HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il) Reynolds number: 100,000 Max Cl/Cd: 57.56 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq258-il-100000-n5.txt Download as CSV file: xf-hq258-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/8 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3205 0.09195 0.08733 -0.0294 1.0000 0.0357
-8.750 -0.3198 0.08812 0.08354 -0.0302 1.0000 0.0344
-8.500 -0.3205 0.08415 0.07962 -0.0311 1.0000 0.0327
-8.250 -0.3230 0.07960 0.07508 -0.0327 1.0000 0.0249
-8.000 -0.3236 0.07577 0.07131 -0.0329 1.0000 0.0233
-7.500 -0.4057 0.08227 0.07763 -0.0322 1.0000 0.0242
-7.250 -0.4095 0.07913 0.07458 -0.0327 1.0000 0.0229
-7.000 -0.4113 0.07544 0.07097 -0.0346 1.0000 0.0216
-6.750 -0.4128 0.07126 0.06685 -0.0376 1.0000 0.0205
-6.500 -0.4136 0.06682 0.06242 -0.0402 1.0000 0.0194
-6.250 -0.4133 0.06154 0.05707 -0.0431 1.0000 0.0180
-5.750 -0.3969 0.05394 0.04914 -0.0443 1.0000 0.0158
-5.500 -0.3904 0.05015 0.04522 -0.0445 1.0000 0.0155
-5.250 -0.3798 0.04629 0.04115 -0.0450 0.9999 0.0152
-5.000 -0.3477 0.04107 0.03552 -0.0495 0.9951 0.0149
-4.750 -0.3148 0.03647 0.03043 -0.0528 0.9903 0.0150
-4.500 -0.2784 0.03248 0.02578 -0.0557 0.9861 0.0165
-4.250 -0.2483 0.02917 0.02208 -0.0581 0.9814 0.0185
-4.000 -0.2143 0.02653 0.01897 -0.0599 0.9771 0.0198
-3.750 -0.1781 0.02385 0.01576 -0.0614 0.9740 0.0205
-3.500 -0.1463 0.02178 0.01329 -0.0619 0.9691 0.0216
-3.250 -0.1129 0.01999 0.01121 -0.0626 0.9649 0.0238
-3.000 -0.0782 0.01861 0.00959 -0.0638 0.9614 0.0290
-2.750 -0.0476 0.01765 0.00855 -0.0645 0.9557 0.0385
-2.500 -0.0143 0.01656 0.00740 -0.0656 0.9509 0.0589
-2.250 0.0218 0.01561 0.00667 -0.0677 0.9474 0.1235
-2.000 0.0468 0.01387 0.00638 -0.0682 0.9409 0.4845
-1.750 0.0752 0.01348 0.00647 -0.0676 0.9360 0.6813
-1.500 0.0989 0.01332 0.00639 -0.0659 0.9291 0.7732
-1.250 0.1208 0.01298 0.00627 -0.0631 0.9231 0.8710
-1.000 0.1686 0.01274 0.00594 -0.0666 0.9208 0.9478
-0.750 0.2066 0.01269 0.00571 -0.0691 0.9131 1.0000
-0.500 0.2439 0.01268 0.00547 -0.0712 0.9077 1.0000
-0.250 0.2737 0.01273 0.00537 -0.0718 0.8986 1.0000
0.000 0.3073 0.01273 0.00524 -0.0731 0.8912 1.0000
0.250 0.3391 0.01274 0.00515 -0.0739 0.8826 1.0000
0.500 0.3686 0.01278 0.00512 -0.0742 0.8730 1.0000
0.750 0.4002 0.01279 0.00505 -0.0749 0.8647 1.0000
1.000 0.4301 0.01281 0.00503 -0.0752 0.8552 1.0000
1.250 0.4586 0.01281 0.00500 -0.0751 0.8432 1.0000
1.500 0.4872 0.01277 0.00492 -0.0749 0.8294 1.0000
1.750 0.5153 0.01273 0.00489 -0.0745 0.8148 1.0000
2.000 0.5429 0.01274 0.00488 -0.0742 0.8006 1.0000
2.250 0.5702 0.01276 0.00490 -0.0738 0.7862 1.0000
2.500 0.5965 0.01281 0.00498 -0.0732 0.7695 1.0000
2.750 0.6231 0.01285 0.00510 -0.0726 0.7516 1.0000
3.000 0.6502 0.01288 0.00515 -0.0721 0.7329 1.0000
3.250 0.6760 0.01296 0.00527 -0.0714 0.7107 1.0000
3.500 0.7020 0.01305 0.00539 -0.0707 0.6865 1.0000
3.750 0.7276 0.01317 0.00551 -0.0699 0.6579 1.0000
4.000 0.7527 0.01333 0.00566 -0.0689 0.6235 1.0000
4.250 0.7772 0.01357 0.00595 -0.0679 0.5818 1.0000
4.500 0.8007 0.01391 0.00618 -0.0668 0.5326 1.0000
4.750 0.8230 0.01439 0.00649 -0.0655 0.4773 1.0000
5.000 0.8443 0.01502 0.00692 -0.0643 0.4203 1.0000
5.250 0.8650 0.01574 0.00746 -0.0631 0.3656 1.0000
5.500 0.8859 0.01650 0.00808 -0.0620 0.3171 1.0000
5.750 0.9059 0.01736 0.00877 -0.0609 0.2668 1.0000
6.000 0.9247 0.01837 0.00948 -0.0597 0.2071 1.0000
6.250 0.9418 0.01968 0.01039 -0.0586 0.1229 1.0000
6.500 0.9550 0.02177 0.01169 -0.0570 0.0301 1.0000
6.750 0.9725 0.02344 0.01333 -0.0553 0.0177 1.0000
7.000 0.9917 0.02480 0.01497 -0.0539 0.0153 1.0000
7.250 1.0094 0.02626 0.01670 -0.0523 0.0138 1.0000
7.500 1.0256 0.02787 0.01856 -0.0506 0.0126 1.0000
7.750 1.0409 0.02956 0.02044 -0.0489 0.0113 1.0000
8.000 1.0509 0.03222 0.02332 -0.0469 0.0096 1.0000
8.250 1.0669 0.03411 0.02542 -0.0454 0.0090 1.0000
8.500 1.0822 0.03632 0.02792 -0.0438 0.0085 1.0000
8.750 1.0961 0.03896 0.03087 -0.0422 0.0084 1.0000
9.000 1.1075 0.04189 0.03416 -0.0404 0.0083 1.0000
9.250 1.1154 0.04500 0.03765 -0.0384 0.0082 1.0000
9.500 1.1186 0.04826 0.04129 -0.0362 0.0082 1.0000
9.750 1.1158 0.05152 0.04492 -0.0335 0.0082 1.0000
10.000 1.1084 0.05476 0.04859 -0.0309 0.0082 1.0000
10.250 1.0975 0.05825 0.05238 -0.0288 0.0082 1.0000
10.500 1.0843 0.06208 0.05648 -0.0275 0.0083 1.0000
10.750 1.0693 0.06636 0.06102 -0.0273 0.0083 1.0000
11.000 1.0533 0.07121 0.06609 -0.0282 0.0084 1.0000
11.250 1.0358 0.07688 0.07197 -0.0305 0.0084 1.0000
11.500 1.0182 0.08335 0.07862 -0.0342 0.0084 1.0000
11.750 1.0000 0.09097 0.08629 -0.0392 0.0085 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/8 AIRFOIL (hq258-il)