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

NACA 66 (p51htip-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA 66 (p51htip-il)
Reynolds number: 1,000,000
Max Cl/Cd: 90.81 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-p51htip-il-1000000.txt
Download as CSV file: xf-p51htip-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66                                         
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5396   0.05460   0.05273  -0.0595   0.8792   0.0120
 -11.000  -0.5485   0.05132   0.04936  -0.0617   0.8744   0.0122
 -10.750  -0.6069   0.05755   0.05545  -0.0689   0.8903   0.0114
 -10.500  -0.6345   0.05301   0.05074  -0.0676   0.8801   0.0113
 -10.250  -0.6583   0.05018   0.04775  -0.0646   0.8710   0.0111
 -10.000  -0.6893   0.04583   0.04318  -0.0594   0.8623   0.0110
  -9.750  -0.7256   0.03834   0.03523  -0.0536   0.8554   0.0110
  -9.500  -0.7598   0.02918   0.02529  -0.0469   0.8490   0.0110
  -9.250  -0.7578   0.02524   0.02084  -0.0437   0.8446   0.0113
  -9.000  -0.7446   0.02290   0.01818  -0.0420   0.8410   0.0115
  -8.750  -0.7234   0.02224   0.01737  -0.0413   0.8375   0.0116
  -8.500  -0.7116   0.01855   0.01325  -0.0394   0.8341   0.0122
  -8.250  -0.6886   0.01818   0.01283  -0.0390   0.8310   0.0127
  -8.000  -0.6654   0.01758   0.01214  -0.0386   0.8281   0.0130
  -7.750  -0.6410   0.01720   0.01171  -0.0383   0.8253   0.0135
  -7.500  -0.6167   0.01659   0.01101  -0.0380   0.8225   0.0141
  -7.250  -0.5924   0.01585   0.01017  -0.0377   0.8197   0.0146
  -7.000  -0.5680   0.01515   0.00935  -0.0373   0.8171   0.0150
  -6.750  -0.5430   0.01467   0.00879  -0.0370   0.8145   0.0153
  -6.500  -0.5212   0.01339   0.00740  -0.0362   0.8123   0.0160
  -6.250  -0.4972   0.01283   0.00684  -0.0358   0.8100   0.0168
  -6.000  -0.4719   0.01249   0.00648  -0.0357   0.8077   0.0175
  -5.750  -0.4463   0.01219   0.00615  -0.0355   0.8055   0.0184
  -5.500  -0.4204   0.01193   0.00586  -0.0353   0.8035   0.0194
  -5.250  -0.3946   0.01166   0.00553  -0.0351   0.8015   0.0200
  -5.000  -0.3723   0.01099   0.00478  -0.0343   0.7993   0.0210
  -4.750  -0.3480   0.01053   0.00432  -0.0339   0.7976   0.0225
  -4.500  -0.3220   0.01027   0.00406  -0.0338   0.7958   0.0237
  -4.250  -0.2956   0.01005   0.00382  -0.0338   0.7939   0.0252
  -4.000  -0.2687   0.00989   0.00365  -0.0338   0.7920   0.0264
  -3.750  -0.2439   0.00946   0.00317  -0.0334   0.7902   0.0284
  -3.500  -0.2178   0.00921   0.00290  -0.0332   0.7886   0.0307
  -3.250  -0.1909   0.00904   0.00271  -0.0332   0.7870   0.0331
  -3.000  -0.1637   0.00892   0.00257  -0.0333   0.7854   0.0349
  -2.750  -0.1373   0.00869   0.00234  -0.0332   0.7839   0.0416
  -2.500  -0.1111   0.00841   0.00217  -0.0331   0.7826   0.0680
  -2.250  -0.0929   0.00730   0.00187  -0.0320   0.7810   0.2898
  -2.000  -0.0755   0.00615   0.00158  -0.0308   0.7792   0.5281
  -1.750  -0.0542   0.00553   0.00152  -0.0298   0.7775   0.6917
  -1.500  -0.0265   0.00548   0.00153  -0.0299   0.7761   0.7216
  -1.250   0.0011   0.00547   0.00157  -0.0298   0.7748   0.7514
  -1.000   0.0288   0.00548   0.00163  -0.0298   0.7735   0.7704
  -0.750   0.0572   0.00552   0.00166  -0.0300   0.7722   0.7821
  -0.500   0.0853   0.00558   0.00174  -0.0301   0.7707   0.7932
  -0.250   0.1133   0.00565   0.00184  -0.0302   0.7693   0.8036
   0.000   0.1420   0.00568   0.00188  -0.0304   0.7678   0.8101
   0.250   0.1705   0.00566   0.00189  -0.0307   0.7655   0.8149
   0.500   0.1992   0.00566   0.00189  -0.0310   0.7629   0.8183
   0.750   0.2282   0.00565   0.00188  -0.0313   0.7602   0.8208
   1.000   0.2571   0.00565   0.00186  -0.0317   0.7574   0.8233
   1.250   0.2859   0.00569   0.00188  -0.0320   0.7542   0.8257
