NACA 23015 (naca23015-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 23015 (naca23015-il) Reynolds number: 100,000 Max Cl/Cd: 33.68 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca23015-il-100000.txt Download as CSV file: xf-naca23015-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 23015
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.5113 0.11291 0.10765 -0.0076 0.9999 0.2035
-10.750 -0.5422 0.10760 0.10244 -0.0120 0.9999 0.2145
-10.500 -0.8014 0.06702 0.06100 -0.0354 0.9999 0.1166
-10.000 -0.8398 0.05529 0.04808 -0.0298 0.9999 0.1041
-9.750 -0.8348 0.05183 0.04427 -0.0279 0.9999 0.1042
-9.500 -0.8268 0.04842 0.04053 -0.0262 0.9999 0.1044
-9.250 -0.8148 0.04487 0.03671 -0.0247 0.9999 0.1048
-9.000 -0.7957 0.04199 0.03388 -0.0241 0.9999 0.1069
-8.750 -0.7781 0.03989 0.03166 -0.0230 0.9999 0.1089
-8.500 -0.7613 0.03774 0.02925 -0.0215 0.9999 0.1102
-8.250 -0.7436 0.03585 0.02711 -0.0201 0.9999 0.1122
-8.000 -0.7252 0.03416 0.02512 -0.0186 0.9999 0.1143
-7.750 -0.7060 0.03261 0.02327 -0.0170 0.9999 0.1160
-7.500 -0.6865 0.03113 0.02153 -0.0155 0.9999 0.1179
-7.250 -0.6653 0.02939 0.01985 -0.0146 0.9999 0.1208
-7.000 -0.6443 0.02820 0.01862 -0.0133 0.9999 0.1235
-6.750 -0.6233 0.02715 0.01751 -0.0120 0.9999 0.1265
-6.500 -0.6028 0.02628 0.01653 -0.0106 0.9999 0.1299
-6.250 -0.5822 0.02526 0.01546 -0.0092 0.9999 0.1329
-6.000 -0.5620 0.02424 0.01459 -0.0079 0.9999 0.1366
-5.750 -0.5428 0.02354 0.01394 -0.0064 0.9999 0.1415
-5.500 -0.5240 0.02301 0.01336 -0.0048 0.9999 0.1464
-5.250 -0.5064 0.02217 0.01271 -0.0032 0.9999 0.1517
-5.000 -0.4896 0.02164 0.01227 -0.0015 0.9999 0.1589
-4.750 -0.4735 0.02107 0.01181 0.0002 0.9999 0.1677
-4.500 -0.4573 0.02059 0.01142 0.0019 0.9999 0.1805
-4.250 -0.4413 0.01989 0.01100 0.0034 0.9999 0.2048
-4.000 -0.4296 0.01808 0.01076 0.0052 0.9999 0.4407
-3.750 -0.4166 0.01785 0.01119 0.0081 0.9999 0.5861
-3.500 -0.4010 0.01807 0.01161 0.0103 0.9999 0.6459
-3.250 -0.3853 0.01838 0.01205 0.0124 0.9999 0.6912
-3.000 -0.3534 0.01866 0.01246 0.0122 0.9945 0.7376
-2.750 -0.3095 0.01886 0.01271 0.0103 0.9840 0.7828
-2.500 -0.2678 0.01910 0.01298 0.0091 0.9725 0.8236
-2.250 -0.2265 0.01947 0.01335 0.0085 0.9608 0.8600
-2.000 -0.1732 0.02002 0.01387 0.0062 0.9518 0.8898
-1.750 -0.1093 0.02042 0.01419 0.0014 0.9420 0.9115
-1.500 -0.0411 0.02042 0.01408 -0.0050 0.9315 0.9244
-1.250 0.0306 0.02012 0.01367 -0.0118 0.9200 0.9350
-1.000 0.0947 0.01985 0.01332 -0.0172 0.9015 0.9431
-0.750 0.1439 0.01961 0.01301 -0.0202 0.8801 0.9554
-0.500 0.1916 0.01944 0.01276 -0.0230 0.8551 0.9665
-0.250 0.2381 0.01922 0.01244 -0.0256 0.8291 0.9770
0.000 0.2829 0.01902 0.01212 -0.0281 0.8012 0.9877
0.250 0.3279 0.01882 0.01179 -0.0307 0.7712 0.9978
0.500 0.3508 0.01854 0.01136 -0.0302 0.7431 1.0001
0.750 0.3678 0.01823 0.01091 -0.0285 0.7162 1.0001
