NACA M7 AIRFOIL (m7-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M7 AIRFOIL (m7-il) Reynolds number: 200,000 Max Cl/Cd: 68.23 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m7-il-200000-n5.txt Download as CSV file: xf-m7-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M7 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5254 0.13535 0.13236 0.0432 0.8990 0.0087
-9.500 -0.5224 0.13280 0.12973 0.0426 0.8641 0.0088
-9.250 -0.5183 0.13006 0.12692 0.0413 0.8433 0.0088
-9.000 -0.5139 0.12724 0.12407 0.0398 0.8270 0.0088
-8.750 -0.5093 0.12438 0.12117 0.0380 0.8131 0.0089
-8.500 -0.5048 0.12157 0.11833 0.0360 0.8008 0.0089
-8.250 -0.5003 0.11875 0.11548 0.0336 0.7895 0.0089
-8.000 -0.4958 0.11586 0.11256 0.0310 0.7790 0.0090
-7.750 -0.4904 0.11284 0.10950 0.0285 0.7688 0.0090
-7.500 -0.4814 0.10952 0.10615 0.0258 0.7597 0.0090
-7.250 -0.4758 0.10434 0.10095 0.0250 0.7520 0.0092
-7.000 -0.4692 0.09932 0.09591 0.0261 0.7438 0.0094
-6.750 -0.4609 0.09547 0.09201 0.0249 0.7362 0.0096
-6.500 -0.4507 0.09180 0.08831 0.0230 0.7281 0.0098
-6.250 -0.4390 0.08826 0.08472 0.0208 0.7210 0.0100
-6.000 -0.4254 0.08472 0.08111 0.0184 0.7133 0.0102
-5.750 -0.4104 0.08123 0.07754 0.0160 0.7066 0.0105
-5.500 -0.3934 0.07773 0.07397 0.0134 0.6991 0.0108
-5.000 -0.3550 0.07084 0.06692 0.0085 0.6856 0.0115
-4.750 -0.3337 0.06749 0.06345 0.0062 0.6793 0.0119
-4.500 -0.3108 0.06418 0.06004 0.0040 0.6725 0.0123
-4.250 -0.2866 0.06100 0.05673 0.0020 0.6661 0.0128
-4.000 -0.2604 0.05798 0.05358 0.0001 0.6599 0.0135
-3.750 -0.2312 0.05535 0.05077 -0.0017 0.6532 0.0142
-3.500 -0.1994 0.05320 0.04841 -0.0032 0.6474 0.0148
-3.250 -0.1709 0.05075 0.04581 -0.0042 0.6407 0.0151
-3.000 -0.1437 0.04813 0.04300 -0.0048 0.6350 0.0152
-2.750 -0.1163 0.04543 0.04013 -0.0053 0.6288 0.0153
-2.500 -0.0891 0.04275 0.03727 -0.0056 0.6228 0.0153
-2.250 -0.0672 0.03775 0.03203 -0.0060 0.6182 0.0164
-2.000 -0.0455 0.03513 0.02932 -0.0063 0.6116 0.0183
-1.750 -0.0188 0.03303 0.02702 -0.0064 0.6060 0.0208
-1.500 0.0103 0.03088 0.02464 -0.0063 0.6002 0.0232
-1.250 0.0472 0.02964 0.02293 -0.0053 0.5944 0.0272
-1.000 0.0685 0.02673 0.01995 -0.0058 0.5893 0.0308
-0.750 0.0975 0.02510 0.01809 -0.0056 0.5829 0.0395
0.250 0.2188 0.01763 0.00924 -0.0026 0.5616 0.0141
0.500 0.2481 0.01645 0.00780 -0.0022 0.5559 0.0142
0.750 0.2769 0.01555 0.00666 -0.0019 0.5500 0.0160
1.000 0.3048 0.01462 0.00555 -0.0015 0.5452 0.0162
1.250 0.3322 0.01385 0.00472 -0.0012 0.5391 0.0166
1.500 0.3589 0.01331 0.00408 -0.0007 0.5336 0.0176
1.750 0.3860 0.01297 0.00365 -0.0005 0.5282 0.0222
2.000 0.4136 0.01279 0.00337 -0.0003 0.5223 0.0251
2.500 0.4674 0.01240 0.00311 0.0001 0.5046 0.1717
3.000 0.5771 0.01101 0.00327 -0.0118 0.4750 1.0000
3.500 0.6298 0.01120 0.00332 -0.0114 0.4590 1.0000
3.750 0.6561 0.01130 0.00339 -0.0112 0.4508 1.0000
4.250 0.7085 0.01153 0.00358 -0.0108 0.4335 1.0000