   1.500   0.3146   0.00563   0.00187  -0.0323   0.7503   0.8282
   1.750   0.3429   0.00556   0.00180  -0.0325   0.7438   0.8305
   2.000   0.3710   0.00551   0.00176  -0.0326   0.7352   0.8330
   2.250   0.3992   0.00548   0.00171  -0.0327   0.7268   0.8357
   2.500   0.4272   0.00543   0.00168  -0.0328   0.7143   0.8387
   2.750   0.4546   0.00543   0.00163  -0.0327   0.6954   0.8415
   3.000   0.4819   0.00542   0.00164  -0.0327   0.6737   0.8444
   3.250   0.5067   0.00558   0.00166  -0.0322   0.6174   0.8474
   3.500   0.5182   0.00670   0.00208  -0.0296   0.4542   0.8516
   3.750   0.5285   0.00792   0.00259  -0.0269   0.2891   0.8560
   4.000   0.5388   0.00902   0.00308  -0.0242   0.1417   0.8606
   4.250   0.5532   0.00987   0.00351  -0.0222   0.0475   0.8652
   4.500   0.5762   0.01018   0.00378  -0.0215   0.0359   0.8694
   4.750   0.5989   0.01047   0.00407  -0.0207   0.0310   0.8737
   5.000   0.6213   0.01075   0.00439  -0.0199   0.0280   0.8781
   5.250   0.6446   0.01099   0.00466  -0.0192   0.0262   0.8828
   5.500   0.6670   0.01127   0.00495  -0.0185   0.0244   0.8875
   5.750   0.6857   0.01172   0.00547  -0.0170   0.0225   0.8927
   6.000   0.7047   0.01217   0.00598  -0.0155   0.0215   0.8986
   6.250   0.7259   0.01243   0.00629  -0.0145   0.0207   0.9042
   6.500   0.7458   0.01271   0.00661  -0.0132   0.0197   0.9105
   6.750   0.7641   0.01300   0.00694  -0.0117   0.0188   0.9175
   7.000   0.7793   0.01334   0.00734  -0.0095   0.0182   0.9259
   7.250   0.7918   0.01385   0.00789  -0.0069   0.0172   0.9361
   7.500   0.7957   0.01494   0.00911  -0.0028   0.0164   0.9518
   7.750   0.8158   0.01531   0.00954  -0.0017   0.0162   0.9666
   8.000   0.8441   0.01587   0.01017  -0.0026   0.0158   0.9793
   8.250   0.8758   0.01650   0.01085  -0.0042   0.0152   0.9914
   8.500   0.8976   0.01715   0.01154  -0.0039   0.0148   1.0000
   8.750   0.9172   0.01766   0.01207  -0.0030   0.0143   1.0000
   9.000   0.9365   0.01812   0.01257  -0.0021   0.0137   1.0000
   9.250   0.9552   0.01874   0.01321  -0.0012   0.0134   1.0000
   9.500   0.9733   0.01943   0.01392  -0.0002   0.0131   1.0000
   9.750   0.9953   0.02170   0.01631   0.0002   0.0124   1.0000
  10.000   1.0137   0.02248   0.01719   0.0011   0.0122   1.0000
  10.250   1.0327   0.02351   0.01834   0.0020   0.0120   1.0000
  10.500   1.0501   0.02452   0.01947   0.0031   0.0118   1.0000
  10.750   1.0665   0.02565   0.02072   0.0043   0.0116   1.0000
  11.000   1.0813   0.02696   0.02217   0.0056   0.0113   1.0000
  11.250   1.0940   0.02824   0.02359   0.0071   0.0110   1.0000
  11.500   1.1046   0.02968   0.02518   0.0088   0.0107   1.0000
  11.750   1.1139   0.03088   0.02649   0.0105   0.0105   1.0000
  12.000   1.1242   0.03130   0.02694   0.0121   0.0102   1.0000
  12.250   1.1301   0.03292   0.02870   0.0140   0.0101   1.0000
  12.500   1.1352   0.03439   0.03030   0.0158   0.0099   1.0000
  12.750   1.1434   0.03506   0.03099   0.0172   0.0097   1.0000
  13.000   1.1419   0.03752   0.03365   0.0193   0.0097   1.0000
  13.250   1.1451   0.03893   0.03511   0.0207   0.0095   1.0000
  13.500   1.1405   0.04151   0.03785   0.0225   0.0094   1.0000
  13.750   1.1213   0.04581   0.04242   0.0247   0.0093   1.0000
  14.000   1.0960   0.05081   0.04770   0.0264   0.0092   1.0000
  14.250   1.0528   0.05830   0.05554   0.0271   0.0091   1.0000
  14.500   1.0356   0.06303   0.06043   0.0266   0.0091   1.0000
  14.750   1.0122   0.06915   0.06674   0.0249   0.0091   1.0000
  15.000   1.0122   0.07234   0.07000   0.0236   0.0092   1.0000
  15.250   0.9797   0.08153   0.07940   0.0190   0.0092   1.0000
  15.500   0.9618   0.08938   0.08740   0.0142   0.0093   1.0000
<< Back to NACA 66 (p51htip-il)

Polar data table (+)

Polar graphs


<< Back to NACA 66 (p51htip-il)