NACA 63-212 AIRFOIL (n63212-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: NACA 63-212 AIRFOIL (n63212-il) Reynolds number: 100,000 Max Cl/Cd: 49.56 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63212-il-100000.txt Download as CSV file: xf-n63212-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63-212 AIRFOIL
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 -0.5387 0.09680 0.09190 -0.0211 1.0000 0.1584
-9.500 -0.5196 0.09404 0.08911 -0.0195 1.0000 0.1630
-9.250 -0.5495 0.08850 0.08372 -0.0253 1.0000 0.1728
-9.000 -0.5253 0.08601 0.08119 -0.0227 1.0000 0.1769
-8.750 -0.5541 0.08061 0.07593 -0.0278 1.0000 0.1880
-8.000 -0.6706 0.04786 0.04169 -0.0429 1.0000 0.0815
-7.750 -0.6700 0.04296 0.03595 -0.0403 1.0000 0.0704
-7.500 -0.6599 0.03939 0.03218 -0.0388 1.0000 0.0684
-7.250 -0.6502 0.03641 0.02889 -0.0367 1.0000 0.0664
-7.000 -0.6399 0.03379 0.02591 -0.0343 1.0000 0.0649
-6.750 -0.6283 0.03154 0.02331 -0.0319 1.0000 0.0642
-6.500 -0.6150 0.02965 0.02113 -0.0297 1.0000 0.0641
-6.250 -0.6001 0.02800 0.01926 -0.0278 1.0000 0.0647
-6.000 -0.5846 0.02671 0.01781 -0.0259 1.0000 0.0672
-5.750 -0.5681 0.02560 0.01648 -0.0242 1.0000 0.0700
-5.500 -0.5505 0.02422 0.01499 -0.0225 1.0000 0.0725
-5.250 -0.5334 0.02281 0.01370 -0.0211 1.0000 0.0762
-5.000 -0.5161 0.02182 0.01271 -0.0195 1.0000 0.0819
-4.750 -0.4996 0.02071 0.01169 -0.0181 1.0000 0.0903
-4.500 -0.4828 0.01971 0.01081 -0.0169 1.0000 0.1074
-4.250 -0.4574 0.01827 0.00971 -0.0177 0.9974 0.1537
-4.000 -0.4307 0.01578 0.00955 -0.0189 0.9907 0.5604
-3.750 -0.3928 0.01644 0.01017 -0.0201 0.9824 0.6256
-3.500 -0.3582 0.01706 0.01076 -0.0206 0.9734 0.6621
-3.250 -0.3272 0.01779 0.01148 -0.0200 0.9643 0.6962
-3.000 -0.2927 0.01861 0.01227 -0.0196 0.9568 0.7273
-2.750 -0.2669 0.01906 0.01266 -0.0183 0.9469 0.7507
-2.500 -0.2338 0.01949 0.01304 -0.0181 0.9395 0.7694
-2.250 -0.2034 0.01965 0.01310 -0.0182 0.9307 0.7846
-2.000 -0.1688 0.01972 0.01307 -0.0194 0.9231 0.7980
-1.750 -0.1339 0.01973 0.01298 -0.0208 0.9152 0.8110
-1.500 -0.1008 0.01980 0.01297 -0.0214 0.9077 0.8213
-1.250 -0.0677 0.01977 0.01288 -0.0223 0.8998 0.8321
-1.000 -0.0382 0.01975 0.01280 -0.0227 0.8915 0.8436
-0.750 -0.0047 0.01970 0.01269 -0.0236 0.8840 0.8547
-0.500 0.0218 0.01971 0.01267 -0.0233 0.8752 0.8650
-0.250 0.0557 0.01963 0.01255 -0.0242 0.8682 0.8756
0.000 0.0792 0.01964 0.01255 -0.0236 0.8586 0.8880
0.250 0.1162 0.01955 0.01243 -0.0249 0.8523 0.8981
