HQ 1.0/12 AIRFOIL (hq1012-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.0/12 AIRFOIL (hq1012-il) Reynolds number: 200,000 Max Cl/Cd: 54.88 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1012-il-200000-n5.txt Download as CSV file: xf-hq1012-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.7659 0.08668 0.08285 -0.0273 1.0000 0.0124
-13.500 -0.8085 0.07286 0.06879 -0.0368 1.0000 0.0121
-13.250 -0.8386 0.06352 0.05919 -0.0432 1.0000 0.0117
-13.000 -0.8580 0.05714 0.05261 -0.0472 1.0000 0.0118
-12.750 -0.8729 0.05210 0.04737 -0.0497 1.0000 0.0118
-12.500 -0.8849 0.04787 0.04294 -0.0513 1.0000 0.0120
-12.250 -0.8942 0.04428 0.03914 -0.0520 1.0000 0.0121
-12.000 -0.9010 0.04121 0.03588 -0.0520 1.0000 0.0122
-11.750 -0.9051 0.03862 0.03309 -0.0514 1.0000 0.0125
-11.500 -0.9074 0.03633 0.03061 -0.0502 1.0000 0.0128
-11.250 -0.9078 0.03438 0.02844 -0.0485 1.0000 0.0133
-11.000 -0.9057 0.03272 0.02656 -0.0463 1.0000 0.0138
-10.750 -0.9012 0.03110 0.02477 -0.0442 1.0000 0.0143
-10.500 -0.8928 0.02974 0.02334 -0.0426 1.0000 0.0150
-10.250 -0.8807 0.02866 0.02217 -0.0411 1.0000 0.0157
-10.000 -0.8677 0.02747 0.02085 -0.0397 1.0000 0.0164
-9.750 -0.8541 0.02622 0.01946 -0.0381 1.0000 0.0173
-9.500 -0.8396 0.02503 0.01808 -0.0366 1.0000 0.0183
-9.250 -0.8254 0.02385 0.01679 -0.0350 1.0000 0.0193
-9.000 -0.8100 0.02297 0.01588 -0.0335 1.0000 0.0207
-8.750 -0.7932 0.02227 0.01512 -0.0321 1.0000 0.0226
-8.500 -0.7767 0.02153 0.01425 -0.0304 1.0000 0.0249
-8.250 -0.7631 0.02060 0.01330 -0.0285 1.0000 0.0269
-8.000 -0.7482 0.01996 0.01263 -0.0266 1.0000 0.0293
-7.750 -0.7211 0.01927 0.01180 -0.0271 0.9946 0.0330
-7.500 -0.6891 0.01850 0.01102 -0.0288 0.9865 0.0379
-7.250 -0.6555 0.01796 0.01036 -0.0304 0.9792 0.0430
-7.000 -0.6230 0.01723 0.00964 -0.0320 0.9715 0.0488
-6.750 -0.5910 0.01678 0.00908 -0.0332 0.9625 0.0547
-6.500 -0.5595 0.01609 0.00839 -0.0346 0.9540 0.0615
-6.250 -0.5281 0.01561 0.00782 -0.0356 0.9445 0.0682
-6.000 -0.4986 0.01505 0.00726 -0.0364 0.9345 0.0770
-5.750 -0.4685 0.01453 0.00673 -0.0371 0.9253 0.0868
-5.500 -0.4397 0.01407 0.00623 -0.0375 0.9157 0.0991
-5.250 -0.4124 0.01362 0.00580 -0.0377 0.9061 0.1165
-5.000 -0.3857 0.01314 0.00541 -0.0377 0.8972 0.1438
-4.750 -0.3607 0.01268 0.00508 -0.0374 0.8874 0.1823
-4.500 -0.3357 0.01228 0.00477 -0.0371 0.8783 0.2228
-4.250 -0.3110 0.01185 0.00446 -0.0367 0.8700 0.2659
-4.000 -0.2876 0.01137 0.00419 -0.0361 0.8609 0.3196
-3.750 -0.2644 0.01090 0.00398 -0.0354 0.8533 0.3941
-3.500 -0.2406 0.01057 0.00385 -0.0346 0.8449 0.4584
-3.250 -0.2157 0.01037 0.00379 -0.0340 0.8376 0.5101
-3.000 -0.1902 0.01026 0.00376 -0.0333 0.8297 0.5514
-2.750 -0.1641 0.01020 0.00373 -0.0328 0.8228 0.5844
-2.500 -0.1379 0.01018 0.00372 -0.0322 0.8151 0.6105
-2.250 -0.1112 0.01018 0.00369 -0.0318 0.8084 0.6328
-2.000 -0.0844 0.01017 0.00365 -0.0314 0.8007 0.6483
-1.750 -0.0572 0.01016 0.00360 -0.0311 0.7940 0.6600
-1.500 -0.0302 0.01015 0.00357 -0.0307 0.7862 0.6709
-1.250 -0.0030 0.01015 0.00354 -0.0304 0.7795 0.6824
-1.000 0.0239 0.01016 0.00352 -0.0301 0.7717 0.6950
-0.500 0.0779 0.01017 0.00353 -0.0294 0.7574 0.7185
-0.250 0.1047 0.01018 0.00352 -0.0289 0.7492 0.7292
0.000 0.1316 0.01017 0.00349 -0.0285 0.7390 0.7389
0.250 0.1584 0.01016 0.00347 -0.0282 0.7281 0.7462
0.500 0.1856 0.01016 0.00344 -0.0279 0.7183 0.7539
0.750 0.2125 0.01016 0.00342 -0.0275 0.7083 0.7612
1.000 0.2395 0.01016 0.00342 -0.0272 0.6970 0.7690
