NACA M5 AIRFOIL (m5-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M5 AIRFOIL (m5-il) Reynolds number: 500,000 Max Cl/Cd: 76.46 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m5-il-500000.txt Download as CSV file: xf-m5-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M5 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6283 0.10560 0.10351 0.0263 1.0000 0.0237
-9.000 -0.6270 0.10142 0.09934 0.0240 1.0000 0.0244
-8.750 -0.6272 0.09682 0.09476 0.0210 1.0000 0.0254
-6.750 -0.6022 0.05202 0.04939 -0.0054 1.0000 0.0296
-6.500 -0.6055 0.03549 0.03202 -0.0048 1.0000 0.0262
-6.250 -0.6003 0.02531 0.02076 -0.0025 1.0000 0.0285
-6.000 -0.5724 0.02627 0.02191 -0.0031 1.0000 0.0297
-5.750 -0.5455 0.02647 0.02215 -0.0034 1.0000 0.0311
-5.500 -0.5229 0.02326 0.01851 -0.0026 1.0000 0.0327
-5.250 -0.4972 0.02143 0.01635 -0.0022 1.0000 0.0344
-5.000 -0.4700 0.02043 0.01506 -0.0021 1.0000 0.0359
-4.750 -0.4415 0.01720 0.01149 -0.0028 0.9712 0.0373
-4.500 -0.4160 0.01670 0.01090 -0.0023 0.9411 0.0388
-4.250 -0.3941 0.01621 0.01028 -0.0008 0.9147 0.0401
-4.000 -0.3708 0.01541 0.00930 0.0005 0.8914 0.0410
-3.750 -0.3463 0.01466 0.00837 0.0015 0.8709 0.0420
-3.500 -0.3207 0.01408 0.00762 0.0022 0.8517 0.0433
-3.250 -0.2944 0.01361 0.00701 0.0027 0.8338 0.0445
-3.000 -0.2677 0.01310 0.00637 0.0032 0.8171 0.0451
-2.750 -0.2407 0.01270 0.00585 0.0035 0.8014 0.0457
-2.500 -0.2144 0.01188 0.00491 0.0039 0.7868 0.0467
-2.250 -0.1882 0.01114 0.00410 0.0044 0.7730 0.0487
-2.000 -0.1611 0.01084 0.00375 0.0046 0.7596 0.0510
-1.750 -0.1337 0.01059 0.00345 0.0047 0.7468 0.0535
-1.500 -0.1062 0.01036 0.00314 0.0049 0.7346 0.0558
-1.250 -0.0784 0.01019 0.00290 0.0050 0.7225 0.0579
-1.000 -0.0511 0.00982 0.00252 0.0052 0.7107 0.0637
-0.750 -0.0231 0.00967 0.00232 0.0052 0.6997 0.0706
-0.500 0.0047 0.00945 0.00213 0.0053 0.6893 0.0880
-0.250 0.0317 0.00904 0.00202 0.0053 0.6788 0.1700
0.000 0.0446 0.00661 0.00184 0.0077 0.6694 0.8140
0.250 0.0717 0.00657 0.00207 0.0091 0.6600 0.9600
0.500 0.1203 0.00684 0.00226 0.0049 0.6492 0.9838
0.750 0.1904 0.00721 0.00253 -0.0042 0.6372 0.9966
1.000 0.2306 0.00732 0.00257 -0.0070 0.6268 1.0000
1.250 0.2578 0.00734 0.00250 -0.0069 0.6172 1.0000
1.500 0.2850 0.00733 0.00246 -0.0068 0.6070 1.0000
1.750 0.3123 0.00734 0.00243 -0.0067 0.5971 1.0000
2.000 0.3395 0.00738 0.00240 -0.0066 0.5875 1.0000
2.250 0.3666 0.00739 0.00240 -0.0065 0.5770 1.0000
