RAE 5215 AIRFOIL (rae5215-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 5215 AIRFOIL (rae5215-il) Reynolds number: 50,000 Max Cl/Cd: 28.5 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae5215-il-50000.txt Download as CSV file: xf-rae5215-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 5215 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5366 0.12839 0.12091 0.0058 1.0000 0.2553
-9.750 -0.5177 0.12342 0.11594 0.0066 1.0000 0.2647
-9.500 -0.5294 0.12145 0.11407 0.0051 1.0000 0.2757
-9.250 -0.5213 0.11839 0.11103 0.0055 1.0000 0.2901
-9.000 -0.5013 0.11367 0.10632 0.0064 1.0000 0.3019
-8.750 -0.4959 0.11038 0.10308 0.0066 1.0000 0.3155
-8.500 -0.4902 0.10728 0.10003 0.0070 1.0000 0.3318
-8.250 -0.4814 0.10414 0.09691 0.0077 1.0000 0.3501
-8.000 -0.4725 0.10104 0.09386 0.0085 1.0000 0.3691
-7.750 -0.4645 0.09799 0.09085 0.0092 1.0000 0.3877
-7.500 -0.4565 0.09486 0.08777 0.0098 1.0000 0.4049
-7.250 -0.4487 0.09171 0.08467 0.0104 1.0000 0.4213
-7.000 -0.4405 0.08863 0.08164 0.0111 1.0000 0.4386
-6.750 -0.4333 0.08554 0.07860 0.0117 1.0000 0.4542
-6.500 -0.5574 0.05726 0.04972 -0.0404 1.0000 0.1678
-6.250 -0.5447 0.05123 0.04317 -0.0431 1.0000 0.1504
-6.000 -0.5274 0.04635 0.03730 -0.0451 1.0000 0.1376
-5.750 -0.5078 0.04254 0.03336 -0.0448 1.0000 0.1348
-5.500 -0.4863 0.03914 0.02958 -0.0448 1.0000 0.1317
-5.250 -0.4636 0.03636 0.02638 -0.0445 1.0000 0.1323
-5.000 -0.4397 0.03389 0.02351 -0.0440 1.0000 0.1335
-4.750 -0.4150 0.03165 0.02090 -0.0432 1.0000 0.1345
-4.500 -0.3897 0.02997 0.01876 -0.0424 1.0000 0.1382
-4.250 -0.3667 0.02793 0.01686 -0.0413 1.0000 0.1443
-4.000 -0.3424 0.02648 0.01530 -0.0399 1.0000 0.1505
-3.750 -0.3195 0.02508 0.01397 -0.0383 1.0000 0.1608
-3.500 -0.2967 0.02382 0.01279 -0.0367 1.0000 0.1750
-3.250 -0.2740 0.02248 0.01163 -0.0355 1.0000 0.2005
-3.000 -0.2478 0.01911 0.01018 -0.0363 1.0000 0.3940
-2.750 -0.2596 0.02058 0.01266 -0.0220 1.0000 0.7135
-2.500 -0.2635 0.02106 0.01316 -0.0117 1.0000 0.7781
-2.250 -0.2658 0.02102 0.01311 -0.0025 1.0000 0.8278
-2.000 -0.0909 0.02130 0.01258 -0.0170 1.0000 0.9848
-1.750 -0.0326 0.02069 0.01171 -0.0245 1.0000 1.0000
-1.500 -0.0321 0.02041 0.01140 -0.0215 1.0000 1.0000
-1.250 -0.0337 0.02014 0.01112 -0.0182 1.0000 1.0000
-1.000 -0.0367 0.01987 0.01085 -0.0147 1.0000 1.0000
-0.750 -0.0405 0.01959 0.01056 -0.0111 1.0000 1.0000
-0.500 -0.0441 0.01929 0.01025 -0.0076 1.0000 1.0000
-0.250 -0.0426 0.01903 0.00996 -0.0050 1.0000 1.0000
0.000 -0.0285 0.01891 0.00979 -0.0045 1.0000 1.0000
0.250 -0.0063 0.01894 0.00976 -0.0054 1.0000 1.0000
0.500 0.0200 0.01908 0.00983 -0.0069 1.0000 1.0000
0.750 0.0483 0.01929 0.00999 -0.0087 1.0000 1.0000
1.000 0.0773 0.01957 0.01024 -0.0105 1.0000 1.0000
1.250 0.1064 0.01992 0.01057 -0.0123 1.0000 1.0000
1.500 0.1352 0.02032 0.01097 -0.0140 1.0000 1.0000
1.750 0.1634 0.02077 0.01144 -0.0156 1.0000 1.0000
2.000 0.1909 0.02128 0.01199 -0.0170 1.0000 1.0000
2.250 0.2176 0.02185 0.01262 -0.0184 1.0000 1.0000
2.500 0.2435 0.02249 0.01334 -0.0196 1.0000 1.0000
2.750 0.3125 0.02336 0.01445 -0.0284 0.9770 1.0000
3.000 0.4343 0.02259 0.01427 -0.0423 0.9053 1.0000
3.250 0.5195 0.01823 0.00947 -0.0369 0.5455 1.0000
3.500 0.5352 0.02069 0.01015 -0.0338 0.3846 1.0000
3.750 0.5643 0.02226 0.01126 -0.0339 0.3380 1.0000
4.000 0.5966 0.02368 0.01234 -0.0345 0.3083 1.0000
4.250 0.6285 0.02501 0.01360 -0.0350 0.2856 1.0000
4.500 0.6596 0.02647 0.01500 -0.0354 0.2675 1.0000
4.750 0.6894 0.02807 0.01654 -0.0356 0.2519 1.0000
5.000 0.7184 0.02984 0.01822 -0.0358 0.2381 1.0000
5.250 0.7458 0.03153 0.02023 -0.0357 0.2261 1.0000
5.500 0.7725 0.03367 0.02253 -0.0355 0.2160 1.0000
5.750 0.7985 0.03573 0.02471 -0.0353 0.2058 1.0000
6.000 0.8217 0.03836 0.02772 -0.0349 0.1981 1.0000
6.250 0.8444 0.04079 0.03044 -0.0344 0.1898 1.0000
6.500 0.8658 0.04390 0.03371 -0.0340 0.1838 1.0000
6.750 0.8817 0.04734 0.03779 -0.0331 0.1797 1.0000
7.000 0.8963 0.05108 0.04197 -0.0324 0.1766 1.0000
7.250 0.9084 0.05521 0.04650 -0.0318 0.1749 1.0000
7.500 0.9128 0.06063 0.05242 -0.0316 0.1780 1.0000
7.750 0.9166 0.06613 0.05826 -0.0316 0.1813 1.0000
8.000 0.9245 0.07148 0.06376 -0.0318 0.1842 1.0000
8.250 0.7202 0.10748 0.10037 -0.0765 0.4261 1.0000
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Polar data table (+)
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