1.000 0.3855 0.01799 0.01049 -0.0270 0.6905 1.0001
1.250 0.4039 0.01781 0.01018 -0.0256 0.6625 1.0001
1.500 0.4228 0.01768 0.00986 -0.0241 0.6367 1.0001
1.750 0.4420 0.01762 0.00965 -0.0228 0.6094 1.0001
2.000 0.4611 0.01762 0.00948 -0.0214 0.5832 1.0001
2.250 0.4803 0.01769 0.00933 -0.0199 0.5590 1.0001
2.500 0.4996 0.01782 0.00934 -0.0186 0.5331 1.0001
2.750 0.5189 0.01799 0.00935 -0.0172 0.5095 1.0001
3.000 0.5384 0.01822 0.00940 -0.0158 0.4879 1.0001
3.250 0.5583 0.01849 0.00952 -0.0145 0.4682 1.0001
3.500 0.5784 0.01881 0.00971 -0.0133 0.4502 1.0001
3.750 0.5988 0.01916 0.00996 -0.0121 0.4336 1.0001
4.000 0.6195 0.01955 0.01026 -0.0109 0.4187 1.0001
4.250 0.6407 0.01998 0.01057 -0.0099 0.4062 1.0001
4.500 0.6624 0.02041 0.01091 -0.0088 0.3945 1.0001
4.750 0.6840 0.02091 0.01142 -0.0079 0.3834 1.0001
5.000 0.7066 0.02142 0.01175 -0.0070 0.3742 1.0001
5.250 0.7283 0.02195 0.01238 -0.0061 0.3642 1.0001
5.500 0.7518 0.02251 0.01281 -0.0054 0.3564 1.0001
5.750 0.7740 0.02314 0.01355 -0.0046 0.3479 1.0001
6.000 0.7976 0.02371 0.01406 -0.0040 0.3402 1.0001
6.250 0.8208 0.02444 0.01482 -0.0033 0.3331 1.0001
6.500 0.8436 0.02508 0.01554 -0.0027 0.3255 1.0001
6.750 0.8689 0.02580 0.01609 -0.0023 0.3192 1.0001
7.000 0.8899 0.02656 0.01711 -0.0015 0.3118 1.0001
7.250 0.9136 0.02719 0.01772 -0.0010 0.3050 1.0001
7.500 0.9368 0.02805 0.01861 -0.0004 0.2987 1.0001
7.750 0.9579 0.02884 0.01957 0.0003 0.2914 1.0001
8.000 0.9830 0.02949 0.02011 0.0007 0.2852 1.0001
8.250 1.0022 0.03052 0.02137 0.0016 0.2783 1.0001
8.500 1.0242 0.03121 0.02213 0.0023 0.2713 1.0001
8.750 1.0478 0.03208 0.02294 0.0027 0.2651 1.0001
9.000 1.0647 0.03305 0.02417 0.0039 0.2575 1.0001
9.250 1.0899 0.03359 0.02459 0.0042 0.2510 1.0001
9.500 1.1044 0.03486 0.02613 0.0056 0.2438 1.0001
9.750 1.1253 0.03550 0.02680 0.0064 0.2366 1.0001
10.000 1.1448 0.03656 0.02790 0.0072 0.2301 1.0001
10.250 1.1584 0.03761 0.02916 0.0087 0.2225 1.0001
10.500 1.1872 0.03814 0.02943 0.0085 0.2160 1.0001
10.750 1.1893 0.03975 0.03151 0.0112 0.2088 1.0001
11.000 1.2122 0.04015 0.03182 0.0117 0.2021 1.0001
11.250 1.2190 0.04182 0.03372 0.0137 0.1958 1.0001
11.500 1.2302 0.04282 0.03484 0.0154 0.1891 1.0001
11.750 1.2534 0.04353 0.03543 0.0158 0.1832 1.0001
12.000 1.2434 0.04576 0.03807 0.0193 0.1778 1.0001
12.250 1.2617 0.04624 0.03850 0.0203 0.1721 1.0001
12.500 1.2672 0.04789 0.04024 0.0222 0.1675 1.0001
12.750 1.2440 0.05066 0.04336 0.0265 0.1642 1.0001
13.000 1.2353 0.05256 0.04539 0.0295 0.1605 1.0001
13.250 1.2769 0.05224 0.04482 0.0285 0.1549 1.0001
13.500 1.2350 0.05627 0.04921 0.0329 0.1536 1.0001
13.750 1.1848 0.06222 0.05553 0.0347 0.1528 1.0001
14.000 1.1013 0.07375 0.06743 0.0322 0.1534 1.0001
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Polar data table (+)
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