NYU/GRUMMAN K-1 AIRFOIL (k1-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NYU/GRUMMAN K-1 AIRFOIL (k1-il) Reynolds number: 100,000 Max Cl/Cd: 39.81 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-k1-il-100000.txt Download as CSV file: xf-k1-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NYU/GRUMMAN K-1 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5522 0.10739 0.10210 -0.0291 1.0000 0.1718
-10.000 -0.7189 0.06406 0.05848 -0.0616 1.0000 0.0846
-9.750 -0.7351 0.05982 0.05414 -0.0617 1.0000 0.0826
-9.500 -0.7854 0.05370 0.04736 -0.0603 1.0000 0.0781
-9.250 -0.7846 0.05043 0.04394 -0.0589 1.0000 0.0771
-9.000 -0.7856 0.04685 0.04005 -0.0575 1.0000 0.0759
-8.750 -0.7830 0.04329 0.03607 -0.0564 1.0000 0.0751
-8.500 -0.7741 0.04039 0.03278 -0.0554 1.0000 0.0754
-8.250 -0.7612 0.03773 0.02974 -0.0546 1.0000 0.0761
-8.000 -0.7451 0.03523 0.02685 -0.0539 1.0000 0.0764
-7.750 -0.7264 0.03294 0.02423 -0.0531 1.0000 0.0768
-7.500 -0.7060 0.03094 0.02190 -0.0524 1.0000 0.0776
-7.250 -0.6843 0.02920 0.01981 -0.0517 1.0000 0.0787
-7.000 -0.6627 0.02770 0.01839 -0.0508 1.0000 0.0807
-6.750 -0.6405 0.02648 0.01713 -0.0501 1.0000 0.0837
-6.500 -0.6181 0.02530 0.01591 -0.0494 1.0000 0.0877
-6.250 -0.5951 0.02428 0.01484 -0.0487 1.0000 0.0929
-6.000 -0.5722 0.02316 0.01386 -0.0482 1.0000 0.0986
-5.750 -0.5481 0.02211 0.01291 -0.0479 1.0000 0.1081
-5.500 -0.5226 0.02093 0.01188 -0.0481 1.0000 0.1246
-5.250 -0.4928 0.01913 0.01060 -0.0499 1.0000 0.1751
-5.000 -0.4572 0.01698 0.00962 -0.0537 1.0000 0.3510
-4.750 -0.4308 0.01678 0.00975 -0.0538 1.0000 0.4373
-4.500 -0.4073 0.01704 0.01012 -0.0527 1.0000 0.4899
-4.250 -0.3852 0.01747 0.01058 -0.0513 1.0000 0.5273
-4.000 -0.3658 0.01809 0.01124 -0.0489 1.0000 0.5551
-3.750 -0.3468 0.01875 0.01189 -0.0465 1.0000 0.5796
-3.500 -0.3286 0.01934 0.01250 -0.0439 1.0000 0.5981
-3.250 -0.3090 0.01979 0.01292 -0.0418 1.0000 0.6139
-3.000 -0.2868 0.02009 0.01314 -0.0406 1.0000 0.6298
-2.750 -0.2666 0.02046 0.01347 -0.0389 1.0000 0.6445
-2.500 -0.2499 0.02089 0.01392 -0.0361 1.0000 0.6569
-2.250 -0.2299 0.02121 0.01422 -0.0344 1.0000 0.6726
-2.000 -0.2070 0.02147 0.01441 -0.0336 1.0000 0.6890
-1.750 -0.1911 0.02166 0.01463 -0.0309 1.0000 0.6981
-1.500 -0.1665 0.02174 0.01465 -0.0308 1.0000 0.7115
-1.250 -0.1471 0.02186 0.01477 -0.0292 1.0000 0.7224
-1.000 -0.1245 0.02191 0.01481 -0.0287 1.0000 0.7339
-0.750 -0.0992 0.02202 0.01487 -0.0289 1.0000 0.7467
