RAF 69 AIRFOIL (raf69-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 69 AIRFOIL (raf69-il) Reynolds number: 100,000 Max Cl/Cd: 40.09 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf69-il-100000-n5.txt Download as CSV file: xf-raf69-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 69 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.750 -0.8093 0.08086 0.07371 -0.0754 1.0000 0.1050
-16.500 -0.7876 0.08120 0.07417 -0.0744 1.0000 0.1069
-16.250 -0.7942 0.07743 0.07036 -0.0757 1.0000 0.1089
-16.000 -0.8154 0.07177 0.06456 -0.0780 1.0000 0.1109
-15.750 -0.8377 0.06626 0.05888 -0.0800 1.0000 0.1130
-15.500 -0.8590 0.06124 0.05363 -0.0815 1.0000 0.1151
-15.250 -0.8629 0.05845 0.05080 -0.0817 1.0000 0.1171
-15.000 -0.8600 0.05658 0.04895 -0.0815 1.0000 0.1193
-14.750 -0.8630 0.05415 0.04649 -0.0814 1.0000 0.1217
-14.500 -0.8687 0.05152 0.04377 -0.0812 1.0000 0.1242
-14.250 -0.8760 0.04887 0.04098 -0.0808 1.0000 0.1268
-14.000 -0.8794 0.04671 0.03874 -0.0801 1.0000 0.1295
-13.750 -0.8777 0.04520 0.03729 -0.0790 1.0000 0.1320
-13.500 -0.8793 0.04354 0.03561 -0.0778 1.0000 0.1347
-13.250 -0.8826 0.04185 0.03386 -0.0762 1.0000 0.1378
-13.000 -0.8881 0.04019 0.03207 -0.0743 1.0000 0.1411
-12.750 -0.8904 0.03896 0.03087 -0.0720 1.0000 0.1439
-12.500 -0.8951 0.03792 0.02987 -0.0693 1.0000 0.1465
-12.250 -0.9045 0.03696 0.02891 -0.0657 1.0000 0.1493
-12.000 -0.9188 0.03614 0.02806 -0.0612 1.0000 0.1519
-11.750 -0.9083 0.03488 0.02664 -0.0612 0.9946 0.1569
-11.500 -0.8856 0.03380 0.02560 -0.0628 0.9857 0.1618
-11.250 -0.8629 0.03271 0.02441 -0.0644 0.9776 0.1679
-11.000 -0.8406 0.03169 0.02329 -0.0655 0.9689 0.1738
-10.750 -0.8189 0.03085 0.02244 -0.0663 0.9603 0.1796
-10.500 -0.7939 0.02995 0.02136 -0.0676 0.9531 0.1867
-10.250 -0.7719 0.02923 0.02066 -0.0679 0.9445 0.1924
-10.000 -0.7454 0.02848 0.01983 -0.0691 0.9379 0.1997
-9.750 -0.7240 0.02783 0.01910 -0.0690 0.9292 0.2062
-9.500 -0.6991 0.02722 0.01846 -0.0695 0.9218 0.2129
-9.250 -0.6797 0.02666 0.01774 -0.0688 0.9128 0.2202
-9.000 -0.6561 0.02613 0.01726 -0.0687 0.9051 0.2265
-8.750 -0.6314 0.02560 0.01663 -0.0689 0.8984 0.2343
-8.500 -0.6132 0.02519 0.01620 -0.0676 0.8887 0.2407
-8.250 -0.5861 0.02470 0.01567 -0.0680 0.8828 0.2486
-8.000 -0.5704 0.02438 0.01525 -0.0662 0.8733 0.2555
-7.750 -0.5469 0.02398 0.01488 -0.0658 0.8663 0.2625
-7.500 -0.5225 0.02358 0.01436 -0.0655 0.8601 0.2709
-7.250 -0.5064 0.02332 0.01414 -0.0636 0.8508 0.2772
-7.000 -0.4818 0.02295 0.01372 -0.0633 0.8446 0.2856
-6.750 -0.4616 0.02266 0.01340 -0.0621 0.8374 0.2932
