EPPLER 636 AIRFOIL (e636-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 636 AIRFOIL (e636-il) Reynolds number: 1,000,000 Max Cl/Cd: 108.25 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e636-il-1000000-n5.txt Download as CSV file: xf-e636-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 636 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4475 0.08633 0.08355 -0.0134 0.6138 0.0071
-9.250 -0.4547 0.08141 0.07862 -0.0161 0.6039 0.0072
-9.000 -0.4679 0.07607 0.07330 -0.0202 0.5978 0.0071
-8.750 -0.4901 0.07072 0.06793 -0.0230 0.5920 0.0071
-8.500 -0.5067 0.06703 0.06419 -0.0219 0.5836 0.0071
-8.250 -0.5199 0.06233 0.05940 -0.0214 0.5772 0.0071
-8.000 -0.5295 0.05706 0.05400 -0.0205 0.5709 0.0073
-7.500 -0.6005 0.02923 0.02492 -0.0089 0.5748 0.0080
-7.250 -0.6021 0.02393 0.01899 -0.0048 0.5668 0.0082
-7.000 -0.5856 0.02204 0.01681 -0.0030 0.5566 0.0083
-6.750 -0.5650 0.02097 0.01556 -0.0018 0.5470 0.0084
-6.500 -0.5442 0.01976 0.01414 -0.0006 0.5372 0.0085
-6.250 -0.5221 0.01871 0.01290 0.0005 0.5271 0.0086
-6.000 -0.4989 0.01792 0.01194 0.0014 0.5174 0.0087
-5.750 -0.4751 0.01713 0.01100 0.0022 0.5077 0.0088
-5.500 -0.4507 0.01645 0.01017 0.0029 0.4994 0.0090
-5.250 -0.4259 0.01579 0.00938 0.0036 0.4910 0.0091
-5.000 -0.4006 0.01522 0.00870 0.0042 0.4832 0.0094
-4.750 -0.3755 0.01449 0.00781 0.0048 0.4748 0.0095
-4.500 -0.3500 0.01383 0.00704 0.0055 0.4669 0.0097
-4.000 -0.2987 0.01281 0.00583 0.0066 0.4525 0.0102
-3.750 -0.2731 0.01238 0.00531 0.0072 0.4454 0.0104
-3.500 -0.2472 0.01205 0.00491 0.0077 0.4387 0.0107
-3.250 -0.2213 0.01176 0.00456 0.0082 0.4314 0.0109
-3.000 -0.1954 0.01148 0.00422 0.0087 0.4251 0.0110
-2.750 -0.1692 0.01125 0.00394 0.0092 0.4193 0.0112
-2.500 -0.1444 0.01088 0.00351 0.0099 0.4133 0.0114
-2.250 -0.1196 0.01050 0.00309 0.0107 0.4078 0.0118
-2.000 -0.0938 0.01028 0.00282 0.0112 0.4017 0.0120
-1.750 -0.0678 0.01010 0.00262 0.0117 0.3966 0.0124
-1.500 -0.0414 0.00995 0.00245 0.0121 0.3918 0.0129
-1.250 -0.0150 0.00983 0.00229 0.0125 0.3863 0.0134
-1.000 0.0114 0.00970 0.00213 0.0129 0.3817 0.0139
-0.750 0.0381 0.00959 0.00199 0.0133 0.3768 0.0144
-0.500 0.0647 0.00951 0.00187 0.0136 0.3722 0.0149
-0.250 0.0915 0.00944 0.00177 0.0140 0.3682 0.0154
0.000 0.1183 0.00935 0.00167 0.0143 0.3644 0.0166
0.250 0.1452 0.00930 0.00160 0.0146 0.3597 0.0179
0.750 0.1977 0.00908 0.00151 0.0154 0.3521 0.0707
1.000 0.2214 0.00873 0.00148 0.0162 0.3488 0.1898
1.250 0.2080 0.00646 0.00133 0.0244 0.3469 0.8330
1.500 0.2554 0.00645 0.00157 0.0206 0.3421 0.9312
1.750 0.2985 0.00666 0.00175 0.0174 0.3378 0.9471
2.000 0.3272 0.00689 0.00195 0.0176 0.3344 0.9586
2.250 0.3684 0.00707 0.00209 0.0148 0.3306 0.9605
2.500 0.4001 0.00717 0.00213 0.0139 0.3268 0.9614
2.750 0.4313 0.00725 0.00218 0.0132 0.3234 0.9624
3.000 0.4622 0.00732 0.00223 0.0125 0.3199 0.9635
3.250 0.4922 0.00740 0.00229 0.0120 0.3165 0.9649
3.500 0.5211 0.00750 0.00235 0.0118 0.3131 0.9666
