N-12 (n12-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: N-12 (n12-il) Reynolds number: 500,000 Max Cl/Cd: 53.28 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n12-il-500000.txt Download as CSV file: xf-n12-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: N-12
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7156 0.08568 0.08325 -0.0242 1.0000 0.0276
-11.750 -0.9603 0.04772 0.04389 -0.0407 1.0000 0.0186
-11.250 -0.9856 0.04345 0.03933 -0.0323 1.0000 0.0188
-11.000 -0.9542 0.04422 0.04038 -0.0331 1.0000 0.0196
-10.750 -0.9583 0.04234 0.03836 -0.0295 1.0000 0.0198
-10.500 -0.9638 0.03993 0.03574 -0.0257 1.0000 0.0200
-10.250 -0.9626 0.03807 0.03370 -0.0225 1.0000 0.0204
-10.000 -0.9613 0.03593 0.03133 -0.0191 1.0000 0.0208
-9.750 -0.9569 0.03409 0.02924 -0.0158 1.0000 0.0214
-9.500 -0.9503 0.03252 0.02739 -0.0126 1.0000 0.0221
-9.250 -0.9397 0.03223 0.02684 -0.0096 1.0000 0.0229
-9.000 -0.9362 0.02903 0.02328 -0.0062 1.0000 0.0236
-8.750 -0.9216 0.02694 0.02109 -0.0045 1.0000 0.0242
-8.500 -0.9036 0.02581 0.01997 -0.0031 1.0000 0.0250
-8.250 -0.8864 0.02474 0.01883 -0.0014 1.0000 0.0257
-8.000 -0.8691 0.02377 0.01775 0.0003 1.0000 0.0266
-7.750 -0.8514 0.02298 0.01683 0.0022 1.0000 0.0276
-7.500 -0.8332 0.02232 0.01604 0.0040 1.0000 0.0285
-7.250 -0.8151 0.02195 0.01553 0.0060 1.0000 0.0291
-7.000 -0.7969 0.02062 0.01404 0.0077 1.0000 0.0297
-6.750 -0.7772 0.01844 0.01184 0.0089 1.0000 0.0307
-6.500 -0.7586 0.01753 0.01094 0.0105 1.0000 0.0318
-6.250 -0.7404 0.01683 0.01023 0.0123 1.0000 0.0328
-6.000 -0.7227 0.01616 0.00954 0.0142 1.0000 0.0334
-5.750 -0.7058 0.01556 0.00891 0.0163 1.0000 0.0341
-5.500 -0.6896 0.01506 0.00837 0.0185 1.0000 0.0348
-5.250 -0.6736 0.01464 0.00792 0.0207 1.0000 0.0356
-5.000 -0.6393 0.01416 0.00741 0.0191 0.9977 0.0365
-4.750 -0.6040 0.01370 0.00690 0.0174 0.9949 0.0370
-4.500 -0.5717 0.01303 0.00620 0.0162 0.9917 0.0377
-4.250 -0.5400 0.01231 0.00545 0.0151 0.9879 0.0393
-4.000 -0.5050 0.01183 0.00495 0.0134 0.9850 0.0409
-3.750 -0.4715 0.01145 0.00455 0.0121 0.9810 0.0421
-3.500 -0.4377 0.01111 0.00418 0.0108 0.9763 0.0436
-3.250 -0.3996 0.01079 0.00383 0.0086 0.9729 0.0454
-3.000 -0.3668 0.01049 0.00351 0.0076 0.9658 0.0473
-2.500 -0.3013 0.00953 0.00290 0.0055 0.9526 0.1228
-2.250 -0.2692 0.00887 0.00257 0.0044 0.9463 0.1929
-2.000 -0.2438 0.00793 0.00230 0.0046 0.9365 0.3582
-1.750 -0.2038 0.00750 0.00209 0.0019 0.9294 0.4167
-1.500 -0.1613 0.00697 0.00186 -0.0015 0.9180 0.4895
-1.250 -0.1294 0.00626 0.00174 -0.0025 0.9036 0.6391
-1.000 -0.0988 0.00601 0.00169 -0.0029 0.8887 0.7085
-0.750 -0.0712 0.00587 0.00170 -0.0025 0.8739 0.7558
-0.500 -0.0444 0.00582 0.00172 -0.0020 0.8582 0.7916
-0.250 -0.0185 0.00581 0.00177 -0.0011 0.8404 0.8259
0.000 0.0057 0.00583 0.00180 0.0001 0.8208 0.8539
0.250 0.0326 0.00586 0.00182 0.0007 0.7978 0.8738
