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HQ 1.5/12 AIRFOIL (hq1512-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: HQ 1.5/12 AIRFOIL (hq1512-il)
Reynolds number: 50,000
Max Cl/Cd: 32.34 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq1512-il-50000.txt
Download as CSV file: xf-hq1512-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 1.5/12 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4272   0.10332   0.09627  -0.0111   1.0000   0.3333
  -9.000  -0.4406   0.10156   0.09466  -0.0113   1.0000   0.3456
  -8.500  -0.5992   0.06904   0.06237  -0.0453   1.0000   0.1379
  -8.250  -0.6017   0.06461   0.05792  -0.0448   1.0000   0.1353
  -8.000  -0.6135   0.05997   0.05313  -0.0441   1.0000   0.1316
  -7.750  -0.6421   0.05479   0.04722  -0.0425   1.0000   0.1246
  -7.500  -0.6398   0.05104   0.04324  -0.0407   1.0000   0.1240
  -7.250  -0.6369   0.04750   0.03937  -0.0386   1.0000   0.1236
  -7.000  -0.6327   0.04421   0.03565  -0.0363   1.0000   0.1239
  -6.750  -0.6239   0.04132   0.03257  -0.0343   1.0000   0.1273
  -6.500  -0.6120   0.03915   0.03025  -0.0323   1.0000   0.1334
  -6.250  -0.6004   0.03656   0.02711  -0.0302   1.0000   0.1377
  -6.000  -0.5849   0.03424   0.02468  -0.0285   1.0000   0.1447
  -5.750  -0.5693   0.03230   0.02241  -0.0267   1.0000   0.1556
  -5.500  -0.5515   0.03049   0.02042  -0.0251   1.0000   0.1689
  -5.250  -0.5338   0.02889   0.01894  -0.0235   1.0000   0.1880
  -5.000  -0.5154   0.02736   0.01737  -0.0218   1.0000   0.2134
  -4.750  -0.4974   0.02579   0.01602  -0.0201   1.0000   0.2476
  -4.500  -0.4805   0.02416   0.01471  -0.0182   1.0000   0.3017
  -4.250  -0.4661   0.02250   0.01389  -0.0159   1.0000   0.3916
  -4.000  -0.4589   0.02188   0.01431  -0.0107   1.0000   0.5270
  -3.750  -0.4546   0.02255   0.01533  -0.0038   1.0000   0.6308
  -3.500  -0.4503   0.02330   0.01614   0.0032   1.0000   0.6952
  -3.250  -0.4461   0.02379   0.01655   0.0099   1.0000   0.7469
  -3.000  -0.4413   0.02407   0.01677   0.0164   1.0000   0.7906
  -2.750  -0.4337   0.02418   0.01679   0.0222   1.0000   0.8323
  -2.500  -0.4146   0.02434   0.01680   0.0259   1.0000   0.8755
  -2.250  -0.3322   0.02543   0.01743   0.0189   1.0000   0.9270
  -2.000  -0.1768   0.02633   0.01761  -0.0036   1.0000   0.9695
  -1.750  -0.0661   0.02605   0.01689  -0.0206   1.0000   1.0000
  -1.500  -0.0741   0.02582   0.01663  -0.0174   1.0000   1.0000
  -1.250  -0.0824   0.02555   0.01632  -0.0142   1.0000   1.0000
  -1.000  -0.0913   0.02523   0.01597  -0.0108   1.0000   1.0000
  -0.750  -0.1010   0.02485   0.01555  -0.0073   1.0000   1.0000
  -0.500  -0.1112   0.02440   0.01507  -0.0036   1.0000   1.0000
  -0.250  -0.1195   0.02395   0.01456  -0.0002   1.0000   1.0000
   0.000  -0.1198   0.02370   0.01421   0.0020   1.0000   1.0000
   0.250  -0.1098   0.02376   0.01414   0.0028   1.0000   1.0000
