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

GOE 701 AIRFOIL (goe701-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 701 AIRFOIL (goe701-il)
Reynolds number: 50,000
Max Cl/Cd: 31.57 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe701-il-50000-n5.txt
Download as CSV file: xf-goe701-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 701 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.3289   0.12586   0.11869  -0.0401   1.0000   0.1339
 -10.000  -0.3382   0.12386   0.11678  -0.0395   1.0000   0.1341
  -9.750  -0.3481   0.12181   0.11482  -0.0386   1.0000   0.1342
  -9.250  -0.3624   0.11141   0.10446  -0.0383   1.0000   0.1029
  -9.000  -0.3602   0.10890   0.10199  -0.0361   1.0000   0.1019
  -8.750  -0.3687   0.10676   0.09992  -0.0344   1.0000   0.1015
  -8.500  -0.3817   0.10476   0.09801  -0.0327   1.0000   0.1014
  -8.250  -0.3981   0.10287   0.09621  -0.0309   1.0000   0.1015
  -8.000  -0.4035   0.09970   0.09308  -0.0327   0.9960   0.1016
  -7.750  -0.3988   0.09539   0.08875  -0.0374   0.9879   0.1015
  -7.500  -0.3890   0.09081   0.08411  -0.0421   0.9803   0.1011
  -7.250  -0.3790   0.08594   0.07917  -0.0475   0.9723   0.1009
  -7.000  -0.3744   0.08036   0.07349  -0.0534   0.9623   0.1019
  -6.750  -0.3689   0.07411   0.06703  -0.0597   0.9532   0.1028
  -6.500  -0.3582   0.06855   0.06127  -0.0643   0.9453   0.1034
  -6.250  -0.3440   0.06536   0.05797  -0.0656   0.9380   0.1041
  -6.000  -0.3255   0.06282   0.05533  -0.0669   0.9311   0.1056
  -5.750  -0.3046   0.06050   0.05288  -0.0684   0.9250   0.1082
  -5.500  -0.2905   0.05745   0.04963  -0.0695   0.9167   0.1107
  -5.250  -0.2667   0.05323   0.04500  -0.0729   0.9114   0.1126
  -5.000  -0.2554   0.04991   0.04129  -0.0730   0.9026   0.1141
  -4.750  -0.2317   0.04561   0.03623  -0.0754   0.8970   0.1183
  -4.500  -0.2123   0.04472   0.03534  -0.0747   0.8897   0.1208
  -4.250  -0.1870   0.04332   0.03376  -0.0752   0.8834   0.1237
  -4.000  -0.1574   0.04141   0.03150  -0.0765   0.8786   0.1273
  -3.750  -0.1401   0.03983   0.02946  -0.0755   0.8703   0.1322
  -3.500  -0.1089   0.03857   0.02797  -0.0766   0.8654   0.1371
  -3.250  -0.0885   0.03784   0.02710  -0.0758   0.8577   0.1411
  -3.000  -0.0600   0.03673   0.02563  -0.0762   0.8518   0.1474
  -2.750  -0.0254   0.03590   0.02466  -0.0776   0.8476   0.1546
  -2.500  -0.0089   0.03552   0.02413  -0.0759   0.8384   0.1605
  -2.250   0.0252   0.03473   0.02312  -0.0770   0.8336   0.1684
  -2.000   0.0462   0.03451   0.02281  -0.0761   0.8256   0.1763
  -1.750   0.0760   0.03406   0.02222  -0.0765   0.8195   0.1854
  -1.250   0.1297   0.03352   0.02153  -0.0764   0.8057   0.2045
  -1.000   0.1659   0.03310   0.02096  -0.0777   0.8010   0.2164
  -0.750   0.1903   0.03302   0.02084  -0.0772   0.7930   0.2251
  -0.500   0.2220   0.03278   0.02055  -0.0779   0.7866   0.2369
  -0.250   0.2626   0.03231   0.02002  -0.0798   0.7827   0.2509
   0.000   0.2786   0.03251   0.02024  -0.0780   0.7725   0.2605
   0.250   0.3147   0.03210   0.01988  -0.0792   0.7675   0.2785
   0.500   0.3353   0.03219   0.02005  -0.0781   0.7588   0.2968
   0.750   0.3654   0.03190   0.01991  -0.0783   0.7523   0.3235
   1.000   0.4045   0.03118   0.01947  -0.0799   0.7485   0.3750
   1.250   0.4590   0.02996   0.01994  -0.0854   0.7400   1.0000
   1.500   0.4938   0.02994   0.01965  -0.0861   0.7340   1.0000
   1.750   0.5091   0.03043   0.01999  -0.0839   0.7236   1.0000