-6.500 -0.4430 0.02241 0.01316 -0.0605 0.8293 0.3009
-6.250 -0.4172 0.02206 0.01273 -0.0603 0.8236 0.3101
-6.000 -0.3994 0.02185 0.01257 -0.0586 0.8158 0.3179
-5.750 -0.3792 0.02162 0.01225 -0.0573 0.8083 0.3276
-5.500 -0.3528 0.02128 0.01196 -0.0570 0.8029 0.3368
-5.000 -0.3150 0.02090 0.01160 -0.0538 0.7875 0.3564
-4.750 -0.2879 0.02060 0.01125 -0.0536 0.7824 0.3682
-4.500 -0.2731 0.02052 0.01121 -0.0512 0.7734 0.3785
-4.250 -0.2501 0.02030 0.01100 -0.0502 0.7668 0.3904
-3.750 -0.2086 0.02001 0.01074 -0.0475 0.7528 0.4154
-3.500 -0.1847 0.01982 0.01055 -0.0467 0.7466 0.4289
-3.250 -0.1590 0.01964 0.01034 -0.0461 0.7413 0.4438
-3.000 -0.1431 0.01962 0.01040 -0.0439 0.7328 0.4567
-2.750 -0.1185 0.01947 0.01026 -0.0431 0.7267 0.4708
-2.500 -0.0928 0.01932 0.01010 -0.0425 0.7213 0.4861
-2.250 -0.0764 0.01934 0.01015 -0.0404 0.7129 0.5002
-2.000 -0.0511 0.01921 0.01006 -0.0397 0.7068 0.5149
-1.750 -0.0254 0.01911 0.00997 -0.0390 0.7012 0.5298
-1.500 -0.0083 0.01914 0.01004 -0.0370 0.6929 0.5444
-1.250 0.0176 0.01903 0.00994 -0.0364 0.6868 0.5598
-1.000 0.0420 0.01899 0.00994 -0.0355 0.6805 0.5744
-0.750 0.0610 0.01903 0.01004 -0.0337 0.6725 0.5888
-0.500 0.0879 0.01894 0.00994 -0.0332 0.6664 0.6046
-0.250 0.1091 0.01898 0.01001 -0.0318 0.6593 0.6200
0.000 0.1308 0.01902 0.01012 -0.0304 0.6515 0.6345
0.250 0.1596 0.01896 0.01006 -0.0302 0.6458 0.6497
0.500 0.1770 0.01907 0.01022 -0.0281 0.6375 0.6647
0.750 0.2006 0.01910 0.01026 -0.0271 0.6302 0.6797
1.000 0.2298 0.01908 0.01023 -0.0269 0.6244 0.6937
1.250 0.2455 0.01924 0.01046 -0.0246 0.6153 0.7076
1.500 0.2710 0.01922 0.01040 -0.0238 0.6086 0.7226
1.750 0.2926 0.01932 0.01055 -0.0224 0.6008 0.7356
2.000 0.3145 0.01939 0.01064 -0.0211 0.5928 0.7487
2.250 0.3431 0.01934 0.01053 -0.0209 0.5868 0.7629
2.500 0.3582 0.01952 0.01079 -0.0185 0.5772 0.7760
2.750 0.3863 0.01953 0.01079 -0.0182 0.5702 0.7884
3.000 0.4053 0.01964 0.01092 -0.0165 0.5619 0.8020
3.250 0.4306 0.01970 0.01101 -0.0158 0.5536 0.8145
3.500 0.4582 0.01978 0.01108 -0.0156 0.5460 0.8264
3.750 0.4779 0.01990 0.01124 -0.0140 0.5370 0.8401
4.000 0.5127 0.01996 0.01127 -0.0151 0.5294 0.8507
4.250 0.5342 0.02018 0.01156 -0.0140 0.5197 0.8635
4.500 0.5694 0.02025 0.01157 -0.0152 0.5122 0.8747
4.750 0.5951 0.02056 0.01197 -0.0151 0.5025 0.8863
5.000 0.6268 0.02068 0.01203 -0.0158 0.4947 0.8985
5.250 0.6599 0.02104 0.01246 -0.0171 0.4850 0.9081
5.500 0.6938 0.02122 0.01260 -0.0183 0.4768 0.9189
5.750 0.7268 0.02159 0.01301 -0.0198 0.4673 0.9291
6.000 0.7635 0.02185 0.01323 -0.0217 0.4585 0.9383
6.250 0.7945 0.02223 0.01365 -0.0228 0.4492 0.9495
