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 HS1-708 AIRFOIL (hs1708-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HAM-STD HS1-708 AIRFOIL (hs1708-il)
Reynolds number: 100,000
Max Cl/Cd: 67.77 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hs1708-il-100000.txt
Download as CSV file: xf-hs1708-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HAM-STD HS1-708 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3853   0.10831   0.10389  -0.0193   1.0000   0.0780
  -7.500  -0.4034   0.10788   0.10355  -0.0168   1.0000   0.0790
  -7.250  -0.4190   0.10733   0.10308  -0.0169   1.0000   0.0799
  -7.000  -0.4277   0.10634   0.10214  -0.0215   1.0000   0.0805
  -6.750  -0.4262   0.10352   0.09935  -0.0269   1.0000   0.0810
  -6.500  -0.4254   0.09858   0.09446  -0.0180   1.0000   0.0823
  -6.250  -0.4222   0.09564   0.09153  -0.0153   1.0000   0.0839
  -6.000  -0.4189   0.09295   0.08887  -0.0148   1.0000   0.0859
  -5.750  -0.4137   0.09011   0.08604  -0.0161   1.0000   0.0886
  -5.500  -0.3763   0.08641   0.08212  -0.0381   1.0000   0.0944
  -5.250  -0.3789   0.08193   0.07776  -0.0323   1.0000   0.0954
  -5.000  -0.3751   0.07891   0.07478  -0.0286   1.0000   0.0970
  -4.750  -0.3637   0.07600   0.07185  -0.0286   1.0000   0.1002
  -4.500  -0.3163   0.07047   0.06606  -0.0443   1.0000   0.1091
  -4.250  -0.3139   0.06784   0.06353  -0.0395   1.0000   0.1119
  -4.000  -0.2739   0.06342   0.05890  -0.0489   1.0000   0.1239
  -3.750  -0.2631   0.06127   0.05679  -0.0465   1.0000   0.1324
  -3.500  -0.2318   0.05733   0.05272  -0.0513   1.0000   0.1427
  -3.250  -0.1677   0.05271   0.04769  -0.0637   0.9961   0.1678
  -3.000  -0.0775   0.04078   0.03466  -0.0807   0.9942   0.1195
  -2.750  -0.0173   0.03550   0.02847  -0.0888   0.9914   0.1207
  -2.500   0.0291   0.03334   0.02593  -0.0934   0.9879   0.1268
  -2.250   0.0817   0.03155   0.02336  -0.0986   0.9852   0.1368
  -2.000   0.1143   0.03101   0.02286  -0.1003   0.9802   0.1456
  -1.750   0.1557   0.03030   0.02186  -0.1034   0.9761   0.1570
  -1.500   0.2016   0.02983   0.02112  -0.1073   0.9727   0.1707
  -1.250   0.2324   0.02956   0.02071  -0.1084   0.9669   0.1843
  -1.000   0.2720   0.02928   0.02030  -0.1110   0.9621   0.1995
  -0.750   0.3159   0.02926   0.02023  -0.1145   0.9580   0.2218
  -0.500   0.3442   0.02912   0.02002  -0.1151   0.9506   0.2393
  -0.250   0.3889   0.02895   0.01992  -0.1187   0.9457   0.2655
   0.000   0.4239   0.02871   0.01975  -0.1203   0.9358   0.2884
   0.250   0.4707   0.02827   0.01942  -0.1238   0.9261   0.3198
   0.500   0.5205   0.02769   0.01898  -0.1276   0.9172   0.3595
   0.750   0.5535   0.02738   0.01895  -0.1287   0.9073   0.4116
   1.000   0.5932   0.02595   0.01868  -0.1302   0.9019   1.0000
   1.250   0.6197   0.02624   0.01881  -0.1299   0.8910   1.0000
   1.500   0.6557   0.02636   0.01882  -0.1312   0.8824   1.0000
   1.750   0.6957   0.02629   0.01868  -0.1331   0.8743   1.0000
   2.000   0.7252   0.02642   0.01877  -0.1332   0.8633   1.0000
   2.250   0.7751   0.02592   0.01827  -0.1364   0.8579   1.0000
   2.500   0.8036   0.02591   0.01828  -0.1361   0.8458   1.0000
   2.750   0.8379   0.02568   0.01807  -0.1366   0.8352   1.0000
