RAF 30 AIRFOIL (raf30-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAF 30 AIRFOIL (raf30-il) Reynolds number: 500,000 Max Cl/Cd: 58.48 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf30-il-500000.txt Download as CSV file: xf-raf30-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: RAF 30 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-17.500 -0.9595 0.10983 0.10673 -0.0170 1.0000 0.0189
-17.250 -0.9836 0.10100 0.09777 -0.0223 1.0000 0.0191
-17.000 -1.0172 0.09058 0.08712 -0.0288 1.0000 0.0190
-16.750 -1.0375 0.08325 0.07962 -0.0331 1.0000 0.0190
-16.500 -1.0562 0.07664 0.07284 -0.0368 1.0000 0.0190
-16.250 -1.0689 0.07135 0.06744 -0.0394 1.0000 0.0192
-16.000 -1.0752 0.06727 0.06328 -0.0414 1.0000 0.0194
-15.750 -1.0859 0.06266 0.05853 -0.0434 1.0000 0.0195
-15.500 -1.0906 0.05913 0.05492 -0.0449 1.0000 0.0197
-15.250 -1.0996 0.05511 0.05075 -0.0462 1.0000 0.0198
-15.000 -1.1016 0.05223 0.04779 -0.0472 1.0000 0.0202
-14.750 -1.1064 0.04905 0.04447 -0.0478 1.0000 0.0203
-14.500 -1.1062 0.04665 0.04198 -0.0481 1.0000 0.0207
-14.250 -1.1069 0.04428 0.03950 -0.0481 1.0000 0.0211
-14.000 -1.1084 0.04193 0.03700 -0.0478 1.0000 0.0216
-13.750 -1.1096 0.03982 0.03473 -0.0472 1.0000 0.0220
-13.500 -1.1098 0.03766 0.03240 -0.0461 1.0000 0.0223
-13.250 -1.1090 0.03617 0.03074 -0.0448 1.0000 0.0226
-13.000 -1.1077 0.03382 0.02822 -0.0432 1.0000 0.0231
-12.750 -1.0993 0.03207 0.02643 -0.0419 1.0000 0.0236
-12.500 -1.0910 0.03084 0.02517 -0.0404 1.0000 0.0241
-12.250 -1.0826 0.02970 0.02397 -0.0387 1.0000 0.0245
-12.000 -1.0745 0.02878 0.02298 -0.0367 1.0000 0.0252
-11.750 -1.0651 0.02773 0.02184 -0.0347 1.0000 0.0258
-11.500 -1.0563 0.02701 0.02103 -0.0323 1.0000 0.0265
-11.250 -1.0481 0.02637 0.02027 -0.0296 1.0000 0.0272
-11.000 -1.0399 0.02581 0.01960 -0.0266 1.0000 0.0276
-10.750 -1.0272 0.02386 0.01760 -0.0249 1.0000 0.0285
-10.500 -1.0168 0.02292 0.01666 -0.0224 1.0000 0.0292
-10.250 -1.0040 0.02218 0.01590 -0.0202 1.0000 0.0299
-10.000 -0.9903 0.02148 0.01515 -0.0181 1.0000 0.0307
-9.750 -0.9757 0.02085 0.01447 -0.0160 1.0000 0.0317
-9.500 -0.9602 0.02029 0.01385 -0.0141 1.0000 0.0325
-9.250 -0.9439 0.01979 0.01328 -0.0121 1.0000 0.0332
-9.000 -0.9324 0.01877 0.01220 -0.0095 1.0000 0.0340
-8.750 -0.9215 0.01784 0.01126 -0.0068 1.0000 0.0350
-8.500 -0.9073 0.01722 0.01063 -0.0045 1.0000 0.0361
-8.250 -0.8918 0.01670 0.01009 -0.0023 1.0000 0.0372
-8.000 -0.8764 0.01618 0.00954 -0.0001 1.0000 0.0382
-7.750 -0.8606 0.01572 0.00902 0.0020 1.0000 0.0391
-7.500 -0.8444 0.01531 0.00857 0.0042 1.0000 0.0399
