NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Reynolds number: 50,000 Max Cl/Cd: 31.86 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc12n1-il-50000-n5.txt Download as CSV file: xf-rc12n1-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-12(N)1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.6059 0.09605 0.08933 -0.0216 1.0000 0.0584
-9.750 -0.6051 0.09140 0.08469 -0.0231 1.0000 0.0580
-9.500 -0.6077 0.08682 0.08012 -0.0248 1.0000 0.0575
-9.250 -0.6136 0.08241 0.07569 -0.0261 1.0000 0.0570
-9.000 -0.6216 0.07829 0.07154 -0.0266 1.0000 0.0564
-8.750 -0.6277 0.07410 0.06726 -0.0271 1.0000 0.0557
-8.500 -0.6321 0.06995 0.06297 -0.0273 1.0000 0.0550
-8.250 -0.6342 0.06588 0.05872 -0.0271 1.0000 0.0543
-8.000 -0.6338 0.06194 0.05454 -0.0266 1.0000 0.0537
-7.750 -0.6307 0.05815 0.05048 -0.0258 1.0000 0.0532
-7.500 -0.6250 0.05458 0.04659 -0.0248 1.0000 0.0528
-7.250 -0.6171 0.05125 0.04292 -0.0236 1.0000 0.0526
-7.000 -0.6079 0.04832 0.03962 -0.0222 1.0000 0.0530
-6.750 -0.5990 0.04585 0.03683 -0.0203 1.0000 0.0542
-6.500 -0.5935 0.04382 0.03447 -0.0176 1.0000 0.0554
-6.250 -0.5896 0.04206 0.03238 -0.0143 1.0000 0.0563
-6.000 -0.5732 0.03988 0.02975 -0.0131 0.9948 0.0571
-5.750 -0.5405 0.03727 0.02660 -0.0145 0.9845 0.0579
-5.500 -0.5066 0.03477 0.02375 -0.0161 0.9755 0.0594
-5.250 -0.4716 0.03282 0.02170 -0.0182 0.9678 0.0633
-5.000 -0.4341 0.03113 0.01976 -0.0201 0.9603 0.0675
-4.750 -0.3949 0.02957 0.01791 -0.0218 0.9536 0.0706
-4.500 -0.3587 0.02812 0.01648 -0.0234 0.9467 0.0768
-4.250 -0.3236 0.02707 0.01523 -0.0247 0.9395 0.0842
-4.000 -0.2935 0.02595 0.01408 -0.0254 0.9314 0.0919
-3.750 -0.2659 0.02502 0.01305 -0.0258 0.9231 0.1048
-3.500 -0.2392 0.02404 0.01206 -0.0261 0.9152 0.1240
-3.250 -0.2182 0.02261 0.01109 -0.0258 0.9065 0.1850
-3.000 -0.2163 0.02051 0.01142 -0.0200 0.8990 0.6715
-2.750 -0.1771 0.02114 0.01231 -0.0176 0.8946 0.8069
-2.500 -0.1041 0.02280 0.01366 -0.0205 0.8941 0.8947
-2.000 -0.0280 0.02269 0.01299 -0.0253 0.8801 0.9178
-1.750 0.0073 0.02263 0.01271 -0.0274 0.8725 0.9285
-1.500 0.0464 0.02254 0.01242 -0.0302 0.8660 0.9376
-1.250 0.0817 0.02250 0.01223 -0.0324 0.8585 0.9485
-1.000 0.1175 0.02246 0.01204 -0.0346 0.8521 0.9598
-0.750 0.1563 0.02242 0.01188 -0.0378 0.8445 0.9702
-0.500 0.1951 0.02235 0.01170 -0.0406 0.8385 0.9807
-0.250 0.2340 0.02234 0.01162 -0.0439 0.8307 0.9920
0.000 0.2693 0.02229 0.01150 -0.0462 0.8242 1.0000
0.250 0.2872 0.02244 0.01162 -0.0456 0.8153 1.0000
0.500 0.3063 0.02256 0.01167 -0.0446 0.8082 1.0000
0.750 0.3238 0.02282 0.01192 -0.0437 0.7990 1.0000
1.000 0.3438 0.02301 0.01206 -0.0427 0.7919 1.0000
1.250 0.3621 0.02332 0.01237 -0.0418 0.7827 1.0000
