NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=200,000 Ncrit=5
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 200,000 Max Cl/Cd: 68.89 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m16-il-200000-n5.txt Download as CSV file: xf-m16-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5290 0.13581 0.13262 0.0285 1.0000 0.0169
-9.750 -0.5240 0.13248 0.12931 0.0264 1.0000 0.0169
-9.500 -0.5191 0.12922 0.12608 0.0242 1.0000 0.0170
-9.250 -0.5136 0.12585 0.12273 0.0221 1.0000 0.0170
-9.000 -0.5076 0.12234 0.11925 0.0199 1.0000 0.0171
-8.750 -0.5015 0.11866 0.11559 0.0179 1.0000 0.0171
-8.500 -0.4954 0.11484 0.11181 0.0158 1.0000 0.0171
-8.250 -0.4895 0.11097 0.10797 0.0135 1.0000 0.0171
-7.750 -0.4756 0.10295 0.10000 0.0081 1.0000 0.0171
-7.250 -0.4614 0.09289 0.08998 0.0059 1.0000 0.0165
-7.000 -0.4548 0.08841 0.08553 0.0055 1.0000 0.0155
-6.750 -0.4415 0.08411 0.08119 0.0015 0.9333 0.0148
-6.500 -0.4303 0.08021 0.07717 -0.0013 0.8901 0.0144
-6.250 -0.4179 0.07628 0.07313 -0.0040 0.8625 0.0141
-6.000 -0.4022 0.07229 0.06902 -0.0071 0.8414 0.0145
-5.750 -0.3822 0.06800 0.06458 -0.0109 0.8238 0.0156
-5.500 -0.3601 0.06344 0.05985 -0.0146 0.8084 0.0160
-5.250 -0.3382 0.05896 0.05519 -0.0174 0.7946 0.0160
-5.000 -0.3132 0.05412 0.05012 -0.0204 0.7825 0.0163
-4.750 -0.2855 0.04906 0.04480 -0.0229 0.7712 0.0169
-4.500 -0.2657 0.04677 0.04239 -0.0237 0.7591 0.0187
-4.250 -0.2433 0.04514 0.04062 -0.0243 0.7478 0.0208
-4.000 -0.2159 0.04147 0.03669 -0.0257 0.7383 0.0217
-3.750 -0.1874 0.03794 0.03283 -0.0267 0.7292 0.0234
-3.500 -0.1552 0.03400 0.02847 -0.0275 0.7207 0.0267
-3.250 -0.1251 0.02966 0.02360 -0.0278 0.7132 0.0291
-3.000 -0.1021 0.02934 0.02327 -0.0283 0.7036 0.0323
-2.750 -0.0740 0.02718 0.02077 -0.0284 0.6959 0.0351
-2.500 -0.0448 0.02461 0.01777 -0.0284 0.6882 0.0361
-2.250 -0.0154 0.02286 0.01557 -0.0282 0.6809 0.0388
-2.000 0.0129 0.02095 0.01329 -0.0282 0.6735 0.0386
-1.750 0.0414 0.01939 0.01136 -0.0281 0.6666 0.0387
-1.500 0.0699 0.01821 0.00985 -0.0280 0.6592 0.0393
-1.250 0.0983 0.01746 0.00881 -0.0279 0.6524 0.0400
-1.000 0.1264 0.01629 0.00740 -0.0279 0.6453 0.0410
-0.750 0.1542 0.01547 0.00641 -0.0279 0.6387 0.0418
-0.500 0.1820 0.01496 0.00585 -0.0280 0.6315 0.0444
-0.250 0.2097 0.01460 0.00540 -0.0280 0.6250 0.0464
0.000 0.2376 0.01410 0.00483 -0.0279 0.6178 0.0460
0.250 0.2652 0.01369 0.00435 -0.0278 0.6117 0.0456
0.500 0.2930 0.01333 0.00396 -0.0278 0.6043 0.0453
0.750 0.3205 0.01306 0.00363 -0.0277 0.5982 0.0451
1.000 0.3484 0.01282 0.00340 -0.0277 0.5906 0.0451