3.750 0.5462 0.00760 0.00243 0.0123 0.3097 0.9694
4.000 0.5709 0.00768 0.00250 0.0130 0.3069 0.9718
4.250 0.6027 0.00775 0.00257 0.0121 0.3034 0.9724
4.500 0.6342 0.00785 0.00265 0.0112 0.3000 0.9730
4.750 0.6651 0.00796 0.00274 0.0104 0.2960 0.9737
5.000 0.6960 0.00804 0.00283 0.0097 0.2931 0.9744
5.250 0.7266 0.00813 0.00292 0.0089 0.2895 0.9753
5.500 0.7566 0.00824 0.00302 0.0083 0.2855 0.9763
5.750 0.7858 0.00837 0.00314 0.0079 0.2815 0.9775
6.000 0.8146 0.00846 0.00325 0.0075 0.2782 0.9788
6.250 0.8423 0.00858 0.00337 0.0074 0.2729 0.9804
6.500 0.8665 0.00876 0.00352 0.0080 0.2670 0.9826
6.750 0.8931 0.00887 0.00366 0.0081 0.2623 0.9839
7.000 0.9239 0.00901 0.00379 0.0073 0.2572 0.9844
7.250 0.9544 0.00916 0.00394 0.0064 0.2521 0.9849
7.500 0.9847 0.00929 0.00409 0.0056 0.2464 0.9856
7.750 1.0141 0.00949 0.00427 0.0050 0.2391 0.9863
8.000 1.0433 0.00965 0.00444 0.0044 0.2319 0.9870
8.250 1.0717 0.00990 0.00466 0.0039 0.2208 0.9879
8.500 1.0993 0.01023 0.00494 0.0034 0.2050 0.9890
8.750 1.1261 0.01059 0.00524 0.0031 0.1895 0.9902
9.000 1.1504 0.01115 0.00567 0.0031 0.1656 0.9916
9.250 1.1729 0.01182 0.00622 0.0034 0.1402 0.9930
9.500 1.1957 0.01249 0.00678 0.0036 0.1196 0.9943
9.750 1.2204 0.01331 0.00746 0.0031 0.0949 0.9952
10.000 1.2450 0.01400 0.00807 0.0028 0.0777 0.9962
10.250 1.2682 0.01480 0.00878 0.0026 0.0610 0.9973
10.500 1.2919 0.01557 0.00950 0.0023 0.0482 0.9985
10.750 1.3145 0.01642 0.01030 0.0020 0.0357 0.9996
11.000 1.3280 0.01738 0.01119 0.0035 0.0249 1.0000
11.250 1.3363 0.01819 0.01200 0.0061 0.0186 1.0000
11.500 1.3428 0.01902 0.01283 0.0090 0.0135 1.0000
11.750 1.3346 0.02035 0.01415 0.0139 0.0058 1.0000
12.000 1.3291 0.02104 0.01490 0.0188 0.0056 1.0000
12.250 1.3225 0.02217 0.01610 0.0227 0.0050 1.0000
12.500 1.3197 0.02367 0.01768 0.0249 0.0046 1.0000
12.750 1.3189 0.02549 0.01958 0.0261 0.0044 1.0000
13.000 1.3174 0.02773 0.02191 0.0267 0.0040 1.0000
13.250 1.3174 0.03005 0.02431 0.0270 0.0039 1.0000
13.500 1.3152 0.03273 0.02708 0.0269 0.0038 1.0000
13.750 1.3118 0.03563 0.03006 0.0268 0.0036 1.0000
14.000 1.3074 0.03867 0.03319 0.0266 0.0035 1.0000
14.250 1.3035 0.04170 0.03630 0.0263 0.0035 1.0000
14.500 1.2953 0.04527 0.03996 0.0258 0.0035 1.0000
14.750 1.2862 0.04896 0.04373 0.0253 0.0033 1.0000
15.000 1.2779 0.05261 0.04747 0.0247 0.0032 1.0000
15.250 1.2692 0.05637 0.05131 0.0240 0.0032 1.0000
15.500 1.2623 0.06002 0.05505 0.0233 0.0032 1.0000
15.750 1.2531 0.06411 0.05923 0.0223 0.0032 1.0000
16.000 1.2432 0.06839 0.06358 0.0211 0.0030 1.0000
16.250 1.2354 0.07246 0.06774 0.0200 0.0029 1.0000
16.500 1.2257 0.07688 0.07224 0.0187 0.0028 1.0000
16.750 1.2185 0.08101 0.07645 0.0175 0.0028 1.0000
17.000 1.2081 0.08566 0.08118 0.0161 0.0027 1.0000
17.250 1.2005 0.08997 0.08557 0.0147 0.0028 1.0000
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Polar data table (+)
Polar graphs
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