HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/8 AIRFOIL (hq158-il) Reynolds number: 100,000 Max Cl/Cd: 49.01 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq158-il-100000-n5.txt Download as CSV file: xf-hq158-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5065 0.07964 0.07491 -0.0266 1.0000 0.0192
-8.000 -0.5097 0.07421 0.06954 -0.0320 1.0000 0.0194
-7.750 -0.5128 0.06949 0.06481 -0.0346 1.0000 0.0190
-7.500 -0.5133 0.06481 0.06007 -0.0371 1.0000 0.0188
-7.250 -0.5124 0.06005 0.05521 -0.0391 1.0000 0.0187
-7.000 -0.5092 0.05547 0.05048 -0.0404 1.0000 0.0186
-6.750 -0.5032 0.05087 0.04563 -0.0412 1.0000 0.0189
-6.500 -0.4957 0.04669 0.04120 -0.0413 1.0000 0.0187
-6.250 -0.4864 0.04250 0.03669 -0.0409 1.0000 0.0185
-6.000 -0.4748 0.03859 0.03241 -0.0402 1.0000 0.0184
-5.750 -0.4610 0.03494 0.02831 -0.0391 1.0000 0.0185
-5.500 -0.4451 0.03165 0.02456 -0.0379 1.0000 0.0186
-5.250 -0.4274 0.02884 0.02119 -0.0365 1.0000 0.0189
-5.000 -0.4083 0.02641 0.01830 -0.0350 1.0000 0.0193
-4.750 -0.3887 0.02400 0.01553 -0.0336 1.0000 0.0199
-4.500 -0.3697 0.02217 0.01356 -0.0325 1.0000 0.0226
-4.250 -0.3494 0.02116 0.01237 -0.0313 1.0000 0.0267
-4.000 -0.3284 0.01970 0.01072 -0.0298 1.0000 0.0289
-3.750 -0.3080 0.01842 0.00929 -0.0284 1.0000 0.0315
-3.500 -0.2837 0.01725 0.00807 -0.0280 0.9982 0.0372
-3.250 -0.2498 0.01636 0.00711 -0.0296 0.9923 0.0570
-3.000 -0.2163 0.01538 0.00622 -0.0312 0.9865 0.0984
-2.750 -0.1851 0.01413 0.00558 -0.0328 0.9803 0.2232
-2.500 -0.1569 0.01275 0.00544 -0.0335 0.9744 0.5265
-2.250 -0.1283 0.01250 0.00549 -0.0332 0.9670 0.6590
-2.000 -0.0968 0.01241 0.00547 -0.0333 0.9606 0.7313
-1.750 -0.0695 0.01232 0.00544 -0.0324 0.9529 0.7930
-1.500 -0.0390 0.01220 0.00530 -0.0321 0.9471 0.8462
-1.250 -0.0067 0.01207 0.00515 -0.0325 0.9403 0.8851
-1.000 0.0356 0.01199 0.00498 -0.0352 0.9352 0.9138
-0.750 0.0818 0.01191 0.00477 -0.0390 0.9302 0.9369
-0.500 0.1279 0.01184 0.00460 -0.0429 0.9232 0.9638
-0.250 0.1747 0.01175 0.00442 -0.0470 0.9161 0.9954
0.000 0.2092 0.01172 0.00431 -0.0485 0.9059 1.0000
0.250 0.2385 0.01173 0.00423 -0.0490 0.8942 1.0000
0.500 0.2675 0.01175 0.00419 -0.0493 0.8824 1.0000
0.750 0.2961 0.01176 0.00416 -0.0494 0.8696 1.0000
1.000 0.3241 0.01178 0.00414 -0.0493 0.8562 1.0000
1.250 0.3514 0.01182 0.00417 -0.0491 0.8427 1.0000
1.500 0.3782 0.01186 0.00420 -0.0487 0.8288 1.0000
1.750 0.4047 0.01191 0.00424 -0.0482 0.8136 1.0000
2.000 0.4311 0.01194 0.00427 -0.0475 0.7971 1.0000
2.250 0.4570 0.01201 0.00440 -0.0469 0.7799 1.0000
2.500 0.4826 0.01208 0.00451 -0.0462 0.7617 1.0000
2.750 0.5086 0.01216 0.00461 -0.0455 0.7428 1.0000
3.000 0.5340 0.01225 0.00474 -0.0447 0.7204 1.0000
3.250 0.5593 0.01235 0.00488 -0.0438 0.6953 1.0000
3.500 0.5842 0.01247 0.00508 -0.0429 0.6645 1.0000
3.750 0.6085 0.01264 0.00522 -0.0417 0.6238 1.0000
4.000 0.6318 0.01289 0.00535 -0.0404 0.5667 1.0000
4.250 0.6535 0.01335 0.00552 -0.0388 0.4901 1.0000
4.500 0.6739 0.01405 0.00585 -0.0374 0.4107 1.0000
4.750 0.6949 0.01479 0.00631 -0.0363 0.3354 1.0000
5.000 0.7153 0.01566 0.00697 -0.0352 0.2893 1.0000
5.250 0.7376 0.01637 0.00757 -0.0345 0.2334 1.0000
5.500 0.7591 0.01722 0.00823 -0.0337 0.1730 1.0000
5.750 0.7798 0.01826 0.00902 -0.0328 0.1191 1.0000
6.000 0.7999 0.01946 0.01000 -0.0319 0.0847 1.0000
6.250 0.8206 0.02062 0.01121 -0.0308 0.0671 1.0000
6.500 0.8403 0.02198 0.01267 -0.0295 0.0541 1.0000
6.750 0.8601 0.02320 0.01382 -0.0288 0.0357 1.0000
7.000 0.8792 0.02463 0.01530 -0.0277 0.0239 1.0000
7.250 0.8983 0.02625 0.01728 -0.0263 0.0193 1.0000
7.500 0.9164 0.02809 0.01930 -0.0249 0.0168 1.0000
7.750 0.9330 0.03044 0.02183 -0.0236 0.0152 1.0000
8.000 0.9494 0.03360 0.02526 -0.0221 0.0144 1.0000
8.250 0.9663 0.03648 0.02856 -0.0207 0.0139 1.0000
8.500 0.9796 0.03978 0.03232 -0.0190 0.0135 1.0000
8.750 0.9887 0.04327 0.03631 -0.0172 0.0130 1.0000
9.000 0.9932 0.04687 0.04039 -0.0153 0.0126 1.0000
9.250 0.9928 0.05057 0.04452 -0.0134 0.0121 1.0000
9.500 0.9868 0.05438 0.04870 -0.0114 0.0117 1.0000
9.750 0.9746 0.05799 0.05260 -0.0093 0.0115 1.0000
10.000 0.9582 0.06184 0.05668 -0.0080 0.0114 1.0000
10.250 0.9393 0.06652 0.06155 -0.0083 0.0122 1.0000
10.500 0.9196 0.07194 0.06716 -0.0107 0.0119 1.0000
10.750 0.9003 0.07852 0.07387 -0.0149 0.0124 1.0000
11.000 0.8823 0.08632 0.08175 -0.0206 0.0131 1.0000
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