   2.000   0.5436   0.03035   0.01973  -0.0845   0.7172   1.0000
   2.250   0.5582   0.03083   0.02012  -0.0823   0.7062   1.0000
   2.500   0.5968   0.03052   0.01968  -0.0833   0.6998   1.0000
   2.750   0.6113   0.03087   0.01995  -0.0809   0.6873   1.0000
   3.000   0.6458   0.03048   0.01947  -0.0810   0.6783   1.0000
   3.250   0.6745   0.03024   0.01915  -0.0803   0.6673   1.0000
   3.500   0.6911   0.03042   0.01929  -0.0780   0.6543   1.0000
   3.750   0.7182   0.03027   0.01909  -0.0772   0.6436   1.0000
   4.250   0.7635   0.03034   0.01909  -0.0744   0.6209   1.0000
   4.500   0.7856   0.03042   0.01917  -0.0730   0.6096   1.0000
   4.750   0.8207   0.02998   0.01870  -0.0733   0.6000   1.0000
   5.000   0.8336   0.03038   0.01911  -0.0706   0.5863   1.0000
   5.250   0.8510   0.03063   0.01937  -0.0686   0.5731   1.0000
   5.500   0.8731   0.03068   0.01941  -0.0672   0.5602   1.0000
   5.750   0.8998   0.03055   0.01925  -0.0663   0.5475   1.0000
   6.000   0.9253   0.03048   0.01919  -0.0653   0.5341   1.0000
   6.250   0.9419   0.03079   0.01949  -0.0632   0.5188   1.0000
   6.500   0.9595   0.03108   0.01978  -0.0613   0.5033   1.0000
   6.750   0.9776   0.03139   0.02008  -0.0595   0.4876   1.0000
   7.000   0.9955   0.03176   0.02045  -0.0577   0.4718   1.0000
   7.250   1.0133   0.03216   0.02082  -0.0560   0.4558   1.0000
   7.500   1.0304   0.03264   0.02128  -0.0542   0.4400   1.0000
   7.750   1.0472   0.03320   0.02182  -0.0525   0.4248   1.0000
   8.000   1.0639   0.03382   0.02241  -0.0508   0.4104   1.0000
   8.250   1.0813   0.03445   0.02301  -0.0493   0.3968   1.0000
   8.500   1.1006   0.03504   0.02355  -0.0480   0.3842   1.0000
   8.750   1.1169   0.03582   0.02434  -0.0465   0.3722   1.0000
   9.000   1.1299   0.03683   0.02539  -0.0448   0.3608   1.0000
   9.250   1.1480   0.03760   0.02614  -0.0436   0.3505   1.0000
   9.500   1.1648   0.03845   0.02701  -0.0422   0.3404   1.0000
   9.750   1.1772   0.03961   0.02826  -0.0407   0.3312   1.0000
  10.000   1.2009   0.04020   0.02882  -0.0400   0.3226   1.0000
  10.250   1.2064   0.04173   0.03050  -0.0380   0.3139   1.0000
  10.500   1.2326   0.04226   0.03099  -0.0376   0.3063   1.0000
  10.750   1.2340   0.04407   0.03302  -0.0353   0.2987   1.0000
  11.000   1.2557   0.04482   0.03378  -0.0346   0.2915   1.0000
  11.250   1.2598   0.04660   0.03574  -0.0327   0.2849   1.0000
  11.500   1.2711   0.04799   0.03725  -0.0314   0.2784   1.0000
  11.750   1.2904   0.04900   0.03831  -0.0306   0.2724   1.0000
  12.000   1.2850   0.05147   0.04104  -0.0284   0.2666   1.0000
  12.250   1.3000   0.05271   0.04238  -0.0275   0.2609   1.0000
  12.500   1.3114   0.05431   0.04410  -0.0264   0.2558   1.0000
  12.750   1.2962   0.05771   0.04776  -0.0243   0.2508   1.0000
  13.000   1.3055   0.05926   0.04941  -0.0232   0.2450   1.0000
  13.250   1.3111   0.06114   0.05139  -0.0220   0.2396   1.0000
  13.500   1.2784   0.06662   0.05716  -0.0206   0.2354   1.0000
  13.750   1.2693   0.07014   0.06084  -0.0198   0.2304   1.0000
  14.000   1.2880   0.07051   0.06125  -0.0189   0.2239   1.0000
  14.250   1.2172   0.08217   0.07319  -0.0203   0.2221   1.0000
  14.500   1.0486   0.11391   0.10495  -0.0328   0.2131   1.0000
  14.750   1.0704   0.11357   0.10476  -0.0314   0.2107   1.0000
<< Back to GOE 701 AIRFOIL (goe701-il)

Polar data table (+)

Polar graphs


<< Back to GOE 701 AIRFOIL (goe701-il)