NACA 64-012A AIRFOIL (n64012a-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-012A AIRFOIL (n64012a-il) Reynolds number: 500,000 Max Cl/Cd: 56.37 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012a-il-500000.txt Download as CSV file: xf-n64012a-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-012A 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.9467 0.05400 0.05061 -0.0321 1.0000 0.0169
-11.750 -0.9627 0.05048 0.04686 -0.0311 1.0000 0.0168
-11.500 -0.9763 0.04741 0.04358 -0.0291 1.0000 0.0168
-11.250 -0.9899 0.04424 0.04017 -0.0261 1.0000 0.0168
-11.000 -1.0004 0.04048 0.03613 -0.0232 1.0000 0.0169
-10.750 -1.0002 0.03740 0.03283 -0.0211 1.0000 0.0172
-10.500 -0.9896 0.03584 0.03120 -0.0198 1.0000 0.0176
-10.250 -0.9748 0.03479 0.03011 -0.0188 1.0000 0.0181
-10.000 -0.9660 0.03221 0.02726 -0.0169 1.0000 0.0182
-9.750 -0.9541 0.02982 0.02459 -0.0152 1.0000 0.0184
-9.500 -0.9390 0.02783 0.02237 -0.0138 1.0000 0.0187
-9.250 -0.9219 0.02608 0.02042 -0.0126 1.0000 0.0190
-9.000 -0.9033 0.02451 0.01867 -0.0114 1.0000 0.0194
-8.750 -0.8837 0.02317 0.01717 -0.0103 1.0000 0.0199
-8.500 -0.8633 0.02217 0.01603 -0.0093 1.0000 0.0206
-8.250 -0.8425 0.02138 0.01511 -0.0082 1.0000 0.0211
-8.000 -0.8219 0.02052 0.01415 -0.0071 1.0000 0.0214
-7.750 -0.8050 0.01844 0.01198 -0.0056 1.0000 0.0221
-7.500 -0.7881 0.01743 0.01095 -0.0039 1.0000 0.0228
-7.250 -0.7709 0.01673 0.01023 -0.0021 1.0000 0.0235
-7.000 -0.7541 0.01612 0.00960 -0.0003 1.0000 0.0243
-6.750 -0.7385 0.01557 0.00901 0.0018 1.0000 0.0252
-6.500 -0.7238 0.01513 0.00854 0.0041 1.0000 0.0263
-6.250 -0.6912 0.01466 0.00802 0.0027 0.9972 0.0274
-6.000 -0.6588 0.01344 0.00676 0.0009 0.9929 0.0299
-5.750 -0.6238 0.01288 0.00618 -0.0010 0.9879 0.0326
-5.500 -0.5871 0.01230 0.00555 -0.0033 0.9841 0.0362
-5.250 -0.5539 0.01174 0.00500 -0.0048 0.9773 0.0432
-5.000 -0.5195 0.01119 0.00448 -0.0065 0.9710 0.0553
-4.750 -0.4903 0.01054 0.00400 -0.0071 0.9604 0.0857
-4.500 -0.4643 0.00966 0.00351 -0.0074 0.9488 0.1697
-4.250 -0.4443 0.00847 0.00298 -0.0066 0.9350 0.3208
-4.000 -0.4256 0.00755 0.00266 -0.0052 0.9208 0.4704
-3.750 -0.4016 0.00728 0.00253 -0.0044 0.9080 0.5282
-3.500 -0.3765 0.00714 0.00242 -0.0037 0.8955 0.5636
-3.250 -0.3509 0.00704 0.00234 -0.0032 0.8829 0.5871
-3.000 -0.3246 0.00699 0.00226 -0.0027 0.8712 0.6065
-2.750 -0.2983 0.00694 0.00219 -0.0023 0.8601 0.6244
-2.250 -0.2458 0.00688 0.00212 -0.0014 0.8384 0.6616
-2.000 -0.2194 0.00689 0.00212 -0.0010 0.8284 0.6810
-1.500 -0.1655 0.00687 0.00207 -0.0005 0.8089 0.7047
-1.250 -0.1381 0.00687 0.00203 -0.0003 0.7996 0.7147
-1.000 -0.1106 0.00686 0.00199 -0.0002 0.7905 0.7228