6.500 0.8323 0.02254 0.01392 -0.0251 0.4400 0.9574
6.750 0.8631 0.02295 0.01437 -0.0263 0.4305 0.9687
7.000 0.8967 0.02326 0.01464 -0.0280 0.4211 0.9795
7.250 0.9282 0.02369 0.01513 -0.0296 0.4108 0.9909
7.500 0.9629 0.02402 0.01538 -0.0316 0.4014 1.0000
7.750 0.9428 0.02418 0.01556 -0.0239 0.3953 1.0000
8.000 0.9360 0.02444 0.01579 -0.0184 0.3881 1.0000
8.250 0.9412 0.02472 0.01598 -0.0148 0.3810 1.0000
8.500 0.9382 0.02524 0.01654 -0.0103 0.3728 1.0000
8.750 0.9463 0.02561 0.01682 -0.0074 0.3651 1.0000
9.000 0.9474 0.02626 0.01751 -0.0038 0.3568 1.0000
9.250 0.9542 0.02681 0.01802 -0.0011 0.3489 1.0000
9.500 0.9596 0.02749 0.01871 0.0017 0.3409 1.0000
9.750 0.9650 0.02821 0.01943 0.0043 0.3327 1.0000
10.000 0.9733 0.02891 0.02010 0.0065 0.3251 1.0000
10.250 0.9778 0.02982 0.02105 0.0090 0.3169 1.0000
10.500 0.9881 0.03052 0.02168 0.0108 0.3097 1.0000
10.750 0.9913 0.03165 0.02288 0.0130 0.3015 1.0000
11.000 1.0007 0.03246 0.02363 0.0147 0.2943 1.0000
11.250 1.0051 0.03366 0.02490 0.0165 0.2867 1.0000
11.500 1.0121 0.03472 0.02594 0.0181 0.2795 1.0000
11.750 1.0192 0.03587 0.02710 0.0196 0.2726 1.0000
12.000 1.0237 0.03720 0.02848 0.0211 0.2654 1.0000
12.250 1.0339 0.03818 0.02938 0.0223 0.2593 1.0000
12.500 1.0360 0.03984 0.03114 0.0237 0.2523 1.0000
12.750 1.0437 0.04108 0.03237 0.0248 0.2463 1.0000
13.000 1.0496 0.04253 0.03384 0.0259 0.2402 1.0000
13.250 1.0530 0.04423 0.03561 0.0269 0.2341 1.0000
13.500 1.0630 0.04538 0.03671 0.0277 0.2288 1.0000
13.750 1.0652 0.04731 0.03873 0.0286 0.2231 1.0000
14.000 1.0689 0.04914 0.04062 0.0293 0.2176 1.0000
14.250 1.0800 0.05029 0.04172 0.0300 0.2129 1.0000
14.500 1.0813 0.05246 0.04399 0.0305 0.2080 1.0000
14.750 1.0830 0.05464 0.04626 0.0310 0.2030 1.0000
15.000 1.0912 0.05616 0.04776 0.0315 0.1986 1.0000
15.250 1.0980 0.05790 0.04952 0.0319 0.1945 1.0000
15.500 1.0945 0.06079 0.05256 0.0320 0.1902 1.0000
15.750 1.0975 0.06302 0.05487 0.0322 0.1862 1.0000
16.000 1.1077 0.06443 0.05626 0.0324 0.1828 1.0000
16.250 1.1135 0.06641 0.05828 0.0326 0.1794 1.0000
16.500 1.1022 0.07045 0.06252 0.0321 0.1755 1.0000
16.750 1.0988 0.07363 0.06580 0.0318 0.1719 1.0000
17.000 1.1049 0.07564 0.06784 0.0317 0.1687 1.0000
17.250 1.1254 0.07582 0.06793 0.0322 0.1656 1.0000
17.500 1.0964 0.08244 0.07484 0.0305 0.1623 1.0000
17.750 1.0712 0.08887 0.08147 0.0287 0.1587 1.0000
18.000 1.0642 0.09286 0.08554 0.0275 0.1551 1.0000
18.250 1.0902 0.09202 0.08459 0.0282 0.1515 1.0000
18.500 1.0530 0.10057 0.09337 0.0252 0.1483 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 69 AIRFOIL (raf69-il)