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Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: Cambered plate C=8% T=5% R=1.6 (cp-080-050-gn)
Reynolds number: 50,000
Max Cl/Cd: 25.45 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-cp-080-050-gn-50000.txt
Download as CSV file: xf-cp-080-050-gn-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Cambered plate C=8% T=5% R=1.6                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -4.000  -0.3877   0.11552   0.10846  -0.0081   1.0000   0.1466
  -3.750  -0.3929   0.11401   0.10701  -0.0071   1.0000   0.1491
  -3.500  -0.3998   0.11300   0.10605  -0.0072   1.0000   0.1517
  -3.250  -0.4090   0.11409   0.10715  -0.0124   1.0000   0.1539
  -3.000  -0.4055   0.10995   0.10309  -0.0099   1.0000   0.1551
  -2.750  -0.4006   0.10625   0.09947  -0.0064   1.0000   0.1574
  -2.500  -0.3974   0.10378   0.09705  -0.0051   1.0000   0.1607
  -2.250  -0.3941   0.10186   0.09515  -0.0056   1.0000   0.1649
  -2.000  -0.3821   0.10233   0.09551  -0.0150   1.0000   0.1698
  -1.750  -0.3809   0.09781   0.09112  -0.0104   1.0000   0.1714
  -1.500  -0.3765   0.09478   0.08816  -0.0079   1.0000   0.1741
  -1.250  -0.3690   0.09246   0.08587  -0.0077   1.0000   0.1784
  -1.000  -0.3422   0.09237   0.08561  -0.0175   1.0000   0.1864
  -0.750  -0.3414   0.08824   0.08161  -0.0134   1.0000   0.1884
  -0.500  -0.3352   0.08548   0.07892  -0.0114   1.0000   0.1928
  -0.250  -0.3051   0.08501   0.07830  -0.0193   1.0000   0.2039
   0.000  -0.3029   0.08128   0.07470  -0.0157   1.0000   0.2068
   0.250  -0.2920   0.07901   0.07248  -0.0151   1.0000   0.2137
   0.500  -0.2671   0.07737   0.07076  -0.0196   1.0000   0.2234
   0.750  -0.2583   0.07480   0.06828  -0.0180   1.0000   0.2293
   1.000  -0.2327   0.07339   0.06679  -0.0220   1.0000   0.2416
   1.250  -0.2224   0.07102   0.06451  -0.0206   1.0000   0.2494
   1.500  -0.2024   0.06922   0.06270  -0.0224   1.0000   0.2629
   1.750  -0.1823   0.06769   0.06118  -0.0240   1.0000   0.2798
   2.000  -0.1645   0.06606   0.05957  -0.0247   1.0000   0.2981
   2.250  -0.1484   0.06432   0.05788  -0.0249   1.0000   0.3174
   2.500  -0.1339   0.06248   0.05611  -0.0244   1.0000   0.3383
   2.750  -0.1169   0.06115   0.05482  -0.0247   1.0000   0.3731
   3.000  -0.1077   0.05900   0.05281  -0.0225   1.0000   0.4013
   3.750  -0.0804   0.05318   0.04737  -0.0153   1.0000   0.5366
   4.000  -0.0641   0.05206   0.04632  -0.0145   1.0000   0.5760
   4.250  -0.0520   0.05035   0.04474  -0.0125   1.0000   0.6037
   4.500  -0.0342   0.04929   0.04377  -0.0120   1.0000   0.6290
   4.750  -0.0110   0.04862   0.04316  -0.0130   1.0000   0.6495
   5.000   0.0146   0.04835   0.04291  -0.0146   1.0000   0.6637
   5.250   0.0424   0.04820   0.04278  -0.0170   1.0000   0.6710
   5.500   0.0714   0.04833   0.04293  -0.0197   1.0000   0.6735
   5.750   0.6448   0.03550   0.02483  -0.0768   0.2116   0.2643
   6.000   0.6601   0.03656   0.02552  -0.0745   0.1979   0.2566
   6.250   0.6741   0.03721   0.02598  -0.0723   0.1896   0.2535
   6.500   0.6952   0.03766   0.02632  -0.0707   0.1833   0.2491
   6.750   0.7204   0.03823   0.02660  -0.0695   0.1778   0.2429
   7.000   0.7551   0.03883   0.02673  -0.0693   0.1722   0.2385
   7.250   0.8019   0.03880   0.02655  -0.0704   0.1667   0.2376
   7.500   0.8874   0.03865   0.02613  -0.0769   0.1612   0.2443
   7.750   0.9920   0.03987   0.02704  -0.0874   0.1583   0.2547
   8.000   1.0495   0.04133   0.02850  -0.0912   0.1574   0.2688
   8.250   1.0955   0.04324   0.03046  -0.0935   0.1562   0.2942
   8.500   1.1287   0.04471   0.03230  -0.0937   0.1554   0.3402
   8.750   1.1631   0.04570   0.03415  -0.0937   0.1552   1.0000
   9.000   1.1946   0.04818   0.03643  -0.0934   0.1559   1.0000
   9.250   1.2133   0.04940   0.03790  -0.0907   0.1579   1.0000
   9.500   1.2268   0.05132   0.04027  -0.0875   0.1612   1.0000
   9.750   1.2419   0.05393   0.04320  -0.0850   0.1646   1.0000
  10.000   1.2572   0.05683   0.04634  -0.0829   0.1674   1.0000
  10.250   1.2719   0.06000   0.04968  -0.0809   0.1697   1.0000
  10.500   1.2890   0.06371   0.05347  -0.0795   0.1717   1.0000
  10.750   1.2774   0.06546   0.05590  -0.0736   0.1767   1.0000
  11.000   1.2623   0.06906   0.06002  -0.0685   0.1831   1.0000
  11.250   1.2844   0.07443   0.06536  -0.0687   0.1894   1.0000
  11.750   1.2338   0.08261   0.07448  -0.0590   0.2098   1.0000
  12.000   1.1458   0.08649   0.07882  -0.0508   0.2129   1.0000
  12.250   1.0641   0.09597   0.08865  -0.0504   0.2225   1.0000
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