NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 100,000 Max Cl/Cd: 44.14 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n6h10-il-100000-n5.txt Download as CSV file: xf-n6h10-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 6-H-10 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3988 0.10994 0.10501 -0.0059 0.6970 0.0283
-8.250 -0.3967 0.10691 0.10199 -0.0086 0.6943 0.0283
-8.000 -0.3948 0.10413 0.09922 -0.0104 0.6913 0.0284
-7.750 -0.3901 0.10113 0.09620 -0.0123 0.6887 0.0284
-7.250 -0.3751 0.09106 0.08616 -0.0116 0.6846 0.0295
-7.000 -0.3679 0.08756 0.08262 -0.0119 0.6828 0.0303
-6.750 -0.3601 0.08428 0.07933 -0.0128 0.6807 0.0316
-6.500 -0.3510 0.08117 0.07620 -0.0145 0.6780 0.0326
-6.250 -0.3408 0.07807 0.07306 -0.0162 0.6754 0.0340
-6.000 -0.3279 0.07526 0.07015 -0.0182 0.6730 0.0366
-5.750 -0.3069 0.07411 0.06869 -0.0214 0.6708 0.0385
-5.250 -0.2862 0.06560 0.06007 -0.0214 0.6673 0.0404
-5.000 -0.2716 0.06235 0.05679 -0.0216 0.6643 0.0421
-4.750 -0.2545 0.05948 0.05382 -0.0221 0.6618 0.0444
-4.500 -0.2346 0.05687 0.05104 -0.0225 0.6596 0.0477
-4.250 -0.2064 0.05611 0.04970 -0.0226 0.6575 0.0517
-4.000 -0.1944 0.05132 0.04499 -0.0224 0.6559 0.0534
-3.750 -0.1771 0.04854 0.04210 -0.0217 0.6544 0.0556
-3.500 -0.1566 0.04618 0.03954 -0.0212 0.6529 0.0587
-3.250 -0.1311 0.04437 0.03732 -0.0210 0.6500 0.0670
-2.750 -0.0766 0.03755 0.02977 -0.0184 0.6453 0.0273
-2.500 -0.0532 0.03532 0.02724 -0.0175 0.6433 0.0268
-2.250 -0.0278 0.03334 0.02486 -0.0164 0.6416 0.0271
-2.000 -0.0021 0.03146 0.02261 -0.0154 0.6402 0.0271
-1.750 0.0244 0.02962 0.02040 -0.0146 0.6390 0.0265
-1.500 0.0538 0.02820 0.01869 -0.0149 0.6363 0.0262
-1.250 0.0841 0.02687 0.01707 -0.0150 0.6339 0.0262
-1.000 0.1148 0.02574 0.01568 -0.0153 0.6319 0.0269
-0.750 0.1447 0.02480 0.01456 -0.0155 0.6300 0.0293
-0.500 0.1743 0.02385 0.01355 -0.0158 0.6283 0.0316
-0.250 0.2021 0.02315 0.01278 -0.0156 0.6267 0.0340
0.250 0.4321 0.02048 0.01189 -0.0469 0.6253 1.0000
0.500 0.4586 0.02081 0.01208 -0.0473 0.6233 1.0000
0.750 0.4860 0.02130 0.01250 -0.0482 0.6207 1.0000
1.000 0.5122 0.02172 0.01284 -0.0486 0.6184 1.0000
1.250 0.5375 0.02204 0.01309 -0.0487 0.6163 1.0000
1.500 0.5621 0.02227 0.01325 -0.0485 0.6142 1.0000
1.750 0.5860 0.02243 0.01334 -0.0480 0.6124 1.0000
2.000 0.6097 0.02255 0.01340 -0.0474 0.6109 1.0000
2.250 0.6354 0.02325 0.01415 -0.0483 0.6071 1.0000
2.500 0.6599 0.02378 0.01473 -0.0487 0.6034 1.0000
2.750 0.6833 0.02402 0.01497 -0.0482 0.6006 1.0000
3.000 0.7063 0.02413 0.01509 -0.0475 0.5984 1.0000
3.250 0.7292 0.02422 0.01520 -0.0466 0.5968 1.0000
3.500 0.7522 0.02515 0.01625 -0.0474 0.5916 1.0000
3.750 0.7745 0.02560 0.01678 -0.0471 0.5877 1.0000
4.000 0.7966 0.02581 0.01707 -0.0464 0.5851 1.0000
