HQ 2.5/9 B AIRFOIL (hq259b-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/9 B AIRFOIL (hq259b-il) Reynolds number: 500,000 Max Cl/Cd: 84.56 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq259b-il-500000-n5.txt Download as CSV file: xf-hq259b-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/9 B AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4920 0.07286 0.07063 -0.0414 1.0000 0.0086
-8.750 -0.5650 0.04192 0.03932 -0.0683 0.9875 0.0085
-8.500 -0.5523 0.03297 0.02986 -0.0742 0.9800 0.0089
-8.250 -0.5315 0.02760 0.02401 -0.0773 0.9736 0.0094
-8.000 -0.5066 0.02305 0.01887 -0.0798 0.9686 0.0102
-7.750 -0.4812 0.02038 0.01573 -0.0808 0.9617 0.0112
-7.500 -0.4514 0.01862 0.01361 -0.0821 0.9567 0.0118
-7.250 -0.4238 0.01733 0.01215 -0.0829 0.9499 0.0124
-7.000 -0.3941 0.01661 0.01132 -0.0837 0.9435 0.0128
-6.750 -0.3657 0.01588 0.01046 -0.0842 0.9366 0.0134
-6.500 -0.3377 0.01514 0.00957 -0.0845 0.9292 0.0142
-6.250 -0.3102 0.01452 0.00880 -0.0846 0.9216 0.0150
-6.000 -0.2828 0.01393 0.00806 -0.0846 0.9139 0.0156
-5.750 -0.2566 0.01311 0.00710 -0.0844 0.9059 0.0161
-5.500 -0.2304 0.01232 0.00622 -0.0843 0.8981 0.0168
-5.250 -0.2039 0.01187 0.00571 -0.0842 0.8902 0.0175
-4.750 -0.1500 0.01115 0.00487 -0.0839 0.8760 0.0194
-4.500 -0.1231 0.01075 0.00438 -0.0838 0.8695 0.0201
-4.250 -0.0960 0.01040 0.00396 -0.0836 0.8631 0.0207
-4.000 -0.0687 0.01011 0.00358 -0.0835 0.8569 0.0212
-3.750 -0.0416 0.00968 0.00310 -0.0835 0.8511 0.0226
-3.500 -0.0142 0.00943 0.00283 -0.0834 0.8449 0.0245
-3.250 0.0134 0.00924 0.00258 -0.0834 0.8395 0.0265
-3.000 0.0410 0.00905 0.00234 -0.0833 0.8322 0.0284
-2.750 0.0683 0.00885 0.00210 -0.0831 0.8232 0.0334
-2.500 0.0956 0.00870 0.00192 -0.0829 0.8129 0.0414
-2.250 0.1231 0.00853 0.00177 -0.0828 0.8046 0.0570
-2.000 0.1505 0.00832 0.00164 -0.0828 0.7969 0.0881
-1.750 0.1779 0.00803 0.00152 -0.0828 0.7895 0.1495
-1.500 0.2043 0.00732 0.00135 -0.0831 0.7819 0.3315
-1.250 0.2305 0.00676 0.00134 -0.0831 0.7740 0.5117
-0.750 0.2846 0.00653 0.00137 -0.0826 0.7576 0.6233
-0.500 0.3118 0.00650 0.00139 -0.0823 0.7498 0.6590
-0.250 0.3391 0.00647 0.00142 -0.0821 0.7422 0.6900
0.000 0.3667 0.00648 0.00142 -0.0819 0.7350 0.7024
0.250 0.3946 0.00650 0.00142 -0.0818 0.7268 0.7114
0.500 0.4222 0.00651 0.00143 -0.0817 0.7181 0.7206
0.750 0.4497 0.00654 0.00145 -0.0815 0.7085 0.7296
1.250 0.5044 0.00659 0.00151 -0.0811 0.6839 0.7488
1.500 0.5313 0.00664 0.00155 -0.0808 0.6670 0.7594
1.750 0.5577 0.00671 0.00159 -0.0804 0.6418 0.7707
2.250 0.6043 0.00731 0.00176 -0.0785 0.5213 0.7967
2.500 0.6260 0.00782 0.00200 -0.0773 0.4523 0.8122
2.750 0.6486 0.00824 0.00221 -0.0764 0.3990 0.8288
3.000 0.6725 0.00853 0.00239 -0.0757 0.3638 0.8468
3.250 0.6969 0.00871 0.00255 -0.0750 0.3423 0.8674
3.500 0.7209 0.00886 0.00270 -0.0742 0.3225 0.8915
3.750 0.7454 0.00897 0.00284 -0.0734 0.3046 0.9291
4.000 0.7771 0.00919 0.00301 -0.0744 0.2840 1.0000
