HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/8 AIRFOIL (hq158-il) Reynolds number: 500,000 Max Cl/Cd: 69.99 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq158-il-500000-n5.txt Download as CSV file: xf-hq158-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5156 0.08247 0.08028 -0.0195 1.0000 0.0044
-8.500 -0.5175 0.07810 0.07594 -0.0220 1.0000 0.0043
-8.250 -0.5233 0.07317 0.07106 -0.0252 1.0000 0.0042
-8.000 -0.5330 0.06747 0.06541 -0.0308 1.0000 0.0042
-7.750 -0.5366 0.06087 0.05874 -0.0369 1.0000 0.0040
-7.500 -0.5374 0.05489 0.05265 -0.0402 1.0000 0.0039
-7.250 -0.5369 0.04865 0.04623 -0.0420 1.0000 0.0038
-7.000 -0.5351 0.04214 0.03947 -0.0424 1.0000 0.0036
-6.750 -0.5287 0.02802 0.02448 -0.0442 0.9909 0.0033
-6.500 -0.5054 0.02070 0.01623 -0.0453 0.9832 0.0032
-6.250 -0.4771 0.01829 0.01342 -0.0461 0.9766 0.0033
-6.000 -0.4473 0.01645 0.01127 -0.0470 0.9706 0.0035
-5.750 -0.4178 0.01509 0.00967 -0.0476 0.9631 0.0038
-5.500 -0.3881 0.01390 0.00828 -0.0481 0.9554 0.0044
-5.250 -0.3588 0.01291 0.00707 -0.0485 0.9466 0.0049
-5.000 -0.3307 0.01220 0.00623 -0.0486 0.9362 0.0054
-4.750 -0.3040 0.01130 0.00520 -0.0485 0.9256 0.0066
-4.500 -0.2769 0.01084 0.00467 -0.0485 0.9154 0.0080
-4.000 -0.2240 0.00992 0.00351 -0.0478 0.8950 0.0105
-3.750 -0.1977 0.00950 0.00298 -0.0474 0.8852 0.0135
-3.500 -0.1711 0.00920 0.00262 -0.0471 0.8758 0.0186
-3.250 -0.1445 0.00892 0.00231 -0.0468 0.8662 0.0334
-3.000 -0.1178 0.00871 0.00210 -0.0466 0.8568 0.0477
-2.750 -0.0911 0.00850 0.00190 -0.0463 0.8475 0.0686
-2.500 -0.0652 0.00810 0.00170 -0.0461 0.8376 0.1305
-2.250 -0.0397 0.00760 0.00151 -0.0459 0.8264 0.2307
-2.000 -0.0141 0.00720 0.00132 -0.0457 0.8129 0.3206
-1.750 0.0111 0.00679 0.00119 -0.0453 0.7977 0.4301
-1.500 0.0367 0.00654 0.00111 -0.0449 0.7830 0.5102
-1.250 0.0624 0.00632 0.00109 -0.0445 0.7704 0.5884
-1.000 0.0888 0.00621 0.00108 -0.0441 0.7599 0.6381
-0.750 0.1150 0.00613 0.00107 -0.0436 0.7488 0.6811
-0.500 0.1416 0.00611 0.00106 -0.0432 0.7348 0.7077
-0.250 0.1683 0.00610 0.00105 -0.0429 0.7206 0.7284
0.000 0.1950 0.00610 0.00105 -0.0426 0.7081 0.7485
0.250 0.2219 0.00611 0.00105 -0.0423 0.6949 0.7637
0.500 0.2489 0.00614 0.00105 -0.0420 0.6805 0.7753
0.750 0.2758 0.00617 0.00107 -0.0418 0.6644 0.7869
1.000 0.3026 0.00622 0.00109 -0.0415 0.6471 0.7992
1.250 0.3292 0.00627 0.00113 -0.0411 0.6281 0.8122
