NACA 5-H-10 AIRFOIL (n5h10-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 5-H-10 AIRFOIL (n5h10-il) Reynolds number: 500,000 Max Cl/Cd: 58.72 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n5h10-il-500000-n5.txt Download as CSV file: xf-n5h10-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 5-H-10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5296 0.08677 0.08404 -0.0135 0.7899 0.0063
-9.250 -0.5424 0.08080 0.07811 -0.0171 0.7862 0.0061
-9.000 -0.5584 0.07546 0.07277 -0.0207 0.7822 0.0060
-8.750 -0.5730 0.07128 0.06857 -0.0215 0.7779 0.0061
-8.500 -0.5887 0.06585 0.06307 -0.0221 0.7737 0.0060
-8.250 -0.6012 0.05942 0.05650 -0.0224 0.7700 0.0060
-8.000 -0.6068 0.05346 0.05035 -0.0219 0.7668 0.0062
-7.750 -0.6124 0.04635 0.04297 -0.0205 0.7633 0.0064
-7.500 -0.6690 0.02220 0.01688 -0.0105 0.7607 0.0073
-7.250 -0.6496 0.01997 0.01423 -0.0092 0.7570 0.0076
-7.000 -0.6273 0.01852 0.01247 -0.0083 0.7535 0.0079
-6.750 -0.6050 0.01708 0.01077 -0.0075 0.7498 0.0084
-6.500 -0.5801 0.01641 0.00999 -0.0071 0.7452 0.0090
-6.250 -0.5551 0.01576 0.00920 -0.0067 0.7409 0.0096
-6.000 -0.5296 0.01526 0.00858 -0.0063 0.7370 0.0105
-5.750 -0.5039 0.01469 0.00789 -0.0059 0.7324 0.0114
-5.500 -0.4792 0.01392 0.00702 -0.0054 0.7279 0.0127
-5.250 -0.4530 0.01366 0.00671 -0.0051 0.7237 0.0140
-5.000 -0.4266 0.01335 0.00635 -0.0049 0.7185 0.0155
-4.750 -0.4004 0.01305 0.00596 -0.0046 0.7132 0.0170
-4.500 -0.3745 0.01270 0.00557 -0.0043 0.7086 0.0191
-4.250 -0.3477 0.01253 0.00539 -0.0042 0.7033 0.0212
-4.000 -0.3213 0.01227 0.00506 -0.0040 0.6983 0.0235
-3.750 -0.2944 0.01214 0.00485 -0.0038 0.6936 0.0256
-3.500 -0.2675 0.01198 0.00464 -0.0037 0.6882 0.0265
-3.250 -0.2430 0.01138 0.00396 -0.0031 0.6831 0.0291
-3.000 -0.2172 0.01108 0.00363 -0.0027 0.6783 0.0313
-2.750 -0.1909 0.01085 0.00335 -0.0025 0.6731 0.0331
-2.500 -0.1646 0.01063 0.00308 -0.0022 0.6685 0.0348
-2.250 -0.1383 0.01042 0.00280 -0.0019 0.6643 0.0357
-2.000 -0.1118 0.01022 0.00257 -0.0017 0.6598 0.0364
-1.750 -0.0853 0.01006 0.00235 -0.0015 0.6556 0.0371
-1.500 -0.0587 0.00993 0.00216 -0.0012 0.6516 0.0378
-1.250 -0.0320 0.00977 0.00197 -0.0010 0.6464 0.0399
-1.000 -0.0053 0.00964 0.00182 -0.0008 0.6412 0.0431
-0.750 0.0212 0.00952 0.00170 -0.0006 0.6366 0.0501
-0.500 0.0474 0.00932 0.00161 -0.0004 0.6320 0.0830
-0.250 0.0694 0.00872 0.00152 0.0005 0.6277 0.2453
0.000 0.0827 0.00758 0.00135 0.0029 0.6239 0.5420
0.250 0.0897 0.00674 0.00154 0.0077 0.6206 0.8494
0.500 0.1152 0.00682 0.00164 0.0084 0.6170 0.8703
0.750 0.1413 0.00689 0.00169 0.0089 0.6132 0.8819
1.000 0.1672 0.00697 0.00172 0.0094 0.6094 0.8911
1.250 0.1946 0.00699 0.00177 0.0095 0.6050 0.8968
1.500 0.2200 0.00705 0.00182 0.0102 0.5994 0.9070
2.000 0.2727 0.00714 0.00194 0.0112 0.5895 0.9233
2.250 0.3006 0.00717 0.00197 0.0111 0.5834 0.9254
2.500 0.3282 0.00720 0.00198 0.0111 0.5725 0.9267
2.750 0.3554 0.00724 0.00197 0.0111 0.5552 0.9274
