HQ 0/10 AIRFOIL (hq010-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 0/10 AIRFOIL (hq010-il) Reynolds number: 100,000 Max Cl/Cd: 36.04 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq010-il-100000-n5.txt Download as CSV file: xf-hq010-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7909 0.07517 0.07012 -0.0179 1.0000 0.0211
-11.250 -0.8098 0.06888 0.06371 -0.0222 1.0000 0.0210
-11.000 -0.8304 0.06326 0.05793 -0.0250 1.0000 0.0208
-10.750 -0.8493 0.05873 0.05320 -0.0263 1.0000 0.0209
-10.500 -0.8676 0.05487 0.04909 -0.0259 1.0000 0.0210
-10.250 -0.8844 0.05153 0.04550 -0.0239 1.0000 0.0214
-10.000 -0.8878 0.04869 0.04256 -0.0226 1.0000 0.0225
-9.750 -0.8830 0.04700 0.04077 -0.0215 1.0000 0.0240
-9.500 -0.8781 0.04484 0.03838 -0.0203 1.0000 0.0259
-9.250 -0.8730 0.04222 0.03535 -0.0187 1.0000 0.0284
-9.000 -0.8659 0.03904 0.03161 -0.0170 1.0000 0.0300
-8.750 -0.8536 0.03626 0.02835 -0.0156 1.0000 0.0311
-8.500 -0.8387 0.03392 0.02593 -0.0148 1.0000 0.0336
-8.250 -0.8208 0.03233 0.02414 -0.0140 1.0000 0.0361
-8.000 -0.8009 0.03027 0.02178 -0.0130 1.0000 0.0378
-7.750 -0.7795 0.02840 0.01959 -0.0121 1.0000 0.0396
-7.500 -0.7575 0.02680 0.01774 -0.0112 1.0000 0.0414
-7.250 -0.7367 0.02523 0.01609 -0.0104 1.0000 0.0440
-7.000 -0.7160 0.02416 0.01500 -0.0096 1.0000 0.0479
-6.750 -0.6946 0.02308 0.01381 -0.0087 1.0000 0.0520
-6.500 -0.6740 0.02193 0.01255 -0.0076 1.0000 0.0555
-6.250 -0.6544 0.02086 0.01149 -0.0065 1.0000 0.0604
-6.000 -0.6335 0.02002 0.01054 -0.0055 1.0000 0.0678
-5.750 -0.6134 0.01914 0.00963 -0.0045 1.0000 0.0797
-5.500 -0.5936 0.01824 0.00878 -0.0033 1.0000 0.0946
-5.250 -0.5746 0.01731 0.00800 -0.0022 1.0000 0.1175
-5.000 -0.5566 0.01635 0.00742 -0.0009 1.0000 0.1682
-4.750 -0.5378 0.01563 0.00691 0.0003 1.0000 0.2279
-4.500 -0.5205 0.01488 0.00644 0.0017 1.0000 0.2865
-4.250 -0.5049 0.01412 0.00607 0.0035 1.0000 0.3593
-4.000 -0.4891 0.01357 0.00584 0.0055 1.0000 0.4400
-3.750 -0.4739 0.01318 0.00578 0.0078 1.0000 0.5101
-3.500 -0.4488 0.01301 0.00580 0.0082 0.9942 0.5749
-3.250 -0.4152 0.01299 0.00587 0.0073 0.9837 0.6342
-3.000 -0.3817 0.01304 0.00591 0.0064 0.9741 0.6769
-2.750 -0.3452 0.01306 0.00586 0.0048 0.9662 0.7014
-2.500 -0.3104 0.01306 0.00577 0.0035 0.9570 0.7185
-2.250 -0.2762 0.01305 0.00570 0.0023 0.9478 0.7340
-2.000 -0.2406 0.01305 0.00561 0.0009 0.9400 0.7483
-1.750 -0.2091 0.01305 0.00557 0.0003 0.9295 0.7626
-1.500 -0.1776 0.01307 0.00556 -0.0001 0.9192 0.7784
-1.250 -0.1465 0.01310 0.00555 -0.0003 0.9084 0.7956
-1.000 -0.1158 0.01311 0.00554 -0.0003 0.8978 0.8117
-0.750 -0.0872 0.01311 0.00553 -0.0001 0.8865 0.8251
