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

NASA SC(2)-0710 AIRFOIL (sc20710-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0710 AIRFOIL (sc20710-il)
Reynolds number: 500,000
Max Cl/Cd: 71.58 at α=0.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20710-il-500000.txt
Download as CSV file: xf-sc20710-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0710 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
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.6698   0.09315   0.09054  -0.0382   1.0000   0.0246
 -12.750  -0.7743   0.06333   0.06047  -0.0582   1.0000   0.0226
 -12.500  -0.8190   0.05489   0.05182  -0.0622   1.0000   0.0223
 -12.250  -0.8594   0.04925   0.04598  -0.0619   1.0000   0.0221
 -12.000  -0.8977   0.04593   0.04252  -0.0576   1.0000   0.0219
 -11.750  -0.9278   0.04126   0.03756  -0.0561   1.0000   0.0217
 -11.500  -0.9385   0.03686   0.03279  -0.0558   1.0000   0.0218
 -11.250  -0.9352   0.03345   0.02904  -0.0556   1.0000   0.0219
 -11.000  -0.9241   0.03070   0.02599  -0.0556   1.0000   0.0221
 -10.750  -0.9081   0.02844   0.02346  -0.0557   1.0000   0.0224
 -10.500  -0.8892   0.02648   0.02125  -0.0559   1.0000   0.0228
 -10.250  -0.8681   0.02485   0.01940  -0.0560   1.0000   0.0232
 -10.000  -0.8456   0.02352   0.01787  -0.0562   1.0000   0.0236
  -9.750  -0.8218   0.02267   0.01685  -0.0563   1.0000   0.0240
  -9.500  -0.7984   0.02073   0.01478  -0.0568   1.0000   0.0249
  -9.250  -0.7738   0.01994   0.01397  -0.0570   1.0000   0.0256
  -9.000  -0.7483   0.01923   0.01321  -0.0573   1.0000   0.0263
  -8.750  -0.7219   0.01854   0.01245  -0.0577   1.0000   0.0271
  -8.500  -0.6949   0.01788   0.01172  -0.0581   1.0000   0.0281
  -8.250  -0.6675   0.01734   0.01109  -0.0586   1.0000   0.0290
  -8.000  -0.6374   0.01627   0.00997  -0.0599   1.0000   0.0306
  -7.750  -0.6064   0.01577   0.00949  -0.0612   0.9997   0.0323
  -7.500  -0.5730   0.01539   0.00907  -0.0628   0.9989   0.0346
  -7.250  -0.5387   0.01472   0.00838  -0.0648   0.9984   0.0376
  -7.000  -0.5050   0.01442   0.00809  -0.0665   0.9976   0.0408
  -6.750  -0.4711   0.01413   0.00776  -0.0681   0.9968   0.0436
  -6.500  -0.4362   0.01361   0.00727  -0.0703   0.9963   0.0478
  -6.250  -0.4025   0.01341   0.00704  -0.0719   0.9957   0.0513
  -6.000  -0.3673   0.01299   0.00664  -0.0740   0.9953   0.0561
  -5.750  -0.3330   0.01279   0.00645  -0.0757   0.9947   0.0612
  -5.500  -0.2976   0.01247   0.00616  -0.0777   0.9944   0.0685
  -5.250  -0.2624   0.01224   0.00595  -0.0797   0.9939   0.0777
  -5.000  -0.2274   0.01190   0.00572  -0.0816   0.9932   0.0948
  -4.750  -0.1912   0.01152   0.00551  -0.0839   0.9926   0.1303
  -4.500  -0.1511   0.01092   0.00532  -0.0875   0.9930   0.2222
  -4.250  -0.1056   0.01002   0.00509  -0.0926   0.9944   0.3918
  -4.000  -0.0644   0.00943   0.00514  -0.0963   0.9951   0.5587
  -3.750  -0.0297   0.00947   0.00527  -0.0978   0.9944   0.6036
  -3.500   0.0032   0.00954   0.00537  -0.0989   0.9929   0.6261
  -3.250   0.0350   0.00963   0.00546  -0.0997   0.9907   0.6445
  -3.000   0.0672   0.00976   0.00561  -0.1005   0.9888   0.6600
  -2.750   0.1005   0.00986   0.00573  -0.1016   0.9871   0.6684
  -2.500   0.1379   0.00994   0.00580  -0.1035   0.9853   0.6791
  -2.250   0.1753   0.01002   0.00594  -0.1052   0.9824   0.6894
  -2.000   0.2084   0.01000   0.00596  -0.1059   0.9770   0.6979
  -1.750   0.2514   0.00980   0.00579  -0.1088   0.9734   0.7020
  -1.500   0.2935   0.00945   0.00544  -0.1113   0.9663   0.7071
