Airfoil Tools
Search 1638 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

MARSKE PIONEER IID TIP AIRFOIL (NACA 431012A*.833 HYBRID) (marske4-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: MARSKE PIONEER IID TIP AIRFOIL (NACA 431012A*.833 HYBRID) (marske4-il)
Reynolds number: 1,000,000
Max Cl/Cd: 100.62 at α=10.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-marske4-il-1000000.txt
Download as CSV file: xf-marske4-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MARSKE PIONEER IID TIP AIRFOIL (NACA 431012A*.83
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5797   0.08997   0.08836   0.0038   1.0000   0.0142
  -9.750  -0.5830   0.08516   0.08357   0.0005   1.0000   0.0144
  -9.500  -0.5974   0.07796   0.07637  -0.0068   1.0000   0.0144
  -9.250  -0.5166   0.06673   0.06527  -0.0082   1.0000   0.0156
  -9.000  -0.5445   0.06033   0.05882  -0.0121   1.0000   0.0154
  -8.750  -0.5703   0.05579   0.05422  -0.0115   1.0000   0.0152
  -8.500  -0.5904   0.05177   0.05012  -0.0090   1.0000   0.0153
  -8.250  -0.6526   0.05700   0.05505  -0.0047   1.0000   0.0150
  -8.000  -0.7156   0.03631   0.03338   0.0047   1.0000   0.0135
  -7.500  -0.7404   0.02123   0.01685   0.0170   1.0000   0.0133
  -7.250  -0.7232   0.01954   0.01488   0.0192   1.0000   0.0135
  -7.000  -0.7085   0.01689   0.01192   0.0216   1.0000   0.0139
  -6.750  -0.6869   0.01622   0.01120   0.0230   1.0000   0.0142
  -6.500  -0.6646   0.01581   0.01076   0.0243   1.0000   0.0145
  -6.250  -0.6426   0.01515   0.01004   0.0257   1.0000   0.0148
  -6.000  -0.6109   0.01442   0.00923   0.0250   0.9962   0.0151
  -5.750  -0.5645   0.01343   0.00814   0.0212   0.9725   0.0158
  -5.500  -0.5044   0.01289   0.00713   0.0146   0.8199   0.0164
  -5.000  -0.4606   0.01206   0.00592   0.0176   0.7312   0.0176
  -4.750  -0.4356   0.01189   0.00567   0.0184   0.7106   0.0183
  -4.500  -0.4108   0.01161   0.00530   0.0193   0.6920   0.0189
  -4.250  -0.3856   0.01134   0.00495   0.0202   0.6728   0.0196
  -4.000  -0.3619   0.01090   0.00441   0.0213   0.6501   0.0204
  -3.750  -0.3378   0.01061   0.00404   0.0223   0.6192   0.0212
  -3.500  -0.3133   0.01051   0.00380   0.0232   0.5714   0.0221
  -3.250  -0.2890   0.01050   0.00357   0.0242   0.5084   0.0232
  -3.000  -0.2655   0.01053   0.00332   0.0252   0.4284   0.0243
  -2.750  -0.2422   0.01051   0.00307   0.0262   0.3537   0.0258
  -2.500  -0.2170   0.01048   0.00292   0.0270   0.3187   0.0272
  -2.250  -0.1912   0.01043   0.00277   0.0276   0.2992   0.0284
  -2.000  -0.1663   0.01021   0.00250   0.0285   0.2865   0.0307
  -1.750  -0.1401   0.01014   0.00240   0.0291   0.2760   0.0328
  -1.500  -0.1140   0.01005   0.00225   0.0297   0.2669   0.0348
  -1.250  -0.0882   0.00991   0.00210   0.0303   0.2595   0.0381
  -1.000  -0.0616   0.00986   0.00201   0.0309   0.2517   0.0410
  -0.750  -0.0356   0.00975   0.00189   0.0315   0.2445   0.0472
  -0.500  -0.0096   0.00962   0.00180   0.0321   0.2380   0.0635
  -0.250  -0.0023   0.00800   0.00154   0.0358   0.2328   0.4894
   0.000  -0.0004   0.00675   0.00145   0.0414   0.2291   0.8076
   0.250   0.0227   0.00668   0.00151   0.0429   0.2240   0.8597
   0.500   0.0481   0.00673   0.00159   0.0439   0.2177   0.8869
   0.750   0.0737   0.00681   0.00169   0.0449   0.2127   0.9069
   1.000   0.1007   0.00691   0.00177   0.0455   0.2082   0.9202
   1.250   0.1277   0.00703   0.00186   0.0461   0.2032   0.9302
   1.500   0.1547   0.00717   0.00197   0.0466   0.1987   0.9378
   1.750   0.1828   0.00727   0.00205   0.0469   0.1961   0.9445
   2.000   0.2133   0.00744   0.00219   0.0467   0.1928   0.9508
   2.250   0.2391   0.00756   0.00227   0.0474   0.1896   0.9559
   2.500   0.2689   0.00771   0.00237   0.0471   0.1857   0.9584
   2.750   0.3006   0.00783   0.00247   0.0464   0.1837   0.9600
   3.000   0.3325   0.00793   0.00256   0.0457   0.1820   0.9615
