M6 (65%) (m665-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: M6 (65%) (m665-il) Reynolds number: 1,000,000 Max Cl/Cd: 78.56 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m665-il-1000000.txt Download as CSV file: xf-m665-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4625 0.08014 0.07859 -0.0120 1.0000 0.0126
-9.000 -0.4679 0.07519 0.07365 -0.0138 1.0000 0.0128
-8.750 -0.4767 0.06952 0.06801 -0.0163 1.0000 0.0138
-8.500 -0.4865 0.06438 0.06288 -0.0188 1.0000 0.0137
-8.250 -0.5049 0.05899 0.05751 -0.0230 1.0000 0.0136
-8.000 -0.5237 0.05480 0.05330 -0.0235 1.0000 0.0132
-7.750 -0.5364 0.04924 0.04768 -0.0245 1.0000 0.0135
-7.250 -0.6359 0.03775 0.03533 -0.0185 1.0000 0.0100
-7.000 -0.6934 0.01921 0.01514 -0.0054 1.0000 0.0094
-6.750 -0.6803 0.01713 0.01272 -0.0026 1.0000 0.0096
-6.500 -0.6625 0.01608 0.01150 -0.0006 1.0000 0.0099
-6.250 -0.6453 0.01492 0.01015 0.0015 1.0000 0.0102
-6.000 -0.6257 0.01437 0.00949 0.0032 1.0000 0.0106
-5.750 -0.5959 0.01252 0.00738 0.0025 0.9980 0.0114
-5.500 -0.5615 0.01195 0.00676 0.0009 0.9955 0.0121
-5.250 -0.5276 0.01159 0.00637 -0.0005 0.9926 0.0131
-5.000 -0.4942 0.01115 0.00587 -0.0018 0.9887 0.0141
-4.750 -0.4607 0.01036 0.00499 -0.0031 0.9848 0.0155
-4.500 -0.4282 0.01005 0.00469 -0.0042 0.9791 0.0170
-4.250 -0.3935 0.00981 0.00445 -0.0057 0.9730 0.0186
-4.000 -0.3570 0.00967 0.00429 -0.0076 0.9660 0.0197
-3.750 -0.3214 0.00892 0.00350 -0.0094 0.9559 0.0221
-3.500 -0.2854 0.00866 0.00321 -0.0112 0.9394 0.0239
-3.250 -0.2533 0.00848 0.00297 -0.0121 0.9153 0.0258
-3.000 -0.2261 0.00839 0.00278 -0.0118 0.8891 0.0267
-2.750 -0.2032 0.00798 0.00221 -0.0106 0.8635 0.0298
-2.500 -0.1789 0.00782 0.00196 -0.0096 0.8401 0.0326
-2.250 -0.1541 0.00772 0.00177 -0.0089 0.8183 0.0355
-2.000 -0.1295 0.00759 0.00155 -0.0080 0.7978 0.0399
-1.750 -0.1046 0.00748 0.00139 -0.0073 0.7798 0.0466
-1.500 -0.0803 0.00727 0.00126 -0.0065 0.7628 0.0808
-1.250 -0.0567 0.00698 0.00116 -0.0057 0.7466 0.1491
-1.000 -0.0335 0.00669 0.00107 -0.0048 0.7313 0.2307
-0.750 -0.0126 0.00619 0.00098 -0.0035 0.7171 0.3735
-0.500 -0.0031 0.00509 0.00085 0.0002 0.7040 0.6769
-0.250 0.0105 0.00443 0.00087 0.0037 0.6918 0.8779
0.000 0.0673 0.00448 0.00100 -0.0026 0.6788 0.9341
0.250 0.1132 0.00468 0.00114 -0.0065 0.6655 0.9524
0.500 0.1448 0.00485 0.00125 -0.0071 0.6516 0.9608
0.750 0.1828 0.00509 0.00141 -0.0091 0.6361 0.9680
1.000 0.2248 0.00535 0.00159 -0.0121 0.6183 0.9741
1.250 0.2736 0.00563 0.00180 -0.0166 0.6025 0.9794
1.500 0.3087 0.00581 0.00194 -0.0181 0.5883 0.9844
1.750 0.3478 0.00588 0.00194 -0.0206 0.5692 0.9859
2.000 0.3811 0.00593 0.00194 -0.0218 0.5524 0.9872
2.250 0.4133 0.00600 0.00194 -0.0229 0.5320 0.9887
