NPL 9615 AIRFOIL (npl9615-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NPL 9615 AIRFOIL (npl9615-il) Reynolds number: 100,000 Max Cl/Cd: 42.06 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-npl9615-il-100000-n5.txt Download as CSV file: xf-npl9615-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NPL 9615 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5988 0.08757 0.08204 -0.0250 1.0000 0.0312
-10.000 -0.6157 0.07903 0.07348 -0.0317 1.0000 0.0307
-9.750 -0.6392 0.07212 0.06648 -0.0360 1.0000 0.0302
-9.500 -0.6630 0.06701 0.06126 -0.0366 1.0000 0.0299
-9.250 -0.6838 0.06266 0.05675 -0.0350 1.0000 0.0296
-9.000 -0.6987 0.05823 0.05209 -0.0331 1.0000 0.0294
-8.750 -0.7083 0.05407 0.04765 -0.0308 1.0000 0.0292
-8.500 -0.7132 0.05021 0.04349 -0.0282 1.0000 0.0291
-8.250 -0.7140 0.04674 0.03970 -0.0254 1.0000 0.0292
-8.000 -0.7112 0.04362 0.03627 -0.0227 1.0000 0.0293
-7.750 -0.7053 0.04080 0.03312 -0.0201 1.0000 0.0296
-7.500 -0.6971 0.03816 0.03015 -0.0175 1.0000 0.0299
-7.250 -0.6867 0.03561 0.02726 -0.0150 1.0000 0.0305
-7.000 -0.6745 0.03305 0.02428 -0.0125 1.0000 0.0312
-6.750 -0.6600 0.03041 0.02117 -0.0102 1.0000 0.0324
-6.500 -0.6425 0.02928 0.01996 -0.0086 1.0000 0.0333
-6.250 -0.6243 0.02786 0.01836 -0.0070 1.0000 0.0350
-6.000 -0.6051 0.02627 0.01651 -0.0054 1.0000 0.0372
-5.750 -0.5860 0.02535 0.01552 -0.0039 1.0000 0.0392
-5.500 -0.5666 0.02442 0.01446 -0.0024 1.0000 0.0423
-5.250 -0.5475 0.02377 0.01365 -0.0009 1.0000 0.0464
-5.000 -0.5286 0.02324 0.01313 0.0005 1.0000 0.0499
-4.750 -0.5089 0.02251 0.01234 0.0019 1.0000 0.0531
-4.500 -0.4892 0.02189 0.01166 0.0032 1.0000 0.0565
-4.250 -0.4698 0.02127 0.01104 0.0046 1.0000 0.0595
-4.000 -0.4504 0.02067 0.01044 0.0060 1.0000 0.0620
-3.750 -0.4313 0.02014 0.00987 0.0075 1.0000 0.0646
-3.500 -0.3945 0.01950 0.00927 0.0052 0.9932 0.0682
-3.250 -0.3583 0.01894 0.00869 0.0032 0.9859 0.0719
-3.000 -0.3225 0.01844 0.00819 0.0012 0.9773 0.0778
-2.750 -0.2867 0.01794 0.00772 -0.0008 0.9689 0.0846
-2.500 -0.2506 0.01745 0.00724 -0.0027 0.9602 0.0922
-2.250 -0.2150 0.01695 0.00677 -0.0044 0.9514 0.1039
-2.000 -0.1795 0.01624 0.00629 -0.0061 0.9427 0.1394
-1.750 -0.1617 0.01389 0.00610 -0.0048 0.9328 0.5911
-1.500 -0.1282 0.01347 0.00620 -0.0048 0.9244 0.7308
-1.250 -0.0889 0.01347 0.00645 -0.0057 0.9154 0.8084
-1.000 -0.0486 0.01360 0.00661 -0.0069 0.9064 0.8697
-0.750 0.0004 0.01389 0.00688 -0.0097 0.8964 0.9109
-0.500 0.0459 0.01394 0.00683 -0.0127 0.8846 0.9286
-0.250 0.0928 0.01388 0.00667 -0.0163 0.8715 0.9351
0.000 0.1323 0.01384 0.00655 -0.0185 0.8547 0.9444
0.250 0.1723 0.01380 0.00645 -0.0208 0.8347 0.9523
0.500 0.2130 0.01372 0.00629 -0.0232 0.8117 0.9597
0.750 0.2501 0.01365 0.00609 -0.0247 0.7896 0.9684
1.000 0.2918 0.01358 0.00589 -0.0273 0.7677 0.9745
1.250 0.3292 0.01359 0.00582 -0.0292 0.7446 0.9828
