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

GOE 241 (MVA PR.1) AIRFOIL (goe241-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 241 (MVA PR.1) AIRFOIL (goe241-il)
Reynolds number: 100,000
Max Cl/Cd: 42.35 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe241-il-100000.txt
Download as CSV file: xf-goe241-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 241 (MVA PR.1) 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250   0.1183   0.10932   0.10453  -0.1105   0.9047   0.1288
  -9.000   0.1495   0.10555   0.10072  -0.1144   0.9019   0.1357
  -8.750   0.1014   0.10764   0.10290  -0.1175   0.8861   0.1408
  -8.500   0.1645   0.10028   0.09548  -0.1176   0.8861   0.1446
  -8.250   0.1923   0.09685   0.09200  -0.1210   0.8831   0.1524
  -8.000   0.1509   0.09751   0.09276  -0.1218   0.8671   0.1575
  -7.750   0.2096   0.09196   0.08714  -0.1221   0.8663   0.1626
  -7.500   0.2087   0.09097   0.08618  -0.1206   0.8542   0.1692
  -7.250   0.2167   0.08790   0.08309  -0.1230   0.8485   0.1765
  -7.000   0.2318   0.08628   0.08148  -0.1211   0.8383   0.1836
  -6.750   0.2280   0.08419   0.07940  -0.1224   0.8303   0.1928
  -6.500   0.2502   0.08219   0.07739  -0.1209   0.8218   0.2008
  -6.250   0.2367   0.08096   0.07619  -0.1207   0.8120   0.2096
  -6.000   0.2681   0.07821   0.07339  -0.1203   0.8060   0.2178
  -5.750   0.2537   0.07741   0.07264  -0.1199   0.7939   0.2270
  -5.500   0.2385   0.07829   0.07356  -0.1237   0.7796   0.2421
  -5.250   0.2656   0.05137   0.04607  -0.1689   0.7730   0.1559
  -5.000   0.3600   0.03846   0.03109  -0.1975   0.7663   0.1114
  -4.750   0.3957   0.03609   0.02851  -0.1997   0.7581   0.1105
  -4.500   0.4442   0.03348   0.02559  -0.2034   0.7542   0.1091
  -4.250   0.4708   0.03256   0.02448  -0.2036   0.7438   0.1085
  -4.000   0.5103   0.03110   0.02278  -0.2053   0.7376   0.1086
  -3.750   0.5553   0.02972   0.02110  -0.2076   0.7336   0.1108
  -3.500   0.5733   0.02944   0.02086  -0.2062   0.7216   0.1126
  -3.250   0.6111   0.02816   0.01960  -0.2071   0.7166   0.1154
  -3.000   0.6333   0.02800   0.01951  -0.2060   0.7067   0.1177
  -2.750   0.6668   0.02735   0.01884  -0.2063   0.7000   0.1217
  -2.500   0.7076   0.02644   0.01788  -0.2078   0.6957   0.1293
  -2.250   0.7271   0.02676   0.01828  -0.2070   0.6844   0.1365
  -2.000   0.7705   0.02591   0.01736  -0.2098   0.6794   0.1530
  -1.750   0.8048   0.02549   0.01781  -0.2123   0.6712   0.3375
  -1.500   0.8273   0.02665   0.01903  -0.2098   0.6643   0.4178
  -1.250   0.8526   0.02770   0.01999  -0.2071   0.6599   0.4538
  -1.000   0.8636   0.02908   0.02143  -0.2037   0.6510   0.4719
  -0.750   0.8810   0.02992   0.02231  -0.2000   0.6452   0.4868
  -0.500   0.9083   0.03025   0.02254  -0.1982   0.6413   0.5036
  -0.250   0.9189   0.03136   0.02372  -0.1956   0.6334   0.5161
   0.000   0.9430   0.03178   0.02408  -0.1946   0.6278   0.5328
   0.250   0.9725   0.03178   0.02396  -0.1937   0.6237   0.5514
