NACA 66-021 AIRFOIL (n66021-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 66-021 AIRFOIL (n66021-il) Reynolds number: 200,000 Max Cl/Cd: 26.46 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n66021-il-200000-n5.txt Download as CSV file: xf-n66021-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 66-021 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-19.250 -0.8616 0.09577 0.09009 -0.0745 0.9480 0.0218
-19.000 -0.8703 0.09031 0.08435 -0.0778 0.9378 0.0219
-18.750 -0.8760 0.08573 0.07953 -0.0801 0.9283 0.0219
-18.500 -0.8751 0.08238 0.07604 -0.0815 0.9193 0.0221
-18.250 -0.8741 0.07928 0.07280 -0.0826 0.9112 0.0222
-18.000 -0.8717 0.07649 0.06988 -0.0834 0.9039 0.0223
-17.750 -0.8695 0.07373 0.06698 -0.0840 0.8975 0.0225
-17.500 -0.8661 0.07117 0.06428 -0.0845 0.8914 0.0226
-17.250 -0.8617 0.06878 0.06175 -0.0848 0.8861 0.0228
-17.000 -0.8566 0.06649 0.05934 -0.0851 0.8812 0.0231
-16.750 -0.8505 0.06435 0.05707 -0.0852 0.8763 0.0233
-16.500 -0.8443 0.06229 0.05487 -0.0853 0.8718 0.0237
-16.250 -0.8375 0.06036 0.05280 -0.0853 0.8678 0.0242
-16.000 -0.8273 0.05866 0.05098 -0.0852 0.8639 0.0244
-15.750 -0.8173 0.05702 0.04921 -0.0851 0.8599 0.0247
-15.500 -0.8076 0.05542 0.04744 -0.0849 0.8560 0.0253
-15.250 -0.7972 0.05394 0.04582 -0.0845 0.8526 0.0258
-15.000 -0.7856 0.05264 0.04449 -0.0842 0.8495 0.0262
-14.750 -0.7739 0.05128 0.04310 -0.0840 0.8463 0.0267
-14.500 -0.7613 0.05001 0.04176 -0.0837 0.8430 0.0269
-14.250 -0.7495 0.04872 0.04042 -0.0834 0.8398 0.0274
-14.000 -0.7381 0.04743 0.03906 -0.0829 0.8368 0.0282
-13.750 -0.7264 0.04622 0.03775 -0.0824 0.8340 0.0287
-13.500 -0.7148 0.04504 0.03646 -0.0818 0.8315 0.0293
-13.250 -0.7024 0.04387 0.03521 -0.0813 0.8287 0.0299
-13.000 -0.6927 0.04254 0.03387 -0.0807 0.8257 0.0306
-12.750 -0.6823 0.04130 0.03260 -0.0801 0.8228 0.0313
-12.500 -0.6715 0.04014 0.03138 -0.0793 0.8201 0.0323
-12.250 -0.6602 0.03903 0.03020 -0.0786 0.8176 0.0332
-12.000 -0.6481 0.03801 0.02908 -0.0778 0.8155 0.0346
-11.750 -0.6387 0.03683 0.02788 -0.0768 0.8135 0.0360
-11.500 -0.6270 0.03583 0.02682 -0.0760 0.8116 0.0374
-11.250 -0.6145 0.03486 0.02580 -0.0752 0.8094 0.0393
-11.000 -0.6031 0.03386 0.02478 -0.0743 0.8071 0.0411
-10.750 -0.5908 0.03293 0.02383 -0.0734 0.8050 0.0436
-10.500 -0.5777 0.03208 0.02294 -0.0725 0.8029 0.0466
-10.250 -0.5651 0.03121 0.02206 -0.0715 0.8009 0.0503
-10.000 -0.5530 0.03035 0.02120 -0.0704 0.7991 0.0559
-9.750 -0.5405 0.02952 0.02039 -0.0693 0.7974 0.0637
-9.250 -0.5211 0.02764 0.01869 -0.0664 0.7943 0.1014
-9.000 -0.5148 0.02658 0.01783 -0.0645 0.7928 0.1374
