HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/9 AIRFOIL (hq109-il) Reynolds number: 100,000 Max Cl/Cd: 45 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq109-il-100000-n5.txt Download as CSV file: xf-hq109-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5687 0.08885 0.08400 -0.0162 1.0000 0.0211
-9.250 -0.5736 0.08309 0.07830 -0.0198 1.0000 0.0209
-9.000 -0.5824 0.07591 0.07119 -0.0252 1.0000 0.0202
-8.750 -0.6241 0.06176 0.05693 -0.0352 1.0000 0.0168
-8.250 -0.6365 0.05421 0.04905 -0.0367 1.0000 0.0177
-8.000 -0.6346 0.05097 0.04564 -0.0367 1.0000 0.0183
-7.750 -0.6300 0.04763 0.04207 -0.0365 1.0000 0.0191
-7.500 -0.6250 0.04349 0.03756 -0.0360 1.0000 0.0194
-7.250 -0.6168 0.03946 0.03309 -0.0352 1.0000 0.0197
-7.000 -0.6054 0.03567 0.02881 -0.0340 1.0000 0.0200
-6.750 -0.5907 0.03238 0.02504 -0.0327 1.0000 0.0205
-6.500 -0.5735 0.02951 0.02171 -0.0313 1.0000 0.0214
-6.250 -0.5545 0.02702 0.01881 -0.0299 1.0000 0.0225
-6.000 -0.5339 0.02523 0.01667 -0.0286 1.0000 0.0247
-5.750 -0.5149 0.02346 0.01476 -0.0274 1.0000 0.0280
-5.500 -0.4950 0.02207 0.01326 -0.0261 1.0000 0.0308
-5.250 -0.4753 0.02069 0.01173 -0.0245 1.0000 0.0343
-5.000 -0.4568 0.01944 0.01034 -0.0230 1.0000 0.0394
-4.750 -0.4373 0.01870 0.00950 -0.0216 1.0000 0.0492
-4.500 -0.4193 0.01776 0.00857 -0.0202 1.0000 0.0610
-4.250 -0.4009 0.01700 0.00783 -0.0188 1.0000 0.0770
-4.000 -0.3825 0.01628 0.00719 -0.0175 1.0000 0.0991
-3.750 -0.3634 0.01559 0.00661 -0.0163 0.9999 0.1285
-3.500 -0.3302 0.01464 0.00605 -0.0183 0.9927 0.2170
-3.250 -0.2996 0.01346 0.00562 -0.0199 0.9856 0.3874
-3.000 -0.2705 0.01287 0.00564 -0.0202 0.9778 0.5473
-2.750 -0.2371 0.01275 0.00566 -0.0211 0.9711 0.6316
-2.500 -0.2067 0.01272 0.00569 -0.0211 0.9628 0.6907
-2.250 -0.1738 0.01273 0.00570 -0.0217 0.9561 0.7336
-2.000 -0.1418 0.01272 0.00560 -0.0222 0.9481 0.7625
-1.750 -0.1102 0.01271 0.00553 -0.0226 0.9404 0.7875
-1.500 -0.0786 0.01268 0.00547 -0.0229 0.9328 0.8127
-1.250 -0.0488 0.01265 0.00538 -0.0229 0.9241 0.8358
-1.000 -0.0148 0.01259 0.00528 -0.0237 0.9174 0.8547
-0.750 0.0162 0.01254 0.00519 -0.0241 0.9080 0.8702
-0.500 0.0504 0.01249 0.00510 -0.0252 0.9000 0.8844
-0.250 0.0853 0.01244 0.00501 -0.0264 0.8917 0.8987
0.000 0.1207 0.01241 0.00496 -0.0278 0.8822 0.9131
0.250 0.1590 0.01236 0.00490 -0.0297 0.8735 0.9270
0.500 0.1974 0.01232 0.00486 -0.0318 0.8640 0.9416
