GOE 525 AIRFOIL (goe525-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 525 AIRFOIL (goe525-il) Reynolds number: 500,000 Max Cl/Cd: 90.61 at α=0.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe525-il-500000-n5.txt Download as CSV file: xf-goe525-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 525 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 0.3511 0.09297 0.08982 -0.1835 0.9020 0.0217
-11.500 0.3641 0.08991 0.08672 -0.1863 0.8977 0.0227
-11.250 0.3594 0.08457 0.08135 -0.1889 0.8920 0.0235
-11.000 0.3706 0.08278 0.07955 -0.1897 0.8864 0.0237
-10.750 0.3872 0.08109 0.07784 -0.1913 0.8817 0.0240
-10.500 0.3973 0.07898 0.07570 -0.1924 0.8761 0.0241
-10.250 0.4073 0.07720 0.07391 -0.1929 0.8694 0.0245
-9.250 0.1551 0.03129 0.02736 -0.2147 0.8224 0.0273
-9.000 0.1676 0.02545 0.02118 -0.2275 0.8141 0.0274
-8.750 0.1993 0.02250 0.01801 -0.2360 0.8078 0.0276
-8.500 0.2334 0.02072 0.01606 -0.2415 0.8000 0.0279
-8.250 0.2660 0.01951 0.01467 -0.2450 0.7916 0.0281
-8.000 0.2968 0.01851 0.01351 -0.2474 0.7812 0.0284
-7.750 0.3275 0.01770 0.01256 -0.2494 0.7715 0.0287
-7.500 0.3584 0.01700 0.01170 -0.2512 0.7621 0.0291
-7.250 0.3887 0.01634 0.01090 -0.2527 0.7523 0.0295
-7.000 0.4187 0.01572 0.01011 -0.2539 0.7419 0.0299
-6.750 0.4480 0.01518 0.00942 -0.2548 0.7302 0.0302
-6.500 0.4765 0.01472 0.00881 -0.2555 0.7178 0.0305
-6.250 0.5039 0.01435 0.00828 -0.2559 0.7043 0.0308
-6.000 0.5305 0.01406 0.00783 -0.2559 0.6908 0.0311
-5.750 0.5577 0.01380 0.00744 -0.2561 0.6785 0.0313
-5.500 0.5849 0.01352 0.00704 -0.2563 0.6674 0.0316
-5.000 0.6393 0.01307 0.00643 -0.2567 0.6467 0.0324
-4.500 0.6941 0.01280 0.00601 -0.2569 0.6309 0.0334
-4.250 0.7207 0.01270 0.00585 -0.2568 0.6235 0.0340
-4.000 0.7472 0.01263 0.00570 -0.2567 0.6168 0.0348
-3.750 0.7741 0.01254 0.00555 -0.2567 0.6098 0.0354
-3.500 0.7999 0.01251 0.00544 -0.2564 0.6029 0.0359
-3.250 0.8280 0.01240 0.00529 -0.2566 0.5974 0.0367
-3.000 0.8555 0.01233 0.00519 -0.2567 0.5915 0.0376
-2.750 0.8813 0.01233 0.00514 -0.2564 0.5853 0.0385
-2.500 0.9072 0.01234 0.00509 -0.2561 0.5795 0.0397
-2.250 0.9335 0.01234 0.00506 -0.2559 0.5730 0.0409
-2.000 0.9588 0.01238 0.00505 -0.2555 0.5663 0.0429
-1.750 0.9846 0.01243 0.00507 -0.2552 0.5606 0.0460
-1.500 1.0108 0.01246 0.00509 -0.2550 0.5540 0.0513
-1.250 1.0352 0.01254 0.00516 -0.2545 0.5468 0.0618
-1.000 1.0615 0.01257 0.00522 -0.2544 0.5400 0.0838
-0.750 1.0868 0.01263 0.00534 -0.2541 0.5323 0.1170
-0.500 1.1103 0.01277 0.00548 -0.2534 0.5254 0.1386
-0.250 1.1347 0.01289 0.00560 -0.2529 0.5174 0.1556
0.000 1.1563 0.01309 0.00578 -0.2519 0.5091 0.1731
0.250 1.1813 0.01318 0.00594 -0.2516 0.5009 0.2127
0.500 1.2036 0.01337 0.00619 -0.2508 0.4926 0.2677
0.750 1.2269 0.01354 0.00641 -0.2501 0.4847 0.3064
1.000 1.2477 0.01380 0.00668 -0.2491 0.4760 0.3379
1.250 1.2698 0.01404 0.00693 -0.2482 0.4684 0.3627
1.500 1.2907 0.01434 0.00721 -0.2471 0.4604 0.3813
1.750 1.3114 0.01465 0.00752 -0.2460 0.4530 0.4028
2.000 1.3320 0.01497 0.00784 -0.2449 0.4450 0.4273
