NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-12(N)1 AIRFOIL (rc12n1-il) Reynolds number: 200,000 Max Cl/Cd: 66.6 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc12n1-il-200000.txt Download as CSV file: xf-rc12n1-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-12(N)1 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5934 0.08971 0.08657 -0.0095 1.0000 0.0510
-9.250 -0.6032 0.08263 0.07951 -0.0166 1.0000 0.0516
-9.000 -0.6180 0.07672 0.07357 -0.0216 1.0000 0.0519
-8.750 -0.6311 0.07227 0.06905 -0.0232 1.0000 0.0527
-8.500 -0.6382 0.06797 0.06464 -0.0243 1.0000 0.0540
-8.250 -0.6479 0.06455 0.06082 -0.0250 1.0000 0.0575
-8.000 -0.6518 0.06371 0.05942 -0.0239 1.0000 0.0585
-7.750 -0.6445 0.05474 0.05061 -0.0257 1.0000 0.0602
-7.500 -0.6198 0.05130 0.04720 -0.0280 0.9744 0.0625
-7.250 -0.5980 0.04858 0.04423 -0.0293 0.9547 0.0679
-7.000 -0.5901 0.04512 0.04021 -0.0282 0.9366 0.0735
-6.750 -0.5728 0.04232 0.03740 -0.0276 0.9241 0.0758
-6.500 -0.5575 0.04032 0.03521 -0.0262 0.9130 0.0798
-6.250 -0.5458 0.03806 0.03249 -0.0242 0.9020 0.0882
-6.000 -0.5270 0.03582 0.03020 -0.0233 0.8933 0.0919
-5.750 -0.5018 0.02985 0.02309 -0.0192 0.8858 0.0552
-5.500 -0.4797 0.02690 0.01974 -0.0178 0.8787 0.0517
-5.250 -0.4539 0.02525 0.01744 -0.0158 0.8715 0.0471
-5.000 -0.4286 0.02397 0.01595 -0.0150 0.8649 0.0468
-4.750 -0.4024 0.02203 0.01381 -0.0145 0.8583 0.0470
-4.500 -0.3770 0.02009 0.01175 -0.0138 0.8531 0.0480
-4.250 -0.3508 0.01899 0.01067 -0.0136 0.8464 0.0508
-4.000 -0.3259 0.01828 0.00992 -0.0129 0.8405 0.0541
-3.750 -0.3006 0.01755 0.00912 -0.0123 0.8349 0.0568
-3.500 -0.2769 0.01663 0.00816 -0.0115 0.8288 0.0601
-3.250 -0.2545 0.01599 0.00755 -0.0104 0.8238 0.0667
-3.000 -0.2309 0.01535 0.00690 -0.0097 0.8177 0.0741
-2.750 -0.2076 0.01482 0.00636 -0.0089 0.8121 0.0871
-2.500 -0.1871 0.01398 0.00573 -0.0075 0.8077 0.1362
-2.250 -0.1888 0.01158 0.00562 -0.0023 0.8012 0.6689
-2.000 -0.1682 0.01133 0.00561 0.0000 0.7962 0.7421
-1.750 -0.1442 0.01125 0.00568 0.0016 0.7918 0.7918
-1.500 -0.1169 0.01132 0.00586 0.0023 0.7859 0.8321
-1.250 -0.0899 0.01149 0.00603 0.0033 0.7811 0.8670
-1.000 -0.0605 0.01177 0.00627 0.0039 0.7769 0.8960
-0.750 -0.0270 0.01205 0.00651 0.0033 0.7710 0.9171
-0.500 0.0075 0.01219 0.00657 0.0022 0.7662 0.9306
-0.250 0.0486 0.01235 0.00662 -0.0004 0.7621 0.9378
0.000 0.0882 0.01249 0.00672 -0.0030 0.7558 0.9477
0.250 0.1311 0.01257 0.00672 -0.0061 0.7509 0.9551
0.500 0.1730 0.01267 0.00675 -0.0092 0.7460 0.9636
0.750 0.2145 0.01274 0.00680 -0.0123 0.7394 0.9736
1.000 0.2585 0.01273 0.00672 -0.0157 0.7344 0.9812
1.250 0.3031 0.01274 0.00672 -0.0196 0.7276 0.9893
1.500 0.3482 0.01266 0.00661 -0.0235 0.7213 0.9972
