HQ 1.5/11 AIRFOIL (hq1511-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/11 AIRFOIL (hq1511-il) Reynolds number: 50,000 Max Cl/Cd: 33.18 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1511-il-50000.txt Download as CSV file: xf-hq1511-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/11 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.4588 0.11862 0.11148 -0.0075 1.0000 0.2651
-10.000 -0.4612 0.11557 0.10850 -0.0083 1.0000 0.2751
-9.750 -0.4765 0.11431 0.10735 -0.0095 1.0000 0.2870
-9.500 -0.4620 0.11003 0.10305 -0.0088 1.0000 0.3013
-9.250 -0.4618 0.10705 0.10013 -0.0086 1.0000 0.3167
-9.000 -0.4447 0.10230 0.09538 -0.0076 1.0000 0.3338
-8.750 -0.4285 0.09849 0.09157 -0.0065 1.0000 0.3547
-8.500 -0.4278 0.09560 0.08873 -0.0059 1.0000 0.3728
-8.250 -0.4216 0.09218 0.08534 -0.0055 1.0000 0.3869
-8.000 -0.5557 0.07051 0.06405 -0.0398 1.0000 0.1502
-7.750 -0.5981 0.06219 0.05526 -0.0440 1.0000 0.1272
-7.500 -0.5955 0.05794 0.05084 -0.0431 1.0000 0.1249
-7.250 -0.5939 0.05377 0.04648 -0.0420 1.0000 0.1236
-7.000 -0.5922 0.04978 0.04218 -0.0406 1.0000 0.1230
-6.750 -0.5893 0.04607 0.03800 -0.0389 1.0000 0.1243
-6.500 -0.5846 0.04276 0.03398 -0.0369 1.0000 0.1265
-6.250 -0.5712 0.04019 0.03158 -0.0352 1.0000 0.1365
-6.000 -0.5597 0.03714 0.02777 -0.0333 1.0000 0.1406
-5.750 -0.5438 0.03483 0.02552 -0.0317 1.0000 0.1508
-5.500 -0.5260 0.03228 0.02249 -0.0302 1.0000 0.1565
-5.250 -0.5075 0.03056 0.02051 -0.0287 1.0000 0.1674
-5.000 -0.4878 0.02885 0.01873 -0.0273 1.0000 0.1784
-4.750 -0.4672 0.02729 0.01707 -0.0259 1.0000 0.1912
-4.500 -0.4468 0.02596 0.01567 -0.0245 1.0000 0.2100
-4.250 -0.4263 0.02456 0.01439 -0.0229 1.0000 0.2328
-4.000 -0.4068 0.02316 0.01319 -0.0214 1.0000 0.2733
-3.750 -0.3910 0.02093 0.01211 -0.0194 1.0000 0.3874
-3.500 -0.3946 0.02058 0.01325 -0.0103 1.0000 0.6437
-3.250 -0.3935 0.02124 0.01396 -0.0022 1.0000 0.7261
-3.000 -0.3909 0.02161 0.01430 0.0050 1.0000 0.7820
-2.750 -0.3861 0.02178 0.01439 0.0117 1.0000 0.8311
-2.500 -0.3697 0.02201 0.01447 0.0167 1.0000 0.8857
-2.250 -0.1359 0.02366 0.01491 -0.0148 1.0000 0.9794
-2.000 -0.0670 0.02312 0.01402 -0.0247 1.0000 1.0000
-1.750 -0.0770 0.02288 0.01374 -0.0210 1.0000 1.0000
-1.500 -0.0871 0.02261 0.01343 -0.0173 1.0000 1.0000
-1.250 -0.0975 0.02228 0.01306 -0.0135 1.0000 1.0000
-1.000 -0.1081 0.02190 0.01264 -0.0095 1.0000 1.0000
-0.750 -0.1168 0.02150 0.01219 -0.0058 1.0000 1.0000
-0.500 -0.1178 0.02125 0.01183 -0.0032 1.0000 1.0000
-0.250 -0.1076 0.02125 0.01168 -0.0022 1.0000 1.0000
