NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 500,000 Max Cl/Cd: 51 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n63010a-il-500000-n5.txt Download as CSV file: xf-n63010a-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A010 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
------ -------- --------- --------- -------- -------- --------
-12.750 -0.9715 0.06212 0.05959 -0.0152 1.0000 0.0066
-12.500 -1.0090 0.05253 0.04968 -0.0210 1.0000 0.0065
-12.250 -1.0246 0.04761 0.04456 -0.0229 1.0000 0.0066
-12.000 -1.0348 0.04398 0.04073 -0.0233 1.0000 0.0066
-11.750 -1.0411 0.04115 0.03773 -0.0226 1.0000 0.0067
-11.500 -1.0423 0.03908 0.03550 -0.0211 1.0000 0.0068
-11.250 -1.0406 0.03727 0.03353 -0.0193 1.0000 0.0070
-11.000 -1.0347 0.03539 0.03144 -0.0178 1.0000 0.0072
-10.750 -1.0266 0.03343 0.02925 -0.0164 1.0000 0.0075
-10.500 -1.0168 0.03141 0.02699 -0.0150 1.0000 0.0079
-10.250 -1.0051 0.02939 0.02470 -0.0137 1.0000 0.0082
-10.000 -0.9911 0.02752 0.02256 -0.0124 1.0000 0.0085
-9.750 -0.9747 0.02593 0.02073 -0.0113 1.0000 0.0088
-9.500 -0.9563 0.02465 0.01924 -0.0104 1.0000 0.0090
-9.250 -0.9409 0.02260 0.01695 -0.0091 1.0000 0.0094
-9.000 -0.9218 0.02136 0.01558 -0.0083 1.0000 0.0097
-8.750 -0.9007 0.02052 0.01464 -0.0076 1.0000 0.0101
-8.500 -0.8785 0.01987 0.01392 -0.0070 1.0000 0.0105
-8.250 -0.8562 0.01920 0.01317 -0.0064 1.0000 0.0110
-8.000 -0.8340 0.01844 0.01232 -0.0057 1.0000 0.0116
-7.750 -0.8121 0.01762 0.01140 -0.0049 1.0000 0.0120
-7.500 -0.7903 0.01683 0.01050 -0.0041 1.0000 0.0123
-7.250 -0.7683 0.01612 0.00969 -0.0033 1.0000 0.0127
-7.000 -0.7461 0.01549 0.00898 -0.0024 1.0000 0.0129
-6.750 -0.7260 0.01462 0.00804 -0.0012 1.0000 0.0136
-6.500 -0.7052 0.01395 0.00734 -0.0002 1.0000 0.0144
-6.250 -0.6836 0.01344 0.00680 0.0008 1.0000 0.0150
-6.000 -0.6620 0.01298 0.00630 0.0019 1.0000 0.0157
-5.750 -0.6353 0.01252 0.00578 0.0018 0.9976 0.0164
-5.500 -0.6026 0.01207 0.00527 0.0005 0.9922 0.0173
-5.250 -0.5699 0.01167 0.00481 -0.0009 0.9859 0.0182
-5.000 -0.5382 0.01123 0.00433 -0.0019 0.9779 0.0200
-4.750 -0.5064 0.01085 0.00395 -0.0030 0.9689 0.0238
-4.500 -0.4745 0.01053 0.00361 -0.0041 0.9587 0.0290
-4.250 -0.4435 0.01022 0.00331 -0.0049 0.9467 0.0374
-4.000 -0.4143 0.00991 0.00302 -0.0053 0.9327 0.0513
-3.750 -0.3870 0.00959 0.00276 -0.0053 0.9170 0.0738
-3.500 -0.3614 0.00920 0.00250 -0.0050 0.9008 0.1153
-3.250 -0.3378 0.00860 0.00220 -0.0045 0.8845 0.2000
-3.000 -0.3162 0.00775 0.00188 -0.0037 0.8689 0.3394
-2.750 -0.2928 0.00723 0.00168 -0.0031 0.8542 0.4370
-2.500 -0.2675 0.00699 0.00156 -0.0026 0.8407 0.4896
-2.250 -0.2424 0.00676 0.00148 -0.0021 0.8277 0.5461
-2.000 -0.2171 0.00657 0.00142 -0.0016 0.8156 0.5958
-1.750 -0.1909 0.00648 0.00138 -0.0012 0.8038 0.6280
-1.500 -0.1642 0.00642 0.00134 -0.0009 0.7924 0.6507