0.750 0.2357 0.01230 0.00483 -0.0339 0.8536 0.9566
1.000 0.2746 0.01223 0.00477 -0.0360 0.8407 0.9712
1.250 0.3132 0.01213 0.00468 -0.0380 0.8237 0.9858
1.500 0.3474 0.01206 0.00459 -0.0392 0.8055 1.0000
1.750 0.3674 0.01209 0.00460 -0.0377 0.7879 1.0000
2.000 0.3893 0.01215 0.00464 -0.0365 0.7721 1.0000
2.250 0.4123 0.01222 0.00474 -0.0355 0.7564 1.0000
2.500 0.4360 0.01229 0.00482 -0.0345 0.7397 1.0000
2.750 0.4598 0.01237 0.00492 -0.0335 0.7209 1.0000
3.000 0.4839 0.01246 0.00503 -0.0326 0.7000 1.0000
3.250 0.5080 0.01255 0.00514 -0.0316 0.6759 1.0000
3.500 0.5319 0.01266 0.00529 -0.0305 0.6467 1.0000
3.750 0.5556 0.01280 0.00541 -0.0293 0.6106 1.0000
4.000 0.5787 0.01301 0.00553 -0.0281 0.5638 1.0000
4.250 0.6008 0.01335 0.00567 -0.0267 0.5018 1.0000
4.500 0.6218 0.01388 0.00591 -0.0253 0.4297 1.0000
4.750 0.6425 0.01455 0.00634 -0.0241 0.3593 1.0000
5.000 0.6632 0.01531 0.00683 -0.0230 0.3006 1.0000
5.250 0.6844 0.01607 0.00741 -0.0222 0.2536 1.0000
5.500 0.7062 0.01682 0.00804 -0.0214 0.2086 1.0000
5.750 0.7278 0.01761 0.00870 -0.0206 0.1632 1.0000
6.000 0.7484 0.01856 0.00944 -0.0198 0.1216 1.0000
6.250 0.7688 0.01961 0.01034 -0.0189 0.0929 1.0000
6.500 0.7894 0.02066 0.01146 -0.0179 0.0718 1.0000
6.750 0.8090 0.02187 0.01267 -0.0169 0.0571 1.0000
7.000 0.8285 0.02312 0.01398 -0.0158 0.0450 1.0000
7.250 0.8477 0.02447 0.01543 -0.0146 0.0353 1.0000
7.500 0.8657 0.02600 0.01706 -0.0134 0.0285 1.0000
7.750 0.8848 0.02727 0.01844 -0.0125 0.0229 1.0000
8.000 0.9005 0.02950 0.02080 -0.0111 0.0205 1.0000
8.250 0.9190 0.03169 0.02329 -0.0097 0.0189 1.0000
8.500 0.9363 0.03422 0.02614 -0.0084 0.0177 1.0000
8.750 0.9515 0.03701 0.02929 -0.0070 0.0169 1.0000
9.000 0.9633 0.03995 0.03270 -0.0055 0.0161 1.0000
9.250 0.9720 0.04269 0.03578 -0.0040 0.0153 1.0000
9.500 0.9782 0.04500 0.03832 -0.0026 0.0143 1.0000
9.750 0.9791 0.04786 0.04141 -0.0011 0.0136 1.0000
10.250 0.9571 0.05509 0.04918 0.0031 0.0130 1.0000
10.500 0.9430 0.05866 0.05302 0.0041 0.0129 1.0000
10.750 0.9274 0.06285 0.05743 0.0038 0.0129 1.0000
11.000 0.9109 0.06773 0.06249 0.0021 0.0130 1.0000
11.250 0.8927 0.07356 0.06850 -0.0011 0.0130 1.0000
11.500 0.7428 0.08574 0.08139 -0.0085 0.0172 1.0000
11.750 0.7159 0.09681 0.09252 -0.0148 0.0179 1.0000
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Polar data table (+)
Polar graphs
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