NACA 63012A AIRFOIL (n63012a-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: NACA 63012A AIRFOIL (n63012a-il) Reynolds number: 200,000 Max Cl/Cd: 51.95 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63012a-il-200000.txt Download as CSV file: xf-n63012a-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63012A AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.5141 0.11904 0.11558 -0.0069 1.0000 0.0712
-12.000 -0.5189 0.11475 0.11130 -0.0085 1.0000 0.0735
-11.500 -0.6962 0.10499 0.10148 -0.0063 1.0000 0.0680
-9.750 -0.8905 0.05026 0.04509 -0.0230 1.0000 0.0416
-9.500 -0.8901 0.04594 0.04049 -0.0212 1.0000 0.0405
-9.250 -0.8863 0.04211 0.03631 -0.0191 1.0000 0.0399
-9.000 -0.8795 0.03833 0.03212 -0.0170 1.0000 0.0391
-8.750 -0.8694 0.03444 0.02774 -0.0148 1.0000 0.0378
-8.500 -0.8542 0.03145 0.02433 -0.0130 1.0000 0.0374
-8.250 -0.8354 0.02919 0.02178 -0.0117 1.0000 0.0377
-8.000 -0.8146 0.02729 0.01967 -0.0106 1.0000 0.0383
-7.750 -0.7933 0.02581 0.01804 -0.0096 1.0000 0.0396
-7.500 -0.7717 0.02459 0.01665 -0.0086 1.0000 0.0414
-7.250 -0.7495 0.02341 0.01530 -0.0075 1.0000 0.0426
-7.000 -0.7277 0.02172 0.01353 -0.0064 1.0000 0.0439
-6.750 -0.7088 0.02021 0.01207 -0.0051 1.0000 0.0461
-6.500 -0.6906 0.01927 0.01116 -0.0035 1.0000 0.0487
-6.250 -0.6725 0.01857 0.01040 -0.0018 1.0000 0.0525
-6.000 -0.6589 0.01745 0.00932 0.0005 1.0000 0.0571
-5.750 -0.6437 0.01672 0.00860 0.0026 1.0000 0.0631
-5.500 -0.6312 0.01583 0.00776 0.0051 1.0000 0.0717
-5.250 -0.6182 0.01508 0.00707 0.0075 1.0000 0.0876
-5.000 -0.6072 0.01404 0.00633 0.0099 1.0000 0.1298
-4.750 -0.6006 0.01233 0.00556 0.0124 1.0000 0.2914
-4.500 -0.5886 0.01133 0.00535 0.0144 1.0000 0.4545
-4.250 -0.5496 0.01104 0.00535 0.0117 0.9938 0.5460
-4.000 -0.5102 0.01097 0.00535 0.0092 0.9864 0.5925
-3.750 -0.4710 0.01097 0.00538 0.0069 0.9794 0.6260
-3.500 -0.4311 0.01100 0.00541 0.0045 0.9724 0.6543
-3.250 -0.3933 0.01105 0.00545 0.0025 0.9650 0.6786
-3.000 -0.3540 0.01110 0.00552 0.0005 0.9585 0.7010
-2.750 -0.3182 0.01114 0.00556 -0.0009 0.9508 0.7196
-2.500 -0.2821 0.01112 0.00553 -0.0022 0.9431 0.7352
-2.250 -0.2498 0.01109 0.00548 -0.0029 0.9337 0.7484
-2.000 -0.2159 0.01101 0.00538 -0.0039 0.9260 0.7605
-1.750 -0.1877 0.01095 0.00527 -0.0037 0.9147 0.7717
-1.500 -0.1585 0.01088 0.00522 -0.0036 0.9050 0.7808
-1.250 -0.1301 0.01080 0.00511 -0.0034 0.8955 0.7911
-1.000 -0.1044 0.01076 0.00505 -0.0028 0.8839 0.8019
-0.750 -0.0778 0.01073 0.00502 -0.0021 0.8735 0.8111
-0.500 -0.0512 0.01068 0.00495 -0.0015 0.8642 0.8216
