NACA 8-H-12 AIRFOIL (n8h12-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 8-H-12 AIRFOIL (n8h12-il) Reynolds number: 1,000,000 Max Cl/Cd: 118.18 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n8h12-il-1000000-n5.txt Download as CSV file: xf-n8h12-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 8-H-12 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
------ -------- --------- --------- -------- -------- --------
-14.000 -1.0843 0.05124 0.04882 -0.0241 0.7843 0.0070
-13.500 -1.1736 0.03636 0.03317 -0.0188 0.7637 0.0070
-13.250 -1.1758 0.03338 0.02992 -0.0167 0.7538 0.0071
-13.000 -1.1703 0.03104 0.02734 -0.0149 0.7441 0.0072
-12.750 -1.1591 0.02924 0.02535 -0.0135 0.7353 0.0074
-12.500 -1.1443 0.02780 0.02374 -0.0122 0.7261 0.0075
-12.250 -1.1281 0.02646 0.02224 -0.0110 0.7182 0.0076
-12.000 -1.1094 0.02541 0.02104 -0.0100 0.7099 0.0077
-11.750 -1.0901 0.02437 0.01987 -0.0090 0.7024 0.0079
-11.500 -1.0698 0.02343 0.01879 -0.0081 0.6948 0.0080
-11.250 -1.0491 0.02251 0.01773 -0.0072 0.6878 0.0082
-11.000 -1.0274 0.02169 0.01678 -0.0064 0.6804 0.0084
-10.750 -1.0054 0.02088 0.01583 -0.0056 0.6740 0.0086
-10.500 -0.9827 0.02013 0.01495 -0.0049 0.6673 0.0088
-10.000 -0.9364 0.01870 0.01325 -0.0035 0.6543 0.0091
-9.750 -0.9127 0.01805 0.01247 -0.0028 0.6475 0.0093
-9.500 -0.8894 0.01729 0.01159 -0.0021 0.6418 0.0096
-9.250 -0.8647 0.01676 0.01099 -0.0015 0.6353 0.0099
-8.750 -0.8139 0.01591 0.00999 -0.0006 0.6225 0.0104
-8.500 -0.7883 0.01550 0.00950 -0.0002 0.6158 0.0107
-8.250 -0.7625 0.01509 0.00901 0.0003 0.6103 0.0110
-8.000 -0.7367 0.01467 0.00850 0.0007 0.6042 0.0113
-7.750 -0.7108 0.01428 0.00801 0.0012 0.5975 0.0116
-7.500 -0.6845 0.01393 0.00757 0.0016 0.5915 0.0119
-7.250 -0.6586 0.01351 0.00710 0.0020 0.5849 0.0123
-7.000 -0.6317 0.01328 0.00683 0.0022 0.5793 0.0128
-6.750 -0.6047 0.01303 0.00654 0.0025 0.5737 0.0133
-6.500 -0.5778 0.01278 0.00623 0.0028 0.5677 0.0137
-6.250 -0.5510 0.01250 0.00589 0.0031 0.5621 0.0142
-6.000 -0.5240 0.01223 0.00555 0.0034 0.5560 0.0147
-5.750 -0.4970 0.01199 0.00524 0.0037 0.5496 0.0150
-5.500 -0.4705 0.01163 0.00485 0.0041 0.5446 0.0156
-5.250 -0.4433 0.01142 0.00460 0.0043 0.5390 0.0161
-5.000 -0.4159 0.01125 0.00439 0.0045 0.5336 0.0168
-4.750 -0.3885 0.01104 0.00415 0.0047 0.5290 0.0174
-4.500 -0.3611 0.01084 0.00390 0.0050 0.5237 0.0180
-4.250 -0.3336 0.01068 0.00368 0.0052 0.5182 0.0183
-4.000 -0.3067 0.01039 0.00335 0.0055 0.5133 0.0191
