RAF 28 AIRFOIL (raf28-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 28 AIRFOIL (raf28-il) Reynolds number: 200,000 Max Cl/Cd: 64.41 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf28-il-200000-n5.txt Download as CSV file: xf-raf28-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 28 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5399 0.08595 0.08234 -0.0310 1.0000 0.0178
-10.000 -0.5547 0.07877 0.07523 -0.0348 1.0000 0.0178
-9.750 -0.5827 0.06809 0.06464 -0.0418 1.0000 0.0175
-9.500 -0.6589 0.05100 0.04735 -0.0517 1.0000 0.0164
-9.250 -0.6873 0.04489 0.04092 -0.0502 1.0000 0.0162
-9.000 -0.6924 0.04162 0.03739 -0.0483 1.0000 0.0167
-8.750 -0.6970 0.03789 0.03331 -0.0459 1.0000 0.0171
-8.500 -0.6961 0.03467 0.02972 -0.0434 1.0000 0.0178
-8.250 -0.6922 0.03160 0.02621 -0.0408 1.0000 0.0187
-8.000 -0.6833 0.02926 0.02339 -0.0384 1.0000 0.0199
-7.750 -0.6709 0.02764 0.02138 -0.0362 1.0000 0.0211
-7.500 -0.6590 0.02580 0.01941 -0.0343 1.0000 0.0222
-7.250 -0.6438 0.02486 0.01838 -0.0325 1.0000 0.0233
-7.000 -0.6200 0.02360 0.01692 -0.0324 0.9979 0.0248
-6.750 -0.5875 0.02248 0.01553 -0.0339 0.9932 0.0274
-6.500 -0.5550 0.02150 0.01426 -0.0352 0.9880 0.0293
-6.250 -0.5252 0.01974 0.01227 -0.0362 0.9827 0.0314
-6.000 -0.4939 0.01874 0.01120 -0.0375 0.9770 0.0339
-5.750 -0.4636 0.01787 0.01022 -0.0383 0.9702 0.0359
-5.500 -0.4312 0.01712 0.00935 -0.0394 0.9645 0.0383
-5.250 -0.4018 0.01643 0.00853 -0.0399 0.9563 0.0398
-5.000 -0.3688 0.01574 0.00772 -0.0411 0.9507 0.0406
-4.750 -0.3407 0.01490 0.00685 -0.0414 0.9414 0.0424
-4.500 -0.3072 0.01428 0.00620 -0.0428 0.9353 0.0453
-4.250 -0.2766 0.01380 0.00564 -0.0434 0.9261 0.0474
-4.000 -0.2431 0.01336 0.00511 -0.0447 0.9187 0.0497
-3.750 -0.2103 0.01297 0.00464 -0.0457 0.9100 0.0531
-3.500 -0.1776 0.01256 0.00423 -0.0467 0.9011 0.0620
-3.250 -0.1442 0.01213 0.00386 -0.0480 0.8923 0.0857
-3.000 -0.1137 0.01173 0.00355 -0.0486 0.8818 0.1212
-2.750 -0.0844 0.01122 0.00326 -0.0491 0.8713 0.1838
-2.500 -0.0563 0.01074 0.00309 -0.0494 0.8610 0.2751
-2.250 -0.0294 0.01038 0.00294 -0.0493 0.8503 0.3447
-2.000 -0.0036 0.01012 0.00282 -0.0489 0.8396 0.4009
-1.750 0.0222 0.00990 0.00272 -0.0484 0.8294 0.4513
-1.500 0.0473 0.00967 0.00264 -0.0477 0.8197 0.5103
-1.250 0.0676 0.00925 0.00266 -0.0460 0.8091 0.6258
-1.000 0.0892 0.00885 0.00274 -0.0441 0.7990 0.7633
-0.750 0.1350 0.00870 0.00286 -0.0470 0.7907 0.8709
-0.500 0.1789 0.00874 0.00287 -0.0501 0.7803 0.9065
-0.250 0.2182 0.00880 0.00287 -0.0523 0.7697 0.9285
0.000 0.2548 0.00886 0.00286 -0.0540 0.7585 0.9461
0.250 0.2904 0.00893 0.00285 -0.0556 0.7472 0.9605
0.500 0.3257 0.00898 0.00285 -0.0571 0.7342 0.9725
0.750 0.3622 0.00903 0.00285 -0.0589 0.7209 0.9827
1.000 0.3996 0.00906 0.00283 -0.0610 0.7071 0.9914
1.250 0.4392 0.00909 0.00281 -0.0636 0.6929 0.9998
