DAE-51 AIRFOIL (dae51-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: DAE-51 AIRFOIL (dae51-il) Reynolds number: 200,000 Max Cl/Cd: 87.56 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dae51-il-200000.txt Download as CSV file: xf-dae51-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: DAE-51 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3424 0.09480 0.09152 -0.0402 1.0000 0.0429
-8.000 -0.3652 0.09460 0.09142 -0.0358 1.0000 0.0429
-7.750 -0.3904 0.09458 0.09149 -0.0307 1.0000 0.0429
-7.500 -0.3682 0.08879 0.08570 -0.0447 0.9937 0.0433
-7.250 -0.3555 0.08207 0.07899 -0.0483 0.9895 0.0441
-7.000 -0.3377 0.07883 0.07574 -0.0474 0.9868 0.0452
-6.750 -0.3123 0.07493 0.07181 -0.0519 0.9837 0.0468
-6.500 -0.2882 0.07049 0.06733 -0.0588 0.9769 0.0489
-6.250 -0.2319 0.06073 0.05715 -0.0851 0.9696 0.0548
-6.000 -0.2130 0.05397 0.05035 -0.0899 0.9634 0.0562
-5.750 -0.1861 0.05118 0.04762 -0.0917 0.9606 0.0580
-5.500 -0.1514 0.04755 0.04388 -0.0969 0.9579 0.0619
-5.250 -0.1182 0.04138 0.03715 -0.1049 0.9497 0.0697
-5.000 -0.0888 0.03880 0.03460 -0.1069 0.9458 0.0727
-4.750 -0.0519 0.03521 0.03063 -0.1117 0.9425 0.0847
-4.500 -0.0291 0.03356 0.02894 -0.1118 0.9345 0.0909
-4.250 0.0029 0.03074 0.02586 -0.1143 0.9299 0.1015
-4.000 0.0471 0.02235 0.01565 -0.1147 0.9251 0.0431
-3.750 0.0750 0.02012 0.01310 -0.1147 0.9184 0.0426
-3.500 0.1066 0.01879 0.01146 -0.1150 0.9141 0.0432
-3.250 0.1323 0.01718 0.00978 -0.1148 0.9070 0.0468
-3.000 0.1607 0.01632 0.00883 -0.1145 0.9012 0.0500
-2.750 0.1886 0.01556 0.00798 -0.1141 0.8956 0.0536
-2.500 0.2148 0.01476 0.00721 -0.1137 0.8883 0.0610
-2.250 0.2429 0.01403 0.00647 -0.1133 0.8835 0.0723
-2.000 0.2693 0.01345 0.00601 -0.1131 0.8754 0.1029
-1.750 0.2958 0.01206 0.00566 -0.1133 0.8698 0.3704
-1.500 0.3190 0.01139 0.00573 -0.1124 0.8625 0.5897
-1.250 0.3376 0.01080 0.00571 -0.1095 0.8559 0.7747
-1.000 0.3670 0.01038 0.00547 -0.1087 0.8494 1.0000
-0.750 0.3946 0.01040 0.00532 -0.1085 0.8420 1.0000
-0.500 0.4222 0.01044 0.00520 -0.1082 0.8354 1.0000
-0.250 0.4495 0.01047 0.00511 -0.1080 0.8274 1.0000
0.000 0.4771 0.01051 0.00501 -0.1076 0.8206 1.0000
0.250 0.5044 0.01054 0.00495 -0.1074 0.8124 1.0000
0.500 0.5319 0.01058 0.00489 -0.1071 0.8049 1.0000
0.750 0.5593 0.01061 0.00483 -0.1068 0.7967 1.0000
1.000 0.5867 0.01067 0.00481 -0.1065 0.7883 1.0000
1.250 0.6144 0.01067 0.00473 -0.1061 0.7805 1.0000
1.500 0.6415 0.01074 0.00476 -0.1059 0.7708 1.0000
1.750 0.6693 0.01075 0.00469 -0.1056 0.7631 1.0000
2.000 0.6964 0.01080 0.00472 -0.1053 0.7527 1.0000
2.250 0.7237 0.01086 0.00475 -0.1050 0.7428 1.0000
2.500 0.7515 0.01088 0.00469 -0.1047 0.7339 1.0000
2.750 0.7785 0.01094 0.00474 -0.1044 0.7226 1.0000
3.000 0.8056 0.01100 0.00481 -0.1040 0.7112 1.0000
3.250 0.8327 0.01107 0.00485 -0.1037 0.6999 1.0000
3.500 0.8600 0.01113 0.00487 -0.1034 0.6885 1.0000
