CHEN AIRFOIL (chen-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: CHEN AIRFOIL (chen-il) Reynolds number: 1,000,000 Max Cl/Cd: 123 at α=13° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-chen-il-1000000-n5.txt Download as CSV file: xf-chen-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: CHEN 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.0278 0.09943 0.09489 -0.0551 0.3150 0.0084
-9.250 -0.0213 0.09604 0.09150 -0.0568 0.3149 0.0078
-9.000 -0.0143 0.09295 0.08842 -0.0584 0.3146 0.0076
-8.750 -0.0013 0.09112 0.08660 -0.0597 0.3144 0.0087
-8.500 0.0067 0.08833 0.08381 -0.0612 0.3143 0.0088
-8.250 0.0142 0.08551 0.08100 -0.0627 0.3143 0.0086
-8.000 0.0221 0.08290 0.07840 -0.0642 0.3141 0.0089
-7.750 0.0282 0.08001 0.07553 -0.0659 0.3141 0.0088
-7.500 0.0343 0.07744 0.07298 -0.0673 0.3139 0.0086
-7.250 0.0351 0.07436 0.06992 -0.0692 0.3139 0.0091
-7.000 0.0358 0.07196 0.06754 -0.0695 0.3138 0.0091
-6.750 0.0306 0.06967 0.06527 -0.0684 0.3137 0.0092
-6.500 0.0353 0.06751 0.06311 -0.0684 0.3137 0.0091
-6.250 0.0351 0.06433 0.05993 -0.0683 0.3136 0.0092
-6.000 0.0384 0.06152 0.05710 -0.0679 0.3136 0.0092
-5.750 0.0461 0.05931 0.05486 -0.0673 0.3135 0.0092
-5.500 0.0441 0.05491 0.05042 -0.0661 0.3135 0.0097
-5.250 0.0518 0.05228 0.04773 -0.0649 0.3135 0.0097
-5.000 0.0693 0.05149 0.04693 -0.0643 0.3133 0.0100
-4.750 0.0821 0.04961 0.04500 -0.0632 0.3132 0.0102
-4.500 0.0944 0.04752 0.04286 -0.0618 0.3133 0.0103
-4.250 0.1087 0.04572 0.04101 -0.0605 0.3131 0.0106
-3.750 0.0302 0.02045 0.01404 -0.0363 0.3132 0.0136
-3.500 0.0487 0.01727 0.01049 -0.0345 0.3132 0.0139
-3.250 0.0764 0.01600 0.00908 -0.0344 0.3130 0.0142
-3.000 0.1039 0.01539 0.00835 -0.0341 0.3129 0.0144
-2.750 0.1314 0.01498 0.00787 -0.0338 0.3128 0.0147
-2.500 0.1591 0.01463 0.00744 -0.0336 0.3128 0.0150
-2.250 0.1865 0.01439 0.00714 -0.0332 0.3127 0.0153
-2.000 0.2139 0.01419 0.00689 -0.0329 0.3126 0.0157
-1.750 0.2416 0.01390 0.00656 -0.0326 0.3125 0.0161
-1.500 0.2690 0.01371 0.00634 -0.0323 0.3124 0.0164
-1.250 0.2962 0.01355 0.00616 -0.0319 0.3123 0.0172
-1.000 0.3230 0.01345 0.00604 -0.0315 0.3122 0.0177
-0.750 0.3494 0.01337 0.00594 -0.0310 0.3122 0.0183
-0.500 0.3756 0.01330 0.00585 -0.0304 0.3121 0.0189
-0.250 0.4015 0.01324 0.00577 -0.0297 0.3120 0.0195
0.000 0.4269 0.01316 0.00570 -0.0290 0.3120 0.0211
0.250 0.4521 0.01314 0.00567 -0.0283 0.3118 0.0223
0.500 0.4771 0.01313 0.00566 -0.0274 0.3118 0.0239
0.750 0.5018 0.01314 0.00569 -0.0266 0.3117 0.0264
1.000 0.5265 0.01315 0.00572 -0.0258 0.3116 0.0307
1.250 0.5512 0.01320 0.00578 -0.0249 0.3116 0.0345
1.500 0.5759 0.01327 0.00587 -0.0241 0.3115 0.0385
1.750 0.6006 0.01333 0.00593 -0.0233 0.3114 0.0416
2.000 0.6253 0.01339 0.00601 -0.0225 0.3113 0.0448
2.250 0.6500 0.01349 0.00612 -0.0218 0.3112 0.0472
