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S1223 RTL (s1223rtl-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: S1223 RTL (s1223rtl-il)
Reynolds number: 50,000
Max Cl/Cd: 31.57 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s1223rtl-il-50000.txt
Download as CSV file: xf-s1223rtl-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S1223 RTL                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -6.000  -0.2273   0.13180   0.12702  -0.0060   1.0000   0.1440
  -5.750  -0.2488   0.13347   0.12885  -0.0046   1.0000   0.1448
  -5.500  -0.2707   0.13495   0.13050  -0.0029   1.0000   0.1451
  -5.250  -0.2302   0.12541   0.12104  -0.0018   1.0000   0.1504
  -5.000  -0.1950   0.12324   0.11888  -0.0114   0.9841   0.1599
  -4.750  -0.1536   0.11823   0.11389  -0.0203   0.9641   0.1660
  -4.500  -0.1147   0.11446   0.11013  -0.0285   0.9407   0.1774
  -4.250  -0.0702   0.10846   0.10415  -0.0363   0.9173   0.1852
  -4.000  -0.0389   0.10696   0.10267  -0.0494   0.8817   0.1982
  -3.750   0.0137   0.09783   0.09354  -0.0495   0.8584   0.2101
  -3.500   0.0482   0.09287   0.08859  -0.0554   0.8259   0.2225
  -3.250   0.0852   0.08869   0.08437  -0.0637   0.7957   0.2386
  -3.000   0.1274   0.08488   0.08048  -0.0726   0.7675   0.2571
  -2.750   0.1617   0.07757   0.07309  -0.0683   0.7478   0.2664
  -2.500   0.2002   0.07324   0.06860  -0.0743   0.7237   0.2822
  -2.250   0.2420   0.06932   0.06448  -0.0810   0.6993   0.3005
  -2.000   0.2872   0.06514   0.06004  -0.0864   0.6777   0.3215
  -1.750   0.3305   0.06150   0.05611  -0.0914   0.6580   0.3429
  -1.500   0.7183   0.03559   0.02752  -0.2026   0.6183   0.1585
  -1.250   0.8342   0.03208   0.02278  -0.2224   0.5976   0.1774
  -1.000   0.8898   0.03142   0.02188  -0.2280   0.5833   0.2067
  -0.750   0.9403   0.03141   0.02180  -0.2320   0.5722   0.2712
  -0.500   0.9664   0.03230   0.02294  -0.2296   0.5639   0.3802
  -0.250   0.9815   0.03329   0.02397  -0.2249   0.5574   0.4589
   0.000   0.9949   0.03395   0.02470  -0.2209   0.5505   0.5144
   0.250   1.0178   0.03436   0.02498  -0.2189   0.5438   0.5634
   0.500   1.0499   0.03479   0.02522  -0.2196   0.5373   0.6010
   0.750   1.0774   0.03526   0.02565  -0.2199   0.5309   0.6274
   1.000   1.1148   0.03582   0.02600  -0.2222   0.5252   0.6547
   1.250   1.1523   0.03650   0.02641  -0.2244   0.5202   0.6807
   1.500   1.1735   0.03734   0.02735  -0.2239   0.5153   0.7028
   1.750   1.2008   0.03824   0.02823  -0.2246   0.5100   0.7291
   2.000   1.2270   0.03898   0.02890  -0.2245   0.5053   0.7567
   2.250   1.2568   0.03984   0.02962  -0.2251   0.5012   0.7911
   2.500   1.2676   0.04102   0.03101  -0.2228   0.4982   0.8225
   2.750   1.2727   0.04220   0.03245  -0.2195   0.4953   0.8620
   3.000   1.2770   0.04339   0.03389  -0.2163   0.4923   1.0000
   3.250   1.3202   0.04575   0.03611  -0.2216   0.4875   1.0000
   3.500   1.3699   0.04767   0.03770  -0.2271   0.4829   1.0000
   3.750   1.3970   0.05014   0.04005  -0.2286   0.4797   1.0000
   4.000   1.3944   0.05346   0.04360  -0.2258   0.4777   1.0000
   4.250   1.3795   0.05764   0.04804  -0.2217   0.4764   1.0000
   4.500   1.3368   0.06377   0.05453  -0.2151   0.4760   1.0000
   4.750   1.2342   0.07517   0.06647  -0.2052   0.4787   1.0000
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