HQ 1.0/8 AIRFOIL (hq108-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.0/8 AIRFOIL (hq108-il) Reynolds number: 100,000 Max Cl/Cd: 47.97 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq108-il-100000.txt Download as CSV file: xf-hq108-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/8 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5546 0.09532 0.09058 -0.0102 1.0000 0.0945
-8.500 -0.5748 0.09111 0.08650 -0.0181 1.0000 0.0963
-8.250 -0.5978 0.08667 0.08203 -0.0262 1.0000 0.0968
-8.000 -0.5571 0.08317 0.07861 -0.0156 1.0000 0.1029
-7.750 -0.5640 0.07880 0.07430 -0.0198 1.0000 0.1063
-7.250 -0.5678 0.06953 0.06501 -0.0261 1.0000 0.1148
-7.000 -0.5795 0.06533 0.06048 -0.0319 1.0000 0.1244
-6.750 -0.5595 0.06156 0.05693 -0.0290 1.0000 0.1297
-6.500 -0.5533 0.05792 0.05323 -0.0297 1.0000 0.1424
-5.750 -0.5108 0.03828 0.03164 -0.0334 1.0000 0.0727
-5.500 -0.4886 0.03416 0.02655 -0.0313 1.0000 0.0578
-5.250 -0.4689 0.03018 0.02235 -0.0302 1.0000 0.0551
-5.000 -0.4477 0.02707 0.01883 -0.0288 1.0000 0.0537
-4.750 -0.4251 0.02452 0.01587 -0.0273 1.0000 0.0538
-4.500 -0.4017 0.02244 0.01343 -0.0258 1.0000 0.0557
-4.250 -0.3802 0.02059 0.01154 -0.0246 1.0000 0.0637
-4.000 -0.3581 0.01890 0.00969 -0.0230 1.0000 0.0725
-3.750 -0.3382 0.01731 0.00826 -0.0214 1.0000 0.0901
-3.500 -0.3197 0.01583 0.00702 -0.0198 1.0000 0.1312
-3.250 -0.3043 0.01348 0.00582 -0.0183 1.0000 0.2921
-3.000 -0.2998 0.01209 0.00614 -0.0127 1.0000 0.6787
-2.750 -0.2878 0.01209 0.00621 -0.0084 1.0000 0.7558
-2.500 -0.2768 0.01209 0.00622 -0.0039 1.0000 0.8118
-2.250 -0.2654 0.01206 0.00618 0.0005 1.0000 0.8618
-2.000 -0.2447 0.01206 0.00609 0.0032 1.0000 0.9141
-1.750 -0.1542 0.01242 0.00609 -0.0069 1.0000 0.9738
-1.500 -0.0650 0.01237 0.00566 -0.0188 1.0000 0.9978
-1.250 -0.0801 0.01213 0.00541 -0.0137 1.0000 1.0000
-1.000 -0.0911 0.01196 0.00517 -0.0089 1.0000 1.0000
-0.750 -0.0794 0.01201 0.00507 -0.0075 1.0000 1.0000
-0.500 -0.0621 0.01215 0.00507 -0.0069 1.0000 1.0000
-0.250 -0.0431 0.01234 0.00515 -0.0064 1.0000 1.0000
0.000 -0.0235 0.01258 0.00529 -0.0061 1.0000 1.0000
0.250 0.0113 0.01296 0.00556 -0.0086 0.9950 1.0000
0.500 0.0579 0.01338 0.00593 -0.0132 0.9852 1.0000
0.750 0.1037 0.01377 0.00629 -0.0176 0.9750 1.0000
1.000 0.1505 0.01413 0.00665 -0.0220 0.9650 1.0000
1.250 0.2016 0.01437 0.00693 -0.0270 0.9529 1.0000
1.500 0.2568 0.01439 0.00702 -0.0323 0.9366 1.0000
1.750 0.3140 0.01427 0.00705 -0.0376 0.9217 1.0000
2.000 0.3610 0.01420 0.00710 -0.0410 0.9083 1.0000
2.250 0.4027 0.01412 0.00714 -0.0432 0.8937 1.0000
2.500 0.4398 0.01401 0.00716 -0.0442 0.8776 1.0000
2.750 0.4736 0.01387 0.00719 -0.0443 0.8604 1.0000
3.000 0.4992 0.01380 0.00723 -0.0428 0.8381 1.0000
3.250 0.5241 0.01364 0.00717 -0.0408 0.8139 1.0000
3.500 0.5470 0.01346 0.00707 -0.0383 0.7852 1.0000
3.750 0.5692 0.01325 0.00691 -0.0356 0.7519 1.0000
4.000 0.5907 0.01309 0.00684 -0.0329 0.7102 1.0000
4.250 0.6119 0.01305 0.00679 -0.0304 0.6565 1.0000
4.500 0.6322 0.01318 0.00680 -0.0278 0.5800 1.0000
4.750 0.6499 0.01376 0.00690 -0.0251 0.4662 1.0000
5.000 0.6656 0.01496 0.00742 -0.0228 0.3387 1.0000
5.250 0.6803 0.01668 0.00840 -0.0209 0.2223 1.0000
5.500 0.6966 0.01859 0.00968 -0.0192 0.1364 1.0000
5.750 0.7150 0.02053 0.01140 -0.0176 0.0933 1.0000
6.000 0.7355 0.02253 0.01329 -0.0163 0.0694 1.0000
6.250 0.7590 0.02494 0.01571 -0.0152 0.0577 1.0000
6.500 0.7818 0.02734 0.01814 -0.0144 0.0483 1.0000
6.750 0.8068 0.03000 0.02123 -0.0131 0.0455 1.0000
7.000 0.8292 0.03325 0.02498 -0.0117 0.0442 1.0000
7.250 0.8480 0.03705 0.02934 -0.0101 0.0445 1.0000
7.500 0.8626 0.04125 0.03411 -0.0084 0.0453 1.0000
7.750 0.8735 0.04549 0.03887 -0.0068 0.0455 1.0000
8.000 0.8804 0.04992 0.04377 -0.0053 0.0457 1.0000
8.250 0.8837 0.05457 0.04881 -0.0040 0.0463 1.0000
8.500 0.8850 0.05954 0.05403 -0.0031 0.0477 1.0000
8.750 0.8727 0.06762 0.06263 -0.0024 0.0602 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.0/8 AIRFOIL (hq108-il)