NACA 23112 (naca23112-jf) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NACA 23112 (naca23112-jf) Reynolds number: 100,000 Max Cl/Cd: 34.73 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca23112-jf-100000.txt Download as CSV file: xf-naca23112-jf-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 23112
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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4418 0.09450 0.08978 -0.0125 1.0000 0.1352
-9.500 -0.4880 0.08730 0.08266 -0.0207 1.0000 0.1369
-9.250 -0.5565 0.09168 0.08678 -0.0077 1.0000 0.1287
-9.000 -0.5721 0.08586 0.08103 -0.0134 1.0000 0.1333
-8.750 -0.6114 0.08022 0.07533 -0.0184 1.0000 0.1356
-8.500 -0.6661 0.07839 0.07295 -0.0193 1.0000 0.1375
-8.250 -0.6027 0.07277 0.06793 -0.0172 1.0000 0.1470
-7.750 -0.6103 0.06542 0.06049 -0.0168 1.0000 0.1627
-7.500 -0.6168 0.06180 0.05675 -0.0161 1.0000 0.1743
-6.500 -0.5877 0.05227 0.04730 -0.0085 1.0000 0.2446
-5.750 -0.5507 0.03631 0.02835 -0.0030 1.0000 0.1032
-5.500 -0.5329 0.03305 0.02487 -0.0011 1.0000 0.0972
-5.250 -0.5148 0.03123 0.02232 0.0019 1.0000 0.0916
-5.000 -0.4955 0.02910 0.02014 0.0033 1.0000 0.0934
-4.750 -0.4755 0.02751 0.01842 0.0048 1.0000 0.0947
-4.500 -0.4547 0.02604 0.01675 0.0064 1.0000 0.0951
-4.250 -0.4337 0.02480 0.01540 0.0078 1.0000 0.0969
-4.000 -0.4131 0.02384 0.01435 0.0092 1.0000 0.1006
-3.750 -0.3919 0.02298 0.01336 0.0106 1.0000 0.1035
-3.500 -0.3704 0.02207 0.01237 0.0119 1.0000 0.1061
-3.250 -0.3494 0.02091 0.01139 0.0127 1.0000 0.1119
-3.000 -0.3293 0.02030 0.01082 0.0139 1.0000 0.1176
-2.750 -0.3096 0.01962 0.01022 0.0150 1.0000 0.1234
-2.500 -0.2775 0.01892 0.00965 0.0135 0.9955 0.1353
-2.250 -0.2223 0.01777 0.00870 0.0079 0.9829 0.1614
-2.000 -0.1901 0.01462 0.00840 0.0074 0.9715 0.6972
-1.750 -0.1291 0.01559 0.00981 0.0073 0.9637 0.9182
-1.500 -0.0087 0.01650 0.01042 -0.0071 0.9664 0.9638
-1.250 0.1241 0.01625 0.00997 -0.0257 0.9663 0.9979
-1.000 0.1819 0.01570 0.00930 -0.0319 0.9336 1.0000
-0.750 0.2166 0.01527 0.00875 -0.0332 0.8928 1.0000
-0.500 0.2375 0.01498 0.00831 -0.0316 0.8504 1.0000
-0.250 0.2554 0.01478 0.00792 -0.0293 0.8103 1.0000
0.000 0.2739 0.01466 0.00759 -0.0271 0.7717 1.0000
0.250 0.2935 0.01460 0.00733 -0.0253 0.7323 1.0000
0.500 0.3139 0.01461 0.00710 -0.0237 0.6942 1.0000
0.750 0.3350 0.01468 0.00691 -0.0222 0.6571 1.0000
1.000 0.3568 0.01480 0.00679 -0.0210 0.6197 1.0000
1.250 0.3790 0.01497 0.00672 -0.0199 0.5835 1.0000
1.500 0.4015 0.01519 0.00670 -0.0189 0.5503 1.0000
1.750 0.4239 0.01545 0.00672 -0.0179 0.5212 1.0000
2.000 0.4465 0.01574 0.00682 -0.0170 0.4940 1.0000
2.250 0.4691 0.01607 0.00693 -0.0161 0.4720 1.0000
2.500 0.4919 0.01639 0.00714 -0.0153 0.4515 1.0000
