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GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)
Reynolds number: 500,000
Max Cl/Cd: 75.9 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe417a-il-500000.txt
Download as CSV file: xf-goe417a-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 417A (GEW. PLATTE) AIRFOIL                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3693   0.11571   0.11341  -0.0135   1.0000   0.0156
  -9.000  -0.3706   0.11417   0.11189  -0.0129   1.0000   0.0157
  -8.750  -0.3719   0.11254   0.11028  -0.0121   1.0000   0.0157
  -8.000  -0.3712   0.10577   0.10357  -0.0112   0.9989   0.0158
  -7.750  -0.3564   0.10186   0.09966  -0.0138   0.9969   0.0159
  -7.500  -0.3395   0.09848   0.09627  -0.0176   0.9945   0.0161
  -7.250  -0.3224   0.09527   0.09305  -0.0212   0.9913   0.0164
  -7.000  -0.3025   0.09189   0.08966  -0.0254   0.9882   0.0168
  -6.750  -0.2787   0.08833   0.08608  -0.0306   0.9856   0.0172
  -6.500  -0.2526   0.08482   0.08256  -0.0362   0.9838   0.0177
  -6.250  -0.2330   0.08188   0.07960  -0.0398   0.9800   0.0182
  -6.000  -0.2090   0.07895   0.07664  -0.0444   0.9763   0.0184
  -5.750  -0.1751   0.07579   0.07343  -0.0518   0.9735   0.0186
  -5.500  -0.1328   0.07262   0.07019  -0.0614   0.9715   0.0186
  -2.750   0.2436   0.02181   0.01899  -0.1054   0.9362   0.0220
  -2.500   0.2656   0.01897   0.01616  -0.1064   0.9316   0.0222
  -2.250   0.2947   0.01656   0.01371  -0.1087   0.9286   0.0224
  -2.000   0.3274   0.01451   0.01160  -0.1113   0.9260   0.0228
  -1.750   0.3469   0.01336   0.01040  -0.1106   0.9180   0.0233
  -1.500   0.3783   0.01179   0.00874  -0.1122   0.9127   0.0240
  -1.250   0.4043   0.01067   0.00754  -0.1124   0.9051   0.0249
  -1.000   0.4418   0.01020   0.00685  -0.1132   0.8981   0.0257
  -0.750   0.4693   0.00945   0.00595  -0.1129   0.8894   0.0258
  -0.500   0.4925   0.00789   0.00432  -0.1130   0.8806   0.0260
  -0.250   0.5281   0.02176   0.01796  -0.1176   0.8924   0.0261
   0.000   0.5504   0.02034   0.01650  -0.1170   0.8817   0.0264
   0.250   0.5732   0.01921   0.01530  -0.1160   0.8688   0.0268
   0.500   0.5986   0.01815   0.01415  -0.1154   0.8528   0.0274
   0.750   0.6254   0.01717   0.01303  -0.1149   0.8331   0.0283
   1.000   0.6533   0.01641   0.01205  -0.1141   0.8038   0.0298
   1.250   0.6832   0.01649   0.01173  -0.1127   0.7669   0.0306
   1.500   0.7013   0.01493   0.00996  -0.1112   0.7346   0.0314
   1.750   0.7215   0.01441   0.00927  -0.1096   0.7074   0.0323
   2.000   0.7428   0.01402   0.00870  -0.1080   0.6823   0.0340
   2.250   0.7679   0.01463   0.00903  -0.1062   0.6577   0.0367
   2.500   0.7858   0.01331   0.00758  -0.1043   0.6315   0.0380
   2.750   0.8046   0.01302   0.00712  -0.1024   0.5978   0.0393
   3.000   0.8240   0.01288   0.00679  -0.1004   0.5665   0.0414
   5.500   1.0232   0.01354   0.00528  -0.0831   0.1810   0.0459
   5.750   1.0436   0.01375   0.00541  -0.0816   0.1602   0.0442
   6.000   1.0576   0.01440   0.00562  -0.0790   0.0795   0.0439
   6.250   1.0705   0.01521   0.00624  -0.0761   0.0444   0.0444
   6.500   1.0893   0.01549   0.00659  -0.0743   0.0402   0.0456
   6.750   1.1048   0.01612   0.00725  -0.0718   0.0350   0.0511
   7.000   1.1233   0.01648   0.00765  -0.0699   0.0333   0.0544
   7.250   1.1405   0.01691   0.00814  -0.0678   0.0313   0.0713
   7.500   1.2962   0.01889   0.01143  -0.0985   0.0249   1.0000
   7.750   1.3114   0.01958   0.01216  -0.0960   0.0239   1.0000
   8.000   1.3260   0.02028   0.01288  -0.0936   0.0228   1.0000
   8.250   1.3386   0.02116   0.01380  -0.0908   0.0221   1.0000
   8.500   1.3465   0.02264   0.01531  -0.0873   0.0212   1.0000
   8.750   1.3591   0.02377   0.01651  -0.0845   0.0205   1.0000
   9.000   1.3740   0.02442   0.01724  -0.0822   0.0199   1.0000
   9.250   1.3868   0.02520   0.01809  -0.0795   0.0191   1.0000
   9.500   1.3989   0.02615   0.01910  -0.0767   0.0186   1.0000
   9.750   1.4106   0.02713   0.02017  -0.0739   0.0181   1.0000
  10.000   1.4222   0.02817   0.02127  -0.0712   0.0176   1.0000
  10.250   1.4331   0.02923   0.02238  -0.0685   0.0172   1.0000
  10.500   1.4455   0.03067   0.02386  -0.0663   0.0166   1.0000
  10.750   1.4744   0.03496   0.02834  -0.0678   0.0161   1.0000
  11.000   1.4762   0.03570   0.02926  -0.0633   0.0159   1.0000
  11.250   1.4773   0.03652   0.03026  -0.0588   0.0155   1.0000
  11.500   1.4822   0.03846   0.03242  -0.0554   0.0153   1.0000
  11.750   1.4831   0.03984   0.03397  -0.0515   0.0147   1.0000
  12.000   1.4836   0.04214   0.03650  -0.0480   0.0146   1.0000
  12.250   1.4799   0.04468   0.03928  -0.0441   0.0145   1.0000
  12.500   1.4731   0.04733   0.04216  -0.0402   0.0144   1.0000
  12.750   1.4639   0.04996   0.04501  -0.0364   0.0143   1.0000
  13.000   1.4485   0.05363   0.04896  -0.0326   0.0144   1.0000
  13.250   1.4343   0.05688   0.05242  -0.0296   0.0143   1.0000
  13.500   1.4153   0.06100   0.05678  -0.0270   0.0144   1.0000
  13.750   1.3965   0.06518   0.06117  -0.0253   0.0143   1.0000
  14.000   1.3738   0.07035   0.06656  -0.0245   0.0146   1.0000
  14.250   1.3510   0.07586   0.07227  -0.0248   0.0147   1.0000
  14.500   1.3283   0.08179   0.07839  -0.0262   0.0148   1.0000
  14.750   1.3050   0.08836   0.08514  -0.0285   0.0148   1.0000
  15.000   1.2824   0.09548   0.09241  -0.0317   0.0150   1.0000
  15.250   1.2576   0.10376   0.10084  -0.0361   0.0151   1.0000
  15.500   1.2339   0.11256   0.10979  -0.0414   0.0152   1.0000
  15.750   1.2113   0.12184   0.11918  -0.0470   0.0154   1.0000
  16.000   1.1881   0.13199   0.12942  -0.0530   0.0158   1.0000
  16.250   1.1778   0.14014   0.13768  -0.0575   0.0167   1.0000
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