HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il) Reynolds number: 500,000 Max Cl/Cd: 79.14 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1010-il-500000.txt Download as CSV file: xf-hq1010-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/10 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.7559 0.03821 0.03489 -0.0406 1.0000 0.0102
-9.000 -0.7629 0.03346 0.02972 -0.0386 1.0000 0.0101
-8.750 -0.7609 0.02951 0.02535 -0.0366 1.0000 0.0101
-8.500 -0.7508 0.02674 0.02222 -0.0349 1.0000 0.0102
-8.250 -0.7359 0.02484 0.02003 -0.0333 1.0000 0.0104
-8.000 -0.7223 0.02245 0.01732 -0.0315 1.0000 0.0106
-7.750 -0.7107 0.01943 0.01396 -0.0294 1.0000 0.0111
-7.500 -0.6944 0.01789 0.01227 -0.0276 1.0000 0.0115
-7.250 -0.6766 0.01704 0.01136 -0.0261 1.0000 0.0122
-7.000 -0.6599 0.01632 0.01058 -0.0243 1.0000 0.0128
-6.750 -0.6430 0.01580 0.00998 -0.0225 0.9998 0.0139
-6.500 -0.6076 0.01482 0.00888 -0.0244 0.9962 0.0149
-6.250 -0.5747 0.01334 0.00725 -0.0260 0.9918 0.0168
-6.000 -0.5399 0.01261 0.00649 -0.0278 0.9871 0.0195
-5.750 -0.5035 0.01189 0.00570 -0.0298 0.9835 0.0238
-5.500 -0.4685 0.01141 0.00522 -0.0316 0.9787 0.0326
-5.250 -0.4334 0.01097 0.00478 -0.0333 0.9733 0.0426
-5.000 -0.3974 0.01049 0.00430 -0.0353 0.9689 0.0538
-4.750 -0.3667 0.01007 0.00391 -0.0360 0.9595 0.0649
-4.500 -0.3352 0.00969 0.00355 -0.0369 0.9511 0.0801
-4.250 -0.3066 0.00919 0.00318 -0.0372 0.9414 0.1134
-4.000 -0.2816 0.00860 0.00287 -0.0369 0.9297 0.1830
-3.750 -0.2567 0.00813 0.00262 -0.0365 0.9187 0.2520
-3.500 -0.2328 0.00756 0.00236 -0.0360 0.9083 0.3434
-3.250 -0.2094 0.00702 0.00217 -0.0353 0.8986 0.4547
-3.000 -0.1847 0.00674 0.00209 -0.0346 0.8881 0.5237
-2.750 -0.1588 0.00661 0.00203 -0.0341 0.8787 0.5683
-2.500 -0.1325 0.00655 0.00198 -0.0336 0.8699 0.6020
-2.250 -0.1062 0.00648 0.00195 -0.0331 0.8602 0.6321
-2.000 -0.0794 0.00646 0.00192 -0.0327 0.8513 0.6545
-1.750 -0.0524 0.00646 0.00189 -0.0324 0.8429 0.6730
-1.500 -0.0257 0.00643 0.00190 -0.0320 0.8340 0.6936
-1.250 0.0012 0.00644 0.00188 -0.0315 0.8253 0.7111
-1.000 0.0281 0.00644 0.00185 -0.0311 0.8150 0.7242
-0.750 0.0552 0.00642 0.00183 -0.0308 0.8042 0.7351
-0.500 0.0824 0.00642 0.00181 -0.0305 0.7943 0.7468
-0.250 0.1095 0.00643 0.00178 -0.0302 0.7848 0.7589
0.000 0.1366 0.00643 0.00178 -0.0299 0.7741 0.7705
0.250 0.1636 0.00641 0.00178 -0.0295 0.7632 0.7820
0.500 0.1905 0.00641 0.00179 -0.0291 0.7527 0.7935
0.750 0.2176 0.00642 0.00178 -0.0288 0.7433 0.8039
1.000 0.2450 0.00642 0.00179 -0.0287 0.7326 0.8140
1.250 0.2721 0.00641 0.00182 -0.0283 0.7220 0.8234
1.500 0.2991 0.00641 0.00183 -0.0280 0.7111 0.8334
1.750 0.3260 0.00642 0.00184 -0.0277 0.6991 0.8443
