USA 26 AIRFOIL (usa26-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 26 AIRFOIL (usa26-il) Reynolds number: 100,000 Max Cl/Cd: 52.3 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa26-il-100000-n5.txt Download as CSV file: xf-usa26-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 26 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3659 0.09269 0.08814 -0.0261 1.0000 0.0615
-7.500 -0.3796 0.09092 0.08646 -0.0244 1.0000 0.0630
-7.250 -0.3925 0.08883 0.08445 -0.0231 1.0000 0.0638
-6.750 -0.4186 0.08366 0.07929 -0.0276 1.0000 0.0675
-6.500 -0.3963 0.07801 0.07346 -0.0349 0.9942 0.0678
-6.250 -0.3729 0.07237 0.06765 -0.0402 0.9877 0.0678
-5.750 -0.3298 0.06014 0.05521 -0.0455 0.9764 0.0534
-5.500 -0.3073 0.05593 0.05087 -0.0480 0.9691 0.0521
-5.250 -0.2789 0.05088 0.04557 -0.0520 0.9628 0.0511
-5.000 -0.2566 0.04555 0.03983 -0.0543 0.9541 0.0523
-4.750 -0.2282 0.04084 0.03466 -0.0568 0.9488 0.0524
-4.500 -0.2086 0.03748 0.03092 -0.0566 0.9397 0.0521
-4.250 -0.1774 0.03415 0.02707 -0.0582 0.9343 0.0521
-4.000 -0.1548 0.03176 0.02424 -0.0575 0.9245 0.0524
-3.750 -0.1182 0.02989 0.02173 -0.0591 0.9184 0.0542
-3.500 -0.0934 0.02826 0.01975 -0.0585 0.9076 0.0547
-3.250 -0.0637 0.02643 0.01759 -0.0587 0.8987 0.0549
-3.000 -0.0304 0.02475 0.01565 -0.0596 0.8909 0.0553
-2.750 -0.0027 0.02343 0.01415 -0.0593 0.8806 0.0559
-2.500 0.0316 0.02216 0.01273 -0.0602 0.8728 0.0568
-2.250 0.0623 0.02116 0.01161 -0.0604 0.8622 0.0577
-2.000 0.0921 0.02033 0.01071 -0.0604 0.8501 0.0597
-1.750 0.1242 0.01959 0.00988 -0.0609 0.8374 0.0629
-1.500 0.1581 0.01885 0.00906 -0.0617 0.8239 0.0650
-1.250 0.1924 0.01816 0.00830 -0.0625 0.8087 0.0663
-1.000 0.2268 0.01756 0.00761 -0.0633 0.7911 0.0679
-0.750 0.2620 0.01683 0.00683 -0.0643 0.7710 0.0711
-0.500 0.2948 0.01638 0.00628 -0.0649 0.7454 0.0755
-0.250 0.3290 0.01603 0.00577 -0.0656 0.7164 0.0821
0.250 0.3930 0.01513 0.00535 -0.0669 0.6521 0.2757
0.500 0.4210 0.01506 0.00522 -0.0667 0.6215 0.3423
0.750 0.4444 0.01468 0.00505 -0.0659 0.5959 0.4504
1.000 0.5813 0.01425 0.00529 -0.0882 0.5582 1.0000
1.250 0.6041 0.01450 0.00531 -0.0871 0.5422 1.0000
1.500 0.6269 0.01474 0.00538 -0.0860 0.5284 1.0000
1.750 0.6497 0.01499 0.00547 -0.0850 0.5160 1.0000
2.000 0.6725 0.01525 0.00559 -0.0839 0.5049 1.0000
2.250 0.6952 0.01553 0.00572 -0.0829 0.4946 1.0000
2.500 0.7177 0.01577 0.00589 -0.0818 0.4844 1.0000
2.750 0.7406 0.01605 0.00607 -0.0808 0.4761 1.0000
3.000 0.7632 0.01631 0.00628 -0.0798 0.4676 1.0000
