HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il) Reynolds number: 100,000 Max Cl/Cd: 45.13 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1010-il-100000-n5.txt Download as CSV file: xf-hq1010-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/10 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5861 0.08911 0.08422 -0.0201 1.0000 0.0211
-10.000 -0.5984 0.08088 0.07606 -0.0253 1.0000 0.0208
-9.750 -0.6252 0.06882 0.06403 -0.0347 1.0000 0.0198
-9.500 -0.6564 0.06097 0.05604 -0.0402 1.0000 0.0193
-9.250 -0.6800 0.05611 0.05100 -0.0407 1.0000 0.0190
-9.000 -0.6929 0.05169 0.04632 -0.0405 1.0000 0.0190
-8.750 -0.6986 0.04757 0.04189 -0.0398 1.0000 0.0191
-8.500 -0.6985 0.04380 0.03776 -0.0388 1.0000 0.0193
-8.250 -0.6933 0.04043 0.03402 -0.0376 1.0000 0.0197
-8.000 -0.6845 0.03726 0.03046 -0.0363 1.0000 0.0203
-7.750 -0.6719 0.03476 0.02756 -0.0351 1.0000 0.0216
-7.500 -0.6570 0.03248 0.02487 -0.0337 1.0000 0.0237
-7.250 -0.6402 0.03025 0.02215 -0.0322 1.0000 0.0255
-7.000 -0.6238 0.02761 0.01927 -0.0308 1.0000 0.0268
-6.750 -0.6060 0.02594 0.01749 -0.0295 1.0000 0.0287
-6.500 -0.5874 0.02455 0.01597 -0.0281 1.0000 0.0314
-6.250 -0.5680 0.02347 0.01462 -0.0266 1.0000 0.0360
-6.000 -0.5511 0.02214 0.01332 -0.0252 1.0000 0.0406
-5.750 -0.5332 0.02115 0.01221 -0.0235 1.0000 0.0465
-5.500 -0.5165 0.02022 0.01124 -0.0219 1.0000 0.0540
-5.250 -0.4995 0.01943 0.01034 -0.0202 1.0000 0.0620
-5.000 -0.4830 0.01873 0.00964 -0.0186 1.0000 0.0725
-4.750 -0.4664 0.01806 0.00897 -0.0170 1.0000 0.0837
-4.500 -0.4327 0.01725 0.00815 -0.0187 0.9930 0.1023
-4.250 -0.3994 0.01644 0.00744 -0.0205 0.9858 0.1304
-4.000 -0.3676 0.01556 0.00686 -0.0221 0.9783 0.1980
-3.750 -0.3350 0.01460 0.00638 -0.0239 0.9718 0.3044
-3.500 -0.3075 0.01376 0.00620 -0.0244 0.9635 0.4527
-3.250 -0.2741 0.01347 0.00622 -0.0254 0.9576 0.5562
-3.000 -0.2444 0.01339 0.00620 -0.0255 0.9490 0.6196
-2.750 -0.2114 0.01340 0.00623 -0.0261 0.9428 0.6673
-2.500 -0.1827 0.01342 0.00624 -0.0258 0.9340 0.7031
-2.250 -0.1512 0.01343 0.00620 -0.0261 0.9268 0.7293
-2.000 -0.1205 0.01342 0.00609 -0.0264 0.9189 0.7485
-1.750 -0.0899 0.01340 0.00602 -0.0266 0.9112 0.7654
-1.500 -0.0595 0.01339 0.00597 -0.0267 0.9037 0.7853
-1.250 -0.0318 0.01339 0.00592 -0.0263 0.8951 0.8051
-1.000 -0.0006 0.01334 0.00584 -0.0265 0.8881 0.8209
-0.750 0.0270 0.01331 0.00578 -0.0263 0.8785 0.8339
-0.500 0.0566 0.01327 0.00569 -0.0264 0.8703 0.8460
-0.250 0.0861 0.01322 0.00561 -0.0264 0.8617 0.8577
0.000 0.1147 0.01319 0.00558 -0.0263 0.8523 0.8696
0.250 0.1457 0.01315 0.00553 -0.0267 0.8445 0.8816
0.500 0.1756 0.01313 0.00553 -0.0269 0.8348 0.8943
0.750 0.2073 0.01311 0.00551 -0.0274 0.8246 0.9070
