HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.5/9 AIRFOIL (hq159-il) Reynolds number: 500,000 Max Cl/Cd: 88.43 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq159-il-500000.txt Download as CSV file: xf-hq159-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/9 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5062 0.08296 0.08076 -0.0250 1.0000 0.0174
-8.750 -0.5097 0.07800 0.07583 -0.0281 1.0000 0.0178
-8.500 -0.5153 0.07281 0.07068 -0.0318 1.0000 0.0177
-8.250 -0.5311 0.06626 0.06416 -0.0384 1.0000 0.0174
-8.000 -0.5410 0.06061 0.05843 -0.0422 1.0000 0.0173
-7.750 -0.5456 0.05544 0.05315 -0.0441 1.0000 0.0179
-6.500 -0.5438 0.02661 0.02256 -0.0382 0.9976 0.0120
-6.250 -0.5142 0.02097 0.01620 -0.0399 0.9940 0.0114
-6.000 -0.4827 0.01840 0.01325 -0.0410 0.9901 0.0119
-5.750 -0.4494 0.01644 0.01100 -0.0423 0.9866 0.0123
-5.500 -0.4129 0.01584 0.01026 -0.0442 0.9839 0.0132
-5.250 -0.3802 0.01356 0.00773 -0.0455 0.9811 0.0140
-5.000 -0.3491 0.01207 0.00609 -0.0464 0.9759 0.0153
-4.750 -0.3141 0.01130 0.00526 -0.0480 0.9722 0.0173
-4.500 -0.2783 0.01072 0.00461 -0.0497 0.9689 0.0202
-4.250 -0.2484 0.01004 0.00390 -0.0502 0.9612 0.0296
-4.000 -0.2160 0.00946 0.00337 -0.0512 0.9557 0.0507
-3.750 -0.1876 0.00915 0.00305 -0.0514 0.9467 0.0662
-3.500 -0.1590 0.00874 0.00275 -0.0516 0.9391 0.0972
-3.250 -0.1340 0.00806 0.00246 -0.0514 0.9294 0.1965
-3.000 -0.1103 0.00728 0.00218 -0.0509 0.9196 0.3413
-2.750 -0.0866 0.00664 0.00203 -0.0503 0.9106 0.4925
-2.500 -0.0614 0.00640 0.00196 -0.0497 0.9013 0.5668
-2.250 -0.0354 0.00629 0.00191 -0.0491 0.8913 0.6088
-2.000 -0.0090 0.00624 0.00186 -0.0486 0.8820 0.6389
-1.750 0.0173 0.00619 0.00182 -0.0480 0.8730 0.6670
-1.500 0.0438 0.00616 0.00180 -0.0476 0.8637 0.6932
-1.250 0.0704 0.00615 0.00178 -0.0471 0.8547 0.7136
-1.000 0.0964 0.00614 0.00177 -0.0464 0.8445 0.7375
-0.750 0.1227 0.00612 0.00175 -0.0458 0.8330 0.7552
-0.500 0.1493 0.00610 0.00173 -0.0454 0.8221 0.7697
-0.250 0.1759 0.00610 0.00171 -0.0449 0.8119 0.7841
0.000 0.2025 0.00609 0.00169 -0.0444 0.8015 0.7990
0.250 0.2289 0.00607 0.00168 -0.0439 0.7901 0.8133
0.500 0.2556 0.00605 0.00166 -0.0435 0.7786 0.8254
0.750 0.2824 0.00603 0.00166 -0.0432 0.7680 0.8373
1.000 0.3090 0.00602 0.00165 -0.0427 0.7573 0.8501
1.250 0.3353 0.00600 0.00165 -0.0422 0.7462 0.8644
1.500 0.3613 0.00597 0.00166 -0.0417 0.7341 0.8812
1.750 0.3871 0.00593 0.00167 -0.0410 0.7212 0.9022