   0.500  -0.0948   0.02402   0.01426   0.0028   1.0000   1.0000
   0.750  -0.0775   0.02441   0.01452   0.0026   1.0000   1.0000
   1.000  -0.0461   0.02526   0.01526  -0.0002   0.9947   1.0000
   1.250   0.0008   0.02658   0.01646  -0.0057   0.9825   1.0000
   1.500   0.0444   0.02778   0.01758  -0.0105   0.9695   1.0000
   1.750   0.0863   0.02891   0.01865  -0.0148   0.9556   1.0000
   2.000   0.1307   0.03005   0.01976  -0.0193   0.9393   1.0000
   2.250   0.1729   0.03095   0.02066  -0.0230   0.9182   1.0000
   2.500   0.2301   0.03198   0.02172  -0.0287   0.8962   1.0000
   2.750   0.2653   0.03267   0.02244  -0.0307   0.8751   1.0000
   3.000   0.3080   0.03341   0.02323  -0.0337   0.8553   1.0000
   3.250   0.3583   0.03403   0.02396  -0.0374   0.8360   1.0000
   3.500   0.3869   0.03460   0.02460  -0.0379   0.8142   1.0000
   3.750   0.4369   0.03486   0.02503  -0.0409   0.7935   1.0000
   4.000   0.4713   0.03519   0.02547  -0.0415   0.7707   1.0000
   4.250   0.5274   0.03481   0.02533  -0.0444   0.7490   1.0000
   4.500   0.5640   0.03470   0.02537  -0.0445   0.7244   1.0000
   4.750   0.6221   0.03341   0.02432  -0.0459   0.7033   1.0000
   5.000   0.6555   0.03287   0.02396  -0.0447   0.6767   1.0000
   5.250   0.6932   0.03199   0.02324  -0.0435   0.6502   1.0000
   5.500   0.7335   0.03080   0.02219  -0.0423   0.6232   1.0000
   5.750   0.7706   0.02971   0.02121  -0.0407   0.5944   1.0000
   6.000   0.8036   0.02885   0.02040  -0.0389   0.5636   1.0000
   6.250   0.8345   0.02815   0.01968  -0.0370   0.5311   1.0000
   6.500   0.8651   0.02755   0.01896  -0.0351   0.4970   1.0000
   6.750   0.8860   0.02773   0.01907  -0.0328   0.4608   1.0000
   7.250   0.9269   0.02866   0.01963  -0.0284   0.3837   1.0000
   7.500   0.9440   0.02964   0.02042  -0.0262   0.3429   1.0000
   7.750   0.9613   0.03101   0.02149  -0.0241   0.3019   1.0000
   8.000   0.9790   0.03286   0.02311  -0.0224   0.2647   1.0000
   8.250   0.9971   0.03486   0.02500  -0.0208   0.2352   1.0000
   8.500   1.0150   0.03688   0.02701  -0.0195   0.2132   1.0000
   8.750   1.0341   0.03922   0.02950  -0.0183   0.1978   1.0000
   9.000   1.0526   0.04149   0.03191  -0.0171   0.1850   1.0000
   9.250   1.0729   0.04358   0.03398  -0.0163   0.1735   1.0000
   9.500   1.0793   0.04662   0.03758  -0.0141   0.1670   1.0000
   9.750   1.1031   0.04935   0.04023  -0.0138   0.1600   1.0000
  10.000   1.0977   0.05298   0.04450  -0.0109   0.1569   1.0000
  10.250   1.0929   0.05652   0.04846  -0.0085   0.1531   1.0000
  10.500   1.1192   0.05905   0.05083  -0.0085   0.1461   1.0000
  10.750   1.1057   0.06333   0.05556  -0.0059   0.1455   1.0000
  11.000   1.0882   0.06781   0.06037  -0.0036   0.1453   1.0000
  11.250   1.0657   0.07232   0.06512  -0.0014   0.1456   1.0000
  11.500   1.0427   0.07709   0.07003   0.0001   0.1461   1.0000
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