   0.500   0.9843   0.03253   0.02479  -0.1909   0.6170   0.5619
   0.750   1.0123   0.03263   0.02480  -0.1916   0.6103   0.5706
   1.000   1.0515   0.03202   0.02399  -0.1934   0.6057   0.5762
   1.250   1.0768   0.03243   0.02433  -0.1940   0.5992   0.5825
   1.500   1.1000   0.03276   0.02464  -0.1939   0.5927   0.5887
   1.750   1.1356   0.03243   0.02417  -0.1951   0.5883   0.5960
   2.000   1.1723   0.03242   0.02398  -0.1973   0.5834   0.6039
   2.250   1.1818   0.03325   0.02494  -0.1949   0.5760   0.6084
   2.500   1.2181   0.03293   0.02447  -0.1964   0.5707   0.6170
   2.750   1.2593   0.03242   0.02376  -0.1985   0.5662   0.6256
   3.000   1.2629   0.03363   0.02514  -0.1955   0.5585   0.6316
   3.250   1.2955   0.03362   0.02503  -0.1967   0.5533   0.6397
   3.500   1.3339   0.03330   0.02458  -0.1983   0.5496   0.6481
   3.750   1.3466   0.03444   0.02579  -0.1968   0.5439   0.6559
   4.000   1.3588   0.03524   0.02670  -0.1949   0.5381   0.6623
   4.250   1.3942   0.03515   0.02650  -0.1963   0.5336   0.6719
   4.500   1.4388   0.03465   0.02583  -0.1988   0.5300   0.6813
   4.750   1.4280   0.03666   0.02811  -0.1939   0.5229   0.6870
   5.000   1.4511   0.03702   0.02848  -0.1935   0.5174   0.6955
   5.250   1.4964   0.03640   0.02769  -0.1961   0.5131   0.7066
   5.500   1.5029   0.03761   0.02901  -0.1934   0.5069   0.7135
   5.750   1.5101   0.03859   0.03010  -0.1908   0.5001   0.7214
   6.000   1.5559   0.03781   0.02915  -0.1933   0.4952   0.7332
   6.250   1.5693   0.03865   0.03003  -0.1916   0.4890   0.7423
   6.500   1.5675   0.03993   0.03144  -0.1877   0.4817   0.7496
   6.750   1.6178   0.03896   0.03031  -0.1908   0.4765   0.7640
   7.000   1.6169   0.04026   0.03175  -0.1870   0.4696   0.7728
   7.250   1.6155   0.04134   0.03292  -0.1832   0.4623   0.7824
   7.500   1.6830   0.03974   0.03112  -0.1885   0.4570   0.8027
   7.750   1.6393   0.04266   0.03436  -0.1792   0.4485   0.8098
   8.000   1.6682   0.04232   0.03403  -0.1792   0.4418   0.8331
   8.250   1.7049   0.04161   0.03331  -0.1802   0.4356   1.0000
   8.500   1.6643   0.04490   0.03683  -0.1724   0.4263   1.0000
   8.750   1.7408   0.04281   0.03441  -0.1783   0.4207   1.0000
   9.000   1.6774   0.04767   0.03961  -0.1685   0.4111   1.0000
   9.250   1.7179   0.04703   0.03882  -0.1700   0.4051   1.0000
   9.500   1.7330   0.04810   0.03984  -0.1690   0.3986   1.0000
   9.750   1.6872   0.05292   0.04490  -0.1626   0.3901   1.0000
  10.000   1.7530   0.05049   0.04222  -0.1659   0.3860   1.0000
  10.250   1.6147   0.06377   0.05605  -0.1546   0.3746   1.0000
  10.500   1.6566   0.06237   0.05455  -0.1552   0.3712   1.0000
  10.750   1.7236   0.05908   0.05104  -0.1575   0.3689   1.0000
  11.500   0.9951   0.18607   0.17996  -0.1780   0.3810   1.0000
  11.750   0.9981   0.18840   0.18227  -0.1797   0.3754   1.0000
<< Back to GOE 241 (MVA PR.1) AIRFOIL (goe241-il)

Polar data table (+)

Polar graphs


<< Back to GOE 241 (MVA PR.1) AIRFOIL (goe241-il)