FX 63-158 AIRFOIL (fx63158-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 63-158 AIRFOIL (fx63158-il) Reynolds number: 1,000,000 Max Cl/Cd: 85 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx63158-il-1000000.txt Download as CSV file: xf-fx63158-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: FX 63-158 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.0570 0.08436 0.08153 -0.1060 0.7212 0.0110
-11.250 -0.0635 0.07711 0.07423 -0.1091 0.7056 0.0115
-11.000 -0.0696 0.07008 0.06720 -0.1121 0.7014 0.0115
-10.750 -0.0766 0.06198 0.05908 -0.1158 0.6941 0.0115
-10.500 -0.0927 0.05345 0.05043 -0.1208 0.6798 0.0115
-10.250 -0.1158 0.04609 0.04299 -0.1255 0.6786 0.0115
-10.000 -0.1292 0.04156 0.03839 -0.1279 0.6741 0.0115
-9.750 -0.1453 0.03759 0.03428 -0.1294 0.6593 0.0115
-9.500 -0.1669 0.03402 0.03056 -0.1302 0.6528 0.0115
-9.250 -0.1803 0.03142 0.02785 -0.1298 0.6461 0.0115
-9.000 -0.2001 0.02943 0.02572 -0.1278 0.6373 0.0115
-8.750 -0.2131 0.02767 0.02380 -0.1261 0.6274 0.0115
-8.500 -0.2182 0.02551 0.02142 -0.1247 0.6204 0.0116
-8.250 -0.2276 0.02087 0.01653 -0.1235 0.6118 0.0118
-8.000 -0.2163 0.01888 0.01444 -0.1229 0.6021 0.0119
-7.750 -0.2032 0.01718 0.01264 -0.1225 0.5961 0.0121
-7.500 -0.1878 0.01569 0.01105 -0.1221 0.5879 0.0123
-7.250 -0.1714 0.01427 0.00946 -0.1218 0.5783 0.0126
-7.000 -0.1529 0.01292 0.00794 -0.1215 0.5695 0.0131
-6.750 -0.1303 0.01264 0.00716 -0.1208 0.5605 0.0146
-6.500 -0.1119 0.01000 0.00423 -0.1209 0.5537 0.0150
-6.250 -0.0890 0.00892 0.00311 -0.1209 0.5463 0.0153
-6.000 -0.0648 0.00809 0.00218 -0.1209 0.5396 0.0156
-5.750 -0.0404 0.00739 0.00133 -0.1208 0.5295 0.0162
-5.500 -0.0203 0.01451 0.00784 -0.1221 0.5254 0.0093
-5.250 0.0047 0.01396 0.00720 -0.1215 0.5083 0.0088
-5.000 0.0284 0.01303 0.00618 -0.1210 0.4944 0.0085
-4.750 0.0544 0.01237 0.00540 -0.1211 0.4804 0.0084
-4.500 0.0828 0.01181 0.00474 -0.1217 0.4698 0.0085
-4.250 0.1119 0.01142 0.00425 -0.1224 0.4580 0.0088
-4.000 0.1413 0.01107 0.00373 -0.1232 0.4286 0.0094
-3.750 0.1692 0.01102 0.00355 -0.1235 0.4042 0.0100
-3.500 0.1960 0.01110 0.00342 -0.1236 0.3586 0.0106
-3.250 0.2233 0.01119 0.00332 -0.1238 0.3201 0.0110
-3.000 0.2494 0.01139 0.00328 -0.1238 0.2700 0.0115
-2.750 0.2762 0.01156 0.00329 -0.1239 0.2277 0.0129
-2.500 0.3050 0.01158 0.00324 -0.1243 0.2116 0.0145
-2.250 0.3333 0.01165 0.00315 -0.1246 0.1802 0.0177
-2.000 0.3627 0.01163 0.00307 -0.1251 0.1720 0.0218
-1.750 0.3918 0.01163 0.00308 -0.1255 0.1682 0.0345
-1.500 0.4204 0.01168 0.00308 -0.1259 0.1581 0.0370
-1.250 0.4482 0.01181 0.00310 -0.1260 0.1347 0.0397
-1.000 0.4769 0.01185 0.00310 -0.1264 0.1302 0.0439
-0.750 0.5061 0.01186 0.00310 -0.1269 0.1267 0.0504
-0.500 0.5362 0.01181 0.00313 -0.1276 0.1254 0.0911
-0.250 0.5652 0.01183 0.00315 -0.1281 0.1235 0.1056
0.000 0.5952 0.01180 0.00319 -0.1288 0.1213 0.1431
0.250 0.6239 0.01185 0.00324 -0.1291 0.1185 0.1545
0.500 0.6533 0.01187 0.00331 -0.1298 0.1158 0.1899
0.750 0.6842 0.01180 0.00336 -0.1307 0.1144 0.2397
1.000 0.7190 0.01153 0.00338 -0.1329 0.1121 0.3627
1.250 0.7554 0.01123 0.00343 -0.1356 0.0964 0.5167
1.500 0.7859 0.01124 0.00360 -0.1365 0.0952 0.5866
1.750 0.8135 0.01139 0.00378 -0.1366 0.0939 0.6133
2.000 0.8408 0.01155 0.00393 -0.1366 0.0926 0.6241
2.250 0.8679 0.01172 0.00408 -0.1366 0.0912 0.6328
2.500 0.8948 0.01189 0.00425 -0.1365 0.0894 0.6382