3.750 0.8871 0.01121 0.00492 -0.1030 0.6763 1.0000
4.000 0.9138 0.01131 0.00504 -0.1027 0.6627 1.0000
4.250 0.9403 0.01142 0.00514 -0.1023 0.6485 1.0000
4.500 0.9668 0.01154 0.00526 -0.1019 0.6337 1.0000
4.750 0.9931 0.01169 0.00541 -0.1014 0.6183 1.0000
5.000 1.0190 0.01186 0.00558 -0.1009 0.6014 1.0000
5.250 1.0446 0.01204 0.00578 -0.1004 0.5826 1.0000
5.500 1.0698 0.01225 0.00597 -0.0998 0.5628 1.0000
5.750 1.0945 0.01250 0.00621 -0.0992 0.5410 1.0000
6.000 1.1184 0.01279 0.00645 -0.0984 0.5160 1.0000
6.250 1.1419 0.01310 0.00674 -0.0976 0.4889 1.0000
6.500 1.1647 0.01348 0.00708 -0.0967 0.4599 1.0000
6.750 1.1863 0.01396 0.00747 -0.0956 0.4270 1.0000
7.000 1.2068 0.01450 0.00792 -0.0945 0.3880 1.0000
7.250 1.2257 0.01519 0.00845 -0.0932 0.3455 1.0000
7.500 1.2432 0.01601 0.00908 -0.0917 0.3006 1.0000
7.750 1.2591 0.01698 0.00984 -0.0901 0.2545 1.0000
8.000 1.2734 0.01809 0.01072 -0.0884 0.2057 1.0000
8.250 1.2843 0.01948 0.01178 -0.0863 0.1521 1.0000
8.500 1.2918 0.02117 0.01311 -0.0837 0.1035 1.0000
8.750 1.2978 0.02291 0.01463 -0.0808 0.0770 1.0000
9.000 1.3016 0.02455 0.01621 -0.0776 0.0639 1.0000
9.250 1.3103 0.02581 0.01754 -0.0750 0.0556 1.0000
9.500 1.3141 0.02751 0.01928 -0.0721 0.0506 1.0000
9.750 1.3223 0.02893 0.02078 -0.0698 0.0463 1.0000
10.000 1.3241 0.03103 0.02284 -0.0671 0.0429 1.0000
10.250 1.3341 0.03248 0.02444 -0.0652 0.0401 1.0000
10.500 1.3434 0.03407 0.02612 -0.0633 0.0377 1.0000
10.750 1.3520 0.03580 0.02789 -0.0615 0.0357 1.0000
11.000 1.3640 0.03846 0.03052 -0.0595 0.0335 1.0000
11.250 1.3746 0.03992 0.03220 -0.0579 0.0322 1.0000
11.500 1.3853 0.04163 0.03408 -0.0563 0.0306 1.0000
11.750 1.3963 0.04355 0.03615 -0.0548 0.0293 1.0000
12.000 1.4066 0.04563 0.03837 -0.0533 0.0283 1.0000
12.250 1.4166 0.04788 0.04075 -0.0518 0.0274 1.0000
12.500 1.4278 0.05062 0.04361 -0.0505 0.0267 1.0000
12.750 1.4367 0.05552 0.04881 -0.0491 0.0258 1.0000
13.000 1.4297 0.05849 0.05206 -0.0470 0.0255 1.0000
13.250 1.4209 0.06145 0.05530 -0.0453 0.0252 1.0000
13.500 1.4109 0.06488 0.05901 -0.0440 0.0250 1.0000
13.750 1.3990 0.06883 0.06323 -0.0431 0.0247 1.0000
14.000 1.3849 0.07326 0.06793 -0.0426 0.0246 1.0000
14.250 1.3685 0.07820 0.07313 -0.0426 0.0246 1.0000
14.500 1.3503 0.08358 0.07876 -0.0433 0.0246 1.0000
14.750 1.3306 0.08945 0.08486 -0.0446 0.0247 1.0000
15.000 1.3099 0.09577 0.09141 -0.0467 0.0248 1.0000
15.250 1.2884 0.10274 0.09858 -0.0496 0.0249 1.0000
15.500 1.2665 0.11032 0.10636 -0.0533 0.0251 1.0000
15.750 1.2444 0.11851 0.11473 -0.0578 0.0253 1.0000
16.000 1.2221 0.12740 0.12377 -0.0630 0.0255 1.0000
16.250 1.1005 0.17249 0.16936 -0.0945 0.0323 1.0000
16.500 1.0952 0.17948 0.17632 -0.0976 0.0332 1.0000
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Polar data table (+)
Polar graphs
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