NACA 23112 (naca23112-jf) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 23112 (naca23112-jf) Reynolds number: 50,000 Max Cl/Cd: 22.35 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca23112-jf-50000.txt Download as CSV file: xf-naca23112-jf-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4638 0.10226 0.09542 0.0252 1.0000 0.4460
-8.250 -0.4369 0.09805 0.09118 0.0261 1.0000 0.4697
-8.000 -0.4516 0.09702 0.09026 0.0282 1.0000 0.4942
-7.500 -0.4117 0.08827 0.08152 0.0277 1.0000 0.5189
-6.750 -0.6083 0.05479 0.04716 -0.0121 1.0000 0.2005
-6.500 -0.5966 0.05015 0.04210 -0.0113 1.0000 0.1819
-6.250 -0.5864 0.04666 0.03790 -0.0095 1.0000 0.1683
-6.000 -0.5700 0.04331 0.03434 -0.0080 1.0000 0.1630
-5.750 -0.5557 0.04132 0.03139 -0.0053 1.0000 0.1537
-5.500 -0.5357 0.03840 0.02831 -0.0040 1.0000 0.1518
-5.250 -0.5161 0.03626 0.02585 -0.0023 1.0000 0.1522
-5.000 -0.4954 0.03438 0.02362 -0.0007 1.0000 0.1530
-4.750 -0.4729 0.03254 0.02150 0.0008 1.0000 0.1534
-4.500 -0.4493 0.03049 0.01934 0.0018 1.0000 0.1559
-4.250 -0.4248 0.02879 0.01767 0.0027 1.0000 0.1616
-4.000 -0.3990 0.02747 0.01623 0.0037 1.0000 0.1664
-3.750 -0.3716 0.02610 0.01487 0.0044 1.0000 0.1737
-3.500 -0.3450 0.02502 0.01381 0.0053 1.0000 0.1843
-3.250 -0.3202 0.02389 0.01283 0.0062 1.0000 0.1966
-3.000 -0.3002 0.02291 0.01199 0.0076 1.0000 0.2150
-2.500 0.0122 0.02174 0.01295 -0.0225 1.0000 1.0000
-2.250 0.0208 0.02135 0.01252 -0.0209 1.0000 1.0000
-2.000 0.0250 0.02108 0.01223 -0.0186 1.0000 1.0000
-1.750 0.0232 0.02096 0.01211 -0.0154 1.0000 1.0000
-1.500 0.0140 0.02101 0.01217 -0.0111 1.0000 1.0000
-1.250 -0.0022 0.02123 0.01240 -0.0060 1.0000 1.0000
-1.000 -0.0206 0.02159 0.01274 -0.0008 1.0000 1.0000
-0.750 -0.0366 0.02205 0.01317 0.0037 1.0000 1.0000
-0.500 0.0374 0.02264 0.01360 -0.0070 0.9773 1.0000
-0.250 0.1472 0.02272 0.01356 -0.0227 0.9434 1.0000
0.000 0.2516 0.02214 0.01293 -0.0361 0.9076 1.0000
0.250 0.3069 0.02168 0.01242 -0.0398 0.8637 1.0000
0.500 0.3414 0.02137 0.01198 -0.0393 0.8220 1.0000
0.750 0.3626 0.02136 0.01182 -0.0366 0.7804 1.0000
1.000 0.3821 0.02141 0.01168 -0.0336 0.7426 1.0000
1.250 0.4011 0.02157 0.01164 -0.0307 0.7070 1.0000
1.500 0.4206 0.02185 0.01173 -0.0283 0.6734 1.0000
1.750 0.4407 0.02224 0.01194 -0.0263 0.6417 1.0000
2.000 0.4611 0.02272 0.01229 -0.0246 0.6131 1.0000
2.250 0.4826 0.02316 0.01253 -0.0228 0.5898 1.0000
2.500 0.5032 0.02380 0.01309 -0.0215 0.5663 1.0000
2.750 0.5244 0.02442 0.01364 -0.0202 0.5463 1.0000
3.000 0.5458 0.02510 0.01423 -0.0190 0.5285 1.0000
3.250 0.5669 0.02583 0.01491 -0.0178 0.5122 1.0000
3.500 0.5879 0.02664 0.01568 -0.0167 0.4974 1.0000
3.750 0.6090 0.02749 0.01652 -0.0156 0.4842 1.0000
4.000 0.6311 0.02824 0.01717 -0.0144 0.4723 1.0000
4.250 0.6506 0.02930 0.01832 -0.0133 0.4596 1.0000
4.500 0.6695 0.03051 0.01963 -0.0123 0.4485 1.0000
4.750 0.6916 0.03138 0.02044 -0.0112 0.4388 1.0000
5.000 0.7084 0.03279 0.02202 -0.0101 0.4280 1.0000
5.250 0.7266 0.03418 0.02349 -0.0090 0.4191 1.0000
5.500 0.7452 0.03548 0.02488 -0.0079 0.4100 1.0000
5.750 0.7594 0.03731 0.02688 -0.0068 0.4013 1.0000
6.000 0.7784 0.03864 0.02825 -0.0056 0.3930 1.0000
6.250 0.7885 0.04105 0.03087 -0.0045 0.3855 1.0000
6.500 0.8035 0.04284 0.03277 -0.0034 0.3774 1.0000
6.750 0.8123 0.04544 0.03555 -0.0023 0.3704 1.0000
7.000 0.8123 0.04900 0.03935 -0.0015 0.3639 1.0000
7.250 0.8372 0.05000 0.04032 -0.0004 0.3568 1.0000
7.500 0.8037 0.05727 0.04791 -0.0004 0.3527 1.0000
7.750 0.7568 0.06644 0.05719 -0.0026 0.3521 1.0000
8.000 0.7171 0.07524 0.06594 -0.0059 0.3551 1.0000
8.250 0.6978 0.08213 0.07282 -0.0088 0.3591 1.0000
8.500 0.7025 0.08646 0.07718 -0.0095 0.3578 1.0000
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Polar data table (+)
Polar graphs
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