HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.0/9 AIRFOIL (hq109-il) Reynolds number: 200,000 Max Cl/Cd: 62.87 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq109-il-200000.txt Download as CSV file: xf-hq109-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/9 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5644 0.09701 0.09356 -0.0203 1.0000 0.0438
-9.500 -0.5710 0.09061 0.08719 -0.0261 1.0000 0.0438
-9.250 -0.5819 0.08530 0.08185 -0.0309 1.0000 0.0439
-9.000 -0.5804 0.07909 0.07577 -0.0292 1.0000 0.0452
-8.750 -0.5749 0.07664 0.07335 -0.0278 1.0000 0.0464
-8.500 -0.5903 0.07115 0.06784 -0.0326 1.0000 0.0463
-8.250 -0.5946 0.06735 0.06401 -0.0338 1.0000 0.0471
-8.000 -0.5975 0.06327 0.05987 -0.0351 1.0000 0.0480
-7.750 -0.5979 0.05927 0.05577 -0.0360 1.0000 0.0495
-7.500 -0.5967 0.05510 0.05144 -0.0367 1.0000 0.0510
-7.000 -0.5959 0.03950 0.03465 -0.0353 1.0000 0.0305
-6.750 -0.5885 0.03444 0.02921 -0.0338 1.0000 0.0281
-6.500 -0.5777 0.02948 0.02365 -0.0317 1.0000 0.0259
-6.250 -0.5605 0.02752 0.02140 -0.0302 1.0000 0.0280
-6.000 -0.5428 0.02523 0.01873 -0.0285 1.0000 0.0293
-5.750 -0.5247 0.02268 0.01581 -0.0266 1.0000 0.0291
-5.500 -0.5056 0.02065 0.01347 -0.0247 1.0000 0.0299
-5.250 -0.4864 0.01913 0.01173 -0.0229 1.0000 0.0311
-5.000 -0.4686 0.01729 0.00975 -0.0210 1.0000 0.0334
-4.750 -0.4515 0.01639 0.00888 -0.0193 1.0000 0.0389
-4.500 -0.4338 0.01536 0.00773 -0.0175 1.0000 0.0454
-4.250 -0.4162 0.01449 0.00684 -0.0158 1.0000 0.0570
-4.000 -0.3985 0.01363 0.00608 -0.0145 1.0000 0.0762
-3.750 -0.3794 0.01306 0.00556 -0.0134 1.0000 0.0989
-3.500 -0.3448 0.01201 0.00495 -0.0156 0.9956 0.1759
-3.250 -0.3138 0.01036 0.00457 -0.0177 0.9901 0.4460
-3.000 -0.2790 0.00994 0.00466 -0.0192 0.9840 0.6034
-2.750 -0.2435 0.00989 0.00469 -0.0206 0.9773 0.6648
-2.500 -0.2062 0.00990 0.00473 -0.0223 0.9712 0.7093
-2.250 -0.1705 0.00992 0.00474 -0.0236 0.9647 0.7427
-2.000 -0.1342 0.00994 0.00476 -0.0250 0.9584 0.7736
-1.750 -0.0961 0.00995 0.00478 -0.0266 0.9536 0.8017
-1.500 -0.0634 0.00990 0.00474 -0.0272 0.9458 0.8209
-1.250 -0.0254 0.00984 0.00464 -0.0288 0.9411 0.8403
-1.000 0.0032 0.00977 0.00458 -0.0286 0.9315 0.8586
-0.750 0.0346 0.00967 0.00450 -0.0287 0.9239 0.8754
-0.500 0.0644 0.00959 0.00441 -0.0285 0.9157 0.8928
-0.250 0.0929 0.00954 0.00437 -0.0281 0.9069 0.9094
0.000 0.1267 0.00946 0.00427 -0.0288 0.8997 0.9224