1.750 0.3802 0.01258 0.00651 -0.0248 0.7155 1.0000
2.000 0.4047 0.01254 0.00647 -0.0246 0.7081 1.0000
2.250 0.4285 0.01247 0.00634 -0.0239 0.7029 1.0000
2.500 0.4536 0.01250 0.00643 -0.0238 0.6946 1.0000
2.750 0.4778 0.01244 0.00633 -0.0231 0.6887 1.0000
3.000 0.5027 0.01246 0.00640 -0.0228 0.6804 1.0000
3.250 0.5272 0.01240 0.00632 -0.0221 0.6738 1.0000
3.500 0.5520 0.01241 0.00640 -0.0216 0.6649 1.0000
3.750 0.5761 0.01226 0.00620 -0.0206 0.6557 1.0000
4.000 0.6000 0.01204 0.00596 -0.0195 0.6429 1.0000
4.250 0.6240 0.01187 0.00582 -0.0186 0.6287 1.0000
4.500 0.6483 0.01175 0.00572 -0.0178 0.6144 1.0000
4.750 0.6726 0.01167 0.00566 -0.0169 0.5995 1.0000
5.000 0.6969 0.01161 0.00563 -0.0161 0.5836 1.0000
5.250 0.7211 0.01160 0.00567 -0.0154 0.5643 1.0000
5.500 0.7449 0.01161 0.00569 -0.0145 0.5407 1.0000
5.750 0.7681 0.01168 0.00574 -0.0136 0.5105 1.0000
6.000 0.7906 0.01187 0.00585 -0.0126 0.4696 1.0000
6.250 0.8110 0.01228 0.00606 -0.0114 0.4120 1.0000
6.500 0.8281 0.01309 0.00650 -0.0099 0.3373 1.0000
6.750 0.8417 0.01428 0.00721 -0.0082 0.2480 1.0000
7.000 0.8520 0.01579 0.00817 -0.0062 0.1622 1.0000
7.250 0.8639 0.01708 0.00918 -0.0042 0.1216 1.0000
7.500 0.8759 0.01827 0.01019 -0.0022 0.1014 1.0000
7.750 0.8887 0.01934 0.01118 -0.0001 0.0879 1.0000
8.000 0.9043 0.02015 0.01202 0.0015 0.0783 1.0000
8.250 0.9173 0.02120 0.01308 0.0035 0.0710 1.0000
8.500 0.9307 0.02220 0.01406 0.0053 0.0648 1.0000
8.750 0.9444 0.02338 0.01528 0.0072 0.0603 1.0000
9.000 0.9599 0.02435 0.01630 0.0088 0.0561 1.0000
9.250 0.9746 0.02607 0.01791 0.0103 0.0521 1.0000
9.500 0.9918 0.02692 0.01892 0.0116 0.0492 1.0000
9.750 1.0091 0.02805 0.02015 0.0129 0.0464 1.0000
10.000 1.0269 0.02933 0.02145 0.0139 0.0442 1.0000
10.250 1.0515 0.03215 0.02430 0.0140 0.0418 1.0000
10.500 1.0660 0.03326 0.02565 0.0155 0.0404 1.0000
10.750 1.0804 0.03471 0.02733 0.0169 0.0387 1.0000
11.000 1.0947 0.03651 0.02932 0.0182 0.0373 1.0000
11.250 1.1075 0.03846 0.03148 0.0195 0.0362 1.0000
11.500 1.1167 0.04049 0.03369 0.0211 0.0353 1.0000
11.750 1.1243 0.04254 0.03587 0.0227 0.0344 1.0000
12.000 1.1305 0.04582 0.03931 0.0239 0.0333 1.0000
12.250 1.1211 0.05052 0.04432 0.0260 0.0328 1.0000
12.500 1.1102 0.05255 0.04662 0.0283 0.0325 1.0000
12.750 1.0982 0.05542 0.04975 0.0297 0.0322 1.0000
13.000 1.0839 0.05901 0.05361 0.0305 0.0321 1.0000
13.250 1.0672 0.06318 0.05802 0.0307 0.0320 1.0000
13.500 1.0485 0.06785 0.06291 0.0303 0.0321 1.0000
13.750 1.0279 0.07307 0.06835 0.0291 0.0322 1.0000
14.000 1.0055 0.07892 0.07439 0.0272 0.0323 1.0000
14.250 0.9819 0.08547 0.08111 0.0247 0.0325 1.0000
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Polar data table (+)
Polar graphs
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