2.000 0.4153 0.00590 0.00169 -0.0408 0.7073 0.9310
2.250 0.4532 0.00592 0.00171 -0.0429 0.6912 0.9652
2.500 0.4947 0.00597 0.00175 -0.0460 0.6695 0.9997
2.750 0.5202 0.00609 0.00179 -0.0455 0.6457 1.0000
3.000 0.5456 0.00624 0.00185 -0.0451 0.6169 1.0000
3.250 0.5704 0.00645 0.00192 -0.0445 0.5775 1.0000
3.500 0.5944 0.00677 0.00204 -0.0437 0.5244 1.0000
3.750 0.6176 0.00720 0.00221 -0.0430 0.4606 1.0000
4.000 0.6409 0.00768 0.00243 -0.0423 0.3984 1.0000
4.250 0.6646 0.00816 0.00268 -0.0417 0.3408 1.0000
4.500 0.6884 0.00865 0.00297 -0.0412 0.2925 1.0000
4.750 0.7126 0.00911 0.00326 -0.0407 0.2542 1.0000
5.000 0.7371 0.00952 0.00355 -0.0403 0.2214 1.0000
5.250 0.7619 0.00992 0.00383 -0.0399 0.1880 1.0000
5.500 0.7857 0.01043 0.00416 -0.0394 0.1456 1.0000
5.750 0.8080 0.01114 0.00461 -0.0387 0.0973 1.0000
6.000 0.8309 0.01179 0.00512 -0.0381 0.0680 1.0000
6.250 0.8544 0.01235 0.00561 -0.0375 0.0529 1.0000
6.500 0.8787 0.01279 0.00606 -0.0370 0.0446 1.0000
6.750 0.9027 0.01326 0.00652 -0.0365 0.0369 1.0000
7.000 0.9269 0.01370 0.00695 -0.0361 0.0306 1.0000
7.250 0.9512 0.01412 0.00742 -0.0356 0.0244 1.0000
7.500 0.9706 0.01525 0.00851 -0.0343 0.0120 1.0000
7.750 0.9931 0.01592 0.00928 -0.0334 0.0102 1.0000
8.000 1.0144 0.01673 0.01017 -0.0325 0.0092 1.0000
8.250 1.0344 0.01768 0.01122 -0.0313 0.0085 1.0000
8.500 1.0518 0.01894 0.01261 -0.0298 0.0080 1.0000
8.750 1.0657 0.02071 0.01455 -0.0278 0.0077 1.0000
9.000 1.0778 0.02293 0.01700 -0.0257 0.0073 1.0000
9.250 1.0974 0.02381 0.01801 -0.0247 0.0071 1.0000
9.500 1.1151 0.02494 0.01929 -0.0234 0.0068 1.0000
9.750 1.1297 0.02656 0.02114 -0.0218 0.0066 1.0000
10.000 1.1420 0.02843 0.02322 -0.0200 0.0064 1.0000
10.250 1.1513 0.03057 0.02561 -0.0179 0.0063 1.0000
10.500 1.1568 0.03292 0.02823 -0.0155 0.0062 1.0000
10.750 1.1565 0.03524 0.03080 -0.0124 0.0061 1.0000
11.000 1.1528 0.03745 0.03323 -0.0092 0.0060 1.0000
11.250 1.1407 0.04066 0.03671 -0.0060 0.0061 1.0000
11.500 1.1295 0.04372 0.04001 -0.0038 0.0060 1.0000
11.750 1.1180 0.04699 0.04349 -0.0025 0.0058 1.0000
12.000 1.0964 0.05197 0.04873 -0.0023 0.0060 1.0000
12.250 1.0802 0.05668 0.05363 -0.0033 0.0059 1.0000
12.500 1.0613 0.06243 0.05959 -0.0057 0.0059 1.0000
12.750 1.0420 0.06910 0.06643 -0.0095 0.0059 1.0000
13.000 1.0190 0.07763 0.07514 -0.0151 0.0060 1.0000
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Polar data table (+)
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