-0.250 -0.0261 0.01067 0.00494 -0.0007 0.8525 0.8326
0.000 0.0000 0.01068 0.00496 0.0000 0.8420 0.8420
0.250 0.0261 0.01067 0.00494 0.0007 0.8326 0.8525
0.500 0.0512 0.01068 0.00495 0.0015 0.8216 0.8642
0.750 0.0778 0.01073 0.00502 0.0021 0.8111 0.8734
1.000 0.1044 0.01076 0.00505 0.0028 0.8019 0.8839
1.250 0.1301 0.01080 0.00511 0.0034 0.7911 0.8955
1.500 0.1585 0.01088 0.00522 0.0037 0.7808 0.9050
1.750 0.1877 0.01094 0.00527 0.0037 0.7716 0.9148
2.000 0.2158 0.01101 0.00538 0.0039 0.7605 0.9260
2.250 0.2497 0.01109 0.00548 0.0029 0.7485 0.9338
2.500 0.2820 0.01112 0.00553 0.0023 0.7352 0.9431
2.750 0.3180 0.01114 0.00556 0.0009 0.7196 0.9509
3.000 0.3539 0.01110 0.00552 -0.0005 0.7010 0.9586
3.250 0.3930 0.01105 0.00544 -0.0025 0.6787 0.9652
3.500 0.4310 0.01100 0.00541 -0.0044 0.6543 0.9725
3.750 0.4708 0.01097 0.00538 -0.0068 0.6261 0.9796
4.000 0.5101 0.01097 0.00534 -0.0092 0.5926 0.9865
4.250 0.5494 0.01103 0.00535 -0.0117 0.5461 0.9940
4.500 0.5881 0.01132 0.00535 -0.0143 0.4560 1.0000
4.750 0.6001 0.01230 0.00555 -0.0123 0.2941 1.0000
5.000 0.6067 0.01402 0.00631 -0.0098 0.1309 1.0000
5.250 0.6176 0.01507 0.00706 -0.0074 0.0879 1.0000
5.500 0.6308 0.01582 0.00775 -0.0050 0.0718 1.0000
5.750 0.6433 0.01671 0.00859 -0.0026 0.0632 1.0000
6.000 0.6586 0.01744 0.00931 -0.0005 0.0571 1.0000
6.250 0.6723 0.01856 0.01039 0.0018 0.0525 1.0000
6.500 0.6904 0.01927 0.01115 0.0035 0.0487 1.0000
6.750 0.7086 0.02020 0.01207 0.0051 0.0461 1.0000
7.000 0.7276 0.02172 0.01352 0.0064 0.0439 1.0000
7.250 0.7495 0.02341 0.01529 0.0075 0.0426 1.0000
7.500 0.7717 0.02459 0.01665 0.0085 0.0414 1.0000
7.750 0.7934 0.02581 0.01803 0.0096 0.0396 1.0000
8.000 0.8148 0.02729 0.01968 0.0106 0.0383 1.0000
8.250 0.8356 0.02919 0.02179 0.0116 0.0377 1.0000
8.500 0.8546 0.03147 0.02435 0.0129 0.0374 1.0000
8.750 0.8698 0.03447 0.02777 0.0147 0.0378 1.0000
9.000 0.8800 0.03837 0.03216 0.0169 0.0391 1.0000
9.250 0.8869 0.04216 0.03636 0.0190 0.0399 1.0000
9.500 0.8907 0.04600 0.04055 0.0210 0.0406 1.0000
9.750 0.8912 0.05035 0.04518 0.0229 0.0416 1.0000
10.000 0.7736 0.06033 0.05641 0.0315 0.0779 1.0000
10.250 0.7385 0.06291 0.05919 0.0346 0.0775 1.0000
10.500 0.7093 0.06724 0.06364 0.0351 0.0775 1.0000
10.750 0.6830 0.07259 0.06907 0.0341 0.0776 1.0000
11.000 0.6562 0.07871 0.07526 0.0321 0.0777 1.0000
11.250 0.6276 0.08575 0.08236 0.0289 0.0778 1.0000
11.500 0.5978 0.09391 0.09055 0.0243 0.0779 1.0000
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Polar data table (+)
Polar graphs
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