HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il) Reynolds number: 1,000,000 Max Cl/Cd: 86.07 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq159b-il-1000000.txt Download as CSV file: xf-hq159b-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/9 B AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5386 0.08031 0.07873 -0.0245 1.0000 0.0099
-9.250 -0.5463 0.07351 0.07195 -0.0288 1.0000 0.0099
-9.000 -0.7299 0.03151 0.02898 -0.0460 1.0000 0.0057
-8.750 -0.7335 0.02564 0.02251 -0.0441 1.0000 0.0059
-8.500 -0.7232 0.02268 0.01918 -0.0424 1.0000 0.0061
-8.250 -0.7085 0.02077 0.01700 -0.0409 1.0000 0.0063
-8.000 -0.6926 0.01937 0.01540 -0.0392 1.0000 0.0064
-7.750 -0.6731 0.01817 0.01402 -0.0382 0.9996 0.0065
-7.500 -0.6415 0.01656 0.01217 -0.0397 0.9972 0.0067
-7.250 -0.6113 0.01415 0.00942 -0.0411 0.9942 0.0069
-7.000 -0.5814 0.01273 0.00780 -0.0421 0.9905 0.0072
-6.750 -0.5490 0.01190 0.00688 -0.0434 0.9872 0.0076
-6.500 -0.5149 0.01138 0.00631 -0.0450 0.9846 0.0083
-6.250 -0.4815 0.01079 0.00565 -0.0464 0.9810 0.0089
-6.000 -0.4502 0.01030 0.00510 -0.0473 0.9750 0.0097
-5.750 -0.4179 0.00951 0.00419 -0.0484 0.9698 0.0114
-5.500 -0.3890 0.00920 0.00387 -0.0487 0.9608 0.0134
-5.250 -0.3609 0.00880 0.00344 -0.0488 0.9516 0.0177
-5.000 -0.3335 0.00870 0.00333 -0.0486 0.9414 0.0207
-4.750 -0.3080 0.00837 0.00298 -0.0481 0.9291 0.0263
-4.500 -0.2815 0.00827 0.00284 -0.0478 0.9170 0.0290
-4.250 -0.2555 0.00800 0.00250 -0.0473 0.9050 0.0337
-4.000 -0.2289 0.00784 0.00231 -0.0470 0.8940 0.0385
-3.750 -0.2023 0.00766 0.00209 -0.0467 0.8837 0.0460
-3.500 -0.1754 0.00748 0.00190 -0.0465 0.8736 0.0561
-3.250 -0.1485 0.00728 0.00173 -0.0463 0.8643 0.0727
-3.000 -0.1217 0.00706 0.00156 -0.0461 0.8557 0.0996
-2.750 -0.0957 0.00661 0.00138 -0.0460 0.8459 0.1841
-2.500 -0.0696 0.00622 0.00122 -0.0459 0.8354 0.2647
-2.250 -0.0443 0.00565 0.00106 -0.0457 0.8254 0.4001
-2.000 -0.0179 0.00534 0.00097 -0.0455 0.8155 0.4814
-1.750 0.0088 0.00510 0.00093 -0.0453 0.8053 0.5548
-1.500 0.0357 0.00497 0.00091 -0.0451 0.7956 0.6089
-1.250 0.0632 0.00493 0.00089 -0.0449 0.7861 0.6380
-0.750 0.1186 0.00488 0.00087 -0.0447 0.7660 0.6781
-0.500 0.1462 0.00489 0.00086 -0.0445 0.7546 0.6945
-0.250 0.1737 0.00490 0.00085 -0.0443 0.7424 0.7108
0.000 0.2014 0.00490 0.00086 -0.0442 0.7315 0.7240
0.250 0.2292 0.00490 0.00087 -0.0441 0.7223 0.7365
0.500 0.2569 0.00492 0.00089 -0.0440 0.7131 0.7477
0.750 0.2849 0.00495 0.00090 -0.0440 0.7031 0.7572
1.000 0.3126 0.00495 0.00093 -0.0439 0.6929 0.7673
1.250 0.3403 0.00497 0.00096 -0.0438 0.6816 0.7776
1.500 0.3678 0.00501 0.00099 -0.0436 0.6677 0.7887
1.750 0.3949 0.00505 0.00103 -0.0434 0.6490 0.8012
2.250 0.4470 0.00533 0.00112 -0.0426 0.5673 0.8269