   3.000   0.8874   0.02474   0.01720  -0.1392   0.8296   1.0000
   3.250   0.9180   0.02438   0.01690  -0.1388   0.8173   1.0000
   3.500   0.9504   0.02389   0.01648  -0.1384   0.8054   1.0000
   3.750   0.9856   0.02318   0.01588  -0.1384   0.7943   1.0000
   4.000   1.0281   0.02201   0.01479  -0.1392   0.7856   1.0000
   4.250   1.0598   0.02124   0.01411  -0.1384   0.7709   1.0000
   4.500   1.0845   0.02073   0.01373  -0.1365   0.7512   1.0000
   4.750   1.1140   0.01997   0.01306  -0.1353   0.7314   1.0000
   5.000   1.1420   0.01928   0.01246  -0.1338   0.7074   1.0000
   5.250   1.1679   0.01868   0.01192  -0.1320   0.6755   1.0000
   5.500   1.1941   0.01812   0.01128  -0.1302   0.6323   1.0000
   5.750   1.2185   0.01798   0.01081  -0.1281   0.5782   1.0000
   6.000   1.2392   0.01848   0.01098  -0.1261   0.5275   1.0000
   6.250   1.2584   0.01920   0.01149  -0.1242   0.4869   1.0000
   6.500   1.2767   0.01995   0.01208  -0.1222   0.4517   1.0000
   6.750   1.2937   0.02072   0.01273  -0.1202   0.4183   1.0000
   7.000   1.3092   0.02152   0.01344  -0.1179   0.3857   1.0000
   7.250   1.3234   0.02237   0.01424  -0.1155   0.3532   1.0000
   7.500   1.3338   0.02332   0.01509  -0.1125   0.3152   1.0000
   7.750   1.3378   0.02447   0.01607  -0.1085   0.2600   1.0000
   8.000   1.3333   0.02637   0.01741  -0.1035   0.1865   1.0000
   8.250   1.3313   0.02843   0.01905  -0.0989   0.1487   1.0000
   8.500   1.3361   0.03028   0.02073  -0.0954   0.1300   1.0000
   8.750   1.3447   0.03204   0.02238  -0.0926   0.1178   1.0000
   9.000   1.3562   0.03371   0.02400  -0.0903   0.1078   1.0000
   9.250   1.3726   0.03531   0.02568  -0.0886   0.0997   1.0000
   9.500   1.4009   0.03746   0.02765  -0.0887   0.0931   1.0000
   9.750   1.4238   0.03918   0.02960  -0.0878   0.0877   1.0000
  10.000   1.4636   0.04207   0.03233  -0.0902   0.0808   1.0000
  10.250   1.4845   0.04416   0.03483  -0.0890   0.0778   1.0000
  10.500   1.5076   0.04673   0.03769  -0.0884   0.0748   1.0000
  10.750   1.5266   0.04925   0.04036  -0.0874   0.0715   1.0000
  11.000   1.5464   0.05344   0.04473  -0.0871   0.0683   1.0000
  11.250   1.5468   0.05605   0.04782  -0.0834   0.0671   1.0000
  11.500   1.5445   0.05915   0.05137  -0.0796   0.0663   1.0000
  11.750   1.5360   0.06227   0.05489  -0.0752   0.0657   1.0000
  12.000   1.5233   0.06547   0.05845  -0.0708   0.0651   1.0000
  12.250   1.5075   0.06884   0.06217  -0.0666   0.0645   1.0000
  12.500   1.4892   0.07250   0.06615  -0.0629   0.0640   1.0000
  12.750   1.4683   0.07651   0.07046  -0.0599   0.0636   1.0000
  13.000   1.4448   0.08103   0.07527  -0.0576   0.0636   1.0000
  13.250   1.4183   0.08619   0.08072  -0.0561   0.0638   1.0000
  13.500   1.3895   0.09205   0.08685  -0.0557   0.0643   1.0000
  13.750   1.3592   0.09868   0.09371  -0.0565   0.0650   1.0000
  14.000   1.3288   0.10608   0.10132  -0.0587   0.0659   1.0000
  14.250   1.2997   0.11427   0.10967  -0.0622   0.0669   1.0000
  14.500   1.2731   0.12308   0.11861  -0.0666   0.0678   1.0000
  14.750   1.2523   0.13210   0.12770  -0.0711   0.0685   1.0000
<< Back to HAM-STD HS1-708 AIRFOIL (hs1708-il)

Polar data table (+)

Polar graphs


<< Back to HAM-STD HS1-708 AIRFOIL (hs1708-il)