-8.750 -0.5111 0.02542 0.01695 -0.0624 0.7910 0.1832
-8.500 -0.5093 0.02426 0.01606 -0.0599 0.7891 0.2295
-8.250 -0.5080 0.02315 0.01519 -0.0572 0.7871 0.2746
-8.000 -0.5087 0.02205 0.01432 -0.0541 0.7850 0.3181
-7.750 -0.5104 0.02099 0.01349 -0.0506 0.7829 0.3634
-7.500 -0.5130 0.01998 0.01274 -0.0468 0.7809 0.4126
-7.250 -0.5091 0.01930 0.01233 -0.0436 0.7793 0.4609
-7.000 -0.4629 0.02072 0.01423 -0.0450 0.7786 0.5349
-6.750 -0.3869 0.02440 0.01812 -0.0493 0.7782 0.5859
-6.250 -0.3672 0.02392 0.01753 -0.0447 0.7758 0.6213
-6.000 -0.3387 0.02447 0.01800 -0.0446 0.7749 0.6272
-5.500 -0.3244 0.02368 0.01714 -0.0396 0.7718 0.6496
-5.000 -0.3089 0.02308 0.01648 -0.0344 0.7684 0.6680
-4.750 -0.2732 0.02371 0.01705 -0.0356 0.7675 0.6691
-4.500 -0.2375 0.02430 0.01759 -0.0369 0.7665 0.6702
-4.000 -0.2218 0.02367 0.01689 -0.0315 0.7634 0.6855
-3.750 -0.1863 0.02413 0.01730 -0.0328 0.7624 0.6863
-3.500 -0.1530 0.02446 0.01760 -0.0338 0.7614 0.6872
-3.000 -0.1560 0.02338 0.01645 -0.0254 0.7586 0.7027
-2.750 -0.1215 0.02366 0.01670 -0.0267 0.7580 0.7033
-2.500 -0.0883 0.02392 0.01692 -0.0278 0.7574 0.7040
-2.000 -0.1202 0.02318 0.01618 -0.0142 0.7532 0.7199
-1.750 -0.0927 0.02357 0.01659 -0.0145 0.7512 0.7207
-1.500 -0.0665 0.02392 0.01696 -0.0146 0.7495 0.7216
-1.250 -0.0424 0.02424 0.01729 -0.0144 0.7479 0.7227
-1.000 -0.0202 0.02452 0.01758 -0.0138 0.7464 0.7240
-0.750 -0.0008 0.02466 0.01773 -0.0127 0.7447 0.7257
-0.250 -0.0277 0.02327 0.01623 0.0003 0.7401 0.7384
0.000 0.0000 0.02327 0.01623 0.0000 0.7392 0.7392
0.250 0.0279 0.02327 0.01623 -0.0003 0.7384 0.7401
0.750 0.0008 0.02466 0.01772 0.0127 0.7257 0.7447
1.000 0.0202 0.02452 0.01758 0.0138 0.7240 0.7464
1.250 0.0425 0.02424 0.01729 0.0144 0.7227 0.7479
1.500 0.0667 0.02392 0.01695 0.0146 0.7216 0.7495
1.750 0.0927 0.02357 0.01659 0.0145 0.7207 0.7512
2.000 0.1202 0.02318 0.01618 0.0142 0.7199 0.7532
2.500 0.0537 0.02478 0.01779 0.0330 0.6979 0.7580
2.750 0.1061 0.02402 0.01706 0.0290 0.7008 0.7583
3.000 0.1504 0.02350 0.01657 0.0262 0.7019 0.7587
4.000 0.1716 0.02516 0.01836 0.0386 0.6736 0.7646
4.250 0.2034 0.02478 0.01803 0.0379 0.6716 0.7656
4.500 0.2376 0.02429 0.01758 0.0368 0.6702 0.7665
4.750 0.2734 0.02370 0.01704 0.0356 0.6691 0.7675
5.000 0.3089 0.02309 0.01648 0.0344 0.6679 0.7685
5.500 0.3345 0.02329 0.01675 0.0383 0.6522 0.7716
6.000 0.3569 0.02369 0.01723 0.0424 0.6334 0.7746
6.500 0.3747 0.02432 0.01799 0.0472 0.6063 0.7770
6.750 0.3864 0.02441 0.01813 0.0493 0.5849 0.7783
7.000 0.4626 0.02075 0.01426 0.0450 0.5355 0.7786