-3.750 -0.2794 0.01018 0.00313 0.0057 0.5079 0.0199
-3.500 -0.2518 0.01003 0.00294 0.0058 0.5029 0.0207
-3.250 -0.2242 0.00987 0.00275 0.0060 0.4988 0.0213
-3.000 -0.1965 0.00971 0.00257 0.0062 0.4944 0.0220
-2.750 -0.1687 0.00959 0.00241 0.0063 0.4898 0.0225
-2.500 -0.1412 0.00943 0.00222 0.0065 0.4856 0.0237
-2.250 -0.1134 0.00928 0.00207 0.0066 0.4819 0.0248
-2.000 -0.0855 0.00917 0.00194 0.0067 0.4776 0.0260
-1.750 -0.0576 0.00908 0.00182 0.0068 0.4732 0.0272
-1.500 -0.0297 0.00898 0.00172 0.0069 0.4691 0.0299
-1.250 -0.0017 0.00887 0.00162 0.0070 0.4656 0.0337
-1.000 0.0260 0.00875 0.00154 0.0071 0.4619 0.0451
-0.750 0.0537 0.00864 0.00148 0.0072 0.4583 0.0618
-0.500 0.0815 0.00856 0.00143 0.0073 0.4548 0.0771
-0.250 0.1094 0.00845 0.00138 0.0074 0.4521 0.0947
0.000 0.1367 0.00828 0.00134 0.0075 0.4491 0.1378
0.250 0.1633 0.00803 0.00130 0.0077 0.4460 0.2108
0.500 0.1886 0.00767 0.00127 0.0081 0.4427 0.3261
0.750 0.2113 0.00710 0.00124 0.0089 0.4395 0.5089
1.000 0.2289 0.00634 0.00120 0.0110 0.4371 0.7290
1.250 0.2657 0.00578 0.00133 0.0094 0.4343 0.9334
1.500 0.3192 0.00595 0.00149 0.0041 0.4310 0.9651
1.750 0.3572 0.00611 0.00162 0.0021 0.4282 0.9765
2.000 0.4001 0.00626 0.00172 -0.0011 0.4252 0.9823
2.250 0.4336 0.00636 0.00179 -0.0022 0.4227 0.9859
2.500 0.4697 0.00640 0.00182 -0.0040 0.4203 0.9873
2.750 0.5029 0.00645 0.00186 -0.0052 0.4176 0.9885
3.000 0.5352 0.00650 0.00190 -0.0062 0.4152 0.9898
3.250 0.5669 0.00656 0.00195 -0.0070 0.4127 0.9912
3.500 0.5980 0.00664 0.00201 -0.0078 0.4102 0.9925
3.750 0.6287 0.00671 0.00208 -0.0084 0.4076 0.9938
4.000 0.6599 0.00676 0.00214 -0.0092 0.4056 0.9948
4.250 0.6920 0.00680 0.00219 -0.0102 0.4029 0.9955
4.500 0.7237 0.00686 0.00225 -0.0111 0.3983 0.9964
4.750 0.7550 0.00696 0.00232 -0.0120 0.3927 0.9973
5.000 0.7865 0.00700 0.00238 -0.0129 0.3872 0.9982
5.250 0.8174 0.00712 0.00247 -0.0137 0.3773 0.9990
5.500 0.8482 0.00726 0.00256 -0.0146 0.3630 0.9997
5.750 0.8769 0.00742 0.00268 -0.0150 0.3481 1.0000
6.000 0.9023 0.00774 0.00287 -0.0148 0.3159 1.0000
6.250 0.9252 0.00867 0.00342 -0.0148 0.2412 1.0000
6.500 0.9484 0.00942 0.00394 -0.0147 0.1956 1.0000
6.750 0.9714 0.01014 0.00446 -0.0145 0.1562 1.0000
7.000 0.9938 0.01085 0.00499 -0.0143 0.1213 1.0000
7.250 1.0162 0.01148 0.00550 -0.0140 0.0951 1.0000
7.500 1.0383 0.01209 0.00599 -0.0136 0.0737 1.0000
7.750 1.0598 0.01270 0.00652 -0.0131 0.0564 1.0000
