RG 8 AIRFOIL (rg8-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RG 8 AIRFOIL (rg8-il) Reynolds number: 200,000 Max Cl/Cd: 68.33 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rg8-il-200000-n5.txt Download as CSV file: xf-rg8-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RG 8 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5113 0.09246 0.08879 -0.0375 1.0000 0.0166
-10.500 -0.5908 0.06625 0.06254 -0.0523 1.0000 0.0152
-10.250 -0.6188 0.05812 0.05429 -0.0583 1.0000 0.0151
-10.000 -0.6431 0.05315 0.04925 -0.0605 1.0000 0.0150
-9.750 -0.6685 0.05002 0.04606 -0.0595 1.0000 0.0149
-9.500 -0.6888 0.04657 0.04243 -0.0582 1.0000 0.0149
-9.250 -0.7027 0.04294 0.03851 -0.0567 1.0000 0.0150
-9.000 -0.7072 0.03988 0.03508 -0.0552 1.0000 0.0153
-8.750 -0.7090 0.03641 0.03124 -0.0538 1.0000 0.0157
-8.500 -0.7027 0.03369 0.02827 -0.0526 1.0000 0.0161
-8.250 -0.6911 0.03180 0.02620 -0.0515 1.0000 0.0165
-8.000 -0.6777 0.02979 0.02393 -0.0505 1.0000 0.0168
-7.750 -0.6530 0.02761 0.02142 -0.0515 0.9982 0.0172
-7.500 -0.6236 0.02569 0.01917 -0.0530 0.9956 0.0177
-7.250 -0.5937 0.02401 0.01723 -0.0544 0.9932 0.0182
-7.000 -0.5643 0.02256 0.01555 -0.0554 0.9903 0.0189
-6.750 -0.5337 0.02131 0.01412 -0.0566 0.9874 0.0195
-6.500 -0.5015 0.02032 0.01296 -0.0580 0.9848 0.0205
-6.250 -0.4699 0.01937 0.01186 -0.0592 0.9821 0.0216
-6.000 -0.4415 0.01825 0.01067 -0.0600 0.9780 0.0227
-5.750 -0.4098 0.01743 0.00977 -0.0612 0.9745 0.0238
-5.500 -0.3758 0.01673 0.00898 -0.0629 0.9717 0.0252
-5.250 -0.3461 0.01610 0.00826 -0.0636 0.9671 0.0271
-5.000 -0.3144 0.01552 0.00756 -0.0646 0.9626 0.0296
-4.750 -0.2801 0.01493 0.00696 -0.0662 0.9594 0.0365
-4.500 -0.2497 0.01432 0.00644 -0.0671 0.9546 0.0546
-4.250 -0.2185 0.01378 0.00601 -0.0681 0.9496 0.0839
-4.000 -0.1838 0.01333 0.00563 -0.0698 0.9462 0.1147
-3.750 -0.1535 0.01286 0.00532 -0.0706 0.9407 0.1550
-3.250 -0.0867 0.01165 0.00469 -0.0737 0.9321 0.2986
-3.000 -0.0576 0.01115 0.00445 -0.0743 0.9254 0.3758
-2.750 -0.0251 0.01071 0.00434 -0.0753 0.9203 0.4674
-2.500 0.0105 0.01048 0.00423 -0.0768 0.9162 0.5281
-2.250 0.0401 0.01036 0.00413 -0.0769 0.9077 0.5597
-2.000 0.0780 0.01022 0.00397 -0.0788 0.9031 0.5848
-1.750 0.1083 0.01013 0.00388 -0.0791 0.8936 0.6048
-1.500 0.1445 0.01002 0.00373 -0.0806 0.8867 0.6239
-1.250 0.1776 0.00992 0.00363 -0.0814 0.8771 0.6399
-1.000 0.2108 0.00983 0.00354 -0.0823 0.8669 0.6546
-0.750 0.2454 0.00974 0.00343 -0.0835 0.8562 0.6688
-0.500 0.2794 0.00967 0.00334 -0.0845 0.8444 0.6827
-0.250 0.3117 0.00961 0.00327 -0.0851 0.8307 0.6952
0.000 0.3425 0.00958 0.00322 -0.0855 0.8155 0.7067
0.250 0.3723 0.00956 0.00319 -0.0856 0.7989 0.7179
0.500 0.4016 0.00958 0.00316 -0.0857 0.7812 0.7286
0.750 0.4303 0.00960 0.00315 -0.0856 0.7628 0.7383
1.000 0.4581 0.00964 0.00315 -0.0854 0.7436 0.7472
1.250 0.4851 0.00971 0.00317 -0.0850 0.7233 0.7569
1.500 0.5116 0.00979 0.00321 -0.0845 0.7028 0.7666
1.750 0.5372 0.00987 0.00326 -0.0839 0.6815 0.7767
2.000 0.5625 0.00998 0.00332 -0.0832 0.6598 0.7878
2.250 0.5873 0.01010 0.00341 -0.0823 0.6375 0.7998
2.500 0.6113 0.01023 0.00350 -0.0814 0.6149 0.8130
