NACA 642-015 AIRFOIL (n64015-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 642-015 AIRFOIL (n64015-il) Reynolds number: 50,000 Max Cl/Cd: 27.61 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64015-il-50000.txt Download as CSV file: xf-n64015-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 642-015 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.7022 0.09685 0.08975 -0.0232 1.0000 0.1450
-11.500 -0.8154 0.08237 0.07508 -0.0332 1.0000 0.1310
-11.250 -0.8229 0.07723 0.06989 -0.0338 1.0000 0.1301
-11.000 -0.8393 0.07256 0.06511 -0.0338 1.0000 0.1290
-10.750 -0.8592 0.06847 0.06088 -0.0327 1.0000 0.1280
-10.500 -0.8769 0.06466 0.05685 -0.0310 1.0000 0.1273
-10.250 -0.8892 0.06093 0.05285 -0.0293 1.0000 0.1270
-10.000 -0.8953 0.05738 0.04898 -0.0275 1.0000 0.1271
-9.750 -0.8962 0.05397 0.04522 -0.0257 1.0000 0.1276
-9.500 -0.8924 0.05071 0.04157 -0.0240 1.0000 0.1283
-9.250 -0.8836 0.04761 0.03805 -0.0224 1.0000 0.1292
-9.000 -0.8746 0.04502 0.03495 -0.0205 1.0000 0.1312
-8.750 -0.8473 0.04196 0.03205 -0.0209 1.0000 0.1373
-8.500 -0.8271 0.03963 0.02943 -0.0200 1.0000 0.1427
-8.250 -0.7993 0.03713 0.02684 -0.0198 1.0000 0.1501
-8.000 -0.7741 0.03528 0.02479 -0.0192 1.0000 0.1607
-7.750 -0.7315 0.03299 0.02279 -0.0205 1.0000 0.1794
-7.500 -0.7011 0.03117 0.02118 -0.0202 1.0000 0.2057
-7.250 -0.6882 0.02928 0.01980 -0.0179 1.0000 0.2408
-7.000 -0.6962 0.02745 0.01865 -0.0130 1.0000 0.2859
-6.750 -0.7170 0.02597 0.01828 -0.0057 1.0000 0.3562
-6.500 -0.7313 0.02777 0.02105 0.0049 1.0000 0.5124
-6.250 -0.4865 0.04727 0.03925 0.0053 1.0000 0.6469
-5.750 -0.3832 0.04903 0.04042 0.0048 1.0000 0.7144
-5.500 -0.3317 0.04855 0.03969 0.0026 1.0000 0.7419
-5.250 -0.3132 0.04786 0.03891 0.0033 1.0000 0.7634
-5.000 -0.3006 0.04722 0.03822 0.0047 1.0000 0.7817
-4.750 -0.2864 0.04657 0.03753 0.0058 1.0000 0.7980
-4.500 -0.2723 0.04593 0.03682 0.0068 1.0000 0.8127
-4.250 -0.2638 0.04540 0.03624 0.0084 1.0000 0.8256
-4.000 -0.2783 0.04518 0.03603 0.0130 1.0000 0.8347
-3.750 -0.2754 0.04474 0.03555 0.0151 1.0000 0.8447
-3.500 -0.2721 0.04424 0.03501 0.0170 1.0000 0.8540
-3.250 -0.2792 0.04389 0.03464 0.0203 1.0000 0.8618
-3.000 -0.2741 0.04336 0.03406 0.0217 1.0000 0.8695
-2.750 -0.2744 0.04292 0.03357 0.0240 1.0000 0.8767
-2.500 -0.2777 0.04245 0.03308 0.0267 1.0000 0.8834
-2.250 -0.2655 0.04196 0.03253 0.0269 1.0000 0.8907
-2.000 -0.2785 0.04155 0.03210 0.0313 1.0000 0.8972
-1.750 -0.2442 0.04108 0.03155 0.0277 0.9979 0.9042
-1.500 -0.2136 0.04081 0.03119 0.0243 0.9902 0.9115
-1.250 -0.1616 0.04054 0.03084 0.0173 0.9831 0.9184
-1.000 -0.1321 0.04031 0.03055 0.0143 0.9757 0.9251
-0.750 -0.0951 0.04015 0.03035 0.0099 0.9695 0.9312
