RAE 2822 AIRFOIL (rae2822-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAE 2822 AIRFOIL (rae2822-il) Reynolds number: 1,000,000 Max Cl/Cd: 79.03 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae2822-il-1000000.txt Download as CSV file: xf-rae2822-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 2822 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-16.000 -0.8978 0.08724 0.08500 -0.0420 1.0000 0.0075
-15.750 -0.9227 0.07820 0.07580 -0.0473 1.0000 0.0075
-15.500 -0.9398 0.07135 0.06882 -0.0513 1.0000 0.0076
-15.250 -0.9585 0.06479 0.06210 -0.0547 1.0000 0.0076
-15.000 -0.9756 0.05901 0.05616 -0.0574 1.0000 0.0076
-14.750 -0.9941 0.05358 0.05054 -0.0595 1.0000 0.0075
-14.500 -0.9968 0.05044 0.04731 -0.0606 1.0000 0.0076
-14.250 -1.0103 0.04610 0.04280 -0.0616 1.0000 0.0076
-14.000 -1.0130 0.04322 0.03979 -0.0621 1.0000 0.0077
-13.750 -1.0167 0.04033 0.03677 -0.0624 1.0000 0.0079
-13.500 -1.0214 0.03742 0.03371 -0.0622 1.0000 0.0078
-13.250 -1.0220 0.03506 0.03122 -0.0619 1.0000 0.0081
-13.000 -1.0205 0.03302 0.02906 -0.0613 1.0000 0.0081
-12.750 -1.0193 0.03104 0.02696 -0.0604 1.0000 0.0082
-12.500 -1.0161 0.02940 0.02519 -0.0592 1.0000 0.0083
-12.250 -1.0120 0.02799 0.02368 -0.0579 1.0000 0.0084
-12.000 -1.0088 0.02668 0.02228 -0.0561 1.0000 0.0085
-11.750 -1.0042 0.02580 0.02132 -0.0540 1.0000 0.0086
-11.500 -1.0158 0.02430 0.01970 -0.0496 1.0000 0.0087
-11.250 -1.0318 0.02297 0.01827 -0.0438 1.0000 0.0088
-11.000 -1.0358 0.02170 0.01689 -0.0396 1.0000 0.0091
-10.750 -1.0254 0.02067 0.01581 -0.0379 0.9996 0.0094
-10.500 -0.9961 0.01985 0.01494 -0.0396 0.9981 0.0098
-10.250 -0.9658 0.01923 0.01428 -0.0413 0.9965 0.0103
-10.000 -0.9354 0.01848 0.01345 -0.0430 0.9951 0.0108
-9.750 -0.9054 0.01775 0.01265 -0.0445 0.9936 0.0113
-9.500 -0.8773 0.01709 0.01193 -0.0454 0.9913 0.0116
-9.250 -0.8503 0.01597 0.01070 -0.0465 0.9887 0.0122
-9.000 -0.8212 0.01497 0.00964 -0.0480 0.9867 0.0130
-8.750 -0.7892 0.01442 0.00906 -0.0497 0.9853 0.0138
-8.500 -0.7564 0.01393 0.00854 -0.0514 0.9842 0.0148
-8.250 -0.7229 0.01349 0.00805 -0.0533 0.9833 0.0155
-8.000 -0.6985 0.01277 0.00728 -0.0533 0.9791 0.0168
-7.750 -0.6681 0.01224 0.00672 -0.0545 0.9767 0.0184
-7.500 -0.6357 0.01185 0.00632 -0.0559 0.9750 0.0200
-7.250 -0.6024 0.01147 0.00590 -0.0576 0.9737 0.0213
-7.000 -0.5690 0.01087 0.00529 -0.0594 0.9725 0.0248
-6.750 -0.5347 0.01057 0.00497 -0.0611 0.9714 0.0274
-6.500 -0.5040 0.01008 0.00450 -0.0622 0.9689 0.0330
-6.250 -0.4766 0.00977 0.00418 -0.0624 0.9644 0.0380
-6.000 -0.4459 0.00941 0.00385 -0.0634 0.9611 0.0468
-5.750 -0.4149 0.00900 0.00351 -0.0645 0.9582 0.0636
-5.500 -0.3852 0.00864 0.00323 -0.0653 0.9544 0.0846
-5.250 -0.3596 0.00821 0.00294 -0.0652 0.9485 0.1175
-5.000 -0.3317 0.00765 0.00260 -0.0658 0.9438 0.1742
-4.750 -0.3068 0.00691 0.00221 -0.0659 0.9379 0.2664
-4.500 -0.2837 0.00590 0.00179 -0.0659 0.9313 0.4204
-4.250 -0.2547 0.00570 0.00167 -0.0663 0.9266 0.4636
-4.000 -0.2275 0.00561 0.00161 -0.0662 0.9201 0.4861
-3.750 -0.1986 0.00555 0.00154 -0.0665 0.9146 0.5009
-3.500 -0.1702 0.00551 0.00149 -0.0666 0.9088 0.5121
-3.250 -0.1421 0.00547 0.00144 -0.0667 0.9025 0.5224
-2.750 -0.0852 0.00544 0.00137 -0.0669 0.8902 0.5394
-2.500 -0.0564 0.00545 0.00133 -0.0671 0.8843 0.5474
-2.250 -0.0283 0.00547 0.00135 -0.0672 0.8778 0.5581
-2.000 0.0000 0.00546 0.00133 -0.0672 0.8714 0.5652
-1.750 0.0284 0.00545 0.00130 -0.0674 0.8653 0.5684