-0.500 -0.0586 0.01312 0.00553 0.0001 0.8762 0.8367
-0.250 -0.0290 0.01311 0.00551 0.0000 0.8671 0.8474
0.000 0.0000 0.01311 0.00549 0.0000 0.8578 0.8578
0.250 0.0290 0.01311 0.00551 0.0000 0.8474 0.8671
0.500 0.0586 0.01312 0.00553 -0.0001 0.8366 0.8762
0.750 0.0872 0.01311 0.00553 0.0001 0.8249 0.8865
1.000 0.1158 0.01311 0.00554 0.0003 0.8117 0.8978
1.250 0.1465 0.01310 0.00555 0.0003 0.7956 0.9084
1.500 0.1776 0.01307 0.00556 0.0001 0.7784 0.9192
1.750 0.2091 0.01305 0.00557 -0.0003 0.7626 0.9295
2.000 0.2406 0.01305 0.00561 -0.0009 0.7483 0.9399
2.250 0.2762 0.01305 0.00570 -0.0023 0.7341 0.9478
2.500 0.3105 0.01306 0.00577 -0.0035 0.7185 0.9570
2.750 0.3452 0.01306 0.00586 -0.0048 0.7014 0.9662
3.000 0.3817 0.01304 0.00590 -0.0064 0.6768 0.9741
3.250 0.4152 0.01299 0.00587 -0.0073 0.6342 0.9837
3.500 0.4488 0.01301 0.00580 -0.0083 0.5751 0.9942
3.750 0.4738 0.01318 0.00578 -0.0078 0.5099 1.0000
4.000 0.4891 0.01357 0.00584 -0.0055 0.4401 1.0000
4.250 0.5049 0.01412 0.00607 -0.0035 0.3599 1.0000
4.500 0.5205 0.01488 0.00644 -0.0018 0.2867 1.0000
4.750 0.5379 0.01563 0.00691 -0.0003 0.2282 1.0000
5.000 0.5566 0.01635 0.00742 0.0009 0.1682 1.0000
5.250 0.5746 0.01731 0.00800 0.0022 0.1173 1.0000
5.500 0.5936 0.01824 0.00878 0.0033 0.0946 1.0000
5.750 0.6134 0.01914 0.00963 0.0045 0.0796 1.0000
6.000 0.6334 0.02003 0.01055 0.0055 0.0676 1.0000
6.250 0.6545 0.02086 0.01148 0.0065 0.0604 1.0000
6.500 0.6739 0.02195 0.01257 0.0076 0.0554 1.0000
6.750 0.6946 0.02308 0.01381 0.0087 0.0520 1.0000
7.000 0.7160 0.02416 0.01500 0.0096 0.0479 1.0000
7.250 0.7366 0.02524 0.01609 0.0104 0.0439 1.0000
7.500 0.7575 0.02680 0.01775 0.0112 0.0414 1.0000
7.750 0.7795 0.02840 0.01959 0.0121 0.0395 1.0000
8.000 0.8008 0.03027 0.02179 0.0130 0.0378 1.0000
8.250 0.8208 0.03233 0.02414 0.0140 0.0361 1.0000
8.500 0.8387 0.03393 0.02594 0.0148 0.0337 1.0000
8.750 0.8536 0.03625 0.02833 0.0156 0.0311 1.0000
9.000 0.8658 0.03906 0.03164 0.0171 0.0300 1.0000
9.250 0.8731 0.04219 0.03532 0.0187 0.0284 1.0000
9.500 0.8778 0.04493 0.03848 0.0203 0.0260 1.0000
9.750 0.8837 0.04685 0.04061 0.0215 0.0239 1.0000
10.000 0.8886 0.04854 0.04239 0.0226 0.0224 1.0000
10.250 0.8843 0.05155 0.04553 0.0239 0.0215 1.0000
10.500 0.8682 0.05496 0.04916 0.0258 0.0211 1.0000
10.750 0.8504 0.05872 0.05317 0.0263 0.0210 1.0000
11.000 0.8307 0.06331 0.05797 0.0250 0.0209 1.0000
11.250 0.8115 0.06868 0.06351 0.0222 0.0210 1.0000
11.500 0.7906 0.07535 0.07029 0.0178 0.0212 1.0000
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Polar data table (+)
Polar graphs
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