2.500 0.6746 0.01358 0.00622 -0.0210 0.3112 0.0501
2.750 0.6993 0.01368 0.00635 -0.0203 0.3111 0.0537
3.000 0.7240 0.01380 0.00648 -0.0195 0.3110 0.0566
3.250 0.7487 0.01390 0.00663 -0.0188 0.3109 0.0629
3.500 0.7735 0.01403 0.00677 -0.0181 0.3108 0.0660
3.750 0.7981 0.01416 0.00691 -0.0174 0.3107 0.0694
4.000 0.8226 0.01431 0.00709 -0.0167 0.3106 0.0730
4.250 0.8472 0.01444 0.00724 -0.0160 0.3105 0.0762
4.500 0.8717 0.01460 0.00742 -0.0154 0.3104 0.0781
4.750 0.8961 0.01473 0.00758 -0.0147 0.3103 0.0819
5.000 0.9204 0.01490 0.00777 -0.0140 0.3102 0.0853
5.250 0.9447 0.01505 0.00795 -0.0133 0.3101 0.0881
5.500 0.9688 0.01524 0.00817 -0.0127 0.3100 0.0908
5.750 0.9930 0.01542 0.00838 -0.0120 0.3099 0.0954
6.000 1.0170 0.01558 0.00858 -0.0114 0.3098 0.1009
6.500 1.0653 0.01586 0.00898 -0.0101 0.3094 0.1289
6.750 1.0893 0.01602 0.00917 -0.0095 0.3092 0.1405
7.000 1.1130 0.01622 0.00942 -0.0088 0.3091 0.1494
7.250 1.1368 0.01640 0.00964 -0.0082 0.3089 0.1569
7.500 1.1604 0.01660 0.00987 -0.0076 0.3087 0.1630
7.750 1.1840 0.01679 0.01012 -0.0070 0.3086 0.1711
8.250 1.2323 0.01697 0.01039 -0.0059 0.3080 0.2048
8.750 1.3687 0.01550 0.01032 -0.0242 0.3063 0.9745
9.250 1.4284 0.01576 0.01059 -0.0256 0.3044 0.9832
9.500 1.4710 0.01597 0.01083 -0.0292 0.3042 0.9847
9.750 1.5057 0.01617 0.01109 -0.0311 0.3040 0.9857
10.000 1.5378 0.01597 0.01093 -0.0322 0.3030 0.9867
10.250 1.5664 0.01624 0.01126 -0.0329 0.3028 0.9877
10.500 1.5942 0.01622 0.01129 -0.0333 0.3019 0.9885
10.750 1.6250 0.01624 0.01135 -0.0343 0.3009 0.9900
11.000 1.6532 0.01613 0.01129 -0.0347 0.2996 0.9910
11.250 1.6786 0.01662 0.01185 -0.0349 0.2999 0.9924
11.500 1.7050 0.01653 0.01180 -0.0349 0.2987 0.9935
11.750 1.7350 0.01642 0.01172 -0.0358 0.2976 0.9944
12.000 1.7681 0.01553 0.01078 -0.0366 0.2947 0.9949
12.250 1.8006 0.01511 0.01036 -0.0378 0.2926 0.9958
12.500 1.8272 0.01543 0.01082 -0.0383 0.2873 0.9965
12.750 1.8610 0.01544 0.01089 -0.0401 0.2826 0.9978
13.000 1.8880 0.01535 0.01083 -0.0404 0.2783 0.9983
13.250 1.8782 0.01740 0.01232 -0.0358 0.2097 1.0000
13.500 1.8051 0.02087 0.01568 -0.0204 0.1745 1.0000
13.750 1.8063 0.02154 0.01642 -0.0167 0.1737 1.0000
14.000 1.7438 0.02568 0.02065 -0.0075 0.1604 1.0000
14.250 1.7546 0.02639 0.02149 -0.0065 0.1647 1.0000
14.500 1.7380 0.03009 0.02541 -0.0064 0.1697 1.0000
14.750 1.6261 0.04712 0.04311 -0.0105 0.1932 1.0000
15.250 1.4638 0.06912 0.06560 -0.0124 0.2099 1.0000
15.500 1.4691 0.07092 0.06749 -0.0124 0.2141 1.0000
15.750 1.3690 0.08547 0.08186 -0.0152 0.1875 1.0000
16.000 1.3697 0.08794 0.08435 -0.0153 0.1845 1.0000
16.250 1.3503 0.09299 0.08930 -0.0162 0.1716 1.0000
16.500 1.3372 0.09735 0.09361 -0.0169 0.1619 1.0000
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Polar data table (+)
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