   3.250   0.3643   0.00803   0.00264   0.0450   0.1796   0.9631
   3.500   0.3954   0.00815   0.00273   0.0444   0.1770   0.9650
   3.750   0.4247   0.00829   0.00283   0.0442   0.1745   0.9675
   4.000   0.4499   0.00846   0.00298   0.0448   0.1717   0.9713
   4.250   0.4836   0.00863   0.00314   0.0436   0.1697   0.9721
   4.500   0.5180   0.00873   0.00324   0.0422   0.1685   0.9727
   4.750   0.5519   0.00885   0.00335   0.0410   0.1668   0.9733
   5.000   0.5853   0.00898   0.00348   0.0398   0.1653   0.9741
   5.250   0.6182   0.00911   0.00361   0.0387   0.1635   0.9751
   5.500   0.6506   0.00926   0.00375   0.0377   0.1618   0.9763
   5.750   0.6823   0.00944   0.00392   0.0369   0.1600   0.9777
   6.000   0.7125   0.00968   0.00415   0.0363   0.1577   0.9796
   6.250   0.7393   0.00992   0.00440   0.0365   0.1559   0.9826
   6.500   0.7706   0.00999   0.00450   0.0357   0.1548   0.9840
   6.750   0.8042   0.01010   0.00464   0.0345   0.1536   0.9847
   7.000   0.8376   0.01021   0.00476   0.0332   0.1517   0.9856
   7.250   0.8705   0.01032   0.00489   0.0320   0.1496   0.9866
   7.500   0.9028   0.01047   0.00504   0.0310   0.1475   0.9878
   7.750   0.9341   0.01070   0.00525   0.0301   0.1444   0.9892
   8.000   0.9648   0.01092   0.00550   0.0293   0.1423   0.9908
   8.250   0.9957   0.01100   0.00563   0.0285   0.1409   0.9924
   8.500   1.0255   0.01112   0.00578   0.0280   0.1389   0.9941
   8.750   1.0583   0.01122   0.00590   0.0268   0.1362   0.9949
   9.000   1.0905   0.01138   0.00607   0.0257   0.1333   0.9959
   9.250   1.1220   0.01163   0.00633   0.0246   0.1299   0.9971
   9.500   1.1545   0.01171   0.00646   0.0235   0.1274   0.9982
   9.750   1.1864   0.01185   0.00662   0.0223   0.1235   0.9993
  10.000   1.2151   0.01209   0.00685   0.0218   0.1192   1.0000
  10.250   1.2356   0.01228   0.00708   0.0231   0.1163   1.0000
  10.500   1.2559   0.01249   0.00732   0.0244   0.1124   1.0000
  10.750   1.2753   0.01278   0.00759   0.0257   0.1079   1.0000
  11.000   1.2950   0.01305   0.00790   0.0271   0.1043   1.0000
  11.250   1.3138   0.01339   0.00824   0.0285   0.0992   1.0000
  11.500   1.3323   0.01377   0.00863   0.0300   0.0948   1.0000
  11.750   1.3506   0.01416   0.00903   0.0314   0.0906   1.0000
  12.000   1.3676   0.01467   0.00953   0.0330   0.0850   1.0000
  12.250   1.3855   0.01509   0.00998   0.0345   0.0808   1.0000
  12.500   1.4015   0.01568   0.01058   0.0362   0.0754   1.0000
  12.750   1.4191   0.01614   0.01107   0.0376   0.0701   1.0000
  13.000   1.4340   0.01683   0.01175   0.0393   0.0618   1.0000
  13.250   1.4415   0.01806   0.01286   0.0418   0.0469   1.0000
  13.500   1.4513   0.01908   0.01388   0.0440   0.0411   1.0000
  13.750   1.4610   0.02005   0.01488   0.0461   0.0382   1.0000
  14.000   1.4710   0.02095   0.01584   0.0482   0.0360   1.0000
  14.250   1.4818   0.02174   0.01671   0.0501   0.0350   1.0000
  14.500   1.4870   0.02266   0.01771   0.0528   0.0339   1.0000
  14.750   1.4866   0.02372   0.01884   0.0560   0.0330   1.0000
  15.000   1.4846   0.02512   0.02032   0.0586   0.0319   1.0000
  15.250   1.4799   0.02706   0.02236   0.0601   0.0309   1.0000
  15.500   1.4794   0.02914   0.02456   0.0603   0.0304   1.0000
  15.750   1.4785   0.03169   0.02723   0.0595   0.0299   1.0000
  16.000   1.4734   0.03505   0.03073   0.0579   0.0294   1.0000
  16.250   1.4598   0.03992   0.03575   0.0551   0.0291   1.0000
  16.500   1.4278   0.04815   0.04420   0.0498   0.0293   1.0000
  16.750   1.3282   0.06743   0.06382   0.0387   0.0303   1.0000
  17.000   1.2607   0.08054   0.07712   0.0319   0.0315   1.0000
<< Back to MARSKE PIONEER IID TIP AIRFOIL (NACA 431012A*.833 HYBRID) (marske4-il)

Polar data table (+)

Polar graphs


<< Back to MARSKE PIONEER IID TIP AIRFOIL (NACA 431012A*.833 HYBRID) (marske4-il)