0.000 -0.0919 0.02142 0.01170 -0.0021 1.0000 1.0000
0.250 -0.0739 0.02171 0.01183 -0.0022 1.0000 1.0000
0.500 -0.0548 0.02208 0.01207 -0.0024 1.0000 1.0000
0.750 -0.0356 0.02253 0.01241 -0.0026 1.0000 1.0000
1.000 -0.0162 0.02304 0.01281 -0.0029 1.0000 1.0000
1.250 0.0029 0.02362 0.01330 -0.0032 1.0000 1.0000
1.500 0.0279 0.02439 0.01401 -0.0046 0.9971 1.0000
1.750 0.0783 0.02589 0.01545 -0.0107 0.9834 1.0000
2.000 0.1254 0.02715 0.01670 -0.0160 0.9665 1.0000
2.250 0.1896 0.02859 0.01816 -0.0238 0.9419 1.0000
2.500 0.2452 0.02948 0.01911 -0.0294 0.9145 1.0000
2.750 0.2921 0.03020 0.01990 -0.0332 0.8903 1.0000
3.000 0.3453 0.03086 0.02069 -0.0377 0.8692 1.0000
3.250 0.3825 0.03144 0.02137 -0.0397 0.8476 1.0000
3.500 0.4283 0.03187 0.02195 -0.0425 0.8266 1.0000
3.750 0.4713 0.03218 0.02245 -0.0446 0.8047 1.0000
4.000 0.5183 0.03222 0.02270 -0.0469 0.7820 1.0000
4.250 0.5613 0.03209 0.02281 -0.0480 0.7577 1.0000
4.500 0.6169 0.03113 0.02211 -0.0495 0.7348 1.0000
4.750 0.6537 0.03058 0.02176 -0.0486 0.7069 1.0000
5.000 0.6925 0.02967 0.02107 -0.0474 0.6777 1.0000
5.250 0.7323 0.02839 0.01994 -0.0457 0.6467 1.0000
5.500 0.7635 0.02743 0.01908 -0.0433 0.6112 1.0000
5.750 0.7947 0.02648 0.01815 -0.0408 0.5750 1.0000
6.000 0.8220 0.02603 0.01766 -0.0385 0.5380 1.0000
6.250 0.8483 0.02588 0.01740 -0.0364 0.5004 1.0000
6.500 0.8707 0.02624 0.01768 -0.0342 0.4603 1.0000
6.750 0.8914 0.02693 0.01819 -0.0320 0.4171 1.0000
7.000 0.9094 0.02785 0.01891 -0.0296 0.3704 1.0000
7.250 0.9275 0.02899 0.01966 -0.0273 0.3222 1.0000
7.500 0.9439 0.03078 0.02115 -0.0252 0.2758 1.0000
7.750 0.9598 0.03267 0.02290 -0.0233 0.2360 1.0000
8.000 0.9777 0.03466 0.02466 -0.0217 0.2037 1.0000
8.250 0.9926 0.03662 0.02679 -0.0198 0.1787 1.0000
8.500 1.0122 0.03875 0.02888 -0.0186 0.1604 1.0000
8.750 1.0293 0.04108 0.03137 -0.0171 0.1461 1.0000
9.000 1.0505 0.04379 0.03411 -0.0163 0.1349 1.0000
9.250 1.0577 0.04740 0.03841 -0.0141 0.1295 1.0000
9.500 1.0792 0.05037 0.04121 -0.0137 0.1211 1.0000
9.750 1.0755 0.05422 0.04577 -0.0110 0.1186 1.0000
10.000 1.0705 0.05851 0.05056 -0.0087 0.1171 1.0000
10.250 1.0607 0.06301 0.05547 -0.0066 0.1168 1.0000
10.500 1.0452 0.06763 0.06042 -0.0047 0.1171 1.0000
10.750 1.0236 0.07216 0.06519 -0.0028 0.1178 1.0000
11.000 0.9994 0.07696 0.07013 -0.0019 0.1188 1.0000
11.250 0.9758 0.08242 0.07570 -0.0023 0.1198 1.0000
11.500 0.9583 0.08841 0.08176 -0.0037 0.1208 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/11 AIRFOIL (hq1511-il)