USA 28 AIRFOIL (usa28-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 28 AIRFOIL (usa28-il) Reynolds number: 100,000 Max Cl/Cd: 54.23 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa28-il-100000.txt Download as CSV file: xf-usa28-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: USA 28 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3178 0.10709 0.10252 -0.0368 1.0000 0.1635
-9.500 -0.3526 0.10523 0.10079 -0.0372 1.0000 0.1641
-8.750 -0.5964 0.07053 0.06602 -0.0520 1.0000 0.0945
-8.500 -0.6110 0.06704 0.06252 -0.0499 1.0000 0.0930
-8.250 -0.6278 0.06324 0.05870 -0.0478 1.0000 0.0920
-8.000 -0.6515 0.05807 0.05338 -0.0462 1.0000 0.0906
-7.750 -0.6746 0.05223 0.04725 -0.0446 1.0000 0.0889
-7.500 -0.6901 0.04701 0.04158 -0.0424 1.0000 0.0885
-7.250 -0.6957 0.04306 0.03717 -0.0400 1.0000 0.0889
-7.000 -0.6954 0.03988 0.03352 -0.0375 1.0000 0.0899
-6.750 -0.6908 0.03718 0.03035 -0.0350 0.9999 0.0907
-6.500 -0.6616 0.03431 0.02676 -0.0367 0.9940 0.0925
-6.250 -0.6328 0.03221 0.02426 -0.0379 0.9876 0.0955
-6.000 -0.6002 0.03109 0.02306 -0.0396 0.9817 0.0993
-5.750 -0.5685 0.02980 0.02148 -0.0407 0.9754 0.1027
-5.500 -0.5379 0.02864 0.01998 -0.0415 0.9688 0.1073
-5.250 -0.5041 0.02803 0.01946 -0.0432 0.9627 0.1145
-5.000 -0.4759 0.02726 0.01850 -0.0435 0.9558 0.1216
-4.750 -0.4432 0.02674 0.01794 -0.0447 0.9503 0.1318
-4.500 -0.4172 0.02627 0.01743 -0.0446 0.9438 0.1477
-4.250 -0.3881 0.02555 0.01685 -0.0452 0.9378 0.1775
-4.000 -0.3624 0.02507 0.01658 -0.0451 0.9309 0.2237
-3.750 -0.3303 0.02525 0.01680 -0.0461 0.9236 0.2585
-3.500 -0.3032 0.02560 0.01719 -0.0460 0.9156 0.2792
-3.250 -0.2688 0.02593 0.01745 -0.0473 0.9088 0.2989
-3.000 -0.2420 0.02628 0.01776 -0.0471 0.9010 0.3154
-2.750 -0.2089 0.02656 0.01793 -0.0480 0.8940 0.3343
-2.500 -0.1803 0.02686 0.01822 -0.0482 0.8869 0.3507
-2.250 -0.1502 0.02689 0.01819 -0.0485 0.8792 0.3666
-2.000 -0.1044 0.02650 0.01776 -0.0516 0.8750 0.3837
-1.750 -0.0903 0.02653 0.01775 -0.0492 0.8648 0.3961
-1.500 -0.0507 0.02613 0.01732 -0.0512 0.8605 0.4160
-1.250 -0.0340 0.02619 0.01735 -0.0493 0.8519 0.4317
-1.000 0.0016 0.02580 0.01696 -0.0506 0.8463 0.4513
-0.750 0.0477 0.02521 0.01633 -0.0537 0.8431 0.4725
-0.500 0.0573 0.02544 0.01658 -0.0507 0.8325 0.4845
-0.250 0.1026 0.02485 0.01598 -0.0536 0.8286 0.5036
0.000 0.1219 0.02492 0.01608 -0.0522 0.8191 0.5164
0.250 0.1661 0.02429 0.01549 -0.0549 0.8141 0.5344
0.500 0.2220 0.02340 0.01464 -0.0596 0.8115 0.5553
0.750 0.2424 0.02357 0.01491 -0.0585 0.8003 0.5692
1.000 0.3080 0.02265 0.01416 -0.0653 0.7974 0.5960
1.250 0.3785 0.02154 0.01338 -0.0729 0.7951 0.6433
1.500 0.5910 0.01944 0.01226 -0.1100 0.7962 1.0000
1.750 0.6125 0.01964 0.01237 -0.1087 0.7843 1.0000
2.000 0.6534 0.01939 0.01203 -0.1108 0.7756 1.0000
2.250 0.6874 0.01924 0.01180 -0.1115 0.7647 1.0000
2.500 0.7114 0.01930 0.01181 -0.1104 0.7517 1.0000
2.750 0.7399 0.01926 0.01173 -0.1100 0.7393 1.0000
3.000 0.7740 0.01910 0.01151 -0.1107 0.7278 1.0000
3.250 0.8102 0.01889 0.01125 -0.1117 0.7159 1.0000