4.250 0.8026 0.00951 0.00322 -0.0741 0.2575 1.0000
4.500 0.8277 0.00986 0.00344 -0.0738 0.2266 1.0000
4.750 0.8514 0.01036 0.00373 -0.0733 0.1829 1.0000
5.000 0.8746 0.01093 0.00408 -0.0727 0.1411 1.0000
5.250 0.8984 0.01144 0.00445 -0.0721 0.1110 1.0000
5.750 0.9472 0.01229 0.00514 -0.0712 0.0777 1.0000
6.250 0.9962 0.01307 0.00587 -0.0703 0.0576 1.0000
6.500 1.0207 0.01346 0.00625 -0.0699 0.0502 1.0000
6.750 1.0446 0.01390 0.00666 -0.0693 0.0418 1.0000
7.000 1.0684 0.01434 0.00708 -0.0688 0.0345 1.0000
7.250 1.0918 0.01481 0.00756 -0.0682 0.0295 1.0000
7.500 1.1148 0.01532 0.00807 -0.0675 0.0259 1.0000
7.750 1.1376 0.01584 0.00862 -0.0668 0.0237 1.0000
8.000 1.1605 0.01631 0.00916 -0.0661 0.0224 1.0000
8.250 1.1828 0.01683 0.00972 -0.0654 0.0209 1.0000
8.500 1.2042 0.01743 0.01037 -0.0645 0.0194 1.0000
8.750 1.2240 0.01819 0.01119 -0.0634 0.0180 1.0000
9.000 1.2456 0.01870 0.01177 -0.0625 0.0173 1.0000
9.250 1.2662 0.01929 0.01244 -0.0616 0.0164 1.0000
9.500 1.2862 0.01990 0.01312 -0.0605 0.0155 1.0000
9.750 1.3054 0.02056 0.01385 -0.0594 0.0146 1.0000
10.000 1.3214 0.02148 0.01482 -0.0579 0.0134 1.0000
10.250 1.3393 0.02217 0.01560 -0.0566 0.0128 1.0000
10.500 1.3565 0.02286 0.01640 -0.0552 0.0122 1.0000
10.750 1.3713 0.02360 0.01723 -0.0534 0.0115 1.0000
11.000 1.3850 0.02434 0.01804 -0.0515 0.0108 1.0000
11.250 1.3977 0.02512 0.01889 -0.0496 0.0102 1.0000
11.500 1.4067 0.02617 0.02001 -0.0472 0.0097 1.0000
11.750 1.4175 0.02711 0.02107 -0.0452 0.0092 1.0000
12.000 1.4273 0.02814 0.02222 -0.0432 0.0088 1.0000
12.250 1.4374 0.02917 0.02335 -0.0414 0.0082 1.0000
12.500 1.4459 0.03035 0.02463 -0.0396 0.0078 1.0000
12.750 1.4541 0.03159 0.02596 -0.0380 0.0075 1.0000
13.000 1.4602 0.03305 0.02752 -0.0363 0.0072 1.0000
13.250 1.4628 0.03488 0.02945 -0.0346 0.0069 1.0000
13.500 1.4639 0.03694 0.03164 -0.0330 0.0067 1.0000
13.750 1.4650 0.03909 0.03394 -0.0317 0.0065 1.0000
14.000 1.4654 0.04143 0.03644 -0.0307 0.0063 1.0000
14.250 1.4641 0.04405 0.03921 -0.0299 0.0061 1.0000
14.500 1.4613 0.04700 0.04232 -0.0294 0.0060 1.0000
14.750 1.4569 0.05029 0.04577 -0.0293 0.0058 1.0000
15.000 1.4505 0.05403 0.04967 -0.0297 0.0057 1.0000
15.250 1.4438 0.05809 0.05388 -0.0305 0.0056 1.0000
15.500 1.4335 0.06296 0.05891 -0.0320 0.0055 1.0000
15.750 1.4234 0.06823 0.06434 -0.0342 0.0054 1.0000
16.000 1.4103 0.07449 0.07078 -0.0372 0.0054 1.0000
16.250 1.3944 0.08170 0.07816 -0.0410 0.0054 1.0000
16.500 1.3777 0.08945 0.08608 -0.0453 0.0053 1.0000
16.750 1.3570 0.09821 0.09502 -0.0503 0.0053 1.0000
17.000 1.3355 0.10738 0.10435 -0.0555 0.0054 1.0000
17.250 1.3118 0.11724 0.11438 -0.0613 0.0054 1.0000
17.500 1.2880 0.12739 0.12468 -0.0672 0.0054 1.0000
17.750 1.2640 0.13807 0.13551 -0.0736 0.0055 1.0000
18.000 1.2391 0.14939 0.14696 -0.0805 0.0055 1.0000
18.250 1.2134 0.16141 0.15913 -0.0878 0.0056 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/9 B AIRFOIL (hq259b-il)