-0.500 -0.0807 0.02205 0.01494 -0.0273 1.0000 0.7561
-0.250 -0.0546 0.02214 0.01500 -0.0278 1.0000 0.7680
0.000 -0.0345 0.02219 0.01508 -0.0268 1.0000 0.7777
0.250 -0.0095 0.02228 0.01518 -0.0271 1.0000 0.7882
0.500 0.0133 0.02239 0.01531 -0.0269 1.0000 0.7982
0.750 0.0369 0.02251 0.01546 -0.0269 1.0000 0.8077
1.000 0.0790 0.02290 0.01588 -0.0305 0.9936 0.8181
1.250 0.1240 0.02325 0.01628 -0.0343 0.9827 0.8273
1.500 0.1888 0.02375 0.01683 -0.0416 0.9653 0.8375
1.750 0.2421 0.02360 0.01678 -0.0454 0.9401 0.8457
2.000 0.3041 0.02334 0.01660 -0.0507 0.9193 0.8547
2.250 0.3404 0.02295 0.01633 -0.0514 0.9042 0.8627
2.500 0.3759 0.02266 0.01615 -0.0524 0.8894 0.8717
2.750 0.4073 0.02219 0.01581 -0.0521 0.8750 0.8801
3.000 0.4512 0.02153 0.01530 -0.0538 0.8602 0.8888
3.250 0.4908 0.02044 0.01438 -0.0536 0.8447 0.8969
3.500 0.5265 0.01925 0.01335 -0.0528 0.8256 0.9061
3.750 0.5529 0.01803 0.01229 -0.0502 0.8014 0.9156
4.000 0.5830 0.01658 0.01099 -0.0478 0.7691 0.9253
4.250 0.6075 0.01526 0.00970 -0.0443 0.6901 0.9369
4.500 0.6268 0.01582 0.00854 -0.0397 0.4057 0.9472
4.750 0.6476 0.01706 0.00912 -0.0388 0.3263 0.9617
5.250 0.7090 0.01893 0.01043 -0.0411 0.2598 1.0000
5.500 0.7395 0.01976 0.01120 -0.0425 0.2357 1.0000
5.750 0.7696 0.02063 0.01196 -0.0438 0.2129 1.0000
6.000 0.7988 0.02162 0.01274 -0.0450 0.1898 1.0000
6.250 0.8272 0.02247 0.01362 -0.0460 0.1650 1.0000
6.500 0.8547 0.02362 0.01468 -0.0467 0.1390 1.0000
6.750 0.8824 0.02497 0.01590 -0.0473 0.1165 1.0000
7.000 0.9112 0.02640 0.01724 -0.0480 0.1014 1.0000
7.250 0.9403 0.02779 0.01865 -0.0486 0.0912 1.0000
7.500 0.9705 0.02948 0.02038 -0.0492 0.0849 1.0000
7.750 1.0003 0.03119 0.02203 -0.0499 0.0803 1.0000
8.000 1.0275 0.03296 0.02403 -0.0500 0.0756 1.0000
8.250 1.0529 0.03512 0.02645 -0.0500 0.0720 1.0000
8.500 1.0786 0.03681 0.02817 -0.0501 0.0686 1.0000
8.750 1.0978 0.03962 0.03148 -0.0492 0.0661 1.0000
9.000 1.1151 0.04280 0.03509 -0.0481 0.0647 1.0000
9.250 1.1298 0.04615 0.03882 -0.0469 0.0638 1.0000
9.500 1.1381 0.05021 0.04335 -0.0451 0.0635 1.0000
9.750 1.1282 0.05634 0.05023 -0.0420 0.0645 1.0000
10.000 1.1142 0.06230 0.05672 -0.0392 0.0659 1.0000
10.250 1.1025 0.06740 0.06213 -0.0371 0.0668 1.0000
10.500 1.0905 0.07190 0.06684 -0.0352 0.0675 1.0000
10.750 1.0788 0.07617 0.07124 -0.0336 0.0678 1.0000
11.000 0.9354 0.09744 0.09322 -0.0427 0.0778 1.0000
11.250 0.9343 0.10278 0.09855 -0.0439 0.0803 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NYU/GRUMMAN K-1 AIRFOIL (k1-il)