4.250 0.8189 0.02600 0.01733 -0.0455 0.5833 1.0000
4.500 0.8411 0.02615 0.01754 -0.0446 0.5819 1.0000
4.750 0.8587 0.02771 0.01933 -0.0454 0.5742 1.0000
5.000 0.8824 0.02671 0.01839 -0.0432 0.5696 1.0000
5.250 0.9031 0.02637 0.01815 -0.0418 0.5598 1.0000
5.500 0.9287 0.02471 0.01649 -0.0390 0.5540 1.0000
5.750 0.9483 0.02483 0.01682 -0.0382 0.5432 1.0000
6.000 0.9702 0.02423 0.01635 -0.0366 0.5329 1.0000
6.250 0.9929 0.02343 0.01567 -0.0349 0.5215 1.0000
6.500 1.0147 0.02299 0.01537 -0.0335 0.5105 1.0000
6.750 1.0307 0.02350 0.01617 -0.0327 0.4930 1.0000
7.000 1.0436 0.02424 0.01715 -0.0316 0.4715 1.0000
7.250 1.0501 0.02534 0.01842 -0.0301 0.4321 1.0000
7.500 1.0607 0.02405 0.01590 -0.0248 0.2977 1.0000
7.750 1.0357 0.02689 0.01834 -0.0203 0.2225 1.0000
8.000 1.0079 0.02895 0.02022 -0.0143 0.2037 1.0000
8.250 0.9841 0.03088 0.02197 -0.0087 0.1805 1.0000
8.500 0.9668 0.03295 0.02382 -0.0044 0.1453 1.0000
8.750 0.9557 0.03498 0.02561 -0.0011 0.1189 1.0000
9.000 0.9520 0.03671 0.02724 0.0013 0.1063 1.0000
9.250 0.9508 0.03839 0.02885 0.0035 0.0981 1.0000
9.500 0.9538 0.03984 0.03037 0.0053 0.0934 1.0000
9.750 0.9567 0.04134 0.03190 0.0070 0.0894 1.0000
10.000 0.9603 0.04282 0.03345 0.0087 0.0864 1.0000
10.250 0.9684 0.04399 0.03476 0.0102 0.0841 1.0000
10.500 0.9774 0.04509 0.03599 0.0117 0.0816 1.0000
10.750 0.9876 0.04608 0.03709 0.0132 0.0788 1.0000
11.000 0.9999 0.04691 0.03802 0.0147 0.0763 1.0000
11.250 1.0104 0.04801 0.03917 0.0161 0.0705 1.0000
11.500 1.0141 0.04984 0.04111 0.0165 0.0622 1.0000
11.750 1.0210 0.05145 0.04287 0.0172 0.0555 1.0000
12.000 1.0261 0.05337 0.04493 0.0175 0.0485 1.0000
12.250 1.0324 0.05524 0.04698 0.0179 0.0405 1.0000
12.500 1.0376 0.05729 0.04915 0.0183 0.0318 1.0000
12.750 1.0425 0.05938 0.05139 0.0190 0.0252 1.0000
13.000 1.0492 0.06131 0.05343 0.0198 0.0207 1.0000
13.250 1.0524 0.06369 0.05584 0.0200 0.0179 1.0000
13.500 1.0570 0.06593 0.05821 0.0211 0.0162 1.0000
13.750 1.0629 0.06819 0.06068 0.0221 0.0154 1.0000
14.000 1.0670 0.07084 0.06356 0.0230 0.0143 1.0000
14.250 1.0689 0.07380 0.06673 0.0234 0.0136 1.0000
14.500 1.0680 0.07715 0.07028 0.0233 0.0128 1.0000
14.750 1.0650 0.08077 0.07407 0.0228 0.0122 1.0000
15.000 1.0606 0.08487 0.07843 0.0225 0.0121 1.0000
15.250 1.0514 0.08953 0.08321 0.0211 0.0113 1.0000
15.500 1.0424 0.09453 0.08845 0.0198 0.0112 1.0000
15.750 1.0311 0.10012 0.09427 0.0180 0.0111 1.0000
16.000 1.0176 0.10630 0.10067 0.0156 0.0110 1.0000
16.250 1.0031 0.11309 0.10768 0.0127 0.0111 1.0000
16.500 0.9869 0.12060 0.11540 0.0091 0.0112 1.0000
16.750 0.9683 0.12928 0.12428 0.0045 0.0113 1.0000
17.000 0.9197 0.14827 0.14358 -0.0055 0.0132 1.0000
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Polar data table (+)
Polar graphs
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