1.500 0.4624 0.00915 0.00281 -0.0626 0.6786 1.0000
1.750 0.4848 0.00923 0.00282 -0.0615 0.6629 1.0000
2.250 0.5296 0.00946 0.00290 -0.0591 0.6283 1.0000
2.500 0.5518 0.00961 0.00297 -0.0579 0.6091 1.0000
2.750 0.5741 0.00977 0.00307 -0.0567 0.5900 1.0000
3.000 0.5966 0.00993 0.00318 -0.0555 0.5728 1.0000
3.250 0.6192 0.01009 0.00331 -0.0544 0.5553 1.0000
3.500 0.6415 0.01025 0.00347 -0.0532 0.5351 1.0000
3.750 0.6633 0.01045 0.00361 -0.0520 0.5130 1.0000
4.250 0.7049 0.01095 0.00394 -0.0491 0.4509 1.0000
4.500 0.7252 0.01126 0.00417 -0.0476 0.4176 1.0000
4.750 0.7456 0.01159 0.00442 -0.0461 0.3857 1.0000
5.000 0.7652 0.01198 0.00471 -0.0446 0.3474 1.0000
5.250 0.7809 0.01265 0.00507 -0.0424 0.2804 1.0000
5.500 0.7946 0.01354 0.00559 -0.0401 0.2071 1.0000
5.750 0.8097 0.01440 0.00615 -0.0381 0.1498 1.0000
6.000 0.8258 0.01519 0.00673 -0.0362 0.1101 1.0000
6.250 0.8422 0.01598 0.00732 -0.0343 0.0650 1.0000
6.500 0.8575 0.01688 0.00803 -0.0323 0.0441 1.0000
6.750 0.8750 0.01758 0.00877 -0.0306 0.0376 1.0000
7.000 0.8918 0.01834 0.00962 -0.0287 0.0325 1.0000
7.250 0.9093 0.01904 0.01042 -0.0270 0.0291 1.0000
7.500 0.9241 0.01995 0.01139 -0.0250 0.0263 1.0000
7.750 0.9379 0.02092 0.01244 -0.0228 0.0241 1.0000
8.000 0.9545 0.02165 0.01329 -0.0210 0.0220 1.0000
8.250 0.9690 0.02254 0.01427 -0.0190 0.0204 1.0000
8.500 0.9826 0.02352 0.01535 -0.0169 0.0192 1.0000
8.750 0.9948 0.02460 0.01648 -0.0147 0.0180 1.0000
9.000 1.0030 0.02624 0.01815 -0.0120 0.0169 1.0000
9.250 1.0175 0.02705 0.01910 -0.0101 0.0159 1.0000
9.500 1.0299 0.02820 0.02038 -0.0080 0.0150 1.0000
9.750 1.0420 0.02951 0.02182 -0.0059 0.0143 1.0000
10.000 1.0539 0.03096 0.02340 -0.0040 0.0138 1.0000
10.250 1.0652 0.03243 0.02501 -0.0020 0.0133 1.0000
10.500 1.0747 0.03389 0.02659 -0.0001 0.0128 1.0000
10.750 1.0828 0.03534 0.02815 0.0019 0.0124 1.0000
11.000 1.0888 0.03712 0.03009 0.0038 0.0119 1.0000
11.250 1.0916 0.03973 0.03288 0.0057 0.0115 1.0000
11.500 1.0938 0.04152 0.03493 0.0077 0.0111 1.0000
11.750 1.0937 0.04396 0.03762 0.0094 0.0109 1.0000
12.000 1.0899 0.04671 0.04065 0.0110 0.0106 1.0000
12.250 1.0830 0.04988 0.04408 0.0122 0.0104 1.0000
12.500 1.0747 0.05329 0.04774 0.0129 0.0104 1.0000
12.750 1.0631 0.05724 0.05193 0.0130 0.0102 1.0000
13.000 1.0491 0.06171 0.05665 0.0124 0.0102 1.0000
13.250 1.0350 0.06647 0.06162 0.0111 0.0102 1.0000
13.500 1.0162 0.07234 0.06772 0.0087 0.0101 1.0000
13.750 0.9976 0.07863 0.07421 0.0057 0.0101 1.0000
14.000 0.9782 0.08560 0.08136 0.0018 0.0102 1.0000
14.250 0.9568 0.09370 0.08964 -0.0031 0.0104 1.0000
14.500 0.9324 0.10341 0.09951 -0.0092 0.0104 1.0000
14.750 0.9063 0.11475 0.11099 -0.0163 0.0106 1.0000
15.000 0.8798 0.12741 0.12373 -0.0237 0.0109 1.0000
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Polar data table (+)
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