-0.750 -0.0829 0.00684 0.00197 -0.0002 0.7814 0.7308
-0.500 -0.0554 0.00686 0.00194 -0.0001 0.7733 0.7392
-0.250 -0.0277 0.00683 0.00193 -0.0001 0.7639 0.7471
0.000 0.0000 0.00685 0.00192 0.0000 0.7557 0.7557
0.250 0.0276 0.00683 0.00193 0.0001 0.7471 0.7639
0.500 0.0553 0.00686 0.00194 0.0001 0.7391 0.7733
0.750 0.0829 0.00684 0.00197 0.0002 0.7308 0.7814
1.000 0.1106 0.00686 0.00199 0.0002 0.7228 0.7905
1.250 0.1380 0.00687 0.00203 0.0004 0.7147 0.7996
1.500 0.1654 0.00687 0.00207 0.0005 0.7047 0.8089
2.000 0.2194 0.00689 0.00212 0.0010 0.6809 0.8284
2.250 0.2457 0.00688 0.00212 0.0015 0.6615 0.8384
2.750 0.2983 0.00694 0.00219 0.0023 0.6244 0.8602
3.000 0.3246 0.00699 0.00226 0.0027 0.6066 0.8712
3.250 0.3509 0.00704 0.00234 0.0032 0.5872 0.8829
3.500 0.3765 0.00714 0.00242 0.0037 0.5636 0.8955
3.750 0.4016 0.00728 0.00253 0.0044 0.5280 0.9081
4.000 0.4256 0.00755 0.00266 0.0052 0.4704 0.9208
4.250 0.4444 0.00846 0.00298 0.0066 0.3212 0.9350
4.500 0.4643 0.00966 0.00351 0.0074 0.1695 0.9489
4.750 0.4903 0.01054 0.00400 0.0071 0.0858 0.9605
5.000 0.5195 0.01119 0.00448 0.0065 0.0553 0.9710
5.250 0.5539 0.01174 0.00500 0.0048 0.0432 0.9773
5.500 0.5872 0.01231 0.00556 0.0032 0.0361 0.9841
5.750 0.6239 0.01288 0.00618 0.0010 0.0326 0.9880
6.000 0.6589 0.01344 0.00676 -0.0010 0.0299 0.9930
6.250 0.6915 0.01466 0.00802 -0.0027 0.0274 0.9973
6.500 0.7235 0.01512 0.00853 -0.0040 0.0263 1.0000
6.750 0.7383 0.01556 0.00901 -0.0018 0.0252 1.0000
7.000 0.7539 0.01611 0.00959 0.0003 0.0243 1.0000
7.250 0.7707 0.01673 0.01022 0.0022 0.0235 1.0000
7.500 0.7880 0.01743 0.01095 0.0039 0.0228 1.0000
7.750 0.8050 0.01844 0.01198 0.0056 0.0221 1.0000
8.000 0.8220 0.02052 0.01414 0.0071 0.0214 1.0000
8.250 0.8426 0.02137 0.01511 0.0082 0.0211 1.0000
8.500 0.8634 0.02218 0.01604 0.0093 0.0206 1.0000
8.750 0.8839 0.02316 0.01717 0.0103 0.0199 1.0000
9.000 0.9036 0.02451 0.01868 0.0114 0.0194 1.0000
9.250 0.9222 0.02608 0.02042 0.0125 0.0190 1.0000
9.500 0.9393 0.02784 0.02238 0.0138 0.0187 1.0000
9.750 0.9544 0.02984 0.02461 0.0152 0.0184 1.0000
10.000 0.9664 0.03222 0.02727 0.0168 0.0182 1.0000
10.250 0.9753 0.03480 0.03012 0.0187 0.0180 1.0000
10.500 0.9902 0.03584 0.03120 0.0197 0.0176 1.0000
10.750 1.0008 0.03742 0.03285 0.0210 0.0172 1.0000
11.000 1.0008 0.04056 0.03621 0.0231 0.0169 1.0000
11.250 0.9900 0.04436 0.04030 0.0260 0.0168 1.0000
11.500 0.9766 0.04751 0.04369 0.0290 0.0167 1.0000
11.750 0.9631 0.05059 0.04699 0.0309 0.0168 1.0000
12.000 0.9471 0.05411 0.05074 0.0318 0.0169 1.0000
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Polar data table (+)
Polar graphs
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