-1.500 0.0567 0.00544 0.00128 -0.0675 0.8584 0.5714
-1.250 0.0853 0.00546 0.00125 -0.0676 0.8523 0.5744
-1.000 0.1134 0.00543 0.00122 -0.0677 0.8453 0.5775
-0.750 0.1418 0.00543 0.00121 -0.0678 0.8387 0.5806
-0.500 0.1697 0.00542 0.00120 -0.0678 0.8306 0.5837
-0.250 0.1979 0.00544 0.00119 -0.0679 0.8224 0.5869
0.000 0.2258 0.00546 0.00119 -0.0679 0.8128 0.5900
0.250 0.2535 0.00545 0.00119 -0.0678 0.8027 0.5936
0.500 0.2810 0.00546 0.00119 -0.0677 0.7917 0.5970
1.000 0.3356 0.00555 0.00121 -0.0674 0.7637 0.6039
1.250 0.3618 0.00560 0.00122 -0.0669 0.7398 0.6075
1.500 0.3882 0.00566 0.00125 -0.0666 0.7199 0.6114
1.750 0.4150 0.00574 0.00130 -0.0664 0.6997 0.6154
2.000 0.4412 0.00586 0.00135 -0.0660 0.6763 0.6192
2.250 0.4667 0.00598 0.00141 -0.0655 0.6437 0.6232
2.500 0.4908 0.00621 0.00150 -0.0647 0.5948 0.6273
2.750 0.5102 0.00680 0.00169 -0.0632 0.4904 0.6315
3.000 0.5276 0.00766 0.00199 -0.0614 0.3569 0.6357
3.250 0.5469 0.00840 0.00230 -0.0601 0.2466 0.6403
3.500 0.5676 0.00907 0.00261 -0.0590 0.1558 0.6452
3.750 0.5903 0.00959 0.00288 -0.0582 0.0986 0.6498
4.000 0.6141 0.00997 0.00314 -0.0576 0.0663 0.6546
4.250 0.6391 0.01027 0.00339 -0.0571 0.0500 0.6599
4.500 0.6641 0.01059 0.00366 -0.0566 0.0403 0.6652
4.750 0.6898 0.01079 0.00390 -0.0563 0.0359 0.6709
5.000 0.7144 0.01113 0.00425 -0.0557 0.0307 0.6769
5.250 0.7402 0.01133 0.00449 -0.0554 0.0282 0.6828
5.500 0.7643 0.01169 0.00486 -0.0548 0.0249 0.6890
5.750 0.7888 0.01204 0.00525 -0.0542 0.0230 0.6953
6.000 0.8137 0.01229 0.00557 -0.0537 0.0215 0.7018
6.250 0.8383 0.01259 0.00588 -0.0532 0.0198 0.7091
6.500 0.8595 0.01322 0.00659 -0.0521 0.0179 0.7162
6.750 0.8844 0.01348 0.00690 -0.0516 0.0172 0.7240
7.000 0.9082 0.01381 0.00730 -0.0509 0.0163 0.7317
7.500 0.9547 0.01459 0.00818 -0.0495 0.0146 0.7483
7.750 0.9712 0.01571 0.00941 -0.0476 0.0135 0.7573
8.000 0.9951 0.01599 0.00977 -0.0470 0.0131 0.7673
8.250 1.0179 0.01637 0.01023 -0.0462 0.0126 0.7776
8.500 1.0400 0.01681 0.01075 -0.0454 0.0121 0.7884
8.750 1.0617 0.01726 0.01129 -0.0445 0.0115 0.8004
9.000 1.0829 0.01774 0.01184 -0.0435 0.0111 0.8137
9.250 1.1024 0.01833 0.01252 -0.0422 0.0107 0.8290
9.500 1.1156 0.01973 0.01410 -0.0400 0.0101 0.8471
9.750 1.1313 0.02076 0.01532 -0.0381 0.0099 0.8701
10.000 1.1467 0.02129 0.01606 -0.0360 0.0097 0.9076
10.250 1.1650 0.02191 0.01687 -0.0346 0.0095 1.0000
10.500 1.1817 0.02280 0.01787 -0.0331 0.0092 1.0000
10.750 1.1964 0.02372 0.01889 -0.0312 0.0090 1.0000
11.000 1.2090 0.02482 0.02010 -0.0291 0.0088 1.0000
11.250 1.2206 0.02589 0.02129 -0.0269 0.0086 1.0000
11.500 1.2306 0.02705 0.02257 -0.0246 0.0085 1.0000
11.750 1.2427 0.02774 0.02332 -0.0227 0.0082 1.0000
12.000 1.2495 0.02912 0.02483 -0.0202 0.0081 1.0000
12.250 1.2593 0.02995 0.02571 -0.0182 0.0079 1.0000
12.500 1.2648 0.03132 0.02720 -0.0159 0.0078 1.0000
12.750 1.2684 0.03281 0.02878 -0.0137 0.0076 1.0000
13.000 1.2644 0.03513 0.03128 -0.0111 0.0075 1.0000
13.250 1.2473 0.03885 0.03527 -0.0080 0.0073 1.0000
13.500 1.2193 0.04389 0.04063 -0.0054 0.0072 1.0000
13.750 1.2049 0.04768 0.04464 -0.0043 0.0071 1.0000
14.000 1.1956 0.05117 0.04828 -0.0039 0.0071 1.0000
14.250 1.1885 0.05465 0.05191 -0.0042 0.0071 1.0000
14.500 1.1750 0.05929 0.05671 -0.0053 0.0071 1.0000
14.750 1.1558 0.06529 0.06289 -0.0076 0.0070 1.0000
15.000 1.1184 0.07537 0.07320 -0.0130 0.0071 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE 2822 AIRFOIL (rae2822-il)