2.750 0.6348 0.01036 0.00361 -0.0803 0.5917 0.8276
3.000 0.6576 0.01051 0.00372 -0.0790 0.5687 0.8439
3.250 0.6798 0.01065 0.00387 -0.0777 0.5449 0.8627
3.500 0.7015 0.01080 0.00399 -0.0762 0.5210 0.8867
3.750 0.7239 0.01095 0.00412 -0.0749 0.4969 0.9192
4.000 0.7547 0.01114 0.00426 -0.0756 0.4699 0.9858
4.250 0.7794 0.01143 0.00448 -0.0751 0.4437 1.0000
4.500 0.8036 0.01176 0.00472 -0.0745 0.4180 1.0000
4.750 0.8273 0.01211 0.00498 -0.0739 0.3916 1.0000
5.000 0.8507 0.01249 0.00526 -0.0731 0.3648 1.0000
5.250 0.8738 0.01288 0.00556 -0.0724 0.3388 1.0000
5.500 0.8968 0.01330 0.00592 -0.0716 0.3124 1.0000
5.750 0.9195 0.01374 0.00628 -0.0708 0.2855 1.0000
6.000 0.9417 0.01421 0.00667 -0.0700 0.2582 1.0000
6.250 0.9637 0.01472 0.00709 -0.0691 0.2317 1.0000
6.500 0.9853 0.01526 0.00754 -0.0682 0.2053 1.0000
6.750 1.0066 0.01583 0.00805 -0.0673 0.1811 1.0000
7.000 1.0279 0.01640 0.00857 -0.0663 0.1588 1.0000
7.250 1.0486 0.01701 0.00913 -0.0653 0.1385 1.0000
7.500 1.0682 0.01773 0.00975 -0.0642 0.1153 1.0000
7.750 1.0870 0.01850 0.01040 -0.0629 0.0912 1.0000
8.000 1.1051 0.01932 0.01113 -0.0616 0.0699 1.0000
8.250 1.1235 0.02010 0.01186 -0.0603 0.0557 1.0000
8.500 1.1417 0.02086 0.01263 -0.0590 0.0455 1.0000
8.750 1.1588 0.02169 0.01347 -0.0575 0.0359 1.0000
9.000 1.1742 0.02264 0.01442 -0.0558 0.0269 1.0000
9.250 1.1872 0.02366 0.01545 -0.0536 0.0203 1.0000
9.500 1.1986 0.02475 0.01657 -0.0513 0.0169 1.0000
9.750 1.2102 0.02578 0.01775 -0.0490 0.0151 1.0000
10.000 1.2209 0.02686 0.01895 -0.0467 0.0138 1.0000
10.250 1.2296 0.02808 0.02027 -0.0443 0.0129 1.0000
10.500 1.2344 0.02958 0.02186 -0.0415 0.0122 1.0000
10.750 1.2409 0.03096 0.02340 -0.0391 0.0117 1.0000
11.000 1.2461 0.03248 0.02508 -0.0368 0.0113 1.0000
11.250 1.2505 0.03411 0.02686 -0.0346 0.0109 1.0000
11.500 1.2543 0.03586 0.02875 -0.0326 0.0104 1.0000
11.750 1.2580 0.03767 0.03070 -0.0309 0.0100 1.0000
12.000 1.2603 0.03968 0.03284 -0.0294 0.0096 1.0000
12.250 1.2613 0.04191 0.03519 -0.0281 0.0092 1.0000
12.500 1.2598 0.04450 0.03794 -0.0271 0.0089 1.0000
12.750 1.2558 0.04754 0.04111 -0.0264 0.0086 1.0000
13.000 1.2500 0.05098 0.04471 -0.0259 0.0085 1.0000
13.250 1.2466 0.05435 0.04822 -0.0259 0.0084 1.0000
13.500 1.2429 0.05794 0.05200 -0.0262 0.0083 1.0000
13.750 1.2380 0.06184 0.05609 -0.0269 0.0082 1.0000
14.000 1.2315 0.06617 0.06062 -0.0280 0.0081 1.0000
14.250 1.2242 0.07083 0.06546 -0.0295 0.0081 1.0000
14.500 1.2158 0.07586 0.07068 -0.0314 0.0080 1.0000
14.750 1.2061 0.08138 0.07640 -0.0338 0.0080 1.0000
15.000 1.1952 0.08736 0.08257 -0.0367 0.0080 1.0000
15.250 1.1832 0.09382 0.08922 -0.0400 0.0079 1.0000
15.500 1.1705 0.10070 0.09628 -0.0438 0.0079 1.0000
15.750 1.1566 0.10815 0.10393 -0.0481 0.0079 1.0000
16.000 1.1423 0.11596 0.11191 -0.0527 0.0079 1.0000
16.250 1.1273 0.12430 0.12042 -0.0578 0.0080 1.0000
16.500 1.1114 0.13318 0.12946 -0.0633 0.0080 1.0000
16.750 1.0949 0.14260 0.13903 -0.0691 0.0080 1.0000
17.000 1.0771 0.15286 0.14944 -0.0756 0.0081 1.0000
17.250 1.0578 0.16416 0.16086 -0.0825 0.0082 1.0000
17.500 1.0340 0.17769 0.17450 -0.0905 0.0084 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RG 8 AIRFOIL (rg8-il)