2.500 0.4443 0.00612 0.00195 -0.0237 0.4995 0.9904
2.750 0.4741 0.00630 0.00197 -0.0243 0.4581 0.9924
3.000 0.5028 0.00653 0.00204 -0.0246 0.4146 0.9942
3.250 0.5334 0.00679 0.00208 -0.0255 0.3577 0.9954
3.500 0.5631 0.00717 0.00217 -0.0263 0.2853 0.9967
3.750 0.5923 0.00764 0.00232 -0.0270 0.2089 0.9981
4.000 0.6209 0.00820 0.00254 -0.0276 0.1311 0.9994
4.250 0.6457 0.00878 0.00282 -0.0274 0.0659 1.0000
4.500 0.6666 0.00923 0.00309 -0.0262 0.0313 1.0000
4.750 0.6887 0.00954 0.00338 -0.0251 0.0232 1.0000
5.000 0.7109 0.00982 0.00366 -0.0240 0.0203 1.0000
5.250 0.7321 0.01027 0.00419 -0.0227 0.0182 1.0000
5.500 0.7544 0.01053 0.00447 -0.0216 0.0175 1.0000
5.750 0.7761 0.01087 0.00486 -0.0205 0.0171 1.0000
6.000 0.7978 0.01120 0.00523 -0.0193 0.0164 1.0000
6.250 0.8190 0.01160 0.00569 -0.0181 0.0158 1.0000
6.500 0.8396 0.01207 0.00620 -0.0167 0.0155 1.0000
6.750 0.8598 0.01257 0.00676 -0.0154 0.0153 1.0000
7.000 0.8795 0.01312 0.00737 -0.0139 0.0150 1.0000
7.250 0.8990 0.01368 0.00799 -0.0124 0.0147 1.0000
7.500 0.9177 0.01433 0.00869 -0.0108 0.0144 1.0000
7.750 0.9356 0.01508 0.00950 -0.0091 0.0141 1.0000
8.000 0.9527 0.01597 0.01045 -0.0073 0.0140 1.0000
8.250 0.9653 0.01764 0.01223 -0.0048 0.0134 1.0000
8.500 0.9811 0.01896 0.01369 -0.0029 0.0130 1.0000
8.750 1.0018 0.01919 0.01399 -0.0017 0.0127 1.0000
9.000 1.0197 0.02001 0.01491 -0.0001 0.0124 1.0000
9.250 1.0374 0.02076 0.01575 0.0015 0.0119 1.0000
9.500 1.0537 0.02175 0.01685 0.0032 0.0116 1.0000
9.750 1.0701 0.02256 0.01775 0.0049 0.0112 1.0000
10.000 1.0855 0.02350 0.01879 0.0067 0.0109 1.0000
10.250 1.0994 0.02454 0.01995 0.0086 0.0107 1.0000
10.500 1.1142 0.02519 0.02066 0.0103 0.0103 1.0000
10.750 1.1265 0.02615 0.02169 0.0123 0.0100 1.0000
11.000 1.1243 0.02930 0.02508 0.0157 0.0095 1.0000
11.250 1.1263 0.03108 0.02710 0.0188 0.0092 1.0000
11.500 1.1340 0.03202 0.02817 0.0213 0.0090 1.0000
11.750 1.1339 0.03311 0.02939 0.0250 0.0088 1.0000
12.000 1.1364 0.03386 0.03024 0.0280 0.0084 1.0000
12.250 1.1365 0.03514 0.03163 0.0305 0.0082 1.0000
12.500 1.1404 0.03618 0.03275 0.0322 0.0079 1.0000
12.750 1.1324 0.03882 0.03556 0.0337 0.0078 1.0000
13.000 1.1256 0.04160 0.03849 0.0343 0.0077 1.0000
13.250 1.1163 0.04503 0.04207 0.0340 0.0076 1.0000
13.500 1.1184 0.04724 0.04436 0.0333 0.0074 1.0000
13.750 1.1121 0.05085 0.04809 0.0319 0.0073 1.0000
14.000 1.0869 0.05772 0.05518 0.0287 0.0074 1.0000
14.250 1.0759 0.06287 0.06045 0.0257 0.0073 1.0000
14.750 0.7583 0.10206 0.10030 0.0107 0.0095 1.0000
15.000 0.7461 0.10897 0.10728 0.0074 0.0096 1.0000
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Polar data table (+)
Polar graphs
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