0.500 0.1435 0.01959 0.01248 -0.0250 0.8429 0.9096
0.750 0.1828 0.01947 0.01234 -0.0269 0.8368 0.9200
1.000 0.2136 0.01953 0.01242 -0.0279 0.8274 0.9323
1.250 0.2616 0.01938 0.01229 -0.0314 0.8214 0.9409
1.500 0.3035 0.01942 0.01238 -0.0346 0.8124 0.9508
1.750 0.3508 0.01925 0.01223 -0.0381 0.8058 0.9599
2.000 0.4002 0.01925 0.01232 -0.0428 0.7964 0.9680
2.250 0.4507 0.01898 0.01210 -0.0469 0.7895 0.9759
2.500 0.4955 0.01898 0.01222 -0.0509 0.7785 0.9860
2.750 0.5426 0.01880 0.01212 -0.0548 0.7691 0.9957
3.000 0.5716 0.01868 0.01206 -0.0552 0.7590 1.0000
3.250 0.5826 0.01888 0.01234 -0.0529 0.7465 1.0000
3.500 0.5945 0.01896 0.01247 -0.0504 0.7344 1.0000
3.750 0.6083 0.01886 0.01240 -0.0476 0.7222 1.0000
4.000 0.6243 0.01860 0.01216 -0.0449 0.7093 1.0000
4.250 0.6410 0.01830 0.01191 -0.0422 0.6947 1.0000
4.500 0.6587 0.01803 0.01167 -0.0398 0.6768 1.0000
4.750 0.6812 0.01748 0.01113 -0.0376 0.6554 1.0000
5.000 0.7057 0.01686 0.01049 -0.0357 0.6311 1.0000
5.250 0.7303 0.01635 0.00997 -0.0340 0.6038 1.0000
5.500 0.7531 0.01608 0.00974 -0.0324 0.5710 1.0000
5.750 0.7750 0.01588 0.00954 -0.0306 0.5267 1.0000
6.000 0.7920 0.01598 0.00941 -0.0281 0.4372 1.0000
6.250 0.7862 0.01861 0.01035 -0.0236 0.1971 1.0000
6.500 0.7919 0.02099 0.01204 -0.0210 0.1313 1.0000
6.750 0.8066 0.02256 0.01341 -0.0193 0.1090 1.0000
7.000 0.8246 0.02410 0.01482 -0.0180 0.0966 1.0000
7.250 0.8460 0.02580 0.01630 -0.0173 0.0869 1.0000
7.500 0.8711 0.02718 0.01781 -0.0167 0.0800 1.0000
7.750 0.9003 0.02949 0.01993 -0.0170 0.0750 1.0000
8.000 0.9264 0.03122 0.02191 -0.0166 0.0712 1.0000
8.250 0.9516 0.03304 0.02389 -0.0162 0.0674 1.0000
8.500 0.9770 0.03535 0.02639 -0.0159 0.0655 1.0000
8.750 1.0006 0.03803 0.02928 -0.0155 0.0644 1.0000
9.000 1.0216 0.04111 0.03264 -0.0148 0.0637 1.0000
9.250 1.0388 0.04509 0.03689 -0.0142 0.0627 1.0000
9.500 1.0489 0.04872 0.04092 -0.0128 0.0620 1.0000
9.750 1.0579 0.05226 0.04485 -0.0114 0.0619 1.0000
10.000 1.0688 0.05580 0.04875 -0.0101 0.0628 1.0000
10.250 1.0400 0.06004 0.05401 -0.0054 0.0675 1.0000
10.500 1.0253 0.06500 0.05933 -0.0034 0.0702 1.0000
10.750 1.0170 0.06982 0.06431 -0.0022 0.0725 1.0000
11.000 0.8970 0.06592 0.06098 0.0041 0.0712 1.0000
11.250 0.8727 0.07169 0.06688 0.0036 0.0723 1.0000
11.500 0.8677 0.07743 0.07273 0.0030 0.0758 1.0000
11.750 0.8097 0.08586 0.08139 -0.0014 0.0763 1.0000
12.000 0.7379 0.10296 0.09865 -0.0113 0.0886 1.0000
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