NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il) Reynolds number: 200,000 Max Cl/Cd: 51.21 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc410-il-200000-n5.txt Download as CSV file: xf-rc410-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(4)-10 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5869 0.09635 0.09288 0.0067 1.0000 0.0242
-9.500 -0.7783 0.04528 0.04098 -0.0253 1.0000 0.0223
-9.250 -0.7899 0.03732 0.03224 -0.0251 1.0000 0.0224
-9.000 -0.7781 0.03433 0.02897 -0.0245 1.0000 0.0227
-8.750 -0.7620 0.03204 0.02646 -0.0240 1.0000 0.0230
-8.500 -0.7425 0.03003 0.02422 -0.0238 0.9750 0.0234
-8.250 -0.7171 0.02818 0.02206 -0.0244 0.9290 0.0238
-8.000 -0.6957 0.02674 0.02034 -0.0237 0.9043 0.0245
-7.750 -0.6752 0.02537 0.01866 -0.0226 0.8868 0.0253
-7.500 -0.6540 0.02402 0.01699 -0.0216 0.8734 0.0260
-7.250 -0.6318 0.02279 0.01546 -0.0206 0.8626 0.0266
-7.000 -0.6086 0.02166 0.01410 -0.0198 0.8530 0.0271
-6.750 -0.5856 0.02073 0.01306 -0.0190 0.8447 0.0278
-6.500 -0.5612 0.01996 0.01220 -0.0184 0.8364 0.0285
-6.250 -0.5369 0.01927 0.01137 -0.0176 0.8297 0.0294
-6.000 -0.5114 0.01856 0.01054 -0.0171 0.8227 0.0306
-5.750 -0.4861 0.01792 0.00974 -0.0165 0.8167 0.0319
-5.500 -0.4607 0.01732 0.00909 -0.0160 0.8109 0.0337
-5.250 -0.4346 0.01683 0.00854 -0.0156 0.8047 0.0361
-5.000 -0.4088 0.01632 0.00793 -0.0150 0.7994 0.0391
-4.750 -0.3825 0.01585 0.00745 -0.0147 0.7939 0.0428
-4.500 -0.3560 0.01541 0.00699 -0.0143 0.7885 0.0480
-4.250 -0.3295 0.01508 0.00661 -0.0138 0.7840 0.0550
-4.000 -0.3024 0.01485 0.00637 -0.0136 0.7792 0.0637
-3.750 -0.2749 0.01468 0.00623 -0.0134 0.7739 0.0720
-3.500 -0.2477 0.01454 0.00605 -0.0131 0.7692 0.0792
-3.250 -0.2202 0.01448 0.00590 -0.0128 0.7650 0.0875
-3.000 -0.1922 0.01443 0.00591 -0.0128 0.7600 0.0954
-2.750 -0.1646 0.01432 0.00574 -0.0126 0.7556 0.1025
-2.500 -0.1373 0.01419 0.00561 -0.0123 0.7518 0.1082
-2.250 -0.1096 0.01406 0.00541 -0.0121 0.7471 0.1140
-2.000 -0.0827 0.01380 0.00517 -0.0118 0.7399 0.1189
-1.750 -0.0561 0.01362 0.00495 -0.0113 0.7322 0.1241
-1.500 -0.0289 0.01345 0.00473 -0.0109 0.7227 0.1296
-1.250 -0.0024 0.01319 0.00448 -0.0104 0.7150 0.1346
-1.000 0.0248 0.01301 0.00431 -0.0101 0.7067 0.1403
-0.750 0.0519 0.01286 0.00411 -0.0097 0.6991 0.1464
-0.500 0.0788 0.01263 0.00394 -0.0094 0.6902 0.1531
-0.250 0.1059 0.01249 0.00376 -0.0090 0.6816 0.1601
0.000 0.1329 0.01230 0.00361 -0.0087 0.6724 0.1689
0.250 0.1601 0.01215 0.00348 -0.0084 0.6628 0.1800
0.500 0.1869 0.01197 0.00333 -0.0079 0.6520 0.1958
0.750 0.2129 0.01164 0.00320 -0.0075 0.6392 0.2486
1.000 0.2250 0.00981 0.00317 -0.0042 0.6279 0.7556
1.250 0.2826 0.00965 0.00333 -0.0092 0.6060 0.9419
1.500 0.3251 0.00975 0.00331 -0.0119 0.5763 0.9710
1.750 0.3696 0.00991 0.00324 -0.0151 0.5264 0.9904
2.000 0.4084 0.01073 0.00322 -0.0180 0.3579 1.0000
2.250 0.4324 0.01136 0.00346 -0.0177 0.2920 1.0000
2.500 0.4569 0.01175 0.00364 -0.0172 0.2625 1.0000
2.750 0.4815 0.01207 0.00382 -0.0167 0.2441 1.0000
3.000 0.5062 0.01234 0.00400 -0.0162 0.2314 1.0000
3.250 0.5307 0.01262 0.00419 -0.0157 0.2211 1.0000