7.250 0.5102 0.01928 0.01234 0.0434 0.4648 0.7793
7.500 0.5134 0.01997 0.01273 0.0468 0.4131 0.7809
7.750 0.5115 0.02095 0.01346 0.0505 0.3647 0.7829
8.000 0.5088 0.02205 0.01432 0.0541 0.3177 0.7850
8.250 0.5075 0.02318 0.01520 0.0573 0.2730 0.7871
8.500 0.5087 0.02429 0.01607 0.0600 0.2270 0.7891
8.750 0.5114 0.02542 0.01695 0.0623 0.1826 0.7910
9.000 0.5151 0.02657 0.01781 0.0645 0.1367 0.7928
9.250 0.5206 0.02768 0.01870 0.0664 0.0989 0.7943
9.500 0.5294 0.02865 0.01955 0.0680 0.0771 0.7958
9.750 0.5411 0.02950 0.02037 0.0692 0.0639 0.7974
10.000 0.5534 0.03034 0.02118 0.0704 0.0556 0.7991
10.250 0.5659 0.03118 0.02203 0.0714 0.0504 0.8009
10.500 0.5783 0.03206 0.02291 0.0724 0.0463 0.8029
10.750 0.5913 0.03292 0.02381 0.0733 0.0433 0.8050
11.000 0.6037 0.03384 0.02475 0.0742 0.0408 0.8072
11.250 0.6152 0.03484 0.02578 0.0751 0.0393 0.8094
11.500 0.6278 0.03580 0.02679 0.0759 0.0375 0.8116
11.750 0.6395 0.03680 0.02784 0.0767 0.0359 0.8135
12.000 0.6492 0.03795 0.02902 0.0777 0.0345 0.8155
12.250 0.6611 0.03899 0.03016 0.0785 0.0330 0.8177
12.500 0.6722 0.04011 0.03135 0.0793 0.0322 0.8201
12.750 0.6834 0.04124 0.03254 0.0799 0.0315 0.8228
13.000 0.6941 0.04246 0.03379 0.0806 0.0307 0.8258
13.250 0.7035 0.04383 0.03515 0.0812 0.0299 0.8287
13.500 0.7157 0.04500 0.03642 0.0817 0.0292 0.8315
13.750 0.7276 0.04616 0.03768 0.0823 0.0288 0.8340
14.000 0.7392 0.04738 0.03900 0.0829 0.0281 0.8369
14.250 0.7511 0.04862 0.04032 0.0833 0.0276 0.8399
14.500 0.7627 0.04993 0.04169 0.0836 0.0269 0.8431
14.750 0.7748 0.05124 0.04304 0.0839 0.0265 0.8465
15.000 0.7868 0.05257 0.04441 0.0841 0.0261 0.8497
15.250 0.7991 0.05383 0.04568 0.0844 0.0257 0.8527
15.500 0.8093 0.05533 0.04735 0.0848 0.0253 0.8562
15.750 0.8195 0.05687 0.04903 0.0850 0.0249 0.8601
16.000 0.8295 0.05853 0.05083 0.0851 0.0245 0.8642
16.250 0.8376 0.06035 0.05279 0.0851 0.0240 0.8682
16.500 0.8449 0.06229 0.05487 0.0852 0.0235 0.8722
16.750 0.8516 0.06429 0.05700 0.0851 0.0232 0.8768
17.000 0.8586 0.06638 0.05922 0.0849 0.0231 0.8817
17.250 0.8639 0.06866 0.06164 0.0847 0.0229 0.8866
17.500 0.8680 0.07105 0.06417 0.0843 0.0226 0.8920
17.750 0.8713 0.07364 0.06689 0.0838 0.0225 0.8982
18.000 0.8741 0.07633 0.06971 0.0832 0.0223 0.9047
18.250 0.8760 0.07921 0.07273 0.0824 0.0222 0.9123
18.500 0.8765 0.08240 0.07609 0.0813 0.0221 0.9204
18.750 0.8782 0.08568 0.07948 0.0798 0.0219 0.9297
19.000 0.8713 0.09051 0.08458 0.0773 0.0218 0.9395
19.250 0.8623 0.09613 0.09049 0.0739 0.0217 0.9498
|
Polar data table (+)
Polar graphs
<< Back to NACA 66-021 AIRFOIL (n66021-il)