2.750 0.5148 0.01675 0.00738 -0.0144 0.4343 1.0000
3.000 0.5376 0.01713 0.00763 -0.0136 0.4192 1.0000
3.250 0.5606 0.01753 0.00789 -0.0128 0.4060 1.0000
3.500 0.5835 0.01792 0.00824 -0.0119 0.3933 1.0000
3.750 0.6063 0.01836 0.00867 -0.0111 0.3818 1.0000
4.000 0.6294 0.01883 0.00903 -0.0103 0.3720 1.0000
4.250 0.6520 0.01928 0.00948 -0.0094 0.3618 1.0000
4.500 0.6749 0.01982 0.01002 -0.0086 0.3532 1.0000
4.750 0.6977 0.02031 0.01048 -0.0078 0.3445 1.0000
5.000 0.7202 0.02090 0.01110 -0.0069 0.3362 1.0000
5.250 0.7430 0.02144 0.01164 -0.0061 0.3284 1.0000
5.500 0.7654 0.02211 0.01235 -0.0052 0.3211 1.0000
5.750 0.7876 0.02268 0.01297 -0.0043 0.3134 1.0000
6.000 0.8103 0.02341 0.01368 -0.0035 0.3068 1.0000
6.250 0.8316 0.02410 0.01453 -0.0025 0.2994 1.0000
6.500 0.8555 0.02481 0.01511 -0.0019 0.2933 1.0000
6.750 0.8748 0.02562 0.01620 -0.0006 0.2859 1.0000
7.000 0.8974 0.02630 0.01688 0.0002 0.2795 1.0000
7.250 0.9177 0.02726 0.01798 0.0012 0.2732 1.0000
7.500 0.9383 0.02805 0.01888 0.0022 0.2661 1.0000
7.750 0.9620 0.02891 0.01966 0.0028 0.2605 1.0000
8.000 0.9784 0.02998 0.02108 0.0043 0.2532 1.0000
8.250 1.0019 0.03061 0.02166 0.0049 0.2470 1.0000
8.500 1.0181 0.03192 0.02324 0.0063 0.2405 1.0000
8.750 1.0388 0.03270 0.02411 0.0072 0.2336 1.0000
9.000 1.0594 0.03372 0.02517 0.0080 0.2275 1.0000
9.250 1.0743 0.03489 0.02662 0.0095 0.2202 1.0000
9.500 1.1020 0.03533 0.02685 0.0095 0.2141 1.0000
9.750 1.1083 0.03704 0.02906 0.0118 0.2067 1.0000
10.000 1.1350 0.03715 0.02903 0.0121 0.2001 1.0000
10.250 1.1417 0.03897 0.03124 0.0142 0.1931 1.0000
10.500 1.1666 0.03900 0.03119 0.0147 0.1860 1.0000
10.750 1.1738 0.04070 0.03318 0.0166 0.1793 1.0000
11.000 1.1961 0.04074 0.03319 0.0174 0.1720 1.0000
11.250 1.2031 0.04232 0.03499 0.0193 0.1655 1.0000
11.500 1.2224 0.04242 0.03510 0.0204 0.1582 1.0000
11.750 1.2282 0.04395 0.03682 0.0223 0.1520 1.0000
12.000 1.2416 0.04438 0.03729 0.0237 0.1450 1.0000
12.250 1.2478 0.04576 0.03877 0.0255 0.1392 1.0000
12.500 1.2443 0.04744 0.04067 0.0280 0.1337 1.0000
12.750 1.2625 0.04769 0.04078 0.0291 0.1273 1.0000
13.000 1.2371 0.05091 0.04435 0.0324 0.1246 1.0000
13.250 1.2269 0.05332 0.04693 0.0339 0.1207 1.0000
13.500 1.2444 0.05376 0.04723 0.0347 0.1148 1.0000
13.750 1.2101 0.05910 0.05290 0.0347 0.1137 1.0000
14.000 1.1658 0.06685 0.06096 0.0320 0.1139 1.0000
14.250 1.1044 0.07884 0.07317 0.0255 0.1155 1.0000
14.750 1.1394 0.07927 0.07354 0.0272 0.1043 1.0000
15.250 0.7872 0.18200 0.17617 -0.0274 0.1556 1.0000
15.500 0.7604 0.18503 0.17914 -0.0332 0.1547 1.0000
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Polar data table (+)
Polar graphs
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