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GOE 391 AIRFOIL (goe391-il) Xfoil prediction polar at RE=200,000 Ncrit=0


Details Polar file
Airfoil: GOE 391 AIRFOIL (goe391-il)
Reynolds number: 200,000
Max Cl/Cd: 84.25 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe391-il-200000.txt
Download as CSV file: xf-goe391-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 391 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
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.4568   0.13997   0.13607  -0.0128   1.0000   0.0171
 -11.500  -0.4518   0.13687   0.13299  -0.0137   1.0000   0.0175
 -11.250  -0.4469   0.13380   0.12993  -0.0145   1.0000   0.0179
 -11.000  -0.4422   0.13068   0.12683  -0.0153   1.0000   0.0185
 -10.750  -0.4376   0.12763   0.12380  -0.0161   1.0000   0.0191
 -10.500  -0.4331   0.12459   0.12078  -0.0169   1.0000   0.0196
 -10.250  -0.4287   0.12158   0.11779  -0.0177   1.0000   0.0202
 -10.000  -0.4244   0.11866   0.11490  -0.0185   1.0000   0.0208
  -9.750  -0.4202   0.11575   0.11198  -0.0192   1.0000   0.0213
  -9.500  -0.4164   0.11294   0.10920  -0.0201   1.0000   0.0222
  -9.250  -0.4126   0.11029   0.10658  -0.0209   1.0000   0.0226
  -9.000  -0.4097   0.10810   0.10443  -0.0221   1.0000   0.0231
  -8.750  -0.4080   0.10586   0.10223  -0.0231   1.0000   0.0234
  -8.500  -0.4070   0.10358   0.09999  -0.0238   1.0000   0.0235
  -8.250  -0.4084   0.10160   0.09807  -0.0245   1.0000   0.0236
  -8.000  -0.4135   0.09997   0.09651  -0.0247   1.0000   0.0237
  -7.750  -0.4180   0.09788   0.09448  -0.0241   1.0000   0.0238
  -7.500  -0.4179   0.09537   0.09201  -0.0247   1.0000   0.0238
  -7.250  -0.4171   0.09316   0.08984  -0.0269   1.0000   0.0239
  -7.000  -0.4143   0.09042   0.08713  -0.0278   1.0000   0.0240
  -6.750  -0.4102   0.08758   0.08429  -0.0288   1.0000   0.0240
  -6.500  -0.4050   0.08462   0.08132  -0.0295   1.0000   0.0241
  -6.250  -0.3986   0.08166   0.07834  -0.0300   1.0000   0.0241
  -6.000  -0.3909   0.07847   0.07513  -0.0307   1.0000   0.0241
  -5.500  -0.3892   0.06887   0.06559  -0.0289   1.0000   0.0249
  -5.250  -0.3845   0.06556   0.06229  -0.0277   1.0000   0.0254
  -5.000  -0.3759   0.06241   0.05912  -0.0275   1.0000   0.0260
  -4.750  -0.3644   0.05922   0.05588  -0.0279   1.0000   0.0267
  -4.500  -0.3498   0.05603   0.05260  -0.0288   1.0000   0.0275
  -4.250  -0.3328   0.05286   0.04934  -0.0298   1.0000   0.0284
  -4.000  -0.3131   0.04954   0.04590  -0.0310   1.0000   0.0299
  -3.750  -0.2906   0.04637   0.04257  -0.0322   1.0000   0.0317
  -3.500  -0.2554   0.04449   0.04026  -0.0336   1.0000   0.0342
  -3.250  -0.2311   0.04221   0.03766  -0.0335   1.0000   0.0345
  -3.000  -0.2120   0.03640   0.03195  -0.0360   0.9984   0.0372
  -2.750  -0.1743   0.03370   0.02898  -0.0388   0.9955   0.0423
  -2.500  -0.1361   0.03042   0.02522  -0.0410   0.9924   0.0475
  -2.250  -0.1024   0.02827   0.02290  -0.0430   0.9892   0.0554
  -2.000  -0.0642   0.02680   0.02088  -0.0449   0.9859   0.0735
  -1.750  -0.0349   0.02426   0.01835  -0.0465   0.9825   0.0907
  -1.250   0.0438   0.01976   0.01280  -0.0482   0.9779   0.0678
  -1.000   0.0778   0.01820   0.01068  -0.0480   0.9735   0.0595
  -0.750   0.1146   0.01693   0.00936  -0.0500   0.9696   0.0671
  -0.500   0.1525   0.01624   0.00852  -0.0518   0.9650   0.0696
  -0.250   0.1890   0.01568   0.00787  -0.0534   0.9588   0.0712
   0.000   0.2315   0.01498   0.00719  -0.0563   0.9551   0.0702
   0.250   0.2613   0.01452   0.00674  -0.0567   0.9471   0.0705
   0.500   0.3025   0.01411   0.00633  -0.0594   0.9426   0.0725
   0.750   0.3351   0.01373   0.00597  -0.0602   0.9349   0.0841
   1.000   0.3979   0.01138   0.00580  -0.0678   0.9351   1.0000
   1.250   0.4453   0.01115   0.00550  -0.0716   0.9293   1.0000
   1.500   0.4834   0.01093   0.00527  -0.0734   0.9203   1.0000
   1.750   0.5323   0.01056   0.00493  -0.0774   0.9152   1.0000
   2.000   0.5639   0.01039   0.00479  -0.0779   0.9051   1.0000
   2.250   0.6012   0.01002   0.00446  -0.0792   0.8939   1.0000
   2.500   0.6388   0.00944   0.00393  -0.0800   0.8748   1.0000
   2.750   0.6690   0.00906   0.00356  -0.0794   0.8481   1.0000
   3.000   0.6980   0.00888   0.00337  -0.0789   0.8190   1.0000
   3.250   0.7238   0.00884   0.00335  -0.0779   0.7846   1.0000
   3.500   0.7490   0.00889   0.00328  -0.0766   0.7299   1.0000
   3.750   0.7701   0.00924   0.00325  -0.0745   0.6440   1.0000
   4.000   0.7875   0.00982   0.00347  -0.0721   0.5649   1.0000
   4.250   0.8015   0.01060   0.00381  -0.0692   0.4572   1.0000
   4.500   0.8117   0.01178   0.00431  -0.0659   0.3262   1.0000
   4.750   0.8205   0.01353   0.00507  -0.0628   0.1442   1.0000
   5.000   0.8367   0.01481   0.00594  -0.0608   0.0809   1.0000
   5.250   0.8570   0.01557   0.00678  -0.0593   0.0735   1.0000
   5.500   0.8764   0.01642   0.00774  -0.0577   0.0673   1.0000
   5.750   0.8973   0.01712   0.00852  -0.0565   0.0597   1.0000
   6.000   0.9162   0.01819   0.00967  -0.0549   0.0515   1.0000
   6.250   0.9335   0.01986   0.01135  -0.0530   0.0411   1.0000
   6.500   0.9530   0.02262   0.01405  -0.0516   0.0335   1.0000
   6.750   0.9756   0.02359   0.01526  -0.0503   0.0279   1.0000
   7.000   0.9982   0.02599   0.01780  -0.0494   0.0249   1.0000
   7.500   1.0381   0.03225   0.02491  -0.0460   0.0221   1.0000
   7.750   1.0542   0.03503   0.02818  -0.0436   0.0210   1.0000
   8.000   1.0653   0.03893   0.03258  -0.0408   0.0212   1.0000
   8.250   1.0712   0.04334   0.03745  -0.0377   0.0218   1.0000
   8.500   1.0170   0.04009   0.03541  -0.0276   0.0307   1.0000
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