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

GOE 531 AIRFOIL (goe531-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 531 AIRFOIL (goe531-il)
Reynolds number: 1,000,000
Max Cl/Cd: 86.71 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe531-il-1000000-n5.txt
Download as CSV file: xf-goe531-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 531 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750   0.6217   0.09261   0.08837  -0.2417   0.6607   0.0207
 -11.500   0.6305   0.09094   0.08670  -0.2423   0.6597   0.0210
 -11.250   0.6378   0.08908   0.08485  -0.2427   0.6588   0.0210
 -11.000   0.6457   0.08735   0.08312  -0.2432   0.6579   0.0213
 -10.750   0.6526   0.08555   0.08132  -0.2435   0.6568   0.0215
 -10.500   0.6588   0.08371   0.07949  -0.2437   0.6558   0.0218
 -10.250   0.6567   0.08076   0.07652  -0.2437   0.6546   0.0227
 -10.000   0.6630   0.07912   0.07489  -0.2438   0.6533   0.0227
  -9.750   0.6739   0.07816   0.07393  -0.2439   0.6518   0.0229
  -9.500   0.6822   0.07679   0.07257  -0.2440   0.6509   0.0230
  -9.250   0.6919   0.07559   0.07139  -0.2441   0.6504   0.0232
  -9.000   0.6997   0.07421   0.07003  -0.2441   0.6498   0.0233
  -8.750   0.7066   0.07279   0.06863  -0.2440   0.6491   0.0235
  -8.500   0.7133   0.07140   0.06726  -0.2437   0.6484   0.0237
  -8.250   0.7191   0.07000   0.06589  -0.2434   0.6476   0.0239
  -3.750   0.7606   0.01633   0.01131  -0.2654   0.6270   0.0454
  -3.500   0.7847   0.01519   0.01002  -0.2661   0.6258   0.0463
  -3.250   0.8075   0.01443   0.00917  -0.2662   0.6246   0.0471
  -3.000   0.8291   0.01387   0.00850  -0.2657   0.6232   0.0480
  -2.750   0.8502   0.01344   0.00798  -0.2651   0.6218   0.0487
  -2.500   0.8715   0.01326   0.00771  -0.2642   0.6205   0.0494
  -2.250   0.8928   0.01315   0.00752  -0.2633   0.6189   0.0500
  -2.000   0.9153   0.01307   0.00738  -0.2626   0.6174   0.0507
  -1.750   0.9383   0.01303   0.00733  -0.2620   0.6166   0.0516
  -1.500   0.9603   0.01304   0.00733  -0.2612   0.6158   0.0523
  -1.250   0.9821   0.01305   0.00733  -0.2603   0.6148   0.0530
  -1.000   1.0041   0.01308   0.00733  -0.2595   0.6138   0.0538
  -0.750   1.0260   0.01310   0.00733  -0.2586   0.6126   0.0546
  -0.500   1.0478   0.01314   0.00734  -0.2578   0.6112   0.0552
  -0.250   1.0691   0.01318   0.00736  -0.2568   0.6097   0.0559
   0.000   1.0891   0.01328   0.00746  -0.2557   0.6080   0.0566
   0.250   1.1087   0.01342   0.00757  -0.2544   0.6066   0.0572
   0.500   1.1284   0.01356   0.00770  -0.2532   0.6050   0.0581
   0.750   1.1481   0.01371   0.00783  -0.2520   0.6035   0.0590
   1.000   1.1680   0.01387   0.00795  -0.2509   0.6019   0.0597
   1.250   1.1892   0.01401   0.00806  -0.2500   0.6006   0.0604
   1.500   1.2093   0.01415   0.00821  -0.2489   0.5996   0.0610
   1.750   1.2293   0.01431   0.00839  -0.2478   0.5984   0.0617
   2.000   1.2496   0.01449   0.00860  -0.2468   0.5971   0.0626
   2.250   1.2700   0.01468   0.00879  -0.2458   0.5957   0.0635
   2.500   1.2900   0.01488   0.00899  -0.2447   0.5941   0.0643
   2.750   1.3093   0.01510   0.00922  -0.2436   0.5924   0.0652
   3.000   1.3262   0.01541   0.00952  -0.2420   0.5905   0.0658
   3.250   1.3433   0.01572   0.00982  -0.2406   0.5887   0.0664
   3.500   1.3605   0.01603   0.01012  -0.2391   0.5868   0.0674
   3.750   1.3790   0.01632   0.01040  -0.2379   0.5849   0.0683
