HAM-STD HS1-708 AIRFOIL (hs1708-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HAM-STD HS1-708 AIRFOIL (hs1708-il) Reynolds number: 1,000,000 Max Cl/Cd: 116.99 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hs1708-il-1000000-n5.txt Download as CSV file: xf-hs1708-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HAM-STD HS1-708 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4592 0.05794 0.05561 -0.0842 0.8830 0.0049
-9.750 -0.4557 0.03122 0.02811 -0.1212 0.8691 0.0047
-9.500 -0.4337 0.02123 0.01751 -0.1294 0.8629 0.0047
-9.250 -0.4076 0.01872 0.01471 -0.1311 0.8578 0.0047
-9.000 -0.3804 0.01714 0.01291 -0.1322 0.8531 0.0048
-8.750 -0.3529 0.01593 0.01152 -0.1330 0.8484 0.0048
-8.500 -0.3252 0.01494 0.01034 -0.1337 0.8443 0.0049
-8.250 -0.2971 0.01409 0.00934 -0.1342 0.8409 0.0049
-8.000 -0.2687 0.01336 0.00848 -0.1347 0.8374 0.0050
-7.750 -0.2403 0.01271 0.00769 -0.1352 0.8341 0.0051
-7.500 -0.2118 0.01215 0.00700 -0.1356 0.8310 0.0052
-7.250 -0.1831 0.01165 0.00638 -0.1359 0.8282 0.0053
-7.000 -0.1543 0.01120 0.00584 -0.1363 0.8254 0.0054
-6.750 -0.1254 0.01079 0.00534 -0.1366 0.8225 0.0056
-6.500 -0.0965 0.01043 0.00489 -0.1369 0.8195 0.0058
-6.250 -0.0676 0.01011 0.00449 -0.1372 0.8169 0.0060
-6.000 -0.0386 0.00980 0.00410 -0.1375 0.8145 0.0064
-5.750 -0.0095 0.00951 0.00376 -0.1378 0.8124 0.0069
-5.500 0.0196 0.00926 0.00347 -0.1381 0.8100 0.0077
-5.250 0.0487 0.00904 0.00320 -0.1383 0.8076 0.0087
-5.000 0.0778 0.00879 0.00293 -0.1386 0.8051 0.0110
-4.750 0.1069 0.00852 0.00267 -0.1389 0.8027 0.0177
-4.500 0.1359 0.00834 0.00249 -0.1391 0.8004 0.0224
-4.250 0.1650 0.00820 0.00233 -0.1394 0.7985 0.0260
-4.000 0.1942 0.00803 0.00219 -0.1397 0.7961 0.0309
-3.750 0.2233 0.00789 0.00206 -0.1399 0.7929 0.0368
-3.500 0.2521 0.00778 0.00195 -0.1401 0.7884 0.0434
-3.250 0.2809 0.00767 0.00184 -0.1403 0.7835 0.0492
-3.000 0.3099 0.00758 0.00174 -0.1405 0.7783 0.0535
-2.750 0.3388 0.00751 0.00165 -0.1406 0.7737 0.0578
-2.500 0.3677 0.00743 0.00158 -0.1408 0.7696 0.0646
-2.250 0.3967 0.00732 0.00152 -0.1410 0.7650 0.0752
-2.000 0.4256 0.00725 0.00145 -0.1412 0.7598 0.0835
-1.750 0.4544 0.00720 0.00140 -0.1414 0.7549 0.0921
-1.500 0.4834 0.00713 0.00136 -0.1416 0.7495 0.1016
-1.250 0.5121 0.00709 0.00133 -0.1417 0.7433 0.1123
-1.000 0.5410 0.00705 0.00131 -0.1419 0.7386 0.1222
-0.750 0.5699 0.00701 0.00130 -0.1421 0.7335 0.1318
-0.500 0.5986 0.00700 0.00129 -0.1422 0.7278 0.1395
-0.250 0.6274 0.00698 0.00130 -0.1424 0.7215 0.1492
0.000 0.6558 0.00698 0.00130 -0.1425 0.7128 0.1586
0.250 0.6844 0.00699 0.00131 -0.1426 0.7031 0.1657
0.500 0.7128 0.00700 0.00133 -0.1427 0.6929 0.1754
0.750 0.7410 0.00703 0.00135 -0.1427 0.6795 0.1827
1.000 0.7685 0.00712 0.00139 -0.1426 0.6557 0.1924
1.250 0.7943 0.00737 0.00147 -0.1422 0.6102 0.2020
1.500 0.8194 0.00773 0.00162 -0.1417 0.5604 0.2111
1.750 0.8450 0.00804 0.00180 -0.1414 0.5214 0.2223
2.000 0.8711 0.00830 0.00195 -0.1411 0.4881 0.2351
2.250 0.8971 0.00855 0.00211 -0.1408 0.4556 0.2491
2.750 0.9487 0.00903 0.00249 -0.1403 0.3914 0.3262
3.250 0.9932 0.00849 0.00302 -0.1385 0.3285 1.0000
3.500 1.0191 0.00877 0.00320 -0.1382 0.3081 1.0000
3.750 1.0447 0.00907 0.00339 -0.1379 0.2880 1.0000
4.000 1.0701 0.00936 0.00358 -0.1375 0.2653 1.0000