0.250 0.1590 0.00939 0.00419 -0.0293 0.8882 0.9357
0.500 0.1945 0.00932 0.00410 -0.0304 0.8763 0.9482
0.750 0.2324 0.00924 0.00399 -0.0321 0.8636 0.9597
1.000 0.2718 0.00917 0.00390 -0.0342 0.8508 0.9707
1.250 0.3136 0.00912 0.00385 -0.0369 0.8388 0.9806
1.500 0.3567 0.00905 0.00378 -0.0401 0.8264 0.9902
1.750 0.3956 0.00899 0.00371 -0.0425 0.8126 1.0000
2.000 0.4114 0.00897 0.00368 -0.0404 0.7975 1.0000
2.250 0.4294 0.00898 0.00369 -0.0386 0.7819 1.0000
2.500 0.4500 0.00901 0.00369 -0.0371 0.7656 1.0000
2.750 0.4721 0.00904 0.00368 -0.0357 0.7478 1.0000
3.000 0.4951 0.00908 0.00372 -0.0346 0.7266 1.0000
3.250 0.5187 0.00913 0.00373 -0.0334 0.7035 1.0000
3.500 0.5424 0.00920 0.00379 -0.0323 0.6745 1.0000
3.750 0.5660 0.00931 0.00384 -0.0312 0.6396 1.0000
4.000 0.5896 0.00947 0.00392 -0.0301 0.5984 1.0000
4.250 0.6124 0.00974 0.00404 -0.0289 0.5434 1.0000
4.500 0.6335 0.01022 0.00421 -0.0275 0.4687 1.0000
4.750 0.6536 0.01091 0.00456 -0.0262 0.3862 1.0000
5.000 0.6741 0.01167 0.00498 -0.0251 0.3128 1.0000
5.250 0.6954 0.01241 0.00549 -0.0243 0.2587 1.0000
5.500 0.7171 0.01315 0.00601 -0.0235 0.2023 1.0000
5.750 0.7377 0.01408 0.00660 -0.0226 0.1408 1.0000
6.000 0.7575 0.01520 0.00750 -0.0216 0.0996 1.0000
6.250 0.7776 0.01633 0.00852 -0.0205 0.0723 1.0000
6.500 0.7984 0.01739 0.00956 -0.0194 0.0534 1.0000
6.750 0.8184 0.01858 0.01074 -0.0183 0.0402 1.0000
7.000 0.8381 0.01998 0.01220 -0.0170 0.0317 1.0000
7.250 0.8550 0.02255 0.01478 -0.0154 0.0271 1.0000
7.500 0.8777 0.02421 0.01664 -0.0143 0.0255 1.0000
7.750 0.8999 0.02607 0.01874 -0.0133 0.0235 1.0000
8.000 0.9204 0.02717 0.01999 -0.0124 0.0206 1.0000
8.250 0.9393 0.02935 0.02236 -0.0114 0.0194 1.0000
8.500 0.9547 0.03283 0.02618 -0.0100 0.0187 1.0000
8.750 0.9674 0.03631 0.03006 -0.0084 0.0186 1.0000
9.000 0.9770 0.03969 0.03387 -0.0065 0.0187 1.0000
9.250 0.9831 0.04308 0.03771 -0.0044 0.0191 1.0000
9.500 0.9627 0.05065 0.04619 -0.0003 0.0219 1.0000
9.750 0.9452 0.05588 0.05179 0.0021 0.0231 1.0000
10.000 0.9261 0.05971 0.05582 0.0045 0.0238 1.0000
10.250 0.9046 0.06391 0.06019 0.0052 0.0242 1.0000
10.500 0.8842 0.06860 0.06503 0.0039 0.0246 1.0000
10.750 0.8609 0.07468 0.07124 0.0004 0.0246 1.0000
11.000 0.8391 0.08210 0.07877 -0.0052 0.0248 1.0000
11.250 0.8181 0.09180 0.08854 -0.0128 0.0251 1.0000
11.500 0.7994 0.10269 0.09938 -0.0197 0.0260 1.0000
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