HQ 1.0/12 AIRFOIL (hq1012-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.0/12 AIRFOIL (hq1012-il) Reynolds number: 1,000,000 Max Cl/Cd: 87.52 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1012-il-1000000.txt Download as CSV file: xf-hq1012-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/12 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
------ -------- --------- --------- -------- -------- --------
-16.750 -0.8408 0.12107 0.11905 -0.0078 1.0000 0.0063
-16.500 -0.8722 0.10969 0.10756 -0.0139 1.0000 0.0063
-16.250 -0.9228 0.09504 0.09272 -0.0219 1.0000 0.0061
-16.000 -0.9462 0.08612 0.08369 -0.0270 1.0000 0.0061
-15.750 -0.9877 0.07393 0.07130 -0.0342 1.0000 0.0060
-15.500 -1.0082 0.06619 0.06340 -0.0390 1.0000 0.0059
-15.250 -1.0278 0.05908 0.05614 -0.0434 1.0000 0.0059
-15.000 -1.0439 0.05301 0.04992 -0.0470 1.0000 0.0060
-14.750 -1.0604 0.04747 0.04421 -0.0499 1.0000 0.0059
-14.500 -1.0671 0.04355 0.04017 -0.0517 1.0000 0.0059
-14.250 -1.0753 0.03976 0.03624 -0.0529 1.0000 0.0060
-14.000 -1.0733 0.03734 0.03373 -0.0536 1.0000 0.0062
-13.750 -1.0760 0.03460 0.03086 -0.0538 1.0000 0.0062
-13.500 -1.0791 0.03202 0.02815 -0.0536 1.0000 0.0062
-13.250 -1.0788 0.02994 0.02595 -0.0530 1.0000 0.0064
-13.000 -1.0780 0.02810 0.02400 -0.0519 1.0000 0.0063
-12.750 -1.0754 0.02657 0.02236 -0.0505 1.0000 0.0065
-12.500 -1.0727 0.02525 0.02093 -0.0484 1.0000 0.0066
-12.250 -1.0691 0.02416 0.01973 -0.0458 1.0000 0.0067
-12.000 -1.0633 0.02321 0.01868 -0.0432 1.0000 0.0068
-11.750 -1.0521 0.02224 0.01762 -0.0413 1.0000 0.0069
-11.500 -1.0373 0.02147 0.01677 -0.0399 1.0000 0.0070
-11.250 -1.0236 0.02054 0.01574 -0.0382 1.0000 0.0072
-11.000 -1.0162 0.01899 0.01404 -0.0358 1.0000 0.0075
-10.750 -1.0018 0.01803 0.01301 -0.0342 1.0000 0.0079
-10.500 -0.9834 0.01743 0.01236 -0.0329 1.0000 0.0082
-10.250 -0.9650 0.01683 0.01171 -0.0316 1.0000 0.0086
-10.000 -0.9463 0.01628 0.01112 -0.0303 1.0000 0.0090
-9.750 -0.9277 0.01576 0.01053 -0.0289 1.0000 0.0094
-9.500 -0.9090 0.01530 0.01002 -0.0275 1.0000 0.0097
-9.250 -0.8907 0.01447 0.00911 -0.0262 0.9993 0.0103
-9.000 -0.8577 0.01370 0.00830 -0.0279 0.9953 0.0112
-8.750 -0.8251 0.01317 0.00773 -0.0293 0.9900 0.0122
-8.500 -0.7908 0.01272 0.00723 -0.0310 0.9852 0.0132
-8.250 -0.7574 0.01204 0.00654 -0.0326 0.9793 0.0157
-8.000 -0.7233 0.01157 0.00604 -0.0343 0.9714 0.0186
-7.750 -0.6908 0.01119 0.00567 -0.0355 0.9601 0.0226
-7.500 -0.6618 0.01080 0.00530 -0.0360 0.9446 0.0280
-7.250 -0.6354 0.01056 0.00499 -0.0358 0.9270 0.0316