4.250 1.0951 0.00970 0.00380 -0.1370 0.2418 1.0000
4.500 1.1194 0.01009 0.00405 -0.1365 0.2148 1.0000
4.750 1.1439 0.01045 0.00430 -0.1360 0.1922 1.0000
5.000 1.1675 0.01087 0.00460 -0.1353 0.1670 1.0000
5.250 1.1906 0.01134 0.00493 -0.1345 0.1413 1.0000
5.500 1.2126 0.01190 0.00532 -0.1336 0.1107 1.0000
5.750 1.2336 0.01254 0.00578 -0.1325 0.0788 1.0000
6.000 1.2560 0.01302 0.00617 -0.1316 0.0621 1.0000
6.250 1.2794 0.01340 0.00651 -0.1309 0.0539 1.0000
6.500 1.3025 0.01379 0.00686 -0.1301 0.0472 1.0000
6.750 1.3260 0.01413 0.00720 -0.1294 0.0427 1.0000
7.000 1.3487 0.01451 0.00757 -0.1286 0.0384 1.0000
7.250 1.3713 0.01489 0.00794 -0.1277 0.0348 1.0000
7.500 1.3942 0.01522 0.00829 -0.1269 0.0333 1.0000
7.750 1.4164 0.01559 0.00867 -0.1260 0.0313 1.0000
8.000 1.4377 0.01601 0.00911 -0.1249 0.0291 1.0000
8.250 1.4582 0.01648 0.00958 -0.1237 0.0264 1.0000
8.500 1.4792 0.01685 0.00998 -0.1226 0.0256 1.0000
8.750 1.4996 0.01724 0.01042 -0.1214 0.0247 1.0000
9.000 1.5191 0.01767 0.01088 -0.1200 0.0237 1.0000
9.250 1.5375 0.01814 0.01138 -0.1185 0.0225 1.0000
9.500 1.5537 0.01865 0.01191 -0.1165 0.0213 1.0000
9.750 1.5684 0.01925 0.01253 -0.1143 0.0197 1.0000
10.000 1.5843 0.01979 0.01311 -0.1124 0.0188 1.0000
10.250 1.6010 0.02028 0.01365 -0.1107 0.0179 1.0000
10.500 1.6167 0.02086 0.01426 -0.1089 0.0169 1.0000
10.750 1.6316 0.02150 0.01493 -0.1070 0.0158 1.0000
11.000 1.6456 0.02222 0.01569 -0.1050 0.0149 1.0000
11.250 1.6587 0.02302 0.01653 -0.1030 0.0139 1.0000
11.500 1.6726 0.02377 0.01734 -0.1012 0.0134 1.0000
11.750 1.6863 0.02454 0.01817 -0.0994 0.0128 1.0000
12.000 1.6993 0.02538 0.01906 -0.0976 0.0121 1.0000
12.250 1.7114 0.02630 0.02002 -0.0957 0.0113 1.0000
12.500 1.7224 0.02732 0.02108 -0.0938 0.0104 1.0000
12.750 1.7324 0.02845 0.02226 -0.0918 0.0096 1.0000
13.000 1.7431 0.02954 0.02343 -0.0900 0.0092 1.0000
13.250 1.7530 0.03070 0.02467 -0.0882 0.0087 1.0000
13.500 1.7619 0.03197 0.02600 -0.0864 0.0081 1.0000
13.750 1.7697 0.03337 0.02746 -0.0846 0.0075 1.0000
14.000 1.7761 0.03491 0.02905 -0.0828 0.0068 1.0000
14.250 1.7823 0.03649 0.03071 -0.0810 0.0063 1.0000
14.500 1.7878 0.03816 0.03246 -0.0793 0.0058 1.0000
14.750 1.7918 0.04002 0.03439 -0.0776 0.0052 1.0000
15.000 1.7938 0.04211 0.03655 -0.0758 0.0046 1.0000
15.250 1.7950 0.04429 0.03884 -0.0742 0.0042 1.0000
15.500 1.7961 0.04658 0.04122 -0.0727 0.0039 1.0000
15.750 1.7962 0.04907 0.04380 -0.0714 0.0036 1.0000
16.000 1.7948 0.05181 0.04663 -0.0702 0.0033 1.0000
16.250 1.7917 0.05481 0.04974 -0.0691 0.0031 1.0000
16.500 1.7875 0.05809 0.05312 -0.0683 0.0028 1.0000
16.750 1.7828 0.06157 0.05671 -0.0677 0.0027 1.0000
17.000 1.7783 0.06517 0.06044 -0.0674 0.0026 1.0000
17.250 1.7725 0.06909 0.06448 -0.0674 0.0025 1.0000
17.500 1.7648 0.07339 0.06890 -0.0677 0.0024 1.0000
17.750 1.7560 0.07805 0.07369 -0.0683 0.0023 1.0000
18.000 1.7456 0.08311 0.07887 -0.0693 0.0023 1.0000
18.250 1.7337 0.08857 0.08447 -0.0706 0.0022 1.0000
18.500 1.7197 0.09445 0.09049 -0.0723 0.0022 1.0000
18.750 1.7045 0.10068 0.09686 -0.0743 0.0021 1.0000
19.000 1.6878 0.10732 0.10364 -0.0768 0.0021 1.0000
19.250 1.6702 0.11429 0.11074 -0.0796 0.0020 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HAM-STD HS1-708 AIRFOIL (hs1708-il)