USA 35 AIRFOIL (usa35-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 35 AIRFOIL (usa35-il) Reynolds number: 500,000 Max Cl/Cd: 92.19 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa35-il-500000-n5.txt Download as CSV file: xf-usa35-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 35 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
------ -------- --------- --------- -------- -------- --------
-13.000 -0.6091 0.03379 0.02941 -0.1352 0.8598 0.0306
-12.750 -0.6156 0.03088 0.02622 -0.1359 0.8460 0.0308
-12.500 -0.6087 0.02887 0.02397 -0.1360 0.8353 0.0309
-12.250 -0.5968 0.02725 0.02214 -0.1359 0.8250 0.0311
-12.000 -0.5816 0.02592 0.02059 -0.1355 0.8157 0.0313
-11.750 -0.5638 0.02477 0.01925 -0.1352 0.8065 0.0315
-11.500 -0.5444 0.02385 0.01822 -0.1348 0.7976 0.0317
-11.250 -0.5229 0.02307 0.01737 -0.1345 0.7900 0.0319
-11.000 -0.5004 0.02240 0.01662 -0.1342 0.7817 0.0322
-10.750 -0.4775 0.02177 0.01589 -0.1338 0.7740 0.0324
-10.500 -0.4537 0.02114 0.01518 -0.1336 0.7663 0.0326
-10.250 -0.4299 0.02054 0.01448 -0.1332 0.7578 0.0329
-10.000 -0.4056 0.01996 0.01379 -0.1329 0.7500 0.0332
-9.750 -0.3809 0.01938 0.01312 -0.1326 0.7414 0.0335
-9.500 -0.3562 0.01885 0.01246 -0.1323 0.7324 0.0338
-9.250 -0.3309 0.01832 0.01182 -0.1320 0.7233 0.0341
-9.000 -0.3057 0.01783 0.01120 -0.1317 0.7127 0.0345
-8.750 -0.2799 0.01736 0.01062 -0.1314 0.7022 0.0348
-8.500 -0.2542 0.01695 0.01006 -0.1310 0.6907 0.0351
-8.250 -0.2282 0.01652 0.00957 -0.1308 0.6798 0.0355
-8.000 -0.2022 0.01620 0.00918 -0.1304 0.6682 0.0359
-7.750 -0.1754 0.01591 0.00882 -0.1302 0.6560 0.0363
-7.500 -0.1489 0.01565 0.00847 -0.1299 0.6441 0.0367
-7.250 -0.1222 0.01540 0.00813 -0.1296 0.6314 0.0372
-7.000 -0.0954 0.01514 0.00778 -0.1294 0.6198 0.0378
-6.500 -0.0416 0.01469 0.00711 -0.1288 0.5974 0.0390
-6.250 -0.0150 0.01449 0.00685 -0.1285 0.5871 0.0396
-6.000 0.0126 0.01431 0.00664 -0.1284 0.5781 0.0403
-5.750 0.0399 0.01417 0.00643 -0.1282 0.5702 0.0410
-5.500 0.0675 0.01400 0.00620 -0.1280 0.5640 0.0419
-5.250 0.0953 0.01382 0.00596 -0.1279 0.5582 0.0427
-5.000 0.1229 0.01366 0.00574 -0.1277 0.5524 0.0435
-4.750 0.1502 0.01355 0.00559 -0.1275 0.5463 0.0444
-4.500 0.1784 0.01341 0.00544 -0.1274 0.5410 0.0454
-4.250 0.2063 0.01330 0.00528 -0.1273 0.5351 0.0465
-4.000 0.2338 0.01322 0.00512 -0.1271 0.5291 0.0477
-3.750 0.2614 0.01312 0.00501 -0.1269 0.5237 0.0489
-3.500 0.2898 0.01303 0.00490 -0.1269 0.5189 0.0504
-3.250 0.3179 0.01295 0.00479 -0.1268 0.5145 0.0519
-3.000 0.3458 0.01286 0.00468 -0.1267 0.5108 0.0533
-2.750 0.3736 0.01281 0.00460 -0.1265 0.5070 0.0548
-2.500 0.4015 0.01276 0.00452 -0.1264 0.5035 0.0564
-2.250 0.4301 0.01267 0.00444 -0.1264 0.5007 0.0582
-2.000 0.4584 0.01260 0.00439 -0.1264 0.4977 0.0602
-1.750 0.4867 0.01256 0.00433 -0.1264 0.4946 0.0623
-1.500 0.5147 0.01251 0.00428 -0.1263 0.4913 0.0643
-1.250 0.5424 0.01249 0.00425 -0.1262 0.4879 0.0667
-1.000 0.5699 0.01251 0.00423 -0.1260 0.4843 0.0692
-0.750 0.5982 0.01245 0.00420 -0.1260 0.4814 0.0721
-0.500 0.6266 0.01242 0.00418 -0.1260 0.4784 0.0752
-0.250 0.6546 0.01238 0.00417 -0.1259 0.4751 0.0788
0.000 0.6824 0.01237 0.00416 -0.1258 0.4717 0.0828
0.250 0.7098 0.01237 0.00417 -0.1257 0.4683 0.0881
0.500 0.7369 0.01239 0.00419 -0.1255 0.4648 0.0948
0.750 0.7647 0.01236 0.00421 -0.1254 0.4617 0.1058
1.250 0.8197 0.01221 0.00431 -0.1253 0.4536 0.1773
1.500 0.8464 0.01219 0.00438 -0.1251 0.4492 0.2170
1.750 0.8725 0.01220 0.00446 -0.1249 0.4451 0.2568
2.000 0.9000 0.01212 0.00455 -0.1249 0.4413 0.3137
