NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Reynolds number: 500,000 Max Cl/Cd: 54.18 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rcsc2-il-500000.txt Download as CSV file: xf-rcsc2-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-SC2 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6452 0.08068 0.07802 -0.0432 1.0000 0.0381
-10.500 -0.6843 0.07168 0.06886 -0.0508 1.0000 0.0381
-10.250 -0.7162 0.06734 0.06439 -0.0505 1.0000 0.0381
-10.000 -0.7473 0.06500 0.06195 -0.0455 1.0000 0.0381
-9.750 -0.7734 0.06241 0.05921 -0.0407 1.0000 0.0382
-9.500 -0.8019 0.05579 0.05253 -0.0370 1.0000 0.0387
-9.250 -0.8004 0.05370 0.05044 -0.0344 1.0000 0.0390
-9.000 -0.8001 0.05180 0.04850 -0.0315 1.0000 0.0392
-8.750 -0.8979 0.03270 0.02762 -0.0165 1.0000 0.0279
-8.500 -0.8878 0.03129 0.02596 -0.0136 1.0000 0.0277
-8.250 -0.8741 0.03026 0.02477 -0.0113 1.0000 0.0275
-8.000 -0.8633 0.02731 0.02162 -0.0090 1.0000 0.0272
-7.750 -0.8496 0.02542 0.01950 -0.0068 1.0000 0.0272
-7.500 -0.8337 0.02391 0.01778 -0.0048 1.0000 0.0272
-7.250 -0.8161 0.02264 0.01633 -0.0031 1.0000 0.0273
-7.000 -0.7973 0.02157 0.01510 -0.0016 1.0000 0.0274
-6.750 -0.7712 0.02062 0.01400 -0.0015 0.9992 0.0275
-6.500 -0.7441 0.01901 0.01226 -0.0017 0.9983 0.0277
-6.250 -0.7163 0.01777 0.01095 -0.0020 0.9972 0.0281
-6.000 -0.6873 0.01691 0.01007 -0.0025 0.9961 0.0285
-5.750 -0.6571 0.01626 0.00941 -0.0032 0.9950 0.0291
-5.500 -0.6266 0.01575 0.00888 -0.0040 0.9937 0.0299
-5.250 -0.5975 0.01525 0.00835 -0.0044 0.9918 0.0307
-5.000 -0.5677 0.01474 0.00781 -0.0050 0.9899 0.0314
-4.750 -0.5367 0.01430 0.00734 -0.0057 0.9882 0.0321
-4.500 -0.5051 0.01367 0.00670 -0.0067 0.9866 0.0332
-4.250 -0.4711 0.01322 0.00627 -0.0082 0.9852 0.0347
-4.000 -0.4386 0.01290 0.00596 -0.0093 0.9832 0.0366
-3.750 -0.4104 0.01259 0.00563 -0.0094 0.9801 0.0385
-3.500 -0.3793 0.01224 0.00533 -0.0102 0.9778 0.0423
-3.250 -0.3454 0.01193 0.00508 -0.0116 0.9758 0.0487
-3.000 -0.3103 0.01158 0.00482 -0.0132 0.9742 0.0652
-2.750 -0.2754 0.01120 0.00463 -0.0149 0.9729 0.0985
-2.500 -0.2500 0.01084 0.00449 -0.0145 0.9690 0.1427
-2.250 -0.2208 0.01040 0.00435 -0.0151 0.9659 0.2111
-2.000 -0.1900 0.00978 0.00421 -0.0161 0.9638 0.3249
-1.750 -0.1604 0.00884 0.00407 -0.0170 0.9620 0.5256
-1.500 -0.1279 0.00815 0.00402 -0.0181 0.9602 0.6817
-1.250 -0.1004 0.00780 0.00400 -0.0175 0.9547 0.7664
-1.000 -0.0633 0.00751 0.00390 -0.0189 0.9502 0.8227
-0.750 -0.0235 0.00730 0.00378 -0.0208 0.9472 0.8612
-0.500 0.0070 0.00716 0.00369 -0.0207 0.9413 0.8810
-0.250 0.0382 0.00701 0.00357 -0.0207 0.9351 0.8990
0.000 0.0710 0.00693 0.00353 -0.0209 0.9315 0.9188