-7.000 -0.6106 0.01031 0.00471 -0.0352 0.9110 0.0357
-6.750 -0.5849 0.01018 0.00451 -0.0348 0.8976 0.0380
-6.500 -0.5601 0.00989 0.00419 -0.0342 0.8853 0.0431
-6.250 -0.5343 0.00970 0.00397 -0.0339 0.8744 0.0469
-6.000 -0.5084 0.00952 0.00373 -0.0335 0.8646 0.0509
-5.750 -0.4826 0.00928 0.00348 -0.0332 0.8551 0.0575
-5.500 -0.4563 0.00908 0.00326 -0.0329 0.8464 0.0646
-5.250 -0.4302 0.00887 0.00304 -0.0326 0.8380 0.0747
-5.000 -0.4041 0.00860 0.00283 -0.0324 0.8298 0.0902
-4.500 -0.3521 0.00803 0.00240 -0.0319 0.8143 0.1403
-4.250 -0.3264 0.00773 0.00221 -0.0316 0.8070 0.1775
-4.000 -0.2998 0.00746 0.00205 -0.0314 0.7994 0.2123
-3.750 -0.2734 0.00723 0.00190 -0.0312 0.7922 0.2448
-3.500 -0.2470 0.00693 0.00174 -0.0311 0.7846 0.2856
-3.250 -0.2216 0.00657 0.00157 -0.0308 0.7774 0.3490
-3.000 -0.1955 0.00623 0.00144 -0.0306 0.7700 0.4132
-2.750 -0.1691 0.00599 0.00134 -0.0303 0.7633 0.4675
-2.500 -0.1422 0.00578 0.00128 -0.0302 0.7566 0.5154
-2.250 -0.1147 0.00569 0.00123 -0.0301 0.7495 0.5438
-2.000 -0.0868 0.00562 0.00119 -0.0300 0.7411 0.5656
-1.500 -0.0311 0.00554 0.00114 -0.0299 0.7249 0.6070
-1.000 0.0252 0.00550 0.00112 -0.0298 0.7099 0.6369
-0.750 0.0532 0.00552 0.00111 -0.0298 0.7023 0.6476
-0.500 0.0819 0.00551 0.00110 -0.0298 0.6948 0.6564
-0.250 0.1099 0.00552 0.00109 -0.0298 0.6869 0.6640
0.000 0.1384 0.00553 0.00109 -0.0298 0.6777 0.6730
0.250 0.1665 0.00553 0.00110 -0.0298 0.6698 0.6817
0.500 0.1948 0.00555 0.00111 -0.0298 0.6617 0.6900
0.750 0.2230 0.00556 0.00113 -0.0298 0.6533 0.6982
1.000 0.2512 0.00559 0.00115 -0.0298 0.6443 0.7060
1.250 0.2794 0.00560 0.00118 -0.0298 0.6349 0.7145
1.500 0.3074 0.00563 0.00121 -0.0298 0.6244 0.7226
1.750 0.3354 0.00568 0.00124 -0.0297 0.6126 0.7303
2.000 0.3632 0.00572 0.00128 -0.0297 0.5981 0.7367
2.250 0.3909 0.00578 0.00132 -0.0296 0.5804 0.7435
2.500 0.4183 0.00587 0.00136 -0.0295 0.5591 0.7502
2.750 0.4454 0.00598 0.00143 -0.0293 0.5361 0.7574
3.000 0.4719 0.00615 0.00151 -0.0290 0.5039 0.7648
3.250 0.4974 0.00640 0.00162 -0.0286 0.4603 0.7726
3.500 0.5230 0.00665 0.00176 -0.0283 0.4218 0.7805
3.750 0.5481 0.00696 0.00192 -0.0278 0.3775 0.7891
4.000 0.5727 0.00730 0.00210 -0.0273 0.3287 0.7973
4.250 0.5978 0.00761 0.00228 -0.0269 0.2933 0.8066
4.500 0.6232 0.00784 0.00247 -0.0265 0.2703 0.8158
4.750 0.6488 0.00806 0.00264 -0.0262 0.2507 0.8259
5.000 0.6746 0.00826 0.00281 -0.0259 0.2347 0.8368
5.250 0.7003 0.00841 0.00298 -0.0255 0.2213 0.8488
5.500 0.7256 0.00856 0.00316 -0.0250 0.2078 0.8625
5.750 0.7503 0.00873 0.00334 -0.0244 0.1923 0.8798