2.500 0.4721 0.00557 0.00121 -0.0421 0.5115 0.8417
2.750 0.4965 0.00587 0.00134 -0.0415 0.4548 0.8592
3.000 0.5206 0.00612 0.00149 -0.0408 0.4049 0.8821
3.250 0.5426 0.00645 0.00166 -0.0396 0.3336 0.9188
3.500 0.5756 0.00681 0.00185 -0.0410 0.2756 0.9738
3.750 0.6068 0.00715 0.00203 -0.0420 0.2386 1.0000
4.000 0.6332 0.00739 0.00220 -0.0418 0.2172 1.0000
4.250 0.6593 0.00766 0.00238 -0.0416 0.1939 1.0000
4.500 0.6845 0.00805 0.00259 -0.0413 0.1493 1.0000
4.750 0.7075 0.00875 0.00297 -0.0407 0.0862 1.0000
5.000 0.7308 0.00943 0.00337 -0.0401 0.0376 1.0000
5.250 0.7562 0.00981 0.00372 -0.0398 0.0303 1.0000
5.500 0.7821 0.01013 0.00408 -0.0394 0.0272 1.0000
5.750 0.8081 0.01042 0.00440 -0.0392 0.0256 1.0000
6.000 0.8337 0.01075 0.00476 -0.0388 0.0238 1.0000
6.250 0.8585 0.01119 0.00523 -0.0384 0.0214 1.0000
6.500 0.8821 0.01182 0.00595 -0.0377 0.0191 1.0000
6.750 0.9084 0.01200 0.00613 -0.0376 0.0180 1.0000
7.000 0.9338 0.01231 0.00647 -0.0373 0.0163 1.0000
7.250 0.9584 0.01272 0.00687 -0.0369 0.0143 1.0000
7.500 0.9807 0.01345 0.00768 -0.0360 0.0125 1.0000
7.750 1.0062 0.01371 0.00796 -0.0358 0.0113 1.0000
8.000 1.0315 0.01396 0.00819 -0.0355 0.0100 1.0000
8.250 1.0512 0.01499 0.00931 -0.0343 0.0086 1.0000
8.500 1.0753 0.01540 0.00977 -0.0338 0.0082 1.0000
8.750 1.0981 0.01595 0.01038 -0.0332 0.0076 1.0000
9.000 1.1205 0.01654 0.01102 -0.0325 0.0071 1.0000
9.250 1.1429 0.01709 0.01162 -0.0318 0.0066 1.0000
9.500 1.1623 0.01799 0.01260 -0.0307 0.0062 1.0000
9.750 1.1735 0.01992 0.01474 -0.0285 0.0057 1.0000
10.000 1.1906 0.02106 0.01601 -0.0271 0.0056 1.0000
10.250 1.2088 0.02196 0.01702 -0.0260 0.0054 1.0000
10.500 1.2253 0.02302 0.01821 -0.0246 0.0052 1.0000
10.750 1.2391 0.02434 0.01968 -0.0229 0.0050 1.0000
11.000 1.2520 0.02562 0.02110 -0.0212 0.0048 1.0000
11.250 1.2641 0.02680 0.02241 -0.0194 0.0046 1.0000
11.500 1.2720 0.02803 0.02377 -0.0170 0.0044 1.0000
11.750 1.2789 0.02903 0.02488 -0.0145 0.0043 1.0000
12.000 1.2863 0.02996 0.02589 -0.0123 0.0042 1.0000
12.250 1.2853 0.03172 0.02780 -0.0096 0.0041 1.0000
12.500 1.2854 0.03347 0.02967 -0.0074 0.0040 1.0000
12.750 1.2810 0.03577 0.03212 -0.0055 0.0039 1.0000
13.000 1.2750 0.03844 0.03497 -0.0040 0.0039 1.0000
13.250 1.2663 0.04162 0.03832 -0.0032 0.0039 1.0000
13.500 1.2548 0.04541 0.04228 -0.0030 0.0039 1.0000
13.750 1.2407 0.04991 0.04695 -0.0038 0.0038 1.0000
14.000 1.2253 0.05511 0.05233 -0.0055 0.0038 1.0000
14.250 1.2056 0.06158 0.05900 -0.0085 0.0038 1.0000
14.500 1.1882 0.06852 0.06609 -0.0126 0.0038 1.0000
14.750 1.1654 0.07734 0.07511 -0.0181 0.0039 1.0000
15.000 1.1361 0.08855 0.08652 -0.0253 0.0039 1.0000
15.250 1.1157 0.09843 0.09652 -0.0314 0.0039 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/9 B AIRFOIL (hq159b-il)