NACA 63A210 (naca63a210-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63A210 (naca63a210-il) Reynolds number: 500,000 Max Cl/Cd: 52.01 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca63a210-il-500000.txt Download as CSV file: xf-naca63a210-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A210
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5795 0.08959 0.08638 -0.0062 0.4960 0.0124
-8.500 -0.5845 0.08183 0.07864 -0.0148 0.4963 0.0123
-8.250 -0.5974 0.07572 0.07248 -0.0207 0.4966 0.0122
-8.000 -0.6044 0.07116 0.06781 -0.0227 0.4969 0.0123
-7.750 -0.6052 0.06680 0.06333 -0.0241 0.4973 0.0124
-7.500 -0.6020 0.06270 0.05908 -0.0249 0.4977 0.0127
-7.250 -0.5954 0.05880 0.05500 -0.0252 0.4981 0.0130
-7.000 -0.5858 0.05509 0.05110 -0.0251 0.4985 0.0135
-6.750 -0.5732 0.05163 0.04741 -0.0248 0.4990 0.0141
-6.500 -0.5574 0.04859 0.04411 -0.0241 0.4994 0.0151
-6.250 -0.5325 0.04805 0.04331 -0.0230 0.4998 0.0166
-6.000 -0.5139 0.04664 0.04160 -0.0221 0.5002 0.0169
-5.750 -0.4965 0.04456 0.03920 -0.0212 0.5003 0.0170
-5.500 -0.4890 0.03735 0.03144 -0.0200 0.5004 0.0182
-5.250 -0.4692 0.03539 0.02931 -0.0195 0.5005 0.0192
-5.000 -0.4473 0.03407 0.02782 -0.0190 0.5006 0.0203
-4.500 -0.3940 0.02839 0.02149 -0.0164 0.5009 0.0154
-4.250 -0.3672 0.02493 0.01777 -0.0152 0.5013 0.0141
-4.000 -0.3404 0.02308 0.01580 -0.0146 0.5018 0.0144
-3.750 -0.3132 0.02232 0.01498 -0.0143 0.5023 0.0161
-3.500 -0.2871 0.02132 0.01390 -0.0138 0.5029 0.0165
-3.250 -0.2626 0.02016 0.01268 -0.0130 0.5036 0.0165
-3.000 -0.2374 0.01947 0.01192 -0.0124 0.5043 0.0169
-2.750 -0.2142 0.01836 0.01068 -0.0115 0.5050 0.0198
-2.500 -0.1881 0.01809 0.01034 -0.0112 0.5058 0.0245
-2.250 -0.1643 0.01743 0.00999 -0.0106 0.5066 0.1224
-2.000 -0.1494 0.01590 0.00999 -0.0094 0.5074 0.5407
-1.750 -0.1271 0.01630 0.01069 -0.0086 0.5080 0.6426
-1.500 -0.1044 0.01693 0.01150 -0.0076 0.5082 0.7143
-1.250 -0.0812 0.01757 0.01222 -0.0067 0.5084 0.7609
-1.000 -0.0521 0.01652 0.01120 -0.0065 0.5087 0.7772
-0.750 -0.0230 0.01575 0.01045 -0.0065 0.5092 0.7902
-0.500 0.0052 0.01532 0.01006 -0.0064 0.5098 0.8036
-0.250 0.0330 0.01505 0.00983 -0.0063 0.5106 0.8168
0.000 0.0604 0.01490 0.00972 -0.0062 0.5114 0.8307
0.250 0.0874 0.01485 0.00972 -0.0060 0.5124 0.8465
0.500 0.1141 0.01487 0.00981 -0.0058 0.5132 0.8645
0.750 0.1405 0.01495 0.00996 -0.0055 0.5139 0.8846
1.250 0.1966 0.01542 0.01053 -0.0056 0.5155 0.9305
1.500 0.2287 0.01594 0.01106 -0.0067 0.5161 0.9517
1.750 0.2641 0.01633 0.01146 -0.0085 0.5164 0.9669