3.250 0.7863 0.01660 0.00648 -0.0788 0.4608 1.0000
3.500 0.8090 0.01687 0.00675 -0.0779 0.4534 1.0000
3.750 0.8322 0.01718 0.00699 -0.0770 0.4475 1.0000
4.000 0.8546 0.01746 0.00729 -0.0759 0.4405 1.0000
4.250 0.8773 0.01777 0.00758 -0.0749 0.4343 1.0000
4.500 0.8997 0.01808 0.00789 -0.0739 0.4279 1.0000
4.750 0.9215 0.01837 0.00820 -0.0728 0.4208 1.0000
5.000 0.9432 0.01868 0.00849 -0.0716 0.4140 1.0000
5.250 0.9639 0.01896 0.00883 -0.0703 0.4061 1.0000
5.500 0.9852 0.01927 0.00913 -0.0690 0.3994 1.0000
5.750 1.0057 0.01957 0.00951 -0.0677 0.3918 1.0000
6.000 1.0269 0.01989 0.00983 -0.0664 0.3855 1.0000
6.250 1.0466 0.02020 0.01026 -0.0649 0.3779 1.0000
6.500 1.0676 0.02053 0.01061 -0.0637 0.3719 1.0000
6.750 1.0880 0.02089 0.01110 -0.0624 0.3658 1.0000
7.000 1.1084 0.02125 0.01158 -0.0610 0.3599 1.0000
7.250 1.1291 0.02162 0.01200 -0.0598 0.3543 1.0000
7.500 1.1462 0.02194 0.01246 -0.0578 0.3450 1.0000
7.750 1.1619 0.02222 0.01282 -0.0556 0.3341 1.0000
8.000 1.1752 0.02247 0.01311 -0.0530 0.3206 1.0000
8.250 1.1878 0.02274 0.01348 -0.0503 0.3060 1.0000
8.500 1.2010 0.02308 0.01388 -0.0477 0.2924 1.0000
8.750 1.2143 0.02345 0.01435 -0.0452 0.2794 1.0000
9.000 1.2272 0.02385 0.01488 -0.0427 0.2646 1.0000
9.250 1.2388 0.02430 0.01544 -0.0400 0.2444 1.0000
9.500 1.2441 0.02488 0.01595 -0.0364 0.2147 1.0000
9.750 1.2479 0.02566 0.01666 -0.0326 0.1913 1.0000
10.000 1.2504 0.02666 0.01756 -0.0289 0.1723 1.0000
10.250 1.2492 0.02800 0.01878 -0.0251 0.1513 1.0000
10.500 1.2479 0.02949 0.02019 -0.0215 0.1263 1.0000
10.750 1.2403 0.03147 0.02195 -0.0177 0.0786 1.0000
11.000 1.2212 0.03434 0.02447 -0.0133 0.0464 1.0000
11.250 1.2115 0.03678 0.02690 -0.0103 0.0346 1.0000
11.500 1.2067 0.03904 0.02924 -0.0080 0.0311 1.0000
11.750 1.1998 0.04168 0.03198 -0.0062 0.0285 1.0000
12.000 1.1925 0.04457 0.03500 -0.0048 0.0270 1.0000
12.250 1.1850 0.04776 0.03832 -0.0040 0.0261 1.0000
12.500 1.1777 0.05118 0.04192 -0.0037 0.0256 1.0000
12.750 1.1680 0.05518 0.04609 -0.0040 0.0248 1.0000
13.000 1.1593 0.05933 0.05042 -0.0047 0.0244 1.0000
13.250 1.1484 0.06401 0.05525 -0.0057 0.0239 1.0000
13.500 1.1374 0.06886 0.06027 -0.0070 0.0236 1.0000
13.750 1.1262 0.07383 0.06541 -0.0084 0.0230 1.0000
14.000 1.1148 0.07894 0.07067 -0.0099 0.0227 1.0000
14.250 1.1038 0.08405 0.07591 -0.0115 0.0222 1.0000
14.500 1.0920 0.08940 0.08139 -0.0132 0.0217 1.0000
14.750 1.0823 0.09449 0.08659 -0.0148 0.0214 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 26 AIRFOIL (usa26-il)