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Cambered plate C=14% T=5% R=0.96 (cp-140-050-gn) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: Cambered plate C=14% T=5% R=0.96 (cp-140-050-gn)
Reynolds number: 500,000
Max Cl/Cd: 53.88 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-cp-140-050-gn-500000.txt
Download as CSV file: xf-cp-140-050-gn-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Cambered plate C=14% T=5% R=0.96                
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000   0.3099   0.10414   0.10095  -0.2115   0.9347   0.0245
 -10.750   0.2933   0.10423   0.10108  -0.2044   0.9269   0.0248
 -10.500   0.3105   0.10156   0.09838  -0.2079   0.9253   0.0250
 -10.000   0.3484   0.09572   0.09249  -0.2157   0.9230   0.0253
  -9.750   0.3690   0.09302   0.08978  -0.2194   0.9220   0.0255
  -9.500   0.3908   0.09044   0.08717  -0.2235   0.9209   0.0259
  -9.250   0.4132   0.08787   0.08457  -0.2278   0.9197   0.0264
  -9.000   0.3997   0.08777   0.08451  -0.2214   0.9127   0.0267
  -8.750   0.4140   0.08570   0.08242  -0.2233   0.9101   0.0272
  -8.500   0.4324   0.08333   0.08003  -0.2266   0.9080   0.0280
  -8.250   0.4532   0.08114   0.07779  -0.2310   0.9060   0.0281
  -7.750   0.4563   0.07892   0.07559  -0.2278   0.8967   0.0282
  -7.500   0.4713   0.07628   0.07294  -0.2287   0.8945   0.0285
  -7.250   0.4894   0.07422   0.07085  -0.2311   0.8922   0.0287
  -7.000   0.5098   0.07215   0.06875  -0.2341   0.8902   0.0291
  -6.750   0.5099   0.07127   0.06789  -0.2315   0.8856   0.0295
  -6.500   0.5079   0.07048   0.06712  -0.2284   0.8805   0.0300
  -6.250   0.5208   0.06877   0.06539  -0.2294   0.8774   0.0307
  -6.000   0.5374   0.06694   0.06353  -0.2318   0.8746   0.0312
  -5.750   0.5237   0.06691   0.06353  -0.2258   0.8687   0.0313
  -5.500   0.5237   0.06635   0.06298  -0.2243   0.8635   0.0313
  -5.250   0.5422   0.06455   0.06114  -0.2276   0.8606   0.0314
  -5.000   0.5630   0.06175   0.05833  -0.2294   0.8589   0.0316
  -4.750   0.5625   0.06082   0.05742  -0.2263   0.8545   0.0317
  -4.500   0.5561   0.06021   0.05684  -0.2217   0.8491   0.0318
  -4.250   0.5734   0.05857   0.05519  -0.2232   0.8463   0.0322
  -4.000   0.5981   0.05670   0.05329  -0.2267   0.8441   0.0328
  -3.750   0.6277   0.05464   0.05117  -0.2315   0.8420   0.0338
  -3.500   0.5982   0.05503   0.05165  -0.2209   0.8347   0.0338
  -3.250   0.6157   0.05352   0.05012  -0.2226   0.8312   0.0344
  -3.000   0.6626   0.05171   0.04822  -0.2330   0.8289   0.0347
  -2.750   0.6925   0.04909   0.04556  -0.2372   0.8270   0.0348
  -2.500   0.6780   0.04852   0.04505  -0.2299   0.8216   0.0349
  -2.250   0.6724   0.04754   0.04410  -0.2247   0.8166   0.0350
  -2.000   0.6980   0.04579   0.04233  -0.2272   0.8143   0.0354
  -1.750   0.7345   0.04387   0.04037  -0.2325   0.8125   0.0360
  -1.500   0.7786   0.04179   0.03823  -0.2396   0.8109   0.0371
