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

HAM-STD HS1-404 AIRFOIL (hs1404-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: HAM-STD HS1-404 AIRFOIL (hs1404-il)
Reynolds number: 1,000,000
Max Cl/Cd: 113.34 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-hs1404-il-1000000-n5.txt
Download as CSV file: xf-hs1404-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HAM-STD HS1-404 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4500   0.09662   0.09499  -0.0075   1.0000   0.0031
  -7.750  -0.4442   0.09327   0.09166  -0.0091   1.0000   0.0031
  -7.500  -0.4391   0.09001   0.08841  -0.0107   1.0000   0.0031
  -7.250  -0.4319   0.08654   0.08497  -0.0131   1.0000   0.0031
  -7.000  -0.4214   0.08277   0.08121  -0.0165   1.0000   0.0031
  -6.750  -0.4101   0.07895   0.07741  -0.0201   1.0000   0.0031
  -6.500  -0.3989   0.07517   0.07363  -0.0234   1.0000   0.0030
  -6.250  -0.3736   0.07001   0.06842  -0.0308   0.9964   0.0030
  -6.000  -0.3446   0.06437   0.06276  -0.0391   0.9918   0.0030
  -5.750  -0.3124   0.05747   0.05580  -0.0493   0.9856   0.0025
  -5.500  -0.2762   0.05124   0.04950  -0.0583   0.9800   0.0022
  -5.250  -0.2379   0.04484   0.04297  -0.0668   0.9727   0.0020
  -5.000  -0.1980   0.03803   0.03599  -0.0744   0.9653   0.0018
  -4.750  -0.1574   0.03019   0.02787  -0.0808   0.9570   0.0017
  -4.500  -0.1165   0.01915   0.01614  -0.0858   0.9483   0.0015
  -4.250  -0.0837   0.01394   0.01019  -0.0871   0.9399   0.0016
  -4.000  -0.0544   0.01182   0.00765  -0.0873   0.9307   0.0019
  -3.750  -0.0263   0.01091   0.00653  -0.0873   0.9210   0.0025
  -3.500   0.0010   0.01059   0.00610  -0.0871   0.9119   0.0029
  -3.250   0.0288   0.00996   0.00529  -0.0871   0.9038   0.0033
  -3.000   0.0574   0.00872   0.00385  -0.0873   0.8962   0.0044
  -2.750   0.0848   0.00852   0.00360  -0.0873   0.8888   0.0057
  -2.500   0.1135   0.00780   0.00272  -0.0872   0.8799   0.0055
  -2.250   0.1419   0.00728   0.00201  -0.0871   0.8688   0.0053
  -2.000   0.1701   0.00696   0.00153  -0.0870   0.8560   0.0052
  -1.750   0.1982   0.00676   0.00120  -0.0868   0.8438   0.0051
  -1.500   0.2262   0.00665   0.00097  -0.0867   0.8329   0.0053
  -1.250   0.2542   0.00658   0.00079  -0.0866   0.8214   0.0065
  -1.000   0.2819   0.00656   0.00066  -0.0866   0.8099   0.0063
  -0.750   0.3102   0.00632   0.00059  -0.0867   0.8000   0.0643
  -0.500   0.3385   0.00609   0.00058  -0.0870   0.7927   0.1397
   0.000   0.3946   0.00595   0.00062  -0.0871   0.7764   0.2141
   0.250   0.4225   0.00595   0.00062  -0.0872   0.7667   0.2259
   0.500   0.4503   0.00595   0.00065  -0.0872   0.7547   0.2391
   0.750   0.4782   0.00593   0.00067  -0.0872   0.7404   0.2640
   1.000   0.5061   0.00568   0.00074  -0.0875   0.7181   0.4229
   1.250   0.5259   0.00464   0.00081  -0.0859   0.6452   1.0000
   1.500   0.5502   0.00542   0.00099  -0.0855   0.5055   1.0000
   1.750   0.5754   0.00608   0.00118  -0.0853   0.3906   1.0000
   2.000   0.6010   0.00665   0.00137  -0.0851   0.2994   1.0000
   2.250   0.6276   0.00699   0.00156  -0.0850   0.2508   1.0000
   2.500   0.6540   0.00738   0.00173  -0.0849   0.1969   1.0000
   2.750   0.6797   0.00790   0.00198  -0.0848   0.1282   1.0000
   3.000   0.7053   0.00846   0.00226  -0.0846   0.0642   1.0000
   3.250   0.7320   0.00874   0.00247  -0.0844   0.0465   1.0000
   3.500   0.7584   0.00906   0.00270  -0.0843   0.0289   1.0000
   3.750   0.7852   0.00931   0.00299  -0.0841   0.0224   1.0000
   4.000   0.8117   0.00960   0.00325  -0.0840   0.0154   1.0000
   4.250   0.8382   0.00988   0.00352  -0.0838   0.0099   1.0000
   4.500   0.8644   0.01021   0.00383  -0.0836   0.0059   1.0000
   4.750   0.8901   0.01063   0.00421  -0.0833   0.0017   1.0000
   5.000   0.9162   0.01099   0.00462  -0.0830   0.0011   1.0000
   5.250   0.9419   0.01144   0.00515  -0.0826   0.0009   1.0000
   5.500   0.9672   0.01201   0.00583  -0.0821   0.0007   1.0000
   5.750   0.9917   0.01273   0.00673  -0.0815   0.0007   1.0000
   6.000   1.0155   0.01365   0.00782  -0.0807   0.0006   1.0000
   6.250   1.0384   0.01483   0.00919  -0.0796   0.0006   1.0000
   6.500   1.0600   0.01644   0.01106  -0.0783   0.0007   1.0000
   6.750   1.0799   0.01893   0.01394  -0.0766   0.0007   1.0000
   7.000   1.0956   0.02356   0.01920  -0.0738   0.0008   1.0000
   7.250   1.0967   0.03424   0.03087  -0.0681   0.0009   1.0000
   7.500   1.0964   0.04352   0.04073  -0.0637   0.0010   1.0000
   7.750   1.0963   0.05061   0.04820  -0.0606   0.0011   1.0000
   8.000   1.0934   0.05712   0.05500  -0.0582   0.0012   1.0000
   8.250   1.0875   0.06321   0.06133  -0.0564   0.0012   1.0000
   8.500   1.0781   0.06904   0.06734  -0.0551   0.0013   1.0000
   8.750   1.0646   0.07445   0.07289  -0.0542   0.0013   1.0000
   9.000   1.0465   0.07904   0.07758  -0.0535   0.0013   1.0000
   9.250   1.0291   0.08488   0.08351  -0.0562   0.0012   1.0000
<< Back to HAM-STD HS1-404 AIRFOIL (hs1404-il)

Polar data table (+)

Polar graphs


<< Back to HAM-STD HS1-404 AIRFOIL (hs1404-il)