-1.250 -0.1372 0.00638 0.00131 -0.0006 0.7813 0.6696
-1.000 -0.1100 0.00635 0.00127 -0.0005 0.7708 0.6846
-0.750 -0.0826 0.00635 0.00125 -0.0003 0.7602 0.6960
-0.500 -0.0550 0.00634 0.00123 -0.0002 0.7493 0.7067
0.000 0.0000 0.00633 0.00121 0.0000 0.7282 0.7283
0.500 0.0551 0.00634 0.00123 0.0002 0.7067 0.7493
0.750 0.0826 0.00635 0.00125 0.0003 0.6959 0.7601
1.000 0.1100 0.00635 0.00127 0.0005 0.6846 0.7708
1.250 0.1372 0.00638 0.00131 0.0006 0.6695 0.7813
1.500 0.1642 0.00642 0.00134 0.0009 0.6507 0.7924
1.750 0.1910 0.00648 0.00138 0.0012 0.6280 0.8039
2.000 0.2171 0.00657 0.00142 0.0016 0.5958 0.8156
2.250 0.2424 0.00676 0.00148 0.0021 0.5453 0.8277
2.500 0.2676 0.00699 0.00156 0.0026 0.4893 0.8407
2.750 0.2929 0.00723 0.00168 0.0031 0.4369 0.8541
3.000 0.3163 0.00774 0.00188 0.0037 0.3402 0.8687
3.250 0.3380 0.00859 0.00220 0.0044 0.2011 0.8843
3.500 0.3615 0.00920 0.00250 0.0050 0.1151 0.9007
3.750 0.3871 0.00959 0.00276 0.0053 0.0738 0.9169
4.000 0.4143 0.00991 0.00302 0.0053 0.0513 0.9328
4.250 0.4435 0.01022 0.00331 0.0049 0.0373 0.9467
4.500 0.4745 0.01053 0.00361 0.0040 0.0290 0.9586
4.750 0.5064 0.01085 0.00395 0.0030 0.0238 0.9688
5.000 0.5382 0.01123 0.00433 0.0019 0.0200 0.9779
5.250 0.5699 0.01167 0.00481 0.0009 0.0182 0.9859
5.500 0.6026 0.01207 0.00527 -0.0005 0.0173 0.9923
5.750 0.6353 0.01252 0.00578 -0.0018 0.0164 0.9977
6.000 0.6620 0.01298 0.00630 -0.0019 0.0156 1.0000
6.250 0.6836 0.01344 0.00680 -0.0008 0.0150 1.0000
6.500 0.7052 0.01395 0.00733 0.0002 0.0144 1.0000
6.750 0.7260 0.01462 0.00804 0.0012 0.0136 1.0000
7.000 0.7461 0.01549 0.00898 0.0024 0.0129 1.0000
7.250 0.7683 0.01612 0.00969 0.0033 0.0127 1.0000
7.500 0.7903 0.01683 0.01051 0.0041 0.0123 1.0000
7.750 0.8122 0.01762 0.01140 0.0049 0.0120 1.0000
8.000 0.8341 0.01845 0.01233 0.0057 0.0116 1.0000
8.250 0.8563 0.01920 0.01318 0.0064 0.0111 1.0000
8.500 0.8788 0.01987 0.01392 0.0070 0.0105 1.0000
8.750 0.9009 0.02053 0.01465 0.0075 0.0101 1.0000
9.000 0.9222 0.02137 0.01560 0.0082 0.0097 1.0000
9.250 0.9414 0.02260 0.01695 0.0090 0.0094 1.0000
9.500 0.9567 0.02467 0.01926 0.0103 0.0090 1.0000
9.750 0.9752 0.02595 0.02075 0.0112 0.0088 1.0000
10.000 0.9916 0.02754 0.02258 0.0123 0.0085 1.0000
10.250 1.0056 0.02941 0.02472 0.0136 0.0082 1.0000
10.500 1.0174 0.03142 0.02699 0.0149 0.0079 1.0000
10.750 1.0273 0.03344 0.02926 0.0163 0.0075 1.0000
11.000 1.0354 0.03542 0.03147 0.0176 0.0072 1.0000
11.250 1.0419 0.03723 0.03349 0.0190 0.0070 1.0000
11.500 1.0441 0.03901 0.03542 0.0209 0.0068 1.0000
11.750 1.0421 0.04119 0.03776 0.0223 0.0067 1.0000
12.000 1.0363 0.04396 0.04071 0.0230 0.0066 1.0000
12.250 1.0266 0.04755 0.04449 0.0227 0.0066 1.0000
12.500 1.0114 0.05241 0.04955 0.0208 0.0065 1.0000
12.750 0.9725 0.06225 0.05972 0.0148 0.0066 1.0000
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