2.250 0.3788 0.00643 0.00190 -0.0268 0.6701 0.8684
2.500 0.4047 0.00643 0.00193 -0.0262 0.6527 0.8826
2.750 0.4302 0.00644 0.00197 -0.0256 0.6336 0.8993
3.000 0.4561 0.00647 0.00202 -0.0249 0.6106 0.9192
3.250 0.4855 0.00657 0.00209 -0.0251 0.5774 0.9442
3.500 0.5211 0.00677 0.00217 -0.0268 0.5337 0.9680
3.750 0.5595 0.00707 0.00230 -0.0293 0.4802 0.9875
4.000 0.5908 0.00754 0.00249 -0.0305 0.4127 1.0000
4.250 0.6121 0.00802 0.00271 -0.0295 0.3510 1.0000
4.500 0.6345 0.00851 0.00297 -0.0288 0.2997 1.0000
4.750 0.6583 0.00890 0.00322 -0.0282 0.2686 1.0000
5.000 0.6828 0.00925 0.00348 -0.0277 0.2432 1.0000
5.250 0.7070 0.00963 0.00376 -0.0272 0.2125 1.0000
5.500 0.7311 0.01005 0.00404 -0.0268 0.1800 1.0000
5.750 0.7549 0.01051 0.00436 -0.0262 0.1439 1.0000
6.000 0.7779 0.01107 0.00474 -0.0256 0.1117 1.0000
6.250 0.8012 0.01159 0.00516 -0.0250 0.0885 1.0000
6.500 0.8247 0.01209 0.00561 -0.0244 0.0705 1.0000
6.750 0.8483 0.01259 0.00604 -0.0239 0.0557 1.0000
7.000 0.8717 0.01309 0.00651 -0.0233 0.0443 1.0000
7.250 0.8948 0.01364 0.00704 -0.0226 0.0349 1.0000
7.500 0.9168 0.01433 0.00771 -0.0218 0.0264 1.0000
7.750 0.9370 0.01525 0.00864 -0.0207 0.0190 1.0000
8.000 0.9599 0.01579 0.00923 -0.0200 0.0153 1.0000
8.250 0.9779 0.01692 0.01044 -0.0186 0.0125 1.0000
8.500 0.9987 0.01767 0.01128 -0.0176 0.0113 1.0000
8.750 1.0186 0.01849 0.01218 -0.0165 0.0104 1.0000
9.000 1.0369 0.01947 0.01322 -0.0153 0.0097 1.0000
9.250 1.0504 0.02099 0.01486 -0.0134 0.0091 1.0000
9.500 1.0607 0.02305 0.01711 -0.0112 0.0086 1.0000
9.750 1.0776 0.02414 0.01835 -0.0098 0.0084 1.0000
10.000 1.0920 0.02555 0.01994 -0.0082 0.0081 1.0000
10.250 1.1045 0.02714 0.02171 -0.0064 0.0078 1.0000
10.500 1.1140 0.02903 0.02381 -0.0043 0.0076 1.0000
10.750 1.1193 0.03101 0.02600 -0.0018 0.0076 1.0000
11.000 1.1202 0.03312 0.02833 0.0012 0.0075 1.0000
11.250 1.1165 0.03560 0.03105 0.0042 0.0074 1.0000
11.500 1.1103 0.03828 0.03397 0.0067 0.0075 1.0000
11.750 1.1012 0.04129 0.03721 0.0087 0.0076 1.0000
12.000 1.0871 0.04499 0.04116 0.0101 0.0076 1.0000
12.250 1.0726 0.04898 0.04538 0.0105 0.0077 1.0000
12.500 1.0562 0.05355 0.05016 0.0098 0.0078 1.0000
12.750 1.0375 0.05892 0.05574 0.0080 0.0078 1.0000
13.000 1.0162 0.06537 0.06239 0.0049 0.0079 1.0000
13.250 0.9970 0.07229 0.06947 0.0006 0.0079 1.0000
13.500 0.9761 0.08048 0.07783 -0.0048 0.0080 1.0000
13.750 0.9549 0.08982 0.08731 -0.0113 0.0080 1.0000
14.000 0.9330 0.10024 0.09787 -0.0182 0.0081 1.0000
14.250 0.9042 0.11322 0.11096 -0.0262 0.0082 1.0000
14.500 0.8658 0.12941 0.12720 -0.0351 0.0085 1.0000
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Polar data table (+)
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