698
   0.250   0.4983   0.01257   0.00509  -0.0955   0.8735   0.0696
   0.500   0.5251   0.01220   0.00469  -0.0947   0.8493   0.0706
   0.750   0.5513   0.01192   0.00438  -0.0938   0.8210   0.0735
   1.000   0.5770   0.01172   0.00409  -0.0927   0.7831   0.0770
   1.250   0.6019   0.01154   0.00370  -0.0915   0.7200   0.0806
   1.500   0.6250   0.01176   0.00343  -0.0899   0.6365   0.0869
   1.750   0.6494   0.01200   0.00338  -0.0889   0.5821   0.1165
   2.000   0.6680   0.01040   0.00341  -0.0869   0.5510   1.0000
   2.250   0.6942   0.01077   0.00351  -0.0865   0.5265   1.0000
   2.500   0.7205   0.01113   0.00366  -0.0861   0.5077   1.0000
   2.750   0.7469   0.01148   0.00384  -0.0858   0.4916   1.0000
   3.000   0.7735   0.01182   0.00406  -0.0856   0.4770   1.0000
   3.250   0.8002   0.01214   0.00429  -0.0854   0.4639   1.0000
   3.500   0.8269   0.01247   0.00455  -0.0852   0.4515   1.0000
   3.750   0.8534   0.01281   0.00481  -0.0850   0.4389   1.0000
   4.000   0.8799   0.01316   0.00512  -0.0847   0.4273   1.0000
   4.250   0.9064   0.01351   0.00541  -0.0845   0.4163   1.0000
   4.500   0.9330   0.01378   0.00572  -0.0843   0.4050   1.0000
   4.750   0.9594   0.01409   0.00606  -0.0841   0.3940   1.0000
   5.000   0.9857   0.01441   0.00637  -0.0838   0.3833   1.0000
   5.250   1.0117   0.01470   0.00667  -0.0835   0.3719   1.0000
   5.500   1.0378   0.01493   0.00699  -0.0832   0.3597   1.0000
   5.750   1.0636   0.01518   0.00732  -0.0828   0.3472   1.0000
   6.000   1.0890   0.01540   0.00759  -0.0824   0.3329   1.0000
   6.250   1.1140   0.01558   0.00780  -0.0819   0.3159   1.0000
   6.500   1.1384   0.01580   0.00802  -0.0813   0.2961   1.0000
   6.750   1.1626   0.01608   0.00833  -0.0808   0.2737   1.0000
   7.000   1.1862   0.01650   0.00872  -0.0801   0.2533   1.0000
   7.250   1.2092   0.01697   0.00916  -0.0795   0.2292   1.0000
   7.500   1.2322   0.01748   0.00964  -0.0788   0.2039   1.0000
   7.750   1.2548   0.01805   0.01019  -0.0781   0.1748   1.0000
   8.000   1.2737   0.01913   0.01096  -0.0770   0.1225   1.0000
   8.250   1.2831   0.02160   0.01277  -0.0746   0.0441   1.0000
   8.750   1.3159   0.02429   0.01571  -0.0709   0.0333   1.0000
   9.000   1.3310   0.02557   0.01713  -0.0691   0.0309   1.0000
   9.250   1.3413   0.02722   0.01889  -0.0666   0.0287   1.0000
   9.500   1.3487   0.02907   0.02084  -0.0637   0.0272   1.0000
   9.750   1.3597   0.03054   0.02244  -0.0613   0.0264   1.0000
  10.000   1.3687   0.03218   0.02420  -0.0587   0.0256   1.0000
  10.250   1.3770   0.03392   0.02606  -0.0560   0.0250   1.0000
  10.500   1.3843   0.03574   0.02800  -0.0531   0.0245   1.0000
  10.750   1.3926   0.03771   0.03009  -0.0506   0.0241   1.0000
  11.000   1.4014   0.03983   0.03237  -0.0482   0.0237   1.0000
  11.250   1.4086   0.04196   0.03465  -0.0459   0.0231   1.0000
  11.500   1.4140   0.04418   0.03699  -0.0437   0.0224   1.0000
  11.750   1.4203   0.04693   0.03985  -0.0418   0.0217   1.0000
  12.000   1.4250   0.05029   0.04341  -0.0399   0.0213   1.0000
  12.250   1.4251   0.05346   0.04683  -0.0377   0.0213   1.0000
  12.500   1.4215   0.05665   0.05027  -0.0356   0.0213   1.0000
  12.750   1.4148   0.05997   0.05384  -0.0337   0.0214   1.0000
  13.000   1.4053   0.06352   0.05766  -0.0323   0.0215   1.0000
  13.250   1.3929   0.06744   0.06185  -0.0315   0.0216   1.0000
  13.500   1.3774   0.07199   0.06669  -0.0316   0.0219   1.0000
  13.750   1.3592   0.07738   0.07238  -0.0326   0.0222   1.0000
  14.000   1.3375   0.08380   0.07910  -0.0349   0.0226   1.0000
  14.250   1.3110   0.09154   0.08715  -0.0391   0.0229   1.0000
  14.500   1.2847   0.10011   0.09599  -0.0443   0.0232   1.0000
  14.750   1.2588   0.10929   0.10540  -0.0503   0.0235   1.0000
  15.000   1.2325   0.11929   0.11561  -0.0570   0.0239   1.0000
  15.250   1.2051   0.13030   0.12680  -0.0646   0.0242   1.0000
  15.500   1.1755   0.14296   0.13962  -0.0733   0.0247   1.0000
  15.750   1.1402   0.15908   0.15585  -0.0841   0.0255   1.0000
  16.000   1.1130   0.17411   0.17087  -0.0929   0.0268   1.0000
  16.250   1.1124   0.17909   0.17586  -0.0946   0.0278   1.0000
  16.500   0.8194   0.17353   0.17043  -0.0699   0.0351   1.0000
<< Back to GOE 276 (DAIMLER VII) AIRFOIL (goe276-il)

Polar data table (+)

Polar graphs


<< Back to GOE 276 (DAIMLER VII) AIRFOIL (goe276-il)