0.250 0.0958 0.00691 0.00354 -0.0194 0.9223 0.9341
0.500 0.1249 0.00678 0.00339 -0.0188 0.9135 0.9446
0.750 0.1595 0.00669 0.00330 -0.0193 0.9020 0.9521
1.000 0.1912 0.00665 0.00324 -0.0195 0.8869 0.9598
1.250 0.2311 0.00666 0.00322 -0.0216 0.8709 0.9635
1.500 0.2687 0.00667 0.00320 -0.0232 0.8450 0.9679
1.750 0.3009 0.00672 0.00311 -0.0237 0.8063 0.9727
2.000 0.3353 0.00694 0.00303 -0.0247 0.7303 0.9745
2.250 0.3630 0.00766 0.00309 -0.0246 0.5841 0.9777
2.500 0.3864 0.00868 0.00330 -0.0241 0.3981 0.9825
2.750 0.4161 0.00945 0.00350 -0.0250 0.2764 0.9854
3.000 0.4501 0.00993 0.00365 -0.0267 0.2043 0.9875
3.250 0.4841 0.01032 0.00381 -0.0283 0.1585 0.9903
3.500 0.5171 0.01068 0.00398 -0.0296 0.1218 0.9933
3.750 0.5516 0.01100 0.00416 -0.0313 0.0960 0.9954
4.000 0.5860 0.01132 0.00434 -0.0329 0.0760 0.9975
4.250 0.6198 0.01169 0.00460 -0.0344 0.0563 0.9995
4.500 0.6437 0.01207 0.00488 -0.0338 0.0440 1.0000
4.750 0.6642 0.01239 0.00515 -0.0324 0.0388 1.0000
5.000 0.6845 0.01269 0.00546 -0.0309 0.0362 1.0000
5.250 0.7044 0.01300 0.00577 -0.0293 0.0342 1.0000
5.500 0.7222 0.01354 0.00629 -0.0274 0.0325 1.0000
5.750 0.7415 0.01388 0.00666 -0.0257 0.0317 1.0000
6.000 0.7602 0.01426 0.00707 -0.0239 0.0309 1.0000
6.250 0.7785 0.01469 0.00752 -0.0220 0.0302 1.0000
6.500 0.7965 0.01514 0.00798 -0.0201 0.0295 1.0000
6.750 0.8143 0.01563 0.00849 -0.0182 0.0290 1.0000
7.000 0.8316 0.01620 0.00907 -0.0161 0.0285 1.0000
7.250 0.8484 0.01691 0.00979 -0.0141 0.0281 1.0000
7.500 0.8650 0.01785 0.01076 -0.0120 0.0277 1.0000
7.750 0.8827 0.01885 0.01182 -0.0101 0.0274 1.0000
8.000 0.9019 0.01945 0.01249 -0.0085 0.0272 1.0000
8.250 0.9211 0.02013 0.01324 -0.0069 0.0269 1.0000
8.500 0.9402 0.02091 0.01410 -0.0053 0.0267 1.0000
8.750 0.9591 0.02183 0.01513 -0.0037 0.0265 1.0000
9.000 0.9778 0.02286 0.01626 -0.0022 0.0264 1.0000
9.250 0.9962 0.02400 0.01753 -0.0006 0.0263 1.0000
9.500 1.0139 0.02527 0.01895 0.0010 0.0262 1.0000
9.750 1.0305 0.02669 0.02053 0.0027 0.0262 1.0000
10.000 1.0458 0.02827 0.02231 0.0045 0.0262 1.0000
10.250 1.0590 0.03008 0.02433 0.0065 0.0262 1.0000
10.500 1.0697 0.03212 0.02662 0.0088 0.0263 1.0000
10.750 1.0773 0.03443 0.02918 0.0113 0.0264 1.0000
11.000 1.0810 0.03705 0.03207 0.0141 0.0265 1.0000
11.250 1.0800 0.03995 0.03524 0.0172 0.0267 1.0000
11.500 1.0740 0.04313 0.03868 0.0206 0.0269 1.0000
11.750 1.0613 0.04656 0.04231 0.0245 0.0272 1.0000
12.000 1.0637 0.04935 0.04517 0.0265 0.0276 1.0000
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Polar data table (+)
Polar graphs
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