M6 (65%) (m665-il) Xfoil prediction polar at RE=200,000 Ncrit=5
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Airfoil: M6 (65%) (m665-il) Reynolds number: 200,000 Max Cl/Cd: 55.39 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m665-il-200000-n5.txt Download as CSV file: xf-m665-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5551 0.08771 0.08408 -0.0105 1.0000 0.0165
-8.750 -0.5593 0.08312 0.07953 -0.0133 1.0000 0.0167
-8.500 -0.5680 0.07817 0.07463 -0.0170 1.0000 0.0161
-8.250 -0.5782 0.07371 0.07017 -0.0189 1.0000 0.0165
-8.000 -0.5844 0.06878 0.06521 -0.0206 1.0000 0.0161
-7.750 -0.5877 0.06379 0.06015 -0.0218 1.0000 0.0161
-7.500 -0.5875 0.05911 0.05535 -0.0223 1.0000 0.0164
-7.250 -0.5863 0.05415 0.05023 -0.0220 1.0000 0.0165
-7.000 -0.5843 0.04898 0.04483 -0.0210 1.0000 0.0165
-6.750 -0.5819 0.04358 0.03913 -0.0192 1.0000 0.0161
-6.500 -0.5788 0.03817 0.03334 -0.0165 1.0000 0.0158
-6.250 -0.5745 0.03310 0.02777 -0.0133 1.0000 0.0156
-6.000 -0.5660 0.02921 0.02340 -0.0102 1.0000 0.0157
-5.750 -0.5537 0.02621 0.01995 -0.0074 1.0000 0.0159
-5.500 -0.5385 0.02380 0.01713 -0.0050 1.0000 0.0162
-5.250 -0.5212 0.02180 0.01474 -0.0029 1.0000 0.0169
-5.000 -0.5024 0.02004 0.01263 -0.0009 1.0000 0.0178
-4.750 -0.4825 0.01865 0.01094 0.0008 1.0000 0.0189
-4.500 -0.4635 0.01769 0.00995 0.0024 1.0000 0.0201
-4.250 -0.4333 0.01693 0.00910 0.0017 0.9958 0.0219
-4.000 -0.3982 0.01631 0.00832 0.0000 0.9884 0.0246
-3.750 -0.3646 0.01550 0.00746 -0.0015 0.9801 0.0282
-3.500 -0.3309 0.01486 0.00679 -0.0029 0.9710 0.0309
-3.250 -0.2968 0.01412 0.00598 -0.0042 0.9613 0.0328
-3.000 -0.2625 0.01346 0.00522 -0.0055 0.9504 0.0347
-2.750 -0.2292 0.01284 0.00454 -0.0067 0.9375 0.0374
-2.500 -0.1952 0.01233 0.00399 -0.0080 0.9235 0.0424
-2.250 -0.1609 0.01195 0.00355 -0.0093 0.9084 0.0498
-2.000 -0.1287 0.01157 0.00321 -0.0101 0.8912 0.0659
-1.750 -0.0998 0.01112 0.00294 -0.0104 0.8723 0.1168
-1.500 -0.0736 0.01061 0.00272 -0.0102 0.8534 0.2052
-1.250 -0.0531 0.00982 0.00250 -0.0089 0.8349 0.3755
-1.000 0.0303 0.00847 0.00299 -0.0194 0.8256 0.9224
-0.750 0.0841 0.00894 0.00331 -0.0240 0.8086 0.9593
-0.500 0.1382 0.00921 0.00341 -0.0292 0.7909 0.9788
-0.250 0.1943 0.00924 0.00329 -0.0351 0.7724 0.9922
0.000 0.2345 0.00918 0.00310 -0.0378 0.7540 0.9983
0.250 0.2626 0.00919 0.00299 -0.0379 0.7363 1.0000
0.500 0.2853 0.00922 0.00293 -0.0369 0.7201 1.0000
0.750 0.3080 0.00925 0.00289 -0.0358 0.7043 1.0000
1.000 0.3310 0.00930 0.00286 -0.0349 0.6892 1.0000
