NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 100,000 Max Cl/Cd: 53.65 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m16-il-100000-n5.txt Download as CSV file: xf-m16-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M16 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.250 -0.4502 0.09546 0.09133 -0.0002 1.0000 0.0290
-7.000 -0.4407 0.09096 0.08686 -0.0004 1.0000 0.0297
-6.750 -0.4294 0.08696 0.08284 -0.0023 1.0000 0.0303
-6.500 -0.4162 0.08308 0.07896 -0.0051 1.0000 0.0311
-6.250 -0.4010 0.07921 0.07509 -0.0083 1.0000 0.0320
-6.000 -0.3836 0.07532 0.07119 -0.0119 1.0000 0.0332
-5.750 -0.3631 0.07155 0.06739 -0.0159 1.0000 0.0360
-5.500 -0.3139 0.06912 0.06460 -0.0274 0.9566 0.0401
-5.250 -0.2974 0.06293 0.05837 -0.0295 0.9235 0.0411
-5.000 -0.2848 0.05911 0.05450 -0.0290 0.8972 0.0430
-4.750 -0.2660 0.05596 0.05121 -0.0297 0.8752 0.0454
-4.250 -0.2165 0.04956 0.04425 -0.0331 0.8417 0.0563
-4.000 -0.1967 0.04684 0.04137 -0.0330 0.8274 0.0609
-3.500 -0.1435 0.04092 0.03483 -0.0349 0.8023 0.0716
-3.250 -0.1085 0.03647 0.02974 -0.0354 0.7921 0.0523
-3.000 -0.0828 0.03352 0.02646 -0.0354 0.7820 0.0511
-2.750 -0.0535 0.03092 0.02330 -0.0355 0.7719 0.0551
-2.250 0.0026 0.02694 0.01844 -0.0351 0.7528 0.0531
-2.000 0.0301 0.02508 0.01622 -0.0349 0.7437 0.0528
-1.750 0.0583 0.02354 0.01429 -0.0347 0.7344 0.0530
-1.500 0.0860 0.02232 0.01265 -0.0343 0.7264 0.0534
-1.250 0.1140 0.02103 0.01110 -0.0343 0.7172 0.0548
-1.000 0.1406 0.02035 0.01036 -0.0343 0.7088 0.0595
-0.750 0.1682 0.01963 0.00943 -0.0340 0.7006 0.0614
-0.500 0.1962 0.01884 0.00844 -0.0338 0.6922 0.0608
-0.250 0.2235 0.01819 0.00761 -0.0334 0.6849 0.0604
0.000 0.2513 0.01763 0.00697 -0.0333 0.6762 0.0602
0.250 0.2781 0.01717 0.00640 -0.0329 0.6691 0.0603
0.500 0.3053 0.01677 0.00597 -0.0327 0.6606 0.0607
0.750 0.3323 0.01649 0.00559 -0.0323 0.6532 0.0614
1.000 0.3597 0.01629 0.00531 -0.0322 0.6451 0.0627
1.250 0.3871 0.01617 0.00509 -0.0320 0.6375 0.0648
1.750 0.4419 0.01599 0.00485 -0.0317 0.6219 0.0783
2.000 0.4621 0.01440 0.00486 -0.0306 0.6149 0.6808
2.250 0.5072 0.01407 0.00490 -0.0337 0.6058 1.0000
2.500 0.5335 0.01423 0.00493 -0.0333 0.5988 1.0000
2.750 0.5605 0.01440 0.00509 -0.0333 0.5900 1.0000
3.000 0.5869 0.01457 0.00522 -0.0329 0.5830 1.0000
3.250 0.6138 0.01476 0.00544 -0.0329 0.5741 1.0000
3.500 0.6405 0.01495 0.00564 -0.0327 0.5665 1.0000
3.750 0.6672 0.01514 0.00590 -0.0325 0.5582 1.0000
4.000 0.6940 0.01536 0.00619 -0.0324 0.5498 1.0000
4.250 0.7206 0.01556 0.00644 -0.0322 0.5422 1.0000
4.500 0.7475 0.01580 0.00682 -0.0322 0.5330 1.0000
4.750 0.7741 0.01602 0.00716 -0.0319 0.5254 1.0000
5.000 0.8009 0.01627 0.00759 -0.0318 0.5162 1.0000
5.250 0.8266 0.01621 0.00762 -0.0312 0.4924 1.0000
5.500 0.8518 0.01616 0.00764 -0.0305 0.4592 1.0000
5.750 0.8755 0.01632 0.00763 -0.0298 0.3906 1.0000
6.000 0.8868 0.01946 0.00893 -0.0294 0.0920 1.0000
6.250 0.9046 0.02166 0.01073 -0.0289 0.0343 1.0000
6.500 0.9262 0.02288 0.01218 -0.0283 0.0299 1.0000
6.750 0.9464 0.02418 0.01370 -0.0277 0.0261 1.0000
7.000 0.9620 0.02600 0.01575 -0.0269 0.0230 1.0000
7.250 0.9786 0.02749 0.01744 -0.0258 0.0221 1.0000
7.500 0.9937 0.02911 0.01920 -0.0244 0.0213 1.0000
7.750 1.0093 0.03080 0.02098 -0.0229 0.0204 1.0000
8.000 1.0268 0.03248 0.02276 -0.0215 0.0189 1.0000
8.250 1.0453 0.03427 0.02466 -0.0203 0.0173 1.0000
8.500 1.0656 0.03641 0.02693 -0.0191 0.0167 1.0000
8.750 1.0865 0.03883 0.02954 -0.0180 0.0163 1.0000
9.000 1.1060 0.04154 0.03258 -0.0169 0.0161 1.0000
9.250 1.1228 0.04449 0.03586 -0.0157 0.0160 1.0000
9.500 1.1360 0.04766 0.03940 -0.0144 0.0160 1.0000
9.750 1.1448 0.05086 0.04295 -0.0131 0.0157 1.0000
10.000 1.1493 0.05408 0.04648 -0.0117 0.0153 1.0000
10.250 1.1495 0.05743 0.05010 -0.0104 0.0148 1.0000
10.500 1.1451 0.06095 0.05388 -0.0090 0.0145 1.0000
10.750 1.1352 0.06435 0.05752 -0.0074 0.0143 1.0000
11.000 1.1234 0.06802 0.06143 -0.0067 0.0142 1.0000
11.250 1.1104 0.07204 0.06567 -0.0068 0.0141 1.0000
11.500 1.0968 0.07634 0.07019 -0.0076 0.0141 1.0000
11.750 1.0827 0.08098 0.07504 -0.0091 0.0142 1.0000
12.000 1.0676 0.08608 0.08035 -0.0113 0.0143 1.0000
12.250 1.0517 0.09172 0.08619 -0.0142 0.0144 1.0000
12.500 1.0348 0.09806 0.09273 -0.0179 0.0146 1.0000
12.750 1.0160 0.10547 0.10032 -0.0225 0.0150 1.0000
13.000 0.9854 0.11714 0.11223 -0.0307 0.0162 1.0000
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