HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/8 AIRFOIL (hq158-il) Reynolds number: 200,000 Max Cl/Cd: 60.74 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq158-il-200000-n5.txt Download as CSV file: xf-hq158-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4194 0.09821 0.09486 -0.0136 1.0000 0.0101
-9.750 -0.5087 0.10323 0.09962 -0.0109 1.0000 0.0104
-9.500 -0.5086 0.09881 0.09524 -0.0126 1.0000 0.0100
-9.250 -0.5078 0.09470 0.09117 -0.0144 1.0000 0.0098
-9.000 -0.5075 0.09050 0.08700 -0.0164 1.0000 0.0096
-8.750 -0.5084 0.08604 0.08258 -0.0186 1.0000 0.0094
-8.500 -0.5089 0.08189 0.07848 -0.0209 1.0000 0.0092
-8.250 -0.5120 0.07731 0.07395 -0.0237 1.0000 0.0090
-8.000 -0.5180 0.07241 0.06912 -0.0273 1.0000 0.0089
-7.750 -0.5226 0.06656 0.06328 -0.0333 1.0000 0.0087
-7.500 -0.5234 0.06104 0.05769 -0.0372 1.0000 0.0085
-7.250 -0.5225 0.05555 0.05208 -0.0398 1.0000 0.0083
-7.000 -0.5190 0.05026 0.04661 -0.0413 1.0000 0.0081
-6.750 -0.5137 0.04495 0.04106 -0.0418 1.0000 0.0080
-6.500 -0.5060 0.04000 0.03581 -0.0414 1.0000 0.0078
-6.250 -0.4965 0.03517 0.03060 -0.0403 1.0000 0.0077
-6.000 -0.4850 0.03079 0.02577 -0.0388 1.0000 0.0077
-5.750 -0.4713 0.02717 0.02169 -0.0371 1.0000 0.0080
-5.500 -0.4486 0.02401 0.01802 -0.0367 0.9978 0.0086
-5.250 -0.4156 0.02216 0.01575 -0.0381 0.9928 0.0103
-5.000 -0.3838 0.01969 0.01276 -0.0391 0.9878 0.0114
-4.750 -0.3524 0.01729 0.01003 -0.0400 0.9834 0.0122
-4.500 -0.3213 0.01585 0.00844 -0.0408 0.9779 0.0134
-4.250 -0.2880 0.01475 0.00721 -0.0420 0.9736 0.0154
-4.000 -0.2566 0.01413 0.00648 -0.0429 0.9670 0.0192
-3.750 -0.2246 0.01305 0.00530 -0.0439 0.9611 0.0242
-3.500 -0.1934 0.01238 0.00455 -0.0446 0.9540 0.0354
-3.250 -0.1613 0.01184 0.00404 -0.0457 0.9475 0.0638
-3.000 -0.1317 0.01122 0.00365 -0.0465 0.9395 0.1209
-2.750 -0.1017 0.01039 0.00327 -0.0475 0.9326 0.2449
-2.500 -0.0757 0.00956 0.00303 -0.0476 0.9236 0.4119
-2.250 -0.0488 0.00905 0.00296 -0.0474 0.9160 0.5507
-2.000 -0.0217 0.00884 0.00290 -0.0470 0.9077 0.6297
-1.750 0.0052 0.00871 0.00286 -0.0465 0.8992 0.6789
-1.500 0.0322 0.00861 0.00282 -0.0459 0.8914 0.7251
-1.250 0.0574 0.00854 0.00280 -0.0449 0.8814 0.7599
-1.000 0.0833 0.00847 0.00273 -0.0441 0.8706 0.7855
-0.750 0.1091 0.00841 0.00263 -0.0433 0.8593 0.8081
-0.500 0.1354 0.00835 0.00254 -0.0426 0.8480 0.8246
-0.250 0.1619 0.00830 0.00245 -0.0421 0.8364 0.8392
