M6 (65%) (m665-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: M6 (65%) (m665-il) Reynolds number: 500,000 Max Cl/Cd: 74.71 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m665-il-500000.txt Download as CSV file: xf-m665-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: M6 (65%)
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.5469 0.11337 0.11096 0.0015 1.0000 0.0176
-10.250 -0.5449 0.10946 0.10705 -0.0002 1.0000 0.0181
-7.750 -0.5898 0.05986 0.05741 -0.0248 1.0000 0.0204
-7.500 -0.5868 0.05677 0.05427 -0.0240 1.0000 0.0207
-7.250 -0.5809 0.05424 0.05166 -0.0231 1.0000 0.0211
-7.000 -0.5734 0.05163 0.04898 -0.0221 1.0000 0.0216
-6.750 -0.5663 0.04841 0.04565 -0.0208 1.0000 0.0221
-6.500 -0.5572 0.04552 0.04263 -0.0194 1.0000 0.0233
-6.250 -0.5485 0.04185 0.03877 -0.0174 1.0000 0.0245
-6.000 -0.5298 0.03994 0.03645 -0.0147 1.0000 0.0267
-5.750 -0.5515 0.02776 0.02354 -0.0075 1.0000 0.0190
-5.500 -0.5472 0.02352 0.01885 -0.0032 1.0000 0.0179
-5.250 -0.5374 0.02029 0.01516 0.0005 1.0000 0.0179
-5.000 -0.5229 0.01825 0.01278 0.0034 1.0000 0.0184
-4.500 -0.4672 0.01479 0.00875 0.0034 0.9960 0.0209
-4.250 -0.4305 0.01444 0.00843 0.0012 0.9918 0.0232
-4.000 -0.3938 0.01372 0.00761 -0.0006 0.9871 0.0254
-3.750 -0.3555 0.01308 0.00685 -0.0028 0.9831 0.0271
-3.500 -0.3242 0.01171 0.00544 -0.0037 0.9756 0.0305
-3.250 -0.2865 0.01110 0.00481 -0.0058 0.9705 0.0331
-3.000 -0.2522 0.01066 0.00433 -0.0071 0.9607 0.0356
-2.750 -0.2178 0.00986 0.00347 -0.0085 0.9500 0.0385
-2.500 -0.1824 0.00935 0.00294 -0.0100 0.9363 0.0431
-2.250 -0.1486 0.00903 0.00256 -0.0112 0.9183 0.0481
-2.000 -0.1190 0.00871 0.00221 -0.0114 0.8976 0.0597
-1.750 -0.0950 0.00824 0.00195 -0.0104 0.8742 0.1219
-1.500 -0.0734 0.00772 0.00178 -0.0092 0.8523 0.2431
-1.250 -0.0595 0.00672 0.00160 -0.0066 0.8308 0.4950
-1.000 -0.0516 0.00546 0.00159 -0.0019 0.8121 0.8483
-0.750 0.0297 0.00570 0.00189 -0.0130 0.7987 0.9352
-0.500 0.0694 0.00606 0.00213 -0.0150 0.7819 0.9534
-0.250 0.1139 0.00640 0.00234 -0.0182 0.7655 0.9643
0.000 0.1664 0.00675 0.00258 -0.0232 0.7490 0.9756
0.250 0.2269 0.00698 0.00270 -0.0301 0.7328 0.9865
0.500 0.2790 0.00701 0.00265 -0.0353 0.7162 0.9942
0.750 0.3232 0.00689 0.00244 -0.0389 0.6982 0.9987
1.000 0.3518 0.00689 0.00235 -0.0391 0.6796 1.0000
1.250 0.3743 0.00691 0.00230 -0.0381 0.6623 1.0000
1.500 0.3971 0.00694 0.00228 -0.0371 0.6469 1.0000
1.750 0.4200 0.00697 0.00227 -0.0361 0.6319 1.0000
2.000 0.4429 0.00701 0.00227 -0.0351 0.6156 1.0000
2.250 0.4657 0.00706 0.00227 -0.0341 0.5970 1.0000
2.500 0.4884 0.00713 0.00229 -0.0331 0.5782 1.0000
2.750 0.5112 0.00720 0.00231 -0.0321 0.5554 1.0000
3.000 0.5334 0.00731 0.00234 -0.0310 0.5238 1.0000
3.250 0.5547 0.00753 0.00236 -0.0297 0.4760 1.0000
3.500 0.5766 0.00775 0.00243 -0.0286 0.4325 1.0000
3.750 0.5984 0.00801 0.00254 -0.0275 0.3870 1.0000
4.000 0.6193 0.00840 0.00270 -0.0263 0.3241 1.0000
4.250 0.6385 0.00905 0.00295 -0.0249 0.2329 1.0000
4.500 0.6557 0.01001 0.00335 -0.0234 0.1172 1.0000
4.750 0.6734 0.01097 0.00391 -0.0217 0.0429 1.0000
5.000 0.6941 0.01152 0.00444 -0.0203 0.0317 1.0000
5.250 0.7155 0.01193 0.00491 -0.0191 0.0285 1.0000
5.500 0.7347 0.01264 0.00567 -0.0175 0.0256 1.0000
5.750 0.7537 0.01336 0.00647 -0.0158 0.0246 1.0000
6.000 0.7736 0.01392 0.00710 -0.0143 0.0240 1.0000
6.250 0.7928 0.01459 0.00784 -0.0127 0.0233 1.0000
6.500 0.8119 0.01527 0.00859 -0.0111 0.0226 1.0000
6.750 0.8309 0.01597 0.00935 -0.0095 0.0219 1.0000
7.000 0.8493 0.01682 0.01026 -0.0078 0.0215 1.0000
7.250 0.8678 0.01772 0.01124 -0.0062 0.0211 1.0000
7.500 0.8863 0.01872 0.01233 -0.0045 0.0207 1.0000
7.750 0.9048 0.01983 0.01354 -0.0029 0.0204 1.0000
8.000 0.9232 0.02114 0.01496 -0.0013 0.0202 1.0000
8.250 0.9413 0.02246 0.01639 0.0002 0.0198 1.0000
8.500 0.9583 0.02393 0.01793 0.0017 0.0190 1.0000
8.750 0.9727 0.02659 0.02079 0.0036 0.0186 1.0000
9.000 0.9857 0.02920 0.02367 0.0056 0.0185 1.0000
9.250 0.9943 0.03240 0.02718 0.0081 0.0184 1.0000
9.500 0.9957 0.03656 0.03172 0.0111 0.0182 1.0000
9.750 1.0122 0.03516 0.03051 0.0133 0.0170 1.0000
10.000 1.0129 0.03821 0.03389 0.0164 0.0168 1.0000
10.250 1.0067 0.04184 0.03783 0.0199 0.0170 1.0000
10.500 0.9970 0.04525 0.04153 0.0234 0.0170 1.0000
10.750 0.9775 0.04904 0.04554 0.0275 0.0173 1.0000
11.000 0.9546 0.05243 0.04908 0.0312 0.0176 1.0000
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Polar data table (+)
Polar graphs
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