2.500 0.3934 0.00742 0.00241 -0.0064 0.5666 1.0000
2.750 0.4202 0.00746 0.00242 -0.0062 0.5561 1.0000
3.000 0.4468 0.00752 0.00243 -0.0060 0.5433 1.0000
3.250 0.4733 0.00756 0.00246 -0.0057 0.5279 1.0000
3.500 0.4996 0.00762 0.00248 -0.0054 0.5098 1.0000
3.750 0.5259 0.00770 0.00252 -0.0052 0.4903 1.0000
4.000 0.5520 0.00781 0.00258 -0.0049 0.4729 1.0000
4.250 0.5781 0.00792 0.00267 -0.0046 0.4529 1.0000
4.500 0.6040 0.00807 0.00276 -0.0043 0.4296 1.0000
4.750 0.6298 0.00829 0.00288 -0.0040 0.3962 1.0000
5.000 0.6553 0.00857 0.00304 -0.0038 0.3584 1.0000
5.250 0.6805 0.00893 0.00325 -0.0035 0.3151 1.0000
5.500 0.7052 0.00949 0.00354 -0.0033 0.2518 1.0000
5.750 0.7278 0.01070 0.00415 -0.0033 0.1321 1.0000
6.000 0.7502 0.01189 0.00490 -0.0030 0.0544 1.0000
6.250 0.7740 0.01260 0.00557 -0.0025 0.0397 1.0000
6.500 0.7982 0.01311 0.00609 -0.0021 0.0331 1.0000
6.750 0.8219 0.01375 0.00679 -0.0015 0.0286 1.0000
7.000 0.8457 0.01428 0.00737 -0.0010 0.0258 1.0000
7.250 0.8663 0.01554 0.00870 -0.0002 0.0230 1.0000
7.500 0.8903 0.01599 0.00921 0.0003 0.0219 1.0000
7.750 0.9136 0.01657 0.00983 0.0008 0.0202 1.0000
8.000 0.9362 0.01727 0.01057 0.0014 0.0190 1.0000
8.250 0.9563 0.01839 0.01175 0.0022 0.0179 1.0000
8.500 0.9732 0.02017 0.01363 0.0034 0.0171 1.0000
8.750 0.9947 0.02107 0.01464 0.0042 0.0166 1.0000
9.000 1.0154 0.02217 0.01584 0.0050 0.0159 1.0000
9.250 1.0361 0.02318 0.01695 0.0058 0.0151 1.0000
9.500 1.0566 0.02409 0.01794 0.0064 0.0144 1.0000
9.750 1.0758 0.02518 0.01910 0.0072 0.0138 1.0000
10.000 1.0929 0.02684 0.02084 0.0081 0.0133 1.0000
10.250 1.1059 0.02986 0.02406 0.0095 0.0129 1.0000
10.500 1.1159 0.03315 0.02767 0.0110 0.0127 1.0000
10.750 1.1274 0.03505 0.02983 0.0123 0.0125 1.0000
11.000 1.1341 0.03757 0.03265 0.0140 0.0124 1.0000
11.250 1.1350 0.04058 0.03596 0.0159 0.0123 1.0000
11.500 1.1278 0.04378 0.03943 0.0183 0.0123 1.0000
11.750 1.1144 0.04719 0.04310 0.0204 0.0123 1.0000
12.000 1.0990 0.05126 0.04739 0.0211 0.0124 1.0000
12.250 1.0816 0.05601 0.05237 0.0204 0.0125 1.0000
12.500 1.0615 0.06168 0.05822 0.0187 0.0126 1.0000
12.750 1.0528 0.06646 0.06310 0.0175 0.0129 1.0000
13.000 1.0406 0.07134 0.06813 0.0148 0.0129 1.0000
13.250 1.0269 0.07686 0.07382 0.0114 0.0130 1.0000
13.500 0.6747 0.12411 0.12189 -0.0045 0.0254 1.0000
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Polar data table (+)
Polar graphs
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