3.500 0.5551 0.01291 0.00440 -0.0151 0.2117 1.0000
3.750 0.5795 0.01317 0.00461 -0.0145 0.2039 1.0000
4.000 0.6037 0.01347 0.00484 -0.0139 0.1968 1.0000
4.250 0.6279 0.01377 0.00509 -0.0132 0.1913 1.0000
4.500 0.6522 0.01403 0.00535 -0.0126 0.1857 1.0000
4.750 0.6760 0.01437 0.00562 -0.0120 0.1803 1.0000
5.000 0.6999 0.01468 0.00591 -0.0113 0.1754 1.0000
5.250 0.7240 0.01496 0.00620 -0.0106 0.1706 1.0000
5.500 0.7476 0.01530 0.00651 -0.0099 0.1664 1.0000
5.750 0.7706 0.01572 0.00687 -0.0092 0.1626 1.0000
6.000 0.7946 0.01601 0.00721 -0.0085 0.1587 1.0000
6.250 0.8183 0.01633 0.00755 -0.0078 0.1545 1.0000
6.500 0.8415 0.01671 0.00790 -0.0071 0.1507 1.0000
6.750 0.8642 0.01717 0.00833 -0.0064 0.1476 1.0000
7.000 0.8879 0.01751 0.00874 -0.0057 0.1443 1.0000
7.250 0.9114 0.01787 0.00915 -0.0050 0.1406 1.0000
7.500 0.9345 0.01826 0.00956 -0.0044 0.1372 1.0000
7.750 0.9567 0.01876 0.01002 -0.0036 0.1341 1.0000
8.000 0.9799 0.01920 0.01053 -0.0030 0.1314 1.0000
8.250 1.0030 0.01961 0.01102 -0.0023 0.1282 1.0000
8.500 1.0258 0.02003 0.01150 -0.0017 0.1249 1.0000
8.750 1.0479 0.02051 0.01198 -0.0010 0.1220 1.0000
9.000 1.0695 0.02111 0.01257 -0.0003 0.1194 1.0000
9.250 1.0920 0.02158 0.01318 0.0004 0.1167 1.0000
9.500 1.1141 0.02207 0.01376 0.0011 0.1135 1.0000
9.750 1.1356 0.02256 0.01431 0.0018 0.1107 1.0000
10.000 1.1561 0.02314 0.01490 0.0025 0.1081 1.0000
10.250 1.1765 0.02379 0.01562 0.0033 0.1056 1.0000
10.500 1.1972 0.02436 0.01634 0.0041 0.1028 1.0000
10.750 1.2172 0.02494 0.01702 0.0049 0.0998 1.0000
11.000 1.2361 0.02554 0.01767 0.0057 0.0972 1.0000
11.250 1.2533 0.02627 0.01842 0.0067 0.0947 1.0000
11.500 1.2714 0.02699 0.01929 0.0076 0.0922 1.0000
11.750 1.2886 0.02771 0.02016 0.0086 0.0894 1.0000
12.000 1.3037 0.02843 0.02098 0.0098 0.0868 1.0000
12.250 1.3160 0.02922 0.02182 0.0112 0.0845 1.0000
12.500 1.3270 0.03023 0.02287 0.0125 0.0824 1.0000
12.750 1.3399 0.03122 0.02406 0.0135 0.0799 1.0000
13.000 1.3515 0.03228 0.02527 0.0145 0.0773 1.0000
13.250 1.3617 0.03344 0.02652 0.0153 0.0750 1.0000
13.500 1.3697 0.03479 0.02791 0.0160 0.0728 1.0000
13.750 1.3782 0.03626 0.02952 0.0167 0.0705 1.0000
14.000 1.3866 0.03780 0.03125 0.0173 0.0680 1.0000
14.250 1.3934 0.03946 0.03304 0.0176 0.0655 1.0000
14.500 1.3975 0.04138 0.03502 0.0179 0.0634 1.0000
14.750 1.4003 0.04355 0.03730 0.0180 0.0613 1.0000
15.000 1.4037 0.04579 0.03973 0.0181 0.0588 1.0000
15.250 1.4047 0.04831 0.04239 0.0179 0.0565 1.0000
15.500 1.4024 0.05125 0.04542 0.0174 0.0546 1.0000
15.750 1.3971 0.05467 0.04893 0.0166 0.0529 1.0000
16.000 1.3921 0.05826 0.05274 0.0157 0.0510 1.0000
16.250 1.3834 0.06247 0.05713 0.0143 0.0494 1.0000
16.500 1.3709 0.06744 0.06226 0.0123 0.0480 1.0000
16.750 1.3542 0.07344 0.06841 0.0095 0.0470 1.0000
17.000 1.3325 0.08075 0.07587 0.0057 0.0463 1.0000
17.250 1.3049 0.08960 0.08490 0.0008 0.0460 1.0000
17.500 1.2710 0.10008 0.09557 -0.0050 0.0459 1.0000
17.750 1.2316 0.11174 0.10741 -0.0113 0.0460 1.0000
18.000 1.1919 0.12361 0.11941 -0.0177 0.0458 1.0000
18.250 1.1633 0.13335 0.12922 -0.0231 0.0451 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA RC(4)-10 AIRFOIL (rc410-il)