3.000 0.3826 0.00730 0.00198 0.0111 0.5383 0.9280
3.250 0.4097 0.00738 0.00201 0.0110 0.5179 0.9287
3.500 0.4361 0.00751 0.00206 0.0111 0.4860 0.9294
3.750 0.4604 0.00784 0.00216 0.0115 0.4246 0.9303
4.000 0.4798 0.00871 0.00255 0.0123 0.3191 0.9317
4.250 0.5043 0.00907 0.00279 0.0126 0.2843 0.9327
4.750 0.5473 0.01043 0.00354 0.0136 0.1333 0.9357
5.000 0.5720 0.01080 0.00383 0.0137 0.1058 0.9368
5.250 0.5969 0.01114 0.00410 0.0139 0.0854 0.9380
5.500 0.6205 0.01160 0.00443 0.0141 0.0514 0.9395
5.750 0.6445 0.01200 0.00478 0.0144 0.0398 0.9411
6.000 0.6689 0.01236 0.00516 0.0146 0.0356 0.9428
6.250 0.6932 0.01272 0.00556 0.0149 0.0326 0.9447
6.500 0.7182 0.01300 0.00591 0.0150 0.0316 0.9467
6.750 0.7432 0.01333 0.00631 0.0151 0.0300 0.9488
7.000 0.7691 0.01371 0.00678 0.0148 0.0279 0.9509
7.250 0.7944 0.01419 0.00729 0.0146 0.0254 0.9535
7.500 0.8186 0.01484 0.00801 0.0145 0.0226 0.9567
7.750 0.8464 0.01511 0.00834 0.0139 0.0210 0.9594
8.000 0.8735 0.01549 0.00878 0.0133 0.0179 0.9626
8.250 0.9007 0.01608 0.00939 0.0124 0.0149 0.9659
8.500 0.9276 0.01669 0.01006 0.0116 0.0132 0.9699
8.750 0.9543 0.01726 0.01065 0.0108 0.0114 0.9748
9.000 0.9790 0.01819 0.01161 0.0100 0.0097 0.9809
9.250 1.0039 0.01899 0.01251 0.0092 0.0090 0.9883
9.500 1.0274 0.01992 0.01353 0.0086 0.0084 0.9978
9.750 1.0348 0.02068 0.01439 0.0112 0.0080 1.0000
10.000 1.0411 0.02146 0.01521 0.0137 0.0075 1.0000
10.250 1.0494 0.02238 0.01617 0.0156 0.0070 1.0000
10.500 1.0538 0.02373 0.01760 0.0178 0.0066 1.0000
10.750 1.0644 0.02473 0.01869 0.0192 0.0064 1.0000
11.000 1.0743 0.02584 0.01990 0.0206 0.0061 1.0000
11.250 1.0839 0.02703 0.02119 0.0219 0.0058 1.0000
11.500 1.0912 0.02847 0.02274 0.0233 0.0056 1.0000
11.750 1.1005 0.02975 0.02412 0.0244 0.0053 1.0000
12.000 1.1078 0.03126 0.02573 0.0256 0.0052 1.0000
12.250 1.1144 0.03287 0.02746 0.0267 0.0051 1.0000
12.500 1.1229 0.03427 0.02894 0.0274 0.0049 1.0000
12.750 1.1269 0.03619 0.03096 0.0283 0.0047 1.0000
13.000 1.1289 0.03834 0.03327 0.0291 0.0046 1.0000
13.250 1.1268 0.04099 0.03607 0.0299 0.0045 1.0000
13.500 1.1267 0.04355 0.03882 0.0305 0.0043 1.0000
13.750 1.1252 0.04628 0.04169 0.0309 0.0044 1.0000
14.000 1.1216 0.04935 0.04496 0.0310 0.0043 1.0000
14.250 1.1171 0.05263 0.04842 0.0309 0.0042 1.0000
14.500 1.1116 0.05613 0.05210 0.0305 0.0042 1.0000
14.750 1.1026 0.06034 0.05650 0.0297 0.0041 1.0000
15.000 1.0981 0.06412 0.06043 0.0285 0.0040 1.0000
15.250 1.0813 0.06994 0.06645 0.0266 0.0040 1.0000
15.500 1.0707 0.07518 0.07186 0.0244 0.0040 1.0000
15.750 1.0576 0.08116 0.07800 0.0216 0.0039 1.0000
16.000 1.0328 0.08983 0.08688 0.0173 0.0039 1.0000
16.250 1.0095 0.09899 0.09621 0.0123 0.0040 1.0000
16.500 0.9861 0.10892 0.10630 0.0068 0.0040 1.0000
16.750 0.9516 0.12207 0.11961 -0.0002 0.0042 1.0000
17.000 0.9086 0.13897 0.13662 -0.0085 0.0043 1.0000
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