1.500 0.3818 0.02360 0.01262 -0.0407 0.7749 1.0000
1.750 0.4012 0.02390 0.01294 -0.0396 0.7663 1.0000
2.000 0.4201 0.02426 0.01331 -0.0385 0.7574 1.0000
2.250 0.4411 0.02451 0.01356 -0.0373 0.7495 1.0000
2.500 0.4591 0.02494 0.01405 -0.0361 0.7394 1.0000
2.750 0.4806 0.02517 0.01430 -0.0348 0.7315 1.0000
3.000 0.4994 0.02556 0.01475 -0.0335 0.7211 1.0000
3.250 0.5184 0.02594 0.01520 -0.0322 0.7109 1.0000
3.500 0.5414 0.02606 0.01540 -0.0309 0.7029 1.0000
3.750 0.5592 0.02651 0.01594 -0.0295 0.6910 1.0000
4.000 0.5785 0.02685 0.01639 -0.0281 0.6799 1.0000
4.250 0.6008 0.02697 0.01659 -0.0266 0.6701 1.0000
4.500 0.6218 0.02713 0.01690 -0.0251 0.6587 1.0000
4.750 0.6410 0.02739 0.01729 -0.0235 0.6456 1.0000
5.000 0.6612 0.02752 0.01756 -0.0219 0.6320 1.0000
5.250 0.6823 0.02745 0.01763 -0.0200 0.6172 1.0000
5.500 0.7049 0.02706 0.01740 -0.0179 0.6001 1.0000
5.750 0.7254 0.02666 0.01713 -0.0155 0.5787 1.0000
6.000 0.7468 0.02609 0.01667 -0.0131 0.5545 1.0000
6.250 0.7655 0.02579 0.01649 -0.0106 0.5254 1.0000
6.500 0.7833 0.02561 0.01643 -0.0082 0.4904 1.0000
6.750 0.8008 0.02551 0.01632 -0.0057 0.4469 1.0000
7.000 0.8166 0.02563 0.01625 -0.0029 0.3915 1.0000
7.250 0.8272 0.02637 0.01662 -0.0001 0.3270 1.0000
7.500 0.8326 0.02772 0.01756 0.0028 0.2646 1.0000
7.750 0.8352 0.02942 0.01888 0.0054 0.2149 1.0000
8.000 0.8383 0.03117 0.02036 0.0079 0.1791 1.0000
8.250 0.8421 0.03294 0.02188 0.0101 0.1558 1.0000
8.500 0.8481 0.03458 0.02344 0.0122 0.1373 1.0000
8.750 0.8551 0.03619 0.02499 0.0142 0.1240 1.0000
9.000 0.8639 0.03779 0.02652 0.0160 0.1127 1.0000
9.250 0.8757 0.03933 0.02810 0.0175 0.1027 1.0000
9.500 0.8938 0.04080 0.02972 0.0189 0.0939 1.0000
9.750 0.9120 0.04234 0.03123 0.0199 0.0863 1.0000
10.000 0.9341 0.04398 0.03312 0.0208 0.0796 1.0000
10.250 0.9563 0.04580 0.03491 0.0214 0.0741 1.0000
10.500 0.9711 0.04790 0.03741 0.0225 0.0692 1.0000
10.750 0.9878 0.05010 0.03979 0.0234 0.0658 1.0000
11.000 1.0053 0.05254 0.04225 0.0239 0.0627 1.0000
11.250 1.0050 0.05535 0.04551 0.0256 0.0605 1.0000
11.500 1.0033 0.05832 0.04883 0.0269 0.0584 1.0000
11.750 1.0005 0.06144 0.05222 0.0278 0.0567 1.0000
12.000 0.9955 0.06491 0.05594 0.0285 0.0557 1.0000
12.250 0.9870 0.06872 0.05999 0.0288 0.0550 1.0000
12.500 0.9750 0.07295 0.06446 0.0285 0.0545 1.0000
12.750 0.9586 0.07779 0.06953 0.0276 0.0543 1.0000
13.000 0.9367 0.08358 0.07555 0.0257 0.0543 1.0000
13.250 0.9046 0.09134 0.08354 0.0219 0.0548 1.0000
13.500 0.8593 0.10274 0.09513 0.0148 0.0561 1.0000
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Polar data table (+)
Polar graphs
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