2.250 1.3512 0.01537 0.00821 -0.2435 0.4379 0.4479
2.500 1.3730 0.01566 0.00852 -0.2426 0.4316 0.4672
2.750 1.3930 0.01604 0.00888 -0.2414 0.4253 0.4842
3.000 1.4122 0.01646 0.00928 -0.2401 0.4201 0.4977
3.250 1.4337 0.01678 0.00961 -0.2392 0.4151 0.5110
3.500 1.4541 0.01716 0.00999 -0.2381 0.4100 0.5234
3.750 1.4734 0.01760 0.01041 -0.2369 0.4055 0.5346
4.000 1.4932 0.01801 0.01083 -0.2358 0.4016 0.5451
4.250 1.5142 0.01838 0.01121 -0.2348 0.3977 0.5545
4.500 1.5342 0.01879 0.01164 -0.2338 0.3935 0.5634
4.750 1.5530 0.01928 0.01211 -0.2326 0.3891 0.5718
5.000 1.5708 0.01982 0.01265 -0.2312 0.3849 0.5785
5.250 1.5912 0.02023 0.01308 -0.2303 0.3811 0.5856
5.500 1.6104 0.02071 0.01357 -0.2292 0.3767 0.5919
5.750 1.6286 0.02125 0.01412 -0.2280 0.3724 0.5971
6.000 1.6452 0.02189 0.01474 -0.2266 0.3685 0.6025
6.250 1.6644 0.02239 0.01527 -0.2255 0.3652 0.6085
6.500 1.6835 0.02290 0.01582 -0.2245 0.3615 0.6139
6.750 1.7013 0.02349 0.01644 -0.2234 0.3575 0.6204
7.000 1.7166 0.02424 0.01718 -0.2219 0.3530 0.6281
7.250 1.7329 0.02494 0.01790 -0.2206 0.3487 0.6351
7.500 1.7504 0.02558 0.01858 -0.2194 0.3435 0.6423
7.750 1.7648 0.02643 0.01942 -0.2180 0.3377 0.6481
8.000 1.7785 0.02733 0.02033 -0.2164 0.3328 0.6528
8.250 1.7958 0.02802 0.02107 -0.2154 0.3285 0.6581
8.500 1.8109 0.02887 0.02195 -0.2141 0.3240 0.6630
8.750 1.8238 0.02990 0.02298 -0.2126 0.3193 0.6671
9.000 1.8385 0.03081 0.02393 -0.2114 0.3151 0.6720
9.250 1.8535 0.03173 0.02488 -0.2102 0.3107 0.6777
9.500 1.8667 0.03279 0.02598 -0.2089 0.3061 0.6828
9.750 1.8783 0.03400 0.02721 -0.2074 0.3018 0.6878
10.000 1.8922 0.03507 0.02833 -0.2063 0.2981 0.6934
10.250 1.9058 0.03617 0.02948 -0.2051 0.2937 0.6991
10.500 1.9164 0.03755 0.03089 -0.2037 0.2886 0.7054
10.750 1.9255 0.03909 0.03245 -0.2023 0.2832 0.7125
11.000 1.9368 0.04048 0.03389 -0.2011 0.2769 0.7200
11.250 1.9426 0.04238 0.03580 -0.1995 0.2705 0.7281
11.500 1.9531 0.04391 0.03739 -0.1984 0.2656 0.7381
11.750 1.9617 0.04562 0.03917 -0.1971 0.2600 0.7519
12.000 1.9662 0.04773 0.04133 -0.1956 0.2542 0.7788
12.250 1.9718 0.04926 0.04303 -0.1939 0.2499 1.0000
12.500 1.9795 0.05116 0.04497 -0.1928 0.2450 1.0000
12.750 1.9829 0.05356 0.04739 -0.1915 0.2398 1.0000
13.000 1.9882 0.05581 0.04968 -0.1904 0.2351 1.0000
13.250 1.9942 0.05801 0.05193 -0.1894 0.2298 1.0000
13.500 1.9955 0.06081 0.05475 -0.1883 0.2247 1.0000
13.750 1.9989 0.06341 0.05739 -0.1874 0.2202 1.0000
14.000 2.0033 0.06591 0.05994 -0.1865 0.2151 1.0000
14.250 2.0026 0.06909 0.06317 -0.1856 0.2099 1.0000
14.500 2.0035 0.07213 0.06625 -0.1849 0.2048 1.0000
14.750 2.0032 0.07537 0.06953 -0.1842 0.1977 1.0000
15.000 1.9986 0.07923 0.07342 -0.1835 0.1916 1.0000
15.250 1.9963 0.08281 0.07705 -0.1830 0.1836 1.0000
15.500 1.9885 0.08721 0.08147 -0.1825 0.1742 1.0000
15.750 1.9761 0.09234 0.08659 -0.1822 0.1615 1.0000
16.000 1.9458 0.10024 0.09442 -0.1821 0.1383 1.0000
16.250 1.9099 0.10916 0.10327 -0.1825 0.1131 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 525 AIRFOIL (goe525-il)