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

USA-35B AIRFOIL (usa35b-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: USA-35B AIRFOIL (usa35b-il)
Reynolds number: 100,000
Max Cl/Cd: 53.35 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa35b-il-100000.txt
Download as CSV file: xf-usa35b-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA-35B 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3138   0.09363   0.08870  -0.0330   1.0000   0.1178
  -9.250  -0.3264   0.09204   0.08722  -0.0319   1.0000   0.1206
  -9.000  -0.3559   0.09153   0.08688  -0.0296   1.0000   0.1224
  -8.750  -0.3988   0.09097   0.08644  -0.0337   1.0000   0.1240
  -8.500  -0.4089   0.08755   0.08308  -0.0333   1.0000   0.1250
  -8.250  -0.4001   0.08485   0.08047  -0.0274   1.0000   0.1263
  -8.000  -0.3992   0.08299   0.07867  -0.0237   1.0000   0.1280
  -7.750  -0.4037   0.08128   0.07703  -0.0214   1.0000   0.1301
  -7.500  -0.3986   0.07786   0.07352  -0.0314   0.9961   0.1390
  -7.250  -0.3692   0.07256   0.06821  -0.0360   0.9890   0.1429
  -6.750  -0.2865   0.06451   0.05973  -0.0558   0.9705   0.1720
  -6.500  -0.2602   0.06004   0.05538  -0.0568   0.9616   0.1758
  -6.250  -0.2194   0.05618   0.05136  -0.0642   0.9534   0.1917
  -6.000  -0.1831   0.05300   0.04804  -0.0693   0.9432   0.2084
  -5.750  -0.1501   0.05028   0.04538  -0.0708   0.9343   0.2177
  -5.250  -0.0399   0.03320   0.02569  -0.0894   0.9179   0.1081
  -5.000   0.0037   0.03013   0.02235  -0.0928   0.9111   0.1045
  -4.750   0.0443   0.02784   0.01965  -0.0951   0.9028   0.1024
  -4.500   0.0844   0.02620   0.01766  -0.0970   0.8941   0.1040
  -4.250   0.1199   0.02499   0.01615  -0.0980   0.8839   0.1068
  -4.000   0.1594   0.02373   0.01460  -0.0994   0.8761   0.1088
  -3.750   0.1895   0.02247   0.01335  -0.0995   0.8647   0.1122
  -3.500   0.2266   0.02145   0.01230  -0.1005   0.8576   0.1199
  -3.250   0.2531   0.02073   0.01164  -0.0999   0.8451   0.1291
  -3.000   0.2825   0.01996   0.01093  -0.0995   0.8352   0.1432
  -2.750   0.3120   0.01914   0.01030  -0.0992   0.8256   0.1762
  -2.500   0.3361   0.01785   0.00979  -0.0984   0.8141   0.3158
  -2.250   0.3812   0.01559   0.00930  -0.0990   0.8071   1.0000
  -2.000   0.4056   0.01573   0.00918  -0.0981   0.7942   1.0000
  -1.750   0.4317   0.01582   0.00906  -0.0973   0.7831   1.0000
  -1.500   0.4602   0.01582   0.00882  -0.0968   0.7740   1.0000
  -1.250   0.4849   0.01604   0.00890  -0.0961   0.7623   1.0000
  -1.000   0.5122   0.01614   0.00882  -0.0956   0.7532   1.0000
  -0.750   0.5386   0.01630   0.00884  -0.0950   0.7429   1.0000
  -0.500   0.5644   0.01652   0.00895  -0.0944   0.7327   1.0000
  -0.250   0.5926   0.01660   0.00887  -0.0940   0.7241   1.0000
   0.000   0.6174   0.01690   0.00910  -0.0933   0.7130   1.0000
   0.250   0.6453   0.01703   0.00910  -0.0929   0.7044   1.0000
   0.500   0.6709   0.01728   0.00929  -0.0923   0.6937   1.0000
   0.750   0.6970   0.01756   0.00950  -0.0917   0.6836   1.0000
   1.000   0.7249   0.01771   0.00954  -0.0913   0.6744   1.0000
   1.250   0.7496   0.01806   0.00987  -0.0906   0.6630   1.0000
   1.500   0.7770   0.01829   0.01001  -0.0901   0.6535   1.0000
   1.750   0.8030   0.01858   0.01026  -0.0895   0.6425   1.0000
   2.000   0.8280   0.01895   0.01061  -0.0888   0.6310   1.0000
   2.250   0.8562   0.01916   0.01073  -0.0884   0.6214   1.0000
   2.500   0.8807   0.01953   0.01110  -0.0877   0.6088   1.0000
   2.750   0.9054   0.01983   0.01138  -0.0868   0.5955   1.0000