1.500 0.3661 0.01361 0.00575 -0.0310 0.7197 0.9904
1.750 0.4036 0.01361 0.00565 -0.0330 0.6899 0.9978
2.000 0.4307 0.01367 0.00557 -0.0328 0.6609 1.0000
2.250 0.4528 0.01375 0.00555 -0.0317 0.6329 1.0000
2.500 0.4748 0.01386 0.00553 -0.0305 0.6063 1.0000
2.750 0.4965 0.01399 0.00557 -0.0293 0.5793 1.0000
3.000 0.5181 0.01416 0.00563 -0.0280 0.5532 1.0000
3.250 0.5396 0.01433 0.00572 -0.0268 0.5271 1.0000
3.500 0.5609 0.01453 0.00584 -0.0256 0.5020 1.0000
3.750 0.5822 0.01474 0.00598 -0.0243 0.4776 1.0000
4.000 0.6032 0.01498 0.00618 -0.0230 0.4516 1.0000
4.250 0.6237 0.01525 0.00636 -0.0216 0.4250 1.0000
4.500 0.6440 0.01554 0.00659 -0.0202 0.3958 1.0000
4.750 0.6639 0.01588 0.00684 -0.0188 0.3659 1.0000
5.000 0.6831 0.01628 0.00712 -0.0172 0.3364 1.0000
5.250 0.7024 0.01670 0.00748 -0.0157 0.3062 1.0000
5.500 0.7211 0.01718 0.00785 -0.0141 0.2788 1.0000
5.750 0.7394 0.01772 0.00827 -0.0125 0.2549 1.0000
6.000 0.7574 0.01830 0.00874 -0.0109 0.2360 1.0000
6.250 0.7764 0.01882 0.00927 -0.0094 0.2184 1.0000
6.500 0.7949 0.01939 0.00980 -0.0078 0.2036 1.0000
6.750 0.8133 0.01996 0.01035 -0.0063 0.1893 1.0000
7.250 0.8493 0.02111 0.01148 -0.0032 0.1592 1.0000
7.500 0.8666 0.02173 0.01204 -0.0016 0.1449 1.0000
7.750 0.8847 0.02231 0.01267 -0.0001 0.1303 1.0000
8.000 0.9021 0.02295 0.01332 0.0014 0.1172 1.0000
8.250 0.9190 0.02364 0.01403 0.0030 0.1048 1.0000
8.500 0.9351 0.02442 0.01481 0.0046 0.0929 1.0000
8.750 0.9511 0.02524 0.01563 0.0063 0.0807 1.0000
9.000 0.9656 0.02618 0.01658 0.0080 0.0708 1.0000
9.250 0.9785 0.02728 0.01771 0.0100 0.0634 1.0000
9.500 0.9905 0.02842 0.01888 0.0121 0.0583 1.0000
9.750 1.0020 0.02959 0.02010 0.0141 0.0540 1.0000
10.000 1.0130 0.03079 0.02138 0.0161 0.0505 1.0000
10.250 1.0231 0.03204 0.02272 0.0181 0.0476 1.0000
10.500 1.0309 0.03330 0.02402 0.0204 0.0456 1.0000
10.750 1.0385 0.03465 0.02551 0.0226 0.0434 1.0000
11.000 1.0445 0.03603 0.02692 0.0247 0.0417 1.0000
11.250 1.0509 0.03758 0.02863 0.0266 0.0399 1.0000
11.500 1.0562 0.03919 0.03033 0.0283 0.0385 1.0000
11.750 1.0608 0.04096 0.03219 0.0298 0.0374 1.0000
12.000 1.0647 0.04294 0.03436 0.0312 0.0361 1.0000
12.250 1.0678 0.04496 0.03650 0.0322 0.0352 1.0000
12.500 1.0709 0.04705 0.03863 0.0330 0.0344 1.0000
12.750 1.0715 0.04967 0.04148 0.0337 0.0338 1.0000
13.000 1.0705 0.05255 0.04458 0.0340 0.0332 1.0000
13.250 1.0680 0.05567 0.04789 0.0340 0.0327 1.0000
13.500 1.0642 0.05903 0.05144 0.0336 0.0323 1.0000
13.750 1.0592 0.06266 0.05524 0.0328 0.0319 1.0000
14.000 1.0528 0.06660 0.05934 0.0317 0.0316 1.0000
14.250 1.0451 0.07088 0.06378 0.0302 0.0314 1.0000
14.500 1.0356 0.07560 0.06865 0.0282 0.0312 1.0000
14.750 1.0233 0.08094 0.07417 0.0256 0.0311 1.0000
15.000 1.0049 0.08761 0.08104 0.0219 0.0311 1.0000
15.250 0.9692 0.09821 0.09197 0.0151 0.0313 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NPL 9615 AIRFOIL (npl9615-il)