1.250 0.3760 0.01267 0.00320 -0.0277 0.5843 0.0453
1.500 0.4041 0.01253 0.00306 -0.0278 0.5767 0.0458
1.750 0.4319 0.01246 0.00294 -0.0278 0.5702 0.0466
2.000 0.4600 0.01239 0.00288 -0.0279 0.5625 0.0481
2.250 0.4878 0.01236 0.00282 -0.0279 0.5558 0.0508
2.500 0.5159 0.01230 0.00284 -0.0280 0.5478 0.0632
3.000 0.5764 0.01045 0.00303 -0.0291 0.5327 1.0000
3.500 0.6310 0.01069 0.00324 -0.0290 0.5172 1.0000
3.750 0.6582 0.01082 0.00339 -0.0290 0.5095 1.0000
4.000 0.6856 0.01095 0.00356 -0.0290 0.5014 1.0000
4.250 0.7129 0.01110 0.00375 -0.0290 0.4932 1.0000
4.750 0.7670 0.01137 0.00408 -0.0289 0.4600 1.0000
5.000 0.7938 0.01156 0.00425 -0.0288 0.4336 1.0000
5.250 0.8198 0.01190 0.00444 -0.0287 0.3859 1.0000
5.500 0.8432 0.01300 0.00494 -0.0288 0.2582 1.0000
5.750 0.8596 0.01623 0.00684 -0.0293 0.0259 1.0000
6.000 0.8846 0.01697 0.00773 -0.0291 0.0208 1.0000
6.250 0.9090 0.01775 0.00868 -0.0289 0.0174 1.0000
6.500 0.9320 0.01879 0.00991 -0.0286 0.0156 1.0000
6.750 0.9513 0.02041 0.01171 -0.0280 0.0140 1.0000
7.000 0.9706 0.02180 0.01324 -0.0273 0.0126 1.0000
7.250 0.9901 0.02308 0.01463 -0.0265 0.0118 1.0000
7.500 1.0082 0.02459 0.01623 -0.0255 0.0112 1.0000
7.750 1.0263 0.02620 0.01794 -0.0243 0.0107 1.0000
8.000 1.0450 0.02794 0.01978 -0.0231 0.0104 1.0000
8.250 1.0645 0.02972 0.02168 -0.0220 0.0099 1.0000
8.500 1.0833 0.03124 0.02330 -0.0213 0.0090 1.0000
8.750 1.1001 0.03324 0.02543 -0.0205 0.0082 1.0000
9.000 1.1165 0.03695 0.02936 -0.0194 0.0079 1.0000
9.250 1.1316 0.03954 0.03225 -0.0182 0.0077 1.0000
9.500 1.1440 0.04226 0.03530 -0.0169 0.0076 1.0000
9.750 1.1529 0.04527 0.03864 -0.0155 0.0076 1.0000
10.000 1.1577 0.04849 0.04219 -0.0139 0.0076 1.0000
10.250 1.1582 0.05186 0.04587 -0.0123 0.0076 1.0000
10.500 1.1530 0.05518 0.04946 -0.0104 0.0076 1.0000
10.750 1.1431 0.05855 0.05305 -0.0087 0.0076 1.0000
11.000 1.1318 0.06233 0.05705 -0.0081 0.0077 1.0000
11.250 1.1192 0.06656 0.06148 -0.0082 0.0077 1.0000
12.250 0.9451 0.07645 0.07210 -0.0011 0.0079 1.0000
12.500 0.9266 0.08164 0.07744 -0.0026 0.0079 1.0000
12.750 0.9078 0.08692 0.08286 -0.0044 0.0080 1.0000
13.000 0.8881 0.09238 0.08846 -0.0065 0.0080 1.0000
13.250 0.8684 0.09793 0.09416 -0.0091 0.0081 1.0000
13.500 0.8455 0.10339 0.09975 -0.0117 0.0081 1.0000
13.750 0.8223 0.10924 0.10574 -0.0149 0.0083 1.0000
14.000 0.7987 0.11634 0.11297 -0.0191 0.0084 1.0000
14.250 0.7673 0.12684 0.12361 -0.0252 0.0089 1.0000
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Polar data table (+)
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