0.500 0.0599 0.00591 0.00182 0.0011 0.7713 0.8883
0.750 0.0871 0.00598 0.00182 0.0015 0.7427 0.9004
1.000 0.1196 0.00610 0.00187 0.0008 0.7105 0.9123
1.250 0.1554 0.00633 0.00195 -0.0007 0.6676 0.9217
1.500 0.1854 0.00660 0.00203 -0.0010 0.6177 0.9308
1.750 0.2218 0.00697 0.00217 -0.0028 0.5603 0.9365
2.000 0.2521 0.00731 0.00231 -0.0032 0.5066 0.9448
2.250 0.2892 0.00780 0.00253 -0.0052 0.4362 0.9491
2.500 0.3185 0.00828 0.00272 -0.0056 0.3669 0.9546
2.750 0.3486 0.00869 0.00292 -0.0062 0.3198 0.9597
3.000 0.3845 0.00900 0.00312 -0.0080 0.2912 0.9626
3.250 0.4177 0.00929 0.00331 -0.0093 0.2665 0.9671
3.500 0.4469 0.00958 0.00349 -0.0096 0.2397 0.9722
3.750 0.4825 0.00989 0.00366 -0.0115 0.1985 0.9746
4.000 0.5125 0.01070 0.00403 -0.0124 0.1040 0.9788
4.250 0.5405 0.01126 0.00436 -0.0127 0.0596 0.9838
4.500 0.5759 0.01168 0.00471 -0.0145 0.0495 0.9860
4.750 0.6117 0.01201 0.00507 -0.0163 0.0464 0.9890
5.000 0.6451 0.01238 0.00549 -0.0176 0.0437 0.9924
5.250 0.6783 0.01287 0.00599 -0.0190 0.0412 0.9953
5.500 0.7123 0.01355 0.00673 -0.0206 0.0389 0.9984
5.750 0.7416 0.01392 0.00715 -0.0212 0.0377 1.0000
6.000 0.7573 0.01430 0.00758 -0.0189 0.0366 1.0000
6.250 0.7726 0.01470 0.00802 -0.0165 0.0355 1.0000
6.500 0.7876 0.01515 0.00849 -0.0141 0.0342 1.0000
6.750 0.8015 0.01574 0.00910 -0.0116 0.0328 1.0000
7.000 0.8117 0.01702 0.01041 -0.0085 0.0314 1.0000
7.250 0.8280 0.01769 0.01116 -0.0063 0.0307 1.0000
7.500 0.8458 0.01816 0.01170 -0.0044 0.0299 1.0000
7.750 0.8636 0.01865 0.01225 -0.0025 0.0287 1.0000
8.000 0.8810 0.01923 0.01288 -0.0007 0.0275 1.0000
8.250 0.8980 0.01988 0.01356 0.0012 0.0265 1.0000
8.500 0.9144 0.02077 0.01447 0.0031 0.0257 1.0000
8.750 0.9284 0.02327 0.01706 0.0050 0.0246 1.0000
9.000 0.9450 0.02355 0.01748 0.0071 0.0239 1.0000
9.250 0.9610 0.02422 0.01829 0.0092 0.0231 1.0000
9.500 0.9764 0.02501 0.01919 0.0112 0.0222 1.0000
9.750 0.9916 0.02576 0.02001 0.0132 0.0214 1.0000
10.000 1.0058 0.02655 0.02087 0.0152 0.0208 1.0000
10.250 1.0192 0.02748 0.02184 0.0172 0.0202 1.0000
10.500 1.0267 0.03030 0.02482 0.0196 0.0195 1.0000
10.750 1.0320 0.03170 0.02646 0.0228 0.0192 1.0000
11.000 1.0370 0.03297 0.02797 0.0260 0.0187 1.0000
11.250 1.0382 0.03459 0.02983 0.0296 0.0182 1.0000
11.500 1.0348 0.03640 0.03185 0.0336 0.0178 1.0000
11.750 1.0263 0.03823 0.03386 0.0381 0.0175 1.0000
12.000 1.0258 0.03927 0.03500 0.0411 0.0171 1.0000
12.250 1.0186 0.04126 0.03714 0.0441 0.0169 1.0000
12.500 1.0246 0.04197 0.03787 0.0455 0.0164 1.0000
12.750 1.0148 0.04459 0.04066 0.0472 0.0163 1.0000
13.000 1.0020 0.04780 0.04404 0.0479 0.0162 1.0000
13.250 0.9912 0.05113 0.04751 0.0477 0.0161 1.0000
13.500 0.9769 0.05537 0.05191 0.0463 0.0160 1.0000
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Polar data table (+)
Polar graphs
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