   4.000   1.3985   0.01659   0.01070  -0.2369   0.5838   0.0694
   4.250   1.4160   0.01694   0.01109  -0.2356   0.5824   0.0703
   4.500   1.4340   0.01729   0.01146  -0.2343   0.5810   0.0711
   4.750   1.4517   0.01766   0.01184  -0.2331   0.5791   0.0721
   5.000   1.4684   0.01807   0.01226  -0.2317   0.5768   0.0729
   5.250   1.4847   0.01850   0.01271  -0.2302   0.5747   0.0742
   5.500   1.4969   0.01911   0.01332  -0.2282   0.5721   0.0752
   5.750   1.5107   0.01966   0.01386  -0.2264   0.5693   0.0764
   6.000   1.5277   0.02012   0.01436  -0.2251   0.5676   0.0773
   6.250   1.5438   0.02064   0.01491  -0.2238   0.5659   0.0786
   6.500   1.5582   0.02125   0.01555  -0.2222   0.5640   0.0797
   6.750   1.5721   0.02188   0.01621  -0.2205   0.5612   0.0809
   7.000   1.5858   0.02253   0.01688  -0.2189   0.5584   0.0831
   7.250   1.5912   0.02362   0.01795  -0.2161   0.5541   0.0848
   7.500   1.6026   0.02445   0.01882  -0.2142   0.5506   0.0867
   7.750   1.6152   0.02525   0.01965  -0.2125   0.5476   0.0884
   8.000   1.6247   0.02623   0.02067  -0.2105   0.5446   0.0920
   8.250   1.6378   0.02700   0.02148  -0.2089   0.5423   0.0972
   8.500   1.6501   0.02782   0.02233  -0.2073   0.5398   0.1091
   8.750   1.6575   0.02893   0.02348  -0.2050   0.5368   0.1305
   9.000   1.6709   0.02974   0.02438  -0.2036   0.5351   0.1543
   9.250   1.6833   0.03063   0.02533  -0.2021   0.5326   0.1669
   9.500   1.6920   0.03177   0.02653  -0.2002   0.5297   0.1770
   9.750   1.6994   0.03300   0.02779  -0.1981   0.5260   0.1849
  10.000   1.7051   0.03432   0.02910  -0.1958   0.5211   0.1929
  10.250   1.7118   0.03567   0.03051  -0.1937   0.5180   0.1995
  10.500   1.7195   0.03695   0.03184  -0.1918   0.5142   0.2066
  10.750   1.7270   0.03822   0.03314  -0.1898   0.5098   0.2120
  11.000   1.7288   0.03993   0.03484  -0.1873   0.5050   0.2161
  11.250   1.7357   0.04130   0.03626  -0.1854   0.5010   0.2206
  11.500   1.7421   0.04273   0.03773  -0.1834   0.4954   0.2262
  11.750   1.7463   0.04432   0.03933  -0.1813   0.4910   0.2316
  12.000   1.7479   0.04611   0.04114  -0.1789   0.4854   0.2352
  12.250   1.7558   0.04746   0.04252  -0.1772   0.4801   0.2383
  12.500   1.7523   0.04969   0.04475  -0.1743   0.4721   0.2420
  12.750   1.7468   0.05217   0.04723  -0.1713   0.4596   0.2469
  13.000   1.7448   0.05434   0.04938  -0.1687   0.4493   0.2509
  13.250   1.7328   0.05743   0.05245  -0.1652   0.4361   0.2524
  13.500   1.7123   0.06131   0.05624  -0.1609   0.4128   0.2537
  13.750   1.6269   0.07136   0.06595  -0.1510   0.3537   0.2493
  14.000   1.6023   0.07611   0.07061  -0.1470   0.3310   0.2509
  14.250   1.5942   0.07930   0.07373  -0.1446   0.3166   0.2541
  14.500   1.5903   0.08212   0.07654  -0.1426   0.3066   0.2562
  14.750   1.5901   0.08460   0.07900  -0.1409   0.2974   0.2596
  15.000   1.5938   0.08666   0.08106  -0.1396   0.2891   0.2640
  15.250   1.5935   0.08919   0.08359  -0.1381   0.2811   0.2676
  15.500   1.5975   0.09129   0.08569  -0.1369   0.2729   0.2711
  15.750   1.5989   0.09365   0.08802  -0.1355   0.2638   0.2745
  16.000   1.6045   0.09554   0.08991  -0.1345   0.2544   0.2784
  16.250   1.6033   0.09823   0.09258  -0.1332   0.2447   0.2832
<< Back to GOE 531 AIRFOIL (goe531-il)

Polar data table (+)

Polar graphs


<< Back to GOE 531 AIRFOIL (goe531-il)