1.000 0.2406 0.01305 0.00546 -0.0282 0.8122 0.9197
1.250 0.2746 0.01297 0.00539 -0.0291 0.7969 0.9327
1.500 0.3095 0.01288 0.00530 -0.0301 0.7803 0.9462
1.750 0.3459 0.01282 0.00525 -0.0317 0.7652 0.9600
2.000 0.3832 0.01279 0.00525 -0.0336 0.7512 0.9741
2.250 0.4210 0.01276 0.00528 -0.0356 0.7364 0.9888
2.500 0.4514 0.01277 0.00530 -0.0363 0.7209 1.0000
2.750 0.4717 0.01282 0.00534 -0.0349 0.7049 1.0000
3.000 0.4929 0.01290 0.00546 -0.0336 0.6862 1.0000
3.250 0.5153 0.01298 0.00554 -0.0324 0.6659 1.0000
3.500 0.5381 0.01307 0.00564 -0.0313 0.6418 1.0000
3.750 0.5610 0.01319 0.00573 -0.0301 0.6127 1.0000
4.000 0.5838 0.01334 0.00582 -0.0289 0.5773 1.0000
4.250 0.6063 0.01356 0.00598 -0.0276 0.5330 1.0000
4.500 0.6278 0.01391 0.00613 -0.0263 0.4784 1.0000
4.750 0.6483 0.01442 0.00637 -0.0249 0.4170 1.0000
5.000 0.6686 0.01504 0.00674 -0.0236 0.3594 1.0000
5.250 0.6890 0.01572 0.00720 -0.0226 0.3123 1.0000
5.500 0.7100 0.01639 0.00777 -0.0216 0.2747 1.0000
5.750 0.7312 0.01708 0.00835 -0.0207 0.2404 1.0000
6.000 0.7527 0.01774 0.00896 -0.0199 0.2063 1.0000
6.250 0.7738 0.01846 0.00959 -0.0190 0.1722 1.0000
6.500 0.7945 0.01926 0.01030 -0.0181 0.1411 1.0000
6.750 0.8143 0.02018 0.01110 -0.0172 0.1151 1.0000
7.000 0.8340 0.02114 0.01207 -0.0162 0.0956 1.0000
7.250 0.8528 0.02220 0.01307 -0.0151 0.0799 1.0000
7.500 0.8717 0.02327 0.01418 -0.0140 0.0675 1.0000
7.750 0.8905 0.02435 0.01534 -0.0129 0.0564 1.0000
8.000 0.9076 0.02572 0.01683 -0.0115 0.0489 1.0000
8.250 0.9237 0.02710 0.01822 -0.0103 0.0412 1.0000
8.500 0.9416 0.02824 0.01952 -0.0092 0.0340 1.0000
8.750 0.9562 0.02988 0.02128 -0.0077 0.0293 1.0000
9.000 0.9717 0.03144 0.02308 -0.0063 0.0252 1.0000
9.250 0.9837 0.03311 0.02481 -0.0050 0.0224 1.0000
9.500 0.9967 0.03514 0.02711 -0.0034 0.0199 1.0000
9.750 1.0085 0.03743 0.02968 -0.0018 0.0187 1.0000
10.000 1.0178 0.03982 0.03235 -0.0002 0.0177 1.0000
10.250 1.0233 0.04230 0.03509 0.0017 0.0170 1.0000
10.500 1.0242 0.04479 0.03782 0.0038 0.0164 1.0000
10.750 1.0219 0.04746 0.04074 0.0059 0.0160 1.0000
11.000 1.0166 0.05035 0.04386 0.0075 0.0157 1.0000
11.250 1.0079 0.05369 0.04743 0.0085 0.0155 1.0000
11.500 0.9970 0.05739 0.05136 0.0088 0.0153 1.0000
11.750 0.9833 0.06166 0.05587 0.0082 0.0153 1.0000
12.000 0.9674 0.06661 0.06104 0.0066 0.0152 1.0000
12.250 0.9482 0.07266 0.06735 0.0036 0.0154 1.0000
12.500 0.9243 0.08042 0.07535 -0.0013 0.0156 1.0000
12.750 0.8936 0.09113 0.08630 -0.0088 0.0160 1.0000
13.000 0.8440 0.10975 0.10510 -0.0213 0.0174 1.0000
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Polar data table (+)
Polar graphs
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