1.250 0.3541 0.00935 0.00285 -0.0339 0.6745 1.0000
1.500 0.3772 0.00940 0.00287 -0.0329 0.6595 1.0000
1.750 0.4002 0.00947 0.00288 -0.0319 0.6425 1.0000
2.000 0.4231 0.00954 0.00290 -0.0309 0.6236 1.0000
2.250 0.4462 0.00962 0.00295 -0.0298 0.6053 1.0000
2.500 0.4694 0.00971 0.00301 -0.0289 0.5885 1.0000
2.750 0.4925 0.00981 0.00309 -0.0279 0.5704 1.0000
3.000 0.5157 0.00991 0.00318 -0.0269 0.5500 1.0000
3.250 0.5384 0.01005 0.00327 -0.0258 0.5251 1.0000
3.500 0.5607 0.01022 0.00339 -0.0247 0.4927 1.0000
3.750 0.5816 0.01050 0.00348 -0.0233 0.4417 1.0000
4.000 0.6012 0.01094 0.00362 -0.0218 0.3708 1.0000
4.250 0.6194 0.01163 0.00387 -0.0203 0.2773 1.0000
4.500 0.6355 0.01270 0.00433 -0.0186 0.1583 1.0000
4.750 0.6527 0.01370 0.00490 -0.0172 0.0769 1.0000
5.000 0.6725 0.01437 0.00544 -0.0158 0.0480 1.0000
5.250 0.6925 0.01500 0.00603 -0.0145 0.0357 1.0000
5.500 0.7131 0.01555 0.00668 -0.0131 0.0311 1.0000
5.750 0.7332 0.01614 0.00734 -0.0117 0.0283 1.0000
6.000 0.7520 0.01688 0.00814 -0.0102 0.0262 1.0000
6.250 0.7699 0.01770 0.00903 -0.0085 0.0245 1.0000
6.500 0.7893 0.01832 0.00976 -0.0071 0.0231 1.0000
6.750 0.8074 0.01912 0.01064 -0.0055 0.0221 1.0000
7.000 0.8252 0.01999 0.01160 -0.0038 0.0215 1.0000
7.250 0.8428 0.02093 0.01262 -0.0021 0.0208 1.0000
7.500 0.8604 0.02193 0.01371 -0.0005 0.0203 1.0000
7.750 0.8780 0.02301 0.01490 0.0012 0.0198 1.0000
8.000 0.8955 0.02417 0.01616 0.0027 0.0193 1.0000
8.250 0.9128 0.02547 0.01757 0.0043 0.0189 1.0000
8.500 0.9297 0.02692 0.01916 0.0059 0.0186 1.0000
8.750 0.9459 0.02858 0.02099 0.0074 0.0183 1.0000
9.000 0.9609 0.03047 0.02310 0.0091 0.0181 1.0000
9.250 0.9738 0.03273 0.02561 0.0109 0.0179 1.0000
9.500 0.9847 0.03509 0.02825 0.0129 0.0178 1.0000
9.750 0.9919 0.03778 0.03125 0.0152 0.0177 1.0000
10.000 0.9940 0.04059 0.03439 0.0178 0.0174 1.0000
10.250 0.9972 0.04228 0.03641 0.0205 0.0169 1.0000
10.500 0.9962 0.04435 0.03880 0.0233 0.0163 1.0000
10.750 0.9898 0.04668 0.04140 0.0264 0.0157 1.0000
11.000 0.9756 0.04934 0.04429 0.0299 0.0156 1.0000
11.250 0.9628 0.05203 0.04719 0.0319 0.0151 1.0000
11.500 0.9461 0.05572 0.05109 0.0325 0.0150 1.0000
11.750 0.9260 0.06048 0.05605 0.0314 0.0152 1.0000
12.000 0.9056 0.06598 0.06174 0.0288 0.0152 1.0000
12.250 0.8829 0.07281 0.06873 0.0246 0.0154 1.0000
12.500 0.8567 0.08151 0.07758 0.0186 0.0157 1.0000
12.750 0.8307 0.09151 0.08766 0.0119 0.0165 1.0000
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Polar data table (+)
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