-0.500 -0.0626 0.03998 0.03015 0.0065 0.9625 0.9375
-0.250 -0.0327 0.04000 0.03015 0.0035 0.9561 0.9438
0.000 0.0003 0.03987 0.03002 -0.0001 0.9496 0.9497
0.250 0.0331 0.04000 0.03015 -0.0036 0.9438 0.9562
0.500 0.0635 0.03997 0.03015 -0.0067 0.9373 0.9627
0.750 0.0954 0.04015 0.03035 -0.0100 0.9312 0.9695
1.000 0.1326 0.04030 0.03054 -0.0144 0.9251 0.9758
1.250 0.1624 0.04053 0.03083 -0.0174 0.9183 0.9832
1.500 0.2144 0.04080 0.03118 -0.0244 0.9115 0.9904
1.750 0.2445 0.04107 0.03154 -0.0277 0.9043 0.9981
2.000 0.2798 0.04154 0.03209 -0.0315 0.8971 1.0000
2.250 0.2655 0.04195 0.03251 -0.0270 0.8907 1.0000
2.500 0.2771 0.04243 0.03306 -0.0266 0.8835 1.0000
2.750 0.2744 0.04290 0.03356 -0.0240 0.8767 1.0000
3.000 0.2746 0.04334 0.03405 -0.0218 0.8694 1.0000
3.250 0.2791 0.04387 0.03462 -0.0203 0.8617 1.0000
3.500 0.2728 0.04422 0.03499 -0.0170 0.8538 1.0000
3.750 0.2755 0.04472 0.03553 -0.0151 0.8448 1.0000
4.000 0.2771 0.04514 0.03599 -0.0129 0.8349 1.0000
4.250 0.2639 0.04537 0.03621 -0.0084 0.8255 1.0000
4.500 0.2725 0.04592 0.03682 -0.0069 0.8129 1.0000
4.750 0.2866 0.04655 0.03751 -0.0059 0.7981 1.0000
5.000 0.3004 0.04719 0.03819 -0.0047 0.7817 1.0000
5.250 0.3133 0.04783 0.03888 -0.0033 0.7633 1.0000
5.500 0.3320 0.04853 0.03966 -0.0026 0.7418 1.0000
5.750 0.3833 0.04901 0.04040 -0.0048 0.7144 1.0000
6.000 0.4398 0.04859 0.04027 -0.0061 0.6825 1.0000
6.250 0.4866 0.04726 0.03925 -0.0053 0.6469 1.0000
6.500 0.7313 0.02776 0.02105 -0.0049 0.5124 1.0000
6.750 0.7170 0.02597 0.01827 0.0057 0.3559 1.0000
7.000 0.6963 0.02744 0.01864 0.0130 0.2859 1.0000
7.250 0.6883 0.02928 0.01980 0.0179 0.2408 1.0000
7.500 0.7012 0.03117 0.02117 0.0202 0.2057 1.0000
7.750 0.7318 0.03301 0.02279 0.0204 0.1790 1.0000
8.000 0.7740 0.03527 0.02479 0.0193 0.1607 1.0000
8.250 0.7993 0.03713 0.02683 0.0198 0.1499 1.0000
8.500 0.8272 0.03963 0.02943 0.0200 0.1427 1.0000
8.750 0.8474 0.04196 0.03205 0.0209 0.1373 1.0000
9.000 0.8748 0.04502 0.03496 0.0205 0.1312 1.0000
9.250 0.8839 0.04761 0.03805 0.0224 0.1292 1.0000
9.500 0.8923 0.05072 0.04159 0.0240 0.1282 1.0000
9.750 0.8963 0.05398 0.04523 0.0257 0.1276 1.0000
10.000 0.8949 0.05738 0.04899 0.0275 0.1270 1.0000
10.250 0.8890 0.06094 0.05286 0.0293 0.1268 1.0000
10.500 0.8769 0.06466 0.05686 0.0310 0.1272 1.0000
10.750 0.8595 0.06848 0.06089 0.0326 0.1280 1.0000
11.000 0.8398 0.07259 0.06514 0.0337 0.1290 1.0000
11.250 0.8203 0.07729 0.06995 0.0338 0.1299 1.0000
11.500 0.8168 0.08245 0.07516 0.0331 0.1311 1.0000
11.750 0.6404 0.10788 0.10068 0.0128 0.1575 1.0000
12.000 0.6620 0.11151 0.10435 0.0138 0.1609 1.0000
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Polar data table (+)
Polar graphs
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