2.250 0.9268 0.01200 0.00467 -0.1248 0.4367 0.4000
2.500 0.9527 0.01187 0.00481 -0.1246 0.4317 0.5093
2.750 0.9762 0.01159 0.00500 -0.1239 0.4268 0.6868
3.000 0.9929 0.01118 0.00521 -0.1210 0.4222 0.9309
3.250 1.0259 0.01128 0.00529 -0.1221 0.4148 1.0000
3.500 1.0496 0.01149 0.00541 -0.1214 0.4074 1.0000
3.750 1.0739 0.01167 0.00554 -0.1207 0.3984 1.0000
4.000 1.0963 0.01191 0.00570 -0.1198 0.3899 1.0000
4.250 1.1192 0.01214 0.00587 -0.1189 0.3812 1.0000
4.500 1.1391 0.01241 0.00607 -0.1175 0.3733 1.0000
4.750 1.1603 0.01267 0.00628 -0.1164 0.3656 1.0000
5.250 1.2013 0.01328 0.00679 -0.1140 0.3520 1.0000
5.500 1.2216 0.01362 0.00708 -0.1128 0.3456 1.0000
5.750 1.2412 0.01400 0.00741 -0.1115 0.3400 1.0000
6.000 1.2625 0.01432 0.00770 -0.1106 0.3346 1.0000
6.250 1.2821 0.01472 0.00807 -0.1094 0.3288 1.0000
6.500 1.3005 0.01518 0.00848 -0.1081 0.3239 1.0000
6.750 1.3222 0.01550 0.00881 -0.1073 0.3205 1.0000
7.000 1.3424 0.01590 0.00920 -0.1064 0.3162 1.0000
7.250 1.3609 0.01638 0.00966 -0.1052 0.3114 1.0000
7.500 1.3778 0.01695 0.01019 -0.1038 0.3069 1.0000
7.750 1.3986 0.01735 0.01060 -0.1030 0.3036 1.0000
8.000 1.4183 0.01780 0.01107 -0.1021 0.3004 1.0000
8.250 1.4368 0.01833 0.01160 -0.1011 0.2971 1.0000
8.500 1.4542 0.01893 0.01219 -0.1000 0.2938 1.0000
8.750 1.4702 0.01961 0.01286 -0.0988 0.2906 1.0000
9.000 1.4892 0.02014 0.01342 -0.0979 0.2880 1.0000
9.250 1.5078 0.02071 0.01401 -0.0971 0.2850 1.0000
9.500 1.5249 0.02136 0.01469 -0.0961 0.2818 1.0000
9.750 1.5406 0.02213 0.01546 -0.0950 0.2783 1.0000
10.000 1.5541 0.02304 0.01637 -0.0938 0.2748 1.0000
10.250 1.5705 0.02378 0.01715 -0.0929 0.2716 1.0000
10.500 1.5865 0.02457 0.01797 -0.0920 0.2680 1.0000
10.750 1.6009 0.02548 0.01890 -0.0910 0.2640 1.0000
11.000 1.6126 0.02660 0.02002 -0.0898 0.2602 1.0000
11.250 1.6252 0.02766 0.02111 -0.0887 0.2568 1.0000
11.500 1.6393 0.02865 0.02214 -0.0879 0.2529 1.0000
11.750 1.6509 0.02983 0.02335 -0.0868 0.2485 1.0000
12.000 1.6587 0.03134 0.02487 -0.0856 0.2440 1.0000
12.250 1.6703 0.03260 0.02616 -0.0847 0.2397 1.0000
12.500 1.6795 0.03406 0.02765 -0.0838 0.2344 1.0000
12.750 1.6837 0.03599 0.02959 -0.0826 0.2291 1.0000
13.000 1.6926 0.03755 0.03119 -0.0817 0.2236 1.0000
13.250 1.6956 0.03969 0.03334 -0.0806 0.2172 1.0000
13.500 1.6986 0.04187 0.03554 -0.0796 0.2109 1.0000
13.750 1.6986 0.04438 0.03806 -0.0786 0.2040 1.0000
14.000 1.6984 0.04698 0.04068 -0.0777 0.1975 1.0000
14.250 1.6955 0.04991 0.04363 -0.0768 0.1913 1.0000
14.500 1.6928 0.05287 0.04662 -0.0761 0.1862 1.0000
14.750 1.6899 0.05590 0.04967 -0.0754 0.1813 1.0000
15.000 1.6835 0.05937 0.05317 -0.0747 0.1769 1.0000
15.250 1.6842 0.06208 0.05593 -0.0743 0.1736 1.0000
15.500 1.6817 0.06518 0.05907 -0.0738 0.1703 1.0000
15.750 1.6776 0.06847 0.06240 -0.0734 0.1674 1.0000
16.000 1.6732 0.07183 0.06580 -0.0731 0.1648 1.0000
16.250 1.6747 0.07455 0.06858 -0.0729 0.1628 1.0000
16.500 1.6755 0.07734 0.07142 -0.0727 0.1606 1.0000
16.750 1.6743 0.08036 0.07450 -0.0726 0.1585 1.0000
17.000 1.6716 0.08359 0.07778 -0.0725 0.1564 1.0000
17.250 1.6680 0.08694 0.08116 -0.0724 0.1543 1.0000
17.500 1.6669 0.09001 0.08428 -0.0725 0.1525 1.0000
17.750 1.6705 0.09251 0.08686 -0.0726 0.1509 1.0000
18.000 1.6714 0.09535 0.08976 -0.0727 0.1493 1.0000
18.250 1.6720 0.09822 0.09270 -0.0729 0.1476 1.0000
18.500 1.6722 0.10114 0.09566 -0.0731 0.1459 1.0000
18.750 1.6698 0.10445 0.09902 -0.0734 0.1442 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 35 AIRFOIL (usa35-il)