   3.000   0.9310   0.02002   0.01152  -0.0859   0.5819   1.0000
   3.250   0.9575   0.02014   0.01155  -0.0852   0.5687   1.0000
   3.500   0.9832   0.02033   0.01169  -0.0844   0.5556   1.0000
   3.750   1.0069   0.02064   0.01205  -0.0835   0.5420   1.0000
   4.000   1.0313   0.02094   0.01236  -0.0826   0.5291   1.0000
   4.250   1.0572   0.02116   0.01254  -0.0820   0.5172   1.0000
   4.500   1.0830   0.02136   0.01272  -0.0813   0.5051   1.0000
   4.750   1.1059   0.02169   0.01312  -0.0803   0.4915   1.0000
   5.000   1.1296   0.02201   0.01348  -0.0794   0.4786   1.0000
   5.250   1.1542   0.02227   0.01376  -0.0786   0.4660   1.0000
   5.500   1.1794   0.02247   0.01391  -0.0778   0.4534   1.0000
   5.750   1.2028   0.02273   0.01418  -0.0768   0.4395   1.0000
   6.000   1.2249   0.02306   0.01458  -0.0757   0.4247   1.0000
   6.250   1.2468   0.02343   0.01497  -0.0745   0.4099   1.0000
   6.500   1.2685   0.02380   0.01534  -0.0733   0.3943   1.0000
   6.750   1.2894   0.02417   0.01567  -0.0720   0.3777   1.0000
   7.000   1.3094   0.02459   0.01604  -0.0706   0.3606   1.0000
   7.250   1.3286   0.02512   0.01649  -0.0691   0.3436   1.0000
   7.500   1.3470   0.02575   0.01709  -0.0676   0.3274   1.0000
   7.750   1.3642   0.02645   0.01781  -0.0660   0.3121   1.0000
   8.000   1.3801   0.02716   0.01858  -0.0643   0.2975   1.0000
   8.250   1.3954   0.02793   0.01943  -0.0625   0.2841   1.0000
   8.500   1.4102   0.02871   0.02027  -0.0607   0.2716   1.0000
   8.750   1.4232   0.02944   0.02104  -0.0586   0.2592   1.0000
   9.000   1.4357   0.03018   0.02178  -0.0566   0.2475   1.0000
   9.250   1.4438   0.03099   0.02272  -0.0540   0.2356   1.0000
   9.500   1.4479   0.03189   0.02374  -0.0508   0.2237   1.0000
   9.750   1.4499   0.03294   0.02486  -0.0476   0.2116   1.0000
  10.000   1.4490   0.03422   0.02621  -0.0445   0.1985   1.0000
  10.250   1.4449   0.03587   0.02789  -0.0415   0.1839   1.0000
  10.500   1.4374   0.03802   0.03005  -0.0388   0.1677   1.0000
  10.750   1.4278   0.04068   0.03266  -0.0364   0.1512   1.0000
  11.000   1.4187   0.04362   0.03554  -0.0345   0.1364   1.0000
  11.250   1.4124   0.04658   0.03847  -0.0329   0.1243   1.0000
  11.500   1.4092   0.04941   0.04127  -0.0316   0.1152   1.0000
  11.750   1.4078   0.05208   0.04389  -0.0306   0.1081   1.0000
  12.000   1.4083   0.05480   0.04671  -0.0297   0.1017   1.0000
  12.250   1.4099   0.05730   0.04921  -0.0289   0.0967   1.0000
  12.500   1.4124   0.05991   0.05190  -0.0280   0.0923   1.0000
  12.750   1.4138   0.06262   0.05472  -0.0274   0.0883   1.0000
  13.000   1.4219   0.06456   0.05653  -0.0264   0.0847   1.0000
  13.250   1.4231   0.06754   0.05972  -0.0258   0.0820   1.0000
  13.500   1.4230   0.07060   0.06297  -0.0255   0.0793   1.0000
  13.750   1.4270   0.07312   0.06553  -0.0250   0.0766   1.0000
  14.000   1.4411   0.07480   0.06711  -0.0237   0.0738   1.0000
  14.250   1.4349   0.07880   0.07140  -0.0239   0.0725   1.0000
  14.500   1.4279   0.08298   0.07584  -0.0242   0.0712   1.0000
  14.750   1.4192   0.08746   0.08055  -0.0249   0.0700   1.0000
  15.000   1.4102   0.09204   0.08533  -0.0259   0.0688   1.0000
  15.250   1.4030   0.09642   0.08985  -0.0269   0.0676   1.0000
  15.750   1.0835   0.17958   0.17433  -0.0787   0.0830   1.0000
  16.000   1.0714   0.18992   0.18464  -0.0843   0.0835   1.0000
  16.250   1.0685   0.19735   0.19206  -0.0880   0.0838   1.0000
<< Back to USA-35B AIRFOIL (usa35b-il)

Polar data table (+)

Polar graphs


<< Back to USA-35B AIRFOIL (usa35b-il)