0.000 0.1884 0.00824 0.00238 -0.0415 0.8238 0.8550
0.250 0.2158 0.00819 0.00232 -0.0411 0.8116 0.8727
0.500 0.2450 0.00814 0.00228 -0.0412 0.7996 0.8933
1.000 0.3166 0.00808 0.00222 -0.0442 0.7739 0.9508
1.250 0.3549 0.00808 0.00221 -0.0465 0.7599 0.9992
1.500 0.3808 0.00816 0.00222 -0.0461 0.7446 1.0000
1.750 0.4066 0.00824 0.00224 -0.0456 0.7276 1.0000
2.000 0.4325 0.00834 0.00230 -0.0452 0.7097 1.0000
2.250 0.4584 0.00844 0.00240 -0.0447 0.6899 1.0000
2.500 0.4842 0.00857 0.00248 -0.0442 0.6675 1.0000
2.750 0.5098 0.00872 0.00257 -0.0436 0.6407 1.0000
3.000 0.5348 0.00892 0.00267 -0.0429 0.6045 1.0000
3.250 0.5588 0.00920 0.00278 -0.0421 0.5529 1.0000
3.500 0.5812 0.00965 0.00297 -0.0410 0.4829 1.0000
3.750 0.6031 0.01023 0.00323 -0.0400 0.4104 1.0000
4.000 0.6257 0.01081 0.00354 -0.0393 0.3403 1.0000
4.250 0.6479 0.01147 0.00392 -0.0385 0.3011 1.0000
4.750 0.6966 0.01238 0.00466 -0.0376 0.2180 1.0000
5.000 0.7201 0.01296 0.00514 -0.0371 0.1715 1.0000
5.250 0.7425 0.01371 0.00564 -0.0365 0.1170 1.0000
5.500 0.7654 0.01444 0.00624 -0.0359 0.0833 1.0000
5.750 0.7881 0.01519 0.00688 -0.0353 0.0587 1.0000
6.000 0.8113 0.01586 0.00750 -0.0347 0.0406 1.0000
6.250 0.8343 0.01655 0.00818 -0.0342 0.0268 1.0000
6.500 0.8564 0.01747 0.00911 -0.0333 0.0157 1.0000
6.750 0.8778 0.01853 0.01033 -0.0322 0.0117 1.0000
7.000 0.8988 0.01961 0.01165 -0.0311 0.0098 1.0000
7.250 0.9183 0.02092 0.01308 -0.0300 0.0083 1.0000
7.500 0.9382 0.02214 0.01445 -0.0290 0.0068 1.0000
7.750 0.9579 0.02346 0.01595 -0.0278 0.0060 1.0000
8.000 0.9764 0.02506 0.01776 -0.0266 0.0056 1.0000
8.250 0.9941 0.02685 0.01978 -0.0253 0.0053 1.0000
8.500 1.0108 0.02884 0.02203 -0.0239 0.0050 1.0000
8.750 1.0257 0.03108 0.02458 -0.0224 0.0049 1.0000
9.000 1.0381 0.03363 0.02747 -0.0208 0.0048 1.0000
9.250 1.0471 0.03649 0.03071 -0.0189 0.0047 1.0000
9.500 1.0513 0.03968 0.03429 -0.0168 0.0046 1.0000
9.750 1.0500 0.04314 0.03814 -0.0144 0.0046 1.0000
10.000 1.0419 0.04658 0.04192 -0.0116 0.0046 1.0000
10.250 1.0270 0.05023 0.04589 -0.0088 0.0047 1.0000
10.500 1.0103 0.05407 0.04997 -0.0074 0.0047 1.0000
10.750 0.9912 0.05883 0.05498 -0.0077 0.0048 1.0000
11.000 0.9702 0.06470 0.06106 -0.0100 0.0048 1.0000
11.250 0.9492 0.07183 0.06838 -0.0145 0.0049 1.0000
11.500 0.9273 0.08128 0.07799 -0.0218 0.0049 1.0000
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Polar data table (+)
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