1.250 0.2662 0.01016 0.00344 -0.0268 0.6858 0.7768
1.500 0.2931 0.01017 0.00344 -0.0265 0.6750 0.7846
1.750 0.3199 0.01019 0.00347 -0.0261 0.6648 0.7928
2.000 0.3467 0.01020 0.00352 -0.0258 0.6536 0.8008
2.250 0.3734 0.01022 0.00357 -0.0254 0.6417 0.8096
2.500 0.3997 0.01024 0.00362 -0.0249 0.6289 0.8178
2.750 0.4261 0.01028 0.00368 -0.0245 0.6146 0.8270
3.000 0.4520 0.01031 0.00374 -0.0240 0.5980 0.8364
3.250 0.4778 0.01035 0.00380 -0.0234 0.5791 0.8461
3.500 0.5032 0.01042 0.00386 -0.0227 0.5557 0.8569
3.750 0.5281 0.01053 0.00394 -0.0220 0.5259 0.8687
4.000 0.5528 0.01071 0.00403 -0.0213 0.4904 0.8818
4.250 0.5773 0.01095 0.00417 -0.0206 0.4496 0.8970
4.500 0.6030 0.01132 0.00439 -0.0203 0.4025 0.9141
4.750 0.6321 0.01179 0.00467 -0.0209 0.3501 0.9349
5.000 0.6640 0.01231 0.00500 -0.0222 0.3050 0.9602
5.250 0.6981 0.01278 0.00536 -0.0240 0.2764 0.9908
5.500 0.7229 0.01320 0.00572 -0.0239 0.2576 1.0000
6.000 0.7672 0.01399 0.00642 -0.0223 0.2223 1.0000
6.250 0.7897 0.01439 0.00678 -0.0216 0.2034 1.0000
6.500 0.8117 0.01483 0.00718 -0.0209 0.1833 1.0000
6.750 0.8338 0.01528 0.00758 -0.0201 0.1628 1.0000
7.000 0.8550 0.01581 0.00803 -0.0193 0.1425 1.0000
7.250 0.8757 0.01638 0.00853 -0.0184 0.1243 1.0000
7.500 0.8962 0.01697 0.00908 -0.0174 0.1090 1.0000
7.750 0.9164 0.01758 0.00969 -0.0164 0.0972 1.0000
8.000 0.9358 0.01825 0.01032 -0.0154 0.0865 1.0000
8.250 0.9557 0.01885 0.01094 -0.0144 0.0771 1.0000
8.500 0.9749 0.01950 0.01162 -0.0133 0.0699 1.0000
8.750 0.9931 0.02020 0.01233 -0.0121 0.0633 1.0000
9.000 1.0113 0.02088 0.01308 -0.0109 0.0580 1.0000
9.250 1.0286 0.02159 0.01386 -0.0096 0.0526 1.0000
9.500 1.0441 0.02241 0.01472 -0.0081 0.0482 1.0000
9.750 1.0610 0.02308 0.01547 -0.0068 0.0430 1.0000
10.000 1.0735 0.02392 0.01630 -0.0050 0.0383 1.0000
10.250 1.0881 0.02462 0.01709 -0.0034 0.0336 1.0000
10.500 1.0978 0.02564 0.01809 -0.0014 0.0300 1.0000
10.750 1.1099 0.02655 0.01914 0.0003 0.0272 1.0000
11.000 1.1204 0.02759 0.02027 0.0019 0.0242 1.0000
11.250 1.1276 0.02889 0.02162 0.0037 0.0219 1.0000
11.500 1.1376 0.03005 0.02290 0.0051 0.0192 1.0000
11.750 1.1440 0.03151 0.02443 0.0066 0.0176 1.0000
12.000 1.1484 0.03319 0.02619 0.0079 0.0160 1.0000
12.250 1.1527 0.03495 0.02810 0.0092 0.0151 1.0000
12.500 1.1558 0.03690 0.03018 0.0102 0.0139 1.0000
12.750 1.1573 0.03905 0.03245 0.0110 0.0132 1.0000
13.000 1.1554 0.04164 0.03513 0.0116 0.0126 1.0000
13.250 1.1514 0.04457 0.03819 0.0119 0.0122 1.0000
13.500 1.1490 0.04750 0.04129 0.0120 0.0118 1.0000
13.750 1.1461 0.05063 0.04458 0.0117 0.0115 1.0000
14.000 1.1408 0.05421 0.04834 0.0111 0.0111 1.0000
14.250 1.1340 0.05816 0.05248 0.0101 0.0108 1.0000
14.500 1.1267 0.06238 0.05686 0.0087 0.0106 1.0000
14.750 1.1181 0.06703 0.06168 0.0069 0.0104 1.0000
15.000 1.1082 0.07215 0.06696 0.0046 0.0103 1.0000
15.250 1.0965 0.07796 0.07292 0.0016 0.0100 1.0000
15.500 1.0854 0.08386 0.07899 -0.0014 0.0101 1.0000
15.750 1.0718 0.09059 0.08587 -0.0051 0.0099 1.0000
16.000 1.0575 0.09770 0.09316 -0.0090 0.0100 1.0000
16.250 1.0420 0.10534 0.10094 -0.0134 0.0098 1.0000
16.500 1.0256 0.11339 0.10915 -0.0179 0.0099 1.0000
16.750 1.0080 0.12193 0.11786 -0.0228 0.0100 1.0000
17.000 0.9888 0.13119 0.12726 -0.0281 0.0100 1.0000
17.250 0.9654 0.14190 0.13812 -0.0342 0.0102 1.0000
17.500 0.9297 0.15680 0.15320 -0.0426 0.0105 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.0/12 AIRFOIL (hq1012-il)