  -1.250   0.7867   0.04087   0.03732  -0.2378   0.8065   0.0378
  -1.000   0.8184   0.03991   0.03632  -0.2421   0.8020   0.0382
  -0.500   0.8782   0.03588   0.03224  -0.2480   0.7972   0.0386
  -0.250   0.9203   0.03408   0.03040  -0.2535   0.7951   0.0391
   0.000   0.9382   0.03302   0.02933  -0.2532   0.7903   0.0397
   0.250   0.9508   0.03212   0.02844  -0.2516   0.7854   0.0403
   0.750   1.0686   0.02811   0.02417  -0.2694   0.7765   0.0426
   1.000   1.0602   0.02767   0.02379  -0.2626   0.7699   0.0427
   1.250   1.0851   0.02655   0.02264  -0.2632   0.7641   0.0433
   1.500   1.1143   0.02553   0.02157  -0.2647   0.7578   0.0441
   1.750   1.1272   0.02489   0.02093  -0.2625   0.7497   0.0451
   2.000   1.1839   0.02355   0.01935  -0.2694   0.7409   0.0470
   2.250   1.1901   0.02290   0.01870  -0.2658   0.7297   0.0473
   2.500   1.1978   0.02256   0.01836  -0.2624   0.7172   0.0477
   2.750   1.2069   0.02240   0.01811  -0.2592   0.6996   0.0483
   3.000   1.2112   0.02251   0.01807  -0.2549   0.6718   0.0490
   3.250   1.2085   0.02301   0.01835  -0.2493   0.6335   0.0498
   3.500   1.1943   0.02417   0.01920  -0.2417   0.5818   0.0502
   3.750   1.1674   0.02620   0.02083  -0.2321   0.5119   0.0504
   4.000   1.1448   0.02846   0.02266  -0.2238   0.4345   0.0509
   4.250   1.1111   0.03170   0.02531  -0.2143   0.3160   0.0506
   4.500   1.0822   0.03508   0.02805  -0.2061   0.1650   0.0506
   5.000   1.1102   0.03685   0.02926  -0.2036   0.0389   0.0528
   5.250   1.1277   0.03710   0.02952  -0.2027   0.0369   0.0534
   5.500   1.1462   0.03746   0.02987  -0.2019   0.0351   0.0546
   6.000   1.1901   0.03816   0.03042  -0.2011   0.0331   0.0591
   6.250   1.2053   0.03876   0.03104  -0.1998   0.0327   0.0600
   6.500   1.2217   0.03941   0.03168  -0.1986   0.0324   0.0615
   6.750   1.2451   0.03979   0.03190  -0.1981   0.0320   0.0660
   7.000   1.2573   0.04072   0.03286  -0.1964   0.0317   0.0669
   7.250   1.2702   0.04171   0.03386  -0.1947   0.0312   0.0686
   9.500   1.3788   0.05319   0.04492  -0.1788   0.0292   0.0717
   9.750   1.3891   0.05473   0.04645  -0.1770   0.0291   0.0699
  10.000   1.4005   0.05622   0.04785  -0.1752   0.0290   0.0658
  10.250   1.4110   0.05780   0.04943  -0.1733   0.0290   0.0650
  10.500   1.4242   0.05906   0.05071  -0.1717   0.0289   0.0646
  10.750   1.4386   0.06022   0.05189  -0.1702   0.0289   0.0644
  11.000   1.4550   0.06121   0.05287  -0.1688   0.0287   0.0649
  11.250   1.4750   0.06185   0.05349  -0.1678   0.0283   0.0658
  11.500   1.5016   0.06197   0.05360  -0.1671   0.0282   0.0675
  11.750   1.5262   0.06233   0.05402  -0.1663   0.0285   0.0705
  15.250   1.6693   0.08775   0.08320  -0.1386   0.0301   1.0000
  16.000   1.6214   0.09787   0.09387  -0.1318   0.0293   1.0000
  16.250   1.6064   0.10135   0.09749  -0.1302   0.0291   1.0000
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