2.000 0.3023 0.01616 0.01134 -0.0108 0.5159 0.9780
2.250 0.3413 0.01589 0.01113 -0.0134 0.5144 0.9857
2.500 0.3789 0.01606 0.01134 -0.0157 0.5121 0.9928
2.750 0.4153 0.01726 0.01252 -0.0182 0.5092 0.9978
3.000 0.4456 0.01733 0.01269 -0.0192 0.5072 1.0000
3.250 0.4691 0.01726 0.01267 -0.0185 0.5029 1.0000
3.500 0.4887 0.01935 0.01469 -0.0177 0.5006 1.0000
3.750 0.5150 0.01818 0.01367 -0.0171 0.4951 1.0000
4.000 0.5385 0.01820 0.01374 -0.0164 0.4859 1.0000
4.250 0.5644 0.01783 0.01341 -0.0159 0.4800 1.0000
4.500 0.5967 0.01618 0.01200 -0.0164 0.4592 1.0000
4.750 0.6290 0.01469 0.01032 -0.0162 0.4509 1.0000
5.000 0.6566 0.01429 0.00990 -0.0160 0.4311 1.0000
5.250 0.6817 0.01436 0.00963 -0.0157 0.3212 1.0000
5.500 0.7060 0.01429 0.00887 -0.0153 0.2022 1.0000
5.750 0.7299 0.01453 0.00879 -0.0149 0.1407 1.0000
6.000 0.7540 0.01481 0.00891 -0.0145 0.1045 1.0000
6.250 0.7781 0.01513 0.00910 -0.0141 0.0776 1.0000
6.500 0.8022 0.01545 0.00933 -0.0137 0.0565 1.0000
6.750 0.8254 0.01587 0.00964 -0.0131 0.0320 1.0000
7.000 0.8479 0.01642 0.01013 -0.0124 0.0194 1.0000
7.250 0.8678 0.01733 0.01112 -0.0112 0.0136 1.0000
7.500 0.8896 0.01803 0.01190 -0.0103 0.0115 1.0000
7.750 0.9097 0.01890 0.01284 -0.0093 0.0100 1.0000
8.000 0.9249 0.02029 0.01435 -0.0076 0.0091 1.0000
8.250 0.9359 0.02212 0.01634 -0.0052 0.0087 1.0000
8.500 0.9457 0.02428 0.01867 -0.0026 0.0085 1.0000
8.750 0.9645 0.02540 0.01992 -0.0015 0.0083 1.0000
9.000 0.9851 0.02620 0.02088 -0.0008 0.0078 1.0000
9.250 1.0026 0.02745 0.02227 0.0003 0.0073 1.0000
9.500 1.0163 0.02938 0.02440 0.0019 0.0072 1.0000
9.750 1.0283 0.03158 0.02682 0.0034 0.0070 1.0000
10.000 1.0380 0.03404 0.02954 0.0050 0.0068 1.0000
10.250 1.0438 0.03689 0.03266 0.0066 0.0066 1.0000
10.500 1.0429 0.04010 0.03616 0.0085 0.0066 1.0000
10.750 1.0350 0.04368 0.04003 0.0101 0.0065 1.0000
11.000 1.0216 0.04816 0.04478 0.0101 0.0066 1.0000
11.250 1.0066 0.05379 0.05069 0.0072 0.0067 1.0000
11.500 0.9900 0.06007 0.05718 0.0034 0.0068 1.0000
11.750 0.9712 0.06612 0.06340 0.0006 0.0069 1.0000
12.000 0.9511 0.07226 0.06970 -0.0023 0.0070 1.0000
12.250 0.9306 0.07873 0.07629 -0.0056 0.0071 1.0000
12.500 0.9101 0.08605 0.08374 -0.0099 0.0072 1.0000
12.750 0.8892 0.09470 0.09252 -0.0154 0.0073 1.0000
13.000 0.8667 0.10610 0.10402 -0.0226 0.0075 1.0000
13.250 0.8342 0.12248 0.12045 -0.0307 0.0080 1.0000
13.500 0.8232 0.12970 0.12765 -0.0337 0.0087 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63A210 (naca63a210-il)