1.500 0.3553 0.00635 0.00116 -0.0407 0.6026 0.8262
1.750 0.3810 0.00645 0.00121 -0.0402 0.5707 0.8414
2.000 0.4062 0.00658 0.00127 -0.0397 0.5345 0.8592
2.250 0.4308 0.00675 0.00137 -0.0390 0.4896 0.8825
2.500 0.4587 0.00698 0.00150 -0.0390 0.4359 0.9227
2.750 0.4995 0.00739 0.00168 -0.0422 0.3699 1.0000
3.000 0.5239 0.00784 0.00186 -0.0418 0.3138 1.0000
3.250 0.5492 0.00817 0.00209 -0.0415 0.2977 1.0000
3.500 0.5765 0.00826 0.00226 -0.0414 0.2766 1.0000
3.750 0.6019 0.00860 0.00243 -0.0411 0.2392 1.0000
4.000 0.6268 0.00902 0.00266 -0.0407 0.1964 1.0000
4.250 0.6519 0.00940 0.00291 -0.0404 0.1609 1.0000
4.500 0.6766 0.00985 0.00319 -0.0400 0.1220 1.0000
4.750 0.6997 0.01053 0.00359 -0.0395 0.0690 1.0000
5.000 0.7243 0.01099 0.00398 -0.0391 0.0467 1.0000
5.250 0.7496 0.01136 0.00432 -0.0388 0.0359 1.0000
5.500 0.7745 0.01178 0.00469 -0.0384 0.0240 1.0000
5.750 0.7994 0.01219 0.00510 -0.0380 0.0162 1.0000
6.000 0.8234 0.01277 0.00564 -0.0374 0.0065 1.0000
6.250 0.8474 0.01335 0.00625 -0.0368 0.0042 1.0000
6.500 0.8717 0.01388 0.00688 -0.0363 0.0037 1.0000
6.750 0.8956 0.01446 0.00760 -0.0357 0.0033 1.0000
7.000 0.9189 0.01514 0.00840 -0.0350 0.0030 1.0000
7.250 0.9415 0.01590 0.00927 -0.0342 0.0028 1.0000
7.500 0.9633 0.01678 0.01029 -0.0333 0.0026 1.0000
7.750 0.9843 0.01778 0.01142 -0.0323 0.0024 1.0000
8.000 1.0046 0.01887 0.01266 -0.0313 0.0023 1.0000
8.250 1.0237 0.02014 0.01410 -0.0301 0.0022 1.0000
8.500 1.0421 0.02153 0.01566 -0.0288 0.0022 1.0000
8.750 1.0593 0.02312 0.01746 -0.0274 0.0021 1.0000
9.000 1.0754 0.02483 0.01941 -0.0260 0.0021 1.0000
9.250 1.0899 0.02674 0.02157 -0.0244 0.0021 1.0000
9.500 1.1015 0.02899 0.02415 -0.0226 0.0021 1.0000
9.750 1.1102 0.03144 0.02692 -0.0205 0.0021 1.0000
10.000 1.1149 0.03411 0.02991 -0.0182 0.0021 1.0000
10.250 1.1157 0.03689 0.03301 -0.0157 0.0021 1.0000
10.500 1.1059 0.03988 0.03628 -0.0120 0.0021 1.0000
10.750 1.0932 0.04288 0.03953 -0.0089 0.0021 1.0000
11.000 1.0776 0.04646 0.04336 -0.0070 0.0022 1.0000
11.250 1.0600 0.05071 0.04784 -0.0066 0.0022 1.0000
11.500 1.0411 0.05584 0.05316 -0.0079 0.0022 1.0000
11.750 1.0198 0.06243 0.05995 -0.0113 0.0022 1.0000
12.000 0.9975 0.07099 0.06869 -0.0174 0.0022 1.0000
12.250 0.9772 0.08170 0.07957 -0.0255 0.0022 1.0000
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Polar data table (+)
Polar graphs
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