3.500 0.8336 0.01899 0.01132 -0.1105 0.7018 1.0000
3.750 0.8576 0.01913 0.01143 -0.1094 0.6879 1.0000
4.000 0.8820 0.01931 0.01160 -0.1084 0.6745 1.0000
4.250 0.9076 0.01951 0.01178 -0.1077 0.6620 1.0000
4.500 0.9384 0.01966 0.01188 -0.1079 0.6510 1.0000
4.750 0.9623 0.01996 0.01219 -0.1069 0.6391 1.0000
5.000 0.9806 0.02041 0.01269 -0.1050 0.6272 1.0000
5.250 1.0052 0.02074 0.01302 -0.1042 0.6167 1.0000
5.500 1.0325 0.02097 0.01324 -0.1038 0.6061 1.0000
5.750 1.0473 0.02142 0.01379 -0.1011 0.5943 1.0000
6.000 1.0686 0.02174 0.01416 -0.0997 0.5835 1.0000
6.250 1.0962 0.02191 0.01434 -0.0993 0.5735 1.0000
6.500 1.1084 0.02242 0.01499 -0.0963 0.5627 1.0000
6.750 1.1304 0.02272 0.01537 -0.0949 0.5531 1.0000
7.000 1.1551 0.02283 0.01551 -0.0939 0.5423 1.0000
7.250 1.1731 0.02270 0.01541 -0.0914 0.5266 1.0000
7.500 1.1836 0.02251 0.01524 -0.0873 0.5073 1.0000
7.750 1.1938 0.02226 0.01494 -0.0832 0.4849 1.0000
8.000 1.2060 0.02224 0.01480 -0.0796 0.4627 1.0000
8.250 1.2128 0.02251 0.01504 -0.0752 0.4408 1.0000
8.500 1.2222 0.02287 0.01535 -0.0715 0.4201 1.0000
8.750 1.2274 0.02330 0.01578 -0.0671 0.3997 1.0000
9.000 1.2237 0.02373 0.01632 -0.0610 0.3787 1.0000
9.250 1.2156 0.02412 0.01675 -0.0543 0.3557 1.0000
9.500 1.2021 0.02468 0.01734 -0.0469 0.3227 1.0000
9.750 1.1833 0.02579 0.01821 -0.0395 0.2620 1.0000
10.000 1.1576 0.02805 0.01987 -0.0324 0.1904 1.0000
10.250 1.1316 0.03115 0.02241 -0.0265 0.1308 1.0000
10.500 1.1130 0.03412 0.02508 -0.0220 0.1033 1.0000
10.750 1.1009 0.03685 0.02769 -0.0185 0.0910 1.0000
11.000 1.0939 0.03933 0.03017 -0.0158 0.0826 1.0000
11.250 1.0892 0.04175 0.03260 -0.0134 0.0773 1.0000
11.500 1.0872 0.04405 0.03492 -0.0114 0.0714 1.0000
11.750 1.0876 0.04621 0.03699 -0.0093 0.0680 1.0000
12.000 1.0971 0.04776 0.03863 -0.0075 0.0647 1.0000
12.250 1.1100 0.04915 0.04004 -0.0059 0.0617 1.0000
12.500 1.1366 0.05026 0.04088 -0.0044 0.0574 1.0000
12.750 1.1501 0.05183 0.04266 -0.0030 0.0551 1.0000
13.000 1.1743 0.05329 0.04423 -0.0019 0.0531 1.0000
13.250 1.1994 0.05503 0.04609 -0.0011 0.0517 1.0000
13.500 1.2227 0.05718 0.04840 -0.0002 0.0505 1.0000
13.750 1.2428 0.05966 0.05099 0.0005 0.0493 1.0000
14.000 1.2693 0.06382 0.05525 0.0004 0.0478 1.0000
14.250 1.2624 0.06674 0.05845 0.0024 0.0476 1.0000
14.500 1.2483 0.06976 0.06178 0.0043 0.0472 1.0000
14.750 1.2419 0.07331 0.06558 0.0056 0.0473 1.0000
15.000 1.2292 0.07722 0.06975 0.0067 0.0473 1.0000
15.250 1.2178 0.08143 0.07420 0.0075 0.0475 1.0000
15.500 1.2007 0.08607 0.07908 0.0078 0.0476 1.0000
15.750 1.1909 0.09073 0.08395 0.0078 0.0479 1.0000
16.000 1.1682 0.09493 0.08839 0.0073 0.0482 1.0000
16.250 1.1386 0.10045 0.09420 0.0056 0.0487 1.0000
16.500 1.0858 0.10952 0.10368 0.0007 0.0498 1.0000
16.750 1.0095 0.12602 0.12061 -0.0100 0.0520 1.0000
17.000 0.9564 0.14217 0.13691 -0.0199 0.0555 1.0000
17.250 0.9374 0.15176 0.14651 -0.0248 0.0568 1.0000
17.500 0.9321 0.15846 0.15324 -0.0275 0.0578 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 28 AIRFOIL (usa28-il)