4.500 0.7347 0.01166 0.00370 -0.0107 0.4252 1.0000
4.750 0.7608 0.01181 0.00381 -0.0105 0.4111 1.0000
5.000 0.7868 0.01197 0.00398 -0.0104 0.3963 1.0000
5.250 0.8128 0.01215 0.00415 -0.0102 0.3813 1.0000
5.500 0.8388 0.01237 0.00435 -0.0101 0.3616 1.0000
5.750 0.8645 0.01267 0.00459 -0.0101 0.3352 1.0000
6.000 0.8894 0.01329 0.00497 -0.0103 0.2779 1.0000
6.250 0.9120 0.01486 0.00589 -0.0110 0.1617 1.0000
6.500 0.9329 0.01671 0.00718 -0.0117 0.0614 1.0000
6.750 0.9569 0.01728 0.00781 -0.0115 0.0550 1.0000
7.000 0.9809 0.01777 0.00840 -0.0113 0.0529 1.0000
7.250 1.0045 0.01829 0.00904 -0.0111 0.0514 1.0000
7.500 1.0276 0.01886 0.00973 -0.0109 0.0502 1.0000
7.750 1.0502 0.01949 0.01052 -0.0106 0.0490 1.0000
8.000 1.0721 0.02020 0.01138 -0.0103 0.0479 1.0000
8.250 1.0930 0.02101 0.01233 -0.0099 0.0469 1.0000
8.500 1.1126 0.02194 0.01341 -0.0095 0.0458 1.0000
8.750 1.1303 0.02302 0.01465 -0.0090 0.0447 1.0000
9.000 1.1454 0.02433 0.01614 -0.0083 0.0436 1.0000
9.250 1.1623 0.02530 0.01726 -0.0078 0.0411 1.0000
9.500 1.1797 0.02612 0.01818 -0.0074 0.0359 1.0000
9.750 1.1917 0.02746 0.01959 -0.0069 0.0320 1.0000
10.000 1.2205 0.02707 0.01942 -0.0068 0.0226 1.0000
10.250 1.2366 0.02797 0.02019 -0.0064 0.0104 1.0000
10.500 1.2390 0.03015 0.02238 -0.0057 0.0083 1.0000
10.750 1.2386 0.03246 0.02484 -0.0048 0.0074 1.0000
11.000 1.2386 0.03500 0.02754 -0.0043 0.0066 1.0000
11.250 1.2374 0.03778 0.03052 -0.0040 0.0060 1.0000
11.500 1.2334 0.04097 0.03386 -0.0038 0.0054 1.0000
11.750 1.2330 0.04382 0.03688 -0.0037 0.0052 1.0000
12.000 1.2322 0.04678 0.04001 -0.0036 0.0049 1.0000
12.250 1.2304 0.04990 0.04331 -0.0036 0.0047 1.0000
12.500 1.2278 0.05316 0.04673 -0.0038 0.0046 1.0000
12.750 1.2246 0.05656 0.05029 -0.0040 0.0044 1.0000
13.000 1.2206 0.06009 0.05400 -0.0043 0.0043 1.0000
13.250 1.2160 0.06379 0.05786 -0.0047 0.0042 1.0000
13.500 1.2103 0.06769 0.06192 -0.0053 0.0041 1.0000
13.750 1.2043 0.07178 0.06618 -0.0061 0.0040 1.0000
14.000 1.1977 0.07617 0.07074 -0.0072 0.0040 1.0000
14.250 1.1903 0.08081 0.07555 -0.0085 0.0039 1.0000
14.500 1.1822 0.08575 0.08065 -0.0100 0.0039 1.0000
14.750 1.1734 0.09097 0.08605 -0.0119 0.0038 1.0000
15.000 1.1636 0.09652 0.09176 -0.0140 0.0038 1.0000
15.250 1.1532 0.10238 0.09778 -0.0164 0.0037 1.0000
15.500 1.1422 0.10858 0.10414 -0.0191 0.0037 1.0000
15.750 1.1309 0.11505 0.11077 -0.0220 0.0037 1.0000
16.000 1.1192 0.12187 0.11776 -0.0253 0.0037 1.0000
16.250 1.1072 0.12908 0.12512 -0.0290 0.0037 1.0000
16.500 1.0949 0.13676 0.13298 -0.0330 0.0037 1.0000
16.750 1.0815 0.14513 0.14152 -0.0374 0.0038 1.0000
17.000 1.0660 0.15473 0.15129 -0.0425 0.0039 1.0000
17.250 1.0494 0.16541 0.16211 -0.0481 0.0041 1.0000
17.500 1.0325 0.17706 0.17385 -0.0539 0.0043 1.0000
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Polar data table (+)
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