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GOE 370 AIRFOIL (goe370-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 370 AIRFOIL (goe370-il)
Reynolds number: 1,000,000
Max Cl/Cd: 154.2 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe370-il-1000000.txt
Download as CSV file: xf-goe370-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 370 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3169   0.10036   0.09879  -0.0263   1.0000   0.0077
  -8.250  -0.3140   0.09786   0.09630  -0.0263   1.0000   0.0082
  -8.000  -0.3130   0.09560   0.09406  -0.0259   1.0000   0.0082
  -7.750  -0.3079   0.09315   0.09164  -0.0271   0.9995   0.0084
  -7.500  -0.2873   0.08913   0.08762  -0.0330   0.9965   0.0085
  -7.250  -0.2685   0.08526   0.08375  -0.0382   0.9909   0.0085
  -7.000  -0.2441   0.08079   0.07928  -0.0451   0.9850   0.0086
  -6.750  -0.2220   0.07664   0.07512  -0.0509   0.9765   0.0086
  -6.250  -0.1614   0.06478   0.06321  -0.0691   0.9590   0.0089
  -6.000  -0.1260   0.06105   0.05943  -0.0763   0.9487   0.0095
  -5.750  -0.0856   0.05740   0.05570  -0.0846   0.9313   0.0109
  -5.500  -0.0458   0.05361   0.05172  -0.0928   0.9014   0.0131
  -5.250  -0.0235   0.05029   0.04824  -0.0960   0.8750   0.0132
  -4.500   0.0462   0.01360   0.00979  -0.1146   0.8283   0.0102
  -4.250   0.0717   0.01277   0.00873  -0.1143   0.8195   0.0112
  -4.000   0.0971   0.01153   0.00718  -0.1140   0.8116   0.0123
  -3.750   0.1231   0.01088   0.00630  -0.1136   0.8044   0.0131
  -3.500   0.1482   0.00946   0.00454  -0.1131   0.7975   0.0149
  -3.000   0.2029   0.00989   0.00498  -0.1128   0.7846   0.0185
  -2.750   0.2301   0.00996   0.00498  -0.1125   0.7787   0.0197
  -2.500   0.2556   0.00873   0.00351  -0.1122   0.7735   0.0227
  -2.250   0.2827   0.00869   0.00345  -0.1120   0.7679   0.0252
  -2.000   0.3095   0.00852   0.00319  -0.1117   0.7628   0.0275
  -1.750   0.3366   0.00834   0.00297  -0.1115   0.7574   0.0292
  -1.500   0.3639   0.00847   0.00306  -0.1113   0.7518   0.0301
  -1.250   0.3898   0.00770   0.00215  -0.1109   0.7468   0.0319
  -1.000   0.4164   0.00740   0.00179  -0.1106   0.7414   0.0325
  -0.750   0.4430   0.00719   0.00151  -0.1103   0.7362   0.0331
  -0.500   0.4697   0.00700   0.00129  -0.1100   0.7298   0.0346
  -0.250   0.4961   0.00688   0.00110  -0.1095   0.7220   0.0361
   0.000   0.5227   0.00675   0.00096  -0.1092   0.7129   0.0396
   0.250   0.5492   0.00670   0.00086  -0.1089   0.7039   0.0425
   0.500   0.5755   0.00665   0.00080  -0.1085   0.6943   0.0542
   0.750   0.6019   0.00653   0.00081  -0.1082   0.6847   0.1016
   1.000   0.6285   0.00652   0.00085  -0.1079   0.6757   0.1283
   1.250   0.6550   0.00654   0.00087  -0.1076   0.6662   0.1398
   1.500   0.6817   0.00656   0.00089  -0.1074   0.6550   0.1499
   1.750   0.7081   0.00657   0.00092  -0.1071   0.6432   0.1619
   2.000   0.7343   0.00658   0.00096  -0.1068   0.6300   0.1740
   2.250   0.7603   0.00660   0.00100  -0.1064   0.6152   0.1946
   2.500   0.7851   0.00651   0.00114  -0.1059   0.5990   0.3314
   2.750   0.8327   0.00540   0.00134  -0.1110   0.5809   1.0000
   3.000   0.8560   0.00563   0.00146  -0.1100   0.5533   1.0000
   3.250   0.8788   0.00592   0.00159  -0.1090   0.5209   1.0000
   3.500   0.9029   0.00612   0.00172  -0.1083   0.5022   1.0000
   3.750   0.9265   0.00636   0.00186  -0.1074   0.4740   1.0000
   4.000   0.9479   0.00679   0.00204  -0.1062   0.4132   1.0000
   4.250   0.9609   0.00797   0.00256  -0.1037   0.2789   1.0000
   4.500   0.9640   0.01022   0.00364  -0.0996   0.0416   1.0000
   4.750   0.9858   0.01073   0.00410  -0.0985   0.0194   1.0000
   5.000   1.0087   0.01111   0.00449  -0.0975   0.0156   1.0000
   5.250   1.0295   0.01174   0.00523  -0.0961   0.0127   1.0000
   5.500   1.0525   0.01210   0.00563  -0.0952   0.0119   1.0000
   5.750   1.0744   0.01257   0.00617  -0.0941   0.0110   1.0000
   6.000   1.0958   0.01308   0.00672  -0.0929   0.0099   1.0000
   6.250   1.1161   0.01368   0.00737  -0.0915   0.0091   1.0000
   6.500   1.1301   0.01489   0.00867  -0.0890   0.0080   1.0000
   6.750   1.1398   0.01658   0.01049  -0.0858   0.0075   1.0000
   7.000   1.1598   0.01721   0.01117  -0.0845   0.0073   1.0000
   7.250   1.1784   0.01804   0.01207  -0.0829   0.0070   1.0000
   7.500   1.1964   0.01905   0.01317  -0.0812   0.0067   1.0000
   7.750   1.2150   0.02006   0.01426  -0.0797   0.0062   1.0000
   8.000   1.2338   0.02119   0.01546  -0.0782   0.0057   1.0000
   8.250   1.2527   0.02266   0.01702  -0.0768   0.0054   1.0000
  16.250   0.8722   0.19622   0.19496  -0.1027   0.0061   1.0000
  16.500   0.8734   0.20065   0.19939  -0.1048   0.0062   1.0000
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