8.000 1.0815 0.01323 0.00699 -0.0126 0.0460 1.0000
8.250 1.1027 0.01375 0.00749 -0.0120 0.0375 1.0000
8.500 1.1230 0.01433 0.00802 -0.0114 0.0297 1.0000
8.750 1.1428 0.01488 0.00856 -0.0107 0.0247 1.0000
9.000 1.1620 0.01545 0.00913 -0.0099 0.0215 1.0000
9.250 1.1810 0.01598 0.00968 -0.0091 0.0197 1.0000
9.500 1.1985 0.01658 0.01029 -0.0082 0.0182 1.0000
9.750 1.2139 0.01728 0.01101 -0.0071 0.0169 1.0000
10.000 1.2289 0.01800 0.01178 -0.0062 0.0164 1.0000
10.250 1.2405 0.01892 0.01275 -0.0052 0.0160 1.0000
10.500 1.2410 0.01998 0.01385 -0.0024 0.0156 1.0000
10.750 1.2462 0.02114 0.01506 -0.0008 0.0151 1.0000
11.000 1.2536 0.02239 0.01635 0.0004 0.0149 1.0000
11.250 1.2607 0.02380 0.01780 0.0013 0.0144 1.0000
11.500 1.2674 0.02534 0.01939 0.0020 0.0141 1.0000
11.750 1.2733 0.02701 0.02111 0.0026 0.0137 1.0000
12.000 1.2784 0.02881 0.02297 0.0032 0.0133 1.0000
12.250 1.2844 0.03057 0.02479 0.0036 0.0133 1.0000
12.500 1.2918 0.03222 0.02649 0.0040 0.0131 1.0000
12.750 1.2970 0.03412 0.02845 0.0043 0.0130 1.0000
13.000 1.3021 0.03605 0.03045 0.0046 0.0128 1.0000
13.250 1.3072 0.03800 0.03245 0.0048 0.0127 1.0000
13.500 1.3111 0.04009 0.03461 0.0049 0.0125 1.0000
13.750 1.3143 0.04228 0.03686 0.0051 0.0124 1.0000
14.000 1.3175 0.04457 0.03921 0.0051 0.0122 1.0000
14.250 1.3209 0.04687 0.04157 0.0050 0.0121 1.0000
14.500 1.3232 0.04933 0.04410 0.0049 0.0120 1.0000
14.750 1.3262 0.05178 0.04660 0.0047 0.0118 1.0000
15.000 1.3286 0.05432 0.04922 0.0045 0.0117 1.0000
15.250 1.3303 0.05697 0.05193 0.0042 0.0115 1.0000
15.500 1.3314 0.05974 0.05475 0.0038 0.0115 1.0000
15.750 1.3325 0.06254 0.05762 0.0034 0.0113 1.0000
16.000 1.3335 0.06539 0.06052 0.0029 0.0111 1.0000
16.250 1.3325 0.06852 0.06371 0.0023 0.0109 1.0000
16.500 1.3317 0.07167 0.06693 0.0017 0.0108 1.0000
16.750 1.3276 0.07528 0.07061 0.0009 0.0106 1.0000
17.000 1.3260 0.07859 0.07399 0.0002 0.0105 1.0000
17.250 1.3228 0.08219 0.07767 -0.0007 0.0104 1.0000
17.500 1.3235 0.08531 0.08087 -0.0015 0.0103 1.0000
17.750 1.3262 0.08813 0.08375 -0.0022 0.0102 1.0000
18.000 1.3245 0.09161 0.08731 -0.0032 0.0102 1.0000
18.250 1.3212 0.09534 0.09113 -0.0042 0.0101 1.0000
18.500 1.3199 0.09886 0.09472 -0.0053 0.0100 1.0000
18.750 1.3189 0.10231 0.09824 -0.0064 0.0099 1.0000
19.000 1.3160 0.10611 0.10212 -0.0077 0.0098 1.0000
19.250 1.3133 0.10990 0.10599 -0.0090 0.0097 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 8-H-12 AIRFOIL (n8h12-il)