  -1.250   0.3407   0.00904   0.00507  -0.1148   0.9614   0.7124
  -1.000   0.3867   0.00857   0.00466  -0.1179   0.9550   0.7160
  -0.750   0.4230   0.00830   0.00442  -0.1190   0.9461   0.7206
  -0.500   0.4559   0.00812   0.00424  -0.1196   0.9345   0.7255
  -0.250   0.4868   0.00795   0.00410  -0.1197   0.9193   0.7281
   0.000   0.5173   0.00783   0.00398  -0.1197   0.8994   0.7300
   0.250   0.5455   0.00781   0.00392  -0.1192   0.8689   0.7320
   0.500   0.5712   0.00798   0.00382  -0.1180   0.7825   0.7342
   0.750   0.5834   0.00943   0.00409  -0.1144   0.5387   0.7367
   1.000   0.6020   0.01068   0.00446  -0.1130   0.3471   0.7393
   1.250   0.6253   0.01164   0.00477  -0.1124   0.2068   0.7420
   1.500   0.6494   0.01231   0.00505  -0.1118   0.1262   0.7440
   1.750   0.6747   0.01276   0.00535  -0.1113   0.0940   0.7457
   2.000   0.7010   0.01312   0.00565  -0.1109   0.0783   0.7475
   2.250   0.7273   0.01351   0.00599  -0.1105   0.0682   0.7495
   2.500   0.7541   0.01380   0.00626  -0.1102   0.0618   0.7520
   2.750   0.7808   0.01422   0.00667  -0.1099   0.0566   0.7548
   3.000   0.8088   0.01449   0.00694  -0.1100   0.0528   0.7576
   3.250   0.8343   0.01502   0.00744  -0.1095   0.0481   0.7595
   3.500   0.8608   0.01525   0.00772  -0.1091   0.0455   0.7611
   3.750   0.8869   0.01555   0.00804  -0.1087   0.0424   0.7628
   4.000   0.9110   0.01623   0.00872  -0.1079   0.0392   0.7648
   4.250   0.9375   0.01655   0.00909  -0.1075   0.0368   0.7669
   4.500   0.9635   0.01697   0.00950  -0.1071   0.0344   0.7693
   4.750   0.9873   0.01791   0.01046  -0.1063   0.0322   0.7718
   5.000   1.0142   0.01834   0.01093  -0.1061   0.0306   0.7743
   5.250   1.0391   0.01877   0.01140  -0.1054   0.0290   0.7761
   5.500   1.0631   0.01933   0.01199  -0.1046   0.0278   0.7780
   5.750   1.0845   0.02071   0.01343  -0.1034   0.0266   0.7802
   6.000   1.1093   0.02132   0.01416  -0.1027   0.0260   0.7827
   6.250   1.1339   0.02212   0.01506  -0.1020   0.0253   0.7851
   6.500   1.1585   0.02299   0.01603  -0.1014   0.0246   0.7876
   6.750   1.1829   0.02392   0.01704  -0.1007   0.0240   0.7900
   7.000   1.2057   0.02480   0.01803  -0.0998   0.0234   0.7917
   7.250   1.2281   0.02574   0.01906  -0.0988   0.0230   0.7935
   7.500   1.2498   0.02686   0.02026  -0.0978   0.0225   0.7956
   7.750   1.2684   0.02923   0.02282  -0.0964   0.0219   0.7977
   8.000   1.2876   0.03071   0.02455  -0.0950   0.0215   0.8002
   8.250   1.3061   0.03233   0.02643  -0.0934   0.0212   0.8027
   8.500   1.1734   0.06493   0.06196  -0.0733   0.0291   0.8025
   8.750   1.1804   0.06608   0.06315  -0.0715   0.0284   0.8049
   9.000   1.1794   0.06785   0.06501  -0.0689   0.0280   0.8067
   9.250   1.1680   0.06982   0.06709  -0.0651   0.0278   0.8087
   9.500   1.1543   0.07245   0.06983  -0.0619   0.0276   0.8108
   9.750   1.1415   0.07534   0.07280  -0.0598   0.0275   0.8130
  10.000   1.1257   0.07910   0.07666  -0.0589   0.0274   0.8152
  10.250   1.1082   0.08376   0.08141  -0.0596   0.0273   0.8173
  10.500   1.0860   0.09013   0.08788  -0.0627   0.0271   0.8191
  10.750   0.8340   0.11667   0.11465  -0.0686   0.0265   0.8078
  11.000   1.0436   0.07314   0.07070  -0.0432   0.0212   0.8186
  11.250   1.0057   0.08236   0.08006  -0.0475   0.0214   0.8192
  11.500   0.9625   0.09449   0.09231  -0.0549   0.0217   0.8192
<< Back to NASA SC(2)-0710 AIRFOIL (sc20710-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0710 AIRFOIL (sc20710-il)