6.000 0.7735 0.00891 0.00354 -0.0235 0.1736 0.9050
6.250 0.8008 0.00915 0.00379 -0.0235 0.1497 0.9534
6.750 0.8679 0.01002 0.00444 -0.0268 0.1025 1.0000
7.000 0.8913 0.01039 0.00473 -0.0262 0.0882 1.0000
7.250 0.9147 0.01077 0.00505 -0.0255 0.0760 1.0000
7.500 0.9380 0.01115 0.00537 -0.0249 0.0661 1.0000
7.750 0.9614 0.01151 0.00569 -0.0243 0.0582 1.0000
8.000 0.9853 0.01183 0.00601 -0.0237 0.0522 1.0000
8.250 1.0083 0.01222 0.00636 -0.0230 0.0461 1.0000
8.500 1.0320 0.01255 0.00670 -0.0224 0.0420 1.0000
8.750 1.0544 0.01296 0.00708 -0.0217 0.0367 1.0000
9.000 1.0773 0.01333 0.00746 -0.0210 0.0330 1.0000
9.250 1.0990 0.01378 0.00788 -0.0202 0.0284 1.0000
9.500 1.1211 0.01419 0.00830 -0.0194 0.0251 1.0000
9.750 1.1412 0.01474 0.00883 -0.0184 0.0205 1.0000
10.000 1.1620 0.01520 0.00929 -0.0174 0.0176 1.0000
10.250 1.1811 0.01577 0.00987 -0.0162 0.0143 1.0000
10.500 1.1981 0.01648 0.01057 -0.0148 0.0114 1.0000
10.750 1.2167 0.01702 0.01115 -0.0135 0.0101 1.0000
11.000 1.2303 0.01789 0.01205 -0.0116 0.0085 1.0000
11.250 1.2462 0.01843 0.01266 -0.0099 0.0079 1.0000
11.500 1.2590 0.01903 0.01332 -0.0077 0.0074 1.0000
11.750 1.2699 0.01974 0.01407 -0.0053 0.0069 1.0000
12.000 1.2793 0.02057 0.01495 -0.0029 0.0066 1.0000
12.250 1.2831 0.02177 0.01625 -0.0001 0.0061 1.0000
12.500 1.2913 0.02277 0.01733 0.0020 0.0060 1.0000
12.750 1.2999 0.02378 0.01842 0.0039 0.0060 1.0000
13.000 1.3092 0.02480 0.01952 0.0055 0.0057 1.0000
13.250 1.3149 0.02613 0.02094 0.0072 0.0056 1.0000
13.500 1.3239 0.02727 0.02215 0.0084 0.0054 1.0000
13.750 1.3271 0.02893 0.02392 0.0099 0.0053 1.0000
14.000 1.3316 0.03057 0.02564 0.0111 0.0052 1.0000
14.250 1.3366 0.03224 0.02740 0.0120 0.0051 1.0000
14.500 1.3349 0.03459 0.02987 0.0129 0.0051 1.0000
14.750 1.3389 0.03654 0.03188 0.0134 0.0048 1.0000
15.000 1.3373 0.03913 0.03458 0.0137 0.0048 1.0000
15.250 1.3331 0.04212 0.03768 0.0138 0.0046 1.0000
15.500 1.3295 0.04523 0.04089 0.0136 0.0046 1.0000
15.750 1.3221 0.04891 0.04470 0.0129 0.0046 1.0000
16.000 1.3085 0.05366 0.04958 0.0116 0.0045 1.0000
16.250 1.3024 0.05772 0.05375 0.0102 0.0045 1.0000
16.500 1.2849 0.06370 0.05988 0.0078 0.0044 1.0000
16.750 1.2758 0.06878 0.06508 0.0054 0.0045 1.0000
17.000 1.2555 0.07593 0.07238 0.0019 0.0044 1.0000
17.250 1.2364 0.08328 0.07988 -0.0020 0.0044 1.0000
17.500 1.2266 0.08920 0.08592 -0.0051 0.0045 1.0000
17.750 1.2061 0.09730 0.09415 -0.0095 0.0045 1.0000
18.000 1.1782 0.10701 0.10400 -0.0148 0.0044 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.0/12 AIRFOIL (hq1012-il)