2.750 0.9217 0.01205 0.00443 -0.1365 0.0883 0.6435
3.000 0.9484 0.01222 0.00460 -0.1364 0.0877 0.6486
3.250 0.9750 0.01241 0.00478 -0.1363 0.0870 0.6536
3.500 1.0014 0.01258 0.00499 -0.1362 0.0863 0.6583
3.750 1.0275 0.01278 0.00519 -0.1360 0.0854 0.6613
4.000 1.0533 0.01299 0.00541 -0.1357 0.0844 0.6657
4.250 1.0788 0.01324 0.00564 -0.1354 0.0833 0.6704
4.500 1.1041 0.01345 0.00587 -0.1352 0.0821 0.6734
4.750 1.1293 0.01367 0.00611 -0.1348 0.0809 0.6753
5.000 1.1543 0.01389 0.00634 -0.1345 0.0799 0.6779
5.250 1.1796 0.01408 0.00656 -0.1342 0.0795 0.6810
5.500 1.2048 0.01427 0.00677 -0.1339 0.0787 0.6837
5.750 1.2299 0.01447 0.00698 -0.1336 0.0765 0.6858
6.000 1.2531 0.01478 0.00725 -0.1330 0.0641 0.6871
6.250 1.2758 0.01510 0.00753 -0.1323 0.0630 0.6880
6.500 1.2983 0.01540 0.00782 -0.1316 0.0614 0.6894
6.750 1.3200 0.01573 0.00816 -0.1307 0.0593 0.6908
7.000 1.3414 0.01606 0.00852 -0.1298 0.0577 0.6922
7.250 1.3615 0.01642 0.00894 -0.1286 0.0567 0.6936
7.500 1.3786 0.01681 0.00939 -0.1269 0.0557 0.6949
7.750 1.3943 0.01726 0.00989 -0.1250 0.0545 0.6963
8.000 1.4093 0.01776 0.01043 -0.1230 0.0533 0.6977
8.250 1.4236 0.01829 0.01101 -0.1209 0.0520 0.6991
8.500 1.4391 0.01878 0.01154 -0.1191 0.0508 0.7006
8.750 1.4590 0.01906 0.01186 -0.1180 0.0494 0.7021
9.000 1.4769 0.01945 0.01229 -0.1167 0.0474 0.7034
9.250 1.4923 0.01996 0.01282 -0.1150 0.0450 0.7046
9.500 1.5088 0.02042 0.01332 -0.1136 0.0431 0.7065
9.750 1.5270 0.02079 0.01374 -0.1125 0.0365 0.7089
10.000 1.5360 0.02166 0.01454 -0.1101 0.0309 0.7110
10.250 1.5459 0.02253 0.01545 -0.1080 0.0284 0.7131
10.500 1.5526 0.02362 0.01657 -0.1056 0.0258 0.7154
10.750 1.5604 0.02472 0.01772 -0.1035 0.0236 0.7175
11.000 1.5704 0.02574 0.01880 -0.1019 0.0225 0.7197
11.250 1.5764 0.02711 0.02020 -0.1000 0.0190 0.7214
11.500 1.5773 0.02895 0.02205 -0.0979 0.0151 0.7234
11.750 1.5796 0.03083 0.02396 -0.0963 0.0124 0.7256
12.000 1.5821 0.03283 0.02601 -0.0949 0.0105 0.7278
12.250 1.5870 0.03475 0.02799 -0.0939 0.0092 0.7297
12.500 1.5889 0.03704 0.03035 -0.0929 0.0070 0.7315
12.750 1.5819 0.04036 0.03371 -0.0918 0.0033 0.7330
13.000 1.5815 0.04314 0.03657 -0.0911 0.0030 0.7348
13.250 1.5825 0.04587 0.03939 -0.0907 0.0028 0.7362
13.500 1.5838 0.04865 0.04225 -0.0903 0.0026 0.7376
13.750 1.5848 0.05152 0.04520 -0.0901 0.0024 0.7387
14.000 1.5858 0.05446 0.04824 -0.0900 0.0024 0.7399
14.250 1.5866 0.05752 0.05141 -0.0901 0.0023 0.7416
14.500 1.5866 0.06074 0.05473 -0.0902 0.0023 0.7433
14.750 1.5861 0.06412 0.05821 -0.0905 0.0022 0.7449
15.000 1.5855 0.06759 0.06179 -0.0909 0.0022 0.7465
15.250 1.5851 0.07110 0.06539 -0.0913 0.0021 0.7480
15.500 1.5839 0.07479 0.06919 -0.0919 0.0021 0.7494
15.750 1.5822 0.07865 0.07315 -0.0927 0.0020 0.7508
16.000 1.5803 0.08262 0.07722 -0.0935 0.0020 0.7522
16.250 1.5787 0.08660 0.08130 -0.0944 0.0020 0.7536
16.500 1.5755 0.09090 0.08570 -0.0955 0.0019 0.7548
16.750 1.5725 0.09522 0.09012 -0.0967 0.0019 0.7559
17.000 1.5696 0.09961 0.09461 -0.0980 0.0019 0.7575
17.250 1.5653 0.10433 0.09945 -0.0996 0.0019 0.7593
17.500 1.5608 0.10911 0.10435 -0.1013 0.0018 0.7610
17.750 1.5568 0.11384 0.10919 -0.1030 0.0018 0.7626
18.000 1.5512 0.11892 0.11438 -0.1051 0.0018 0.7642
18.250 1.5456 0.12404 0.11961 -0.1072 0.0018 0.7659
18.500 1.5400 0.12919 0.12488 -0.1095 0.0017 0.7676
|
Polar data table (+)
Polar graphs
<< Back to FX 63-158 AIRFOIL (fx63158-il)