-7.250 -0.8321 0.01468 0.00791 0.0069 1.0000 0.0413
-7.000 -0.8192 0.01415 0.00738 0.0096 1.0000 0.0429
-6.750 -0.8053 0.01377 0.00698 0.0121 1.0000 0.0445
-6.500 -0.7758 0.01336 0.00654 0.0114 0.9980 0.0465
-6.250 -0.7389 0.01290 0.00606 0.0092 0.9946 0.0498
-6.000 -0.7041 0.01239 0.00558 0.0074 0.9905 0.0562
-5.750 -0.6702 0.01168 0.00508 0.0057 0.9858 0.0855
-5.500 -0.6354 0.01100 0.00466 0.0038 0.9820 0.1320
-5.250 -0.6044 0.01040 0.00427 0.0027 0.9752 0.1731
-5.000 -0.5696 0.00978 0.00393 0.0008 0.9710 0.2286
-4.750 -0.5387 0.00932 0.00365 -0.0001 0.9634 0.2737
-4.250 -0.4696 0.00852 0.00315 -0.0032 0.9504 0.3513
-4.000 -0.4343 0.00814 0.00293 -0.0049 0.9428 0.3932
-3.750 -0.4019 0.00779 0.00274 -0.0060 0.9321 0.4354
-3.500 -0.3683 0.00754 0.00258 -0.0072 0.9211 0.4708
-3.250 -0.3367 0.00734 0.00244 -0.0079 0.9083 0.5028
-3.000 -0.3083 0.00721 0.00234 -0.0079 0.8937 0.5297
-2.750 -0.2819 0.00712 0.00228 -0.0074 0.8786 0.5555
-2.500 -0.2563 0.00705 0.00222 -0.0067 0.8636 0.5781
-2.250 -0.2306 0.00701 0.00216 -0.0060 0.8475 0.5953
-2.000 -0.2053 0.00698 0.00210 -0.0053 0.8308 0.6114
-1.750 -0.1804 0.00696 0.00206 -0.0044 0.8154 0.6309
-1.500 -0.1553 0.00693 0.00204 -0.0036 0.8017 0.6498
-1.250 -0.1298 0.00691 0.00201 -0.0029 0.7888 0.6653
-1.000 -0.1040 0.00689 0.00199 -0.0023 0.7763 0.6780
-0.750 -0.0780 0.00688 0.00196 -0.0017 0.7636 0.6898
-0.500 -0.0521 0.00687 0.00195 -0.0011 0.7509 0.7021
-0.250 -0.0260 0.00687 0.00194 -0.0006 0.7387 0.7146
0.000 0.0000 0.00686 0.00194 0.0000 0.7266 0.7265
0.250 0.0260 0.00687 0.00194 0.0006 0.7146 0.7386
0.500 0.0520 0.00687 0.00195 0.0011 0.7021 0.7509
0.750 0.0780 0.00688 0.00196 0.0017 0.6899 0.7637
1.000 0.1040 0.00689 0.00199 0.0023 0.6780 0.7763
1.250 0.1298 0.00691 0.00201 0.0029 0.6653 0.7889
1.500 0.1553 0.00693 0.00204 0.0036 0.6496 0.8017
1.750 0.1804 0.00696 0.00206 0.0044 0.6311 0.8153
2.000 0.2053 0.00698 0.00210 0.0053 0.6111 0.8309
2.250 0.2306 0.00701 0.00216 0.0060 0.5953 0.8476
2.500 0.2563 0.00705 0.00222 0.0067 0.5784 0.8636
2.750 0.2818 0.00712 0.00227 0.0074 0.5552 0.8786
3.000 0.3083 0.00721 0.00234 0.0079 0.5300 0.8937
3.250 0.3366 0.00735 0.00244 0.0079 0.5021 0.9083
3.500 0.3683 0.00754 0.00257 0.0072 0.4704 0.9211
3.750 0.4020 0.00779 0.00274 0.0060 0.4359 0.9321
4.000 0.4344 0.00814 0.00293 0.0049 0.3930 0.9428
4.250 0.4696 0.00852 0.00315 0.0032 0.3511 0.9506
4.500 0.5032 0.00891 0.00339 0.0019 0.3146 0.9582
4.750 0.5387 0.00932 0.00365 0.0001 0.2730 0.9634
5.000 0.5696 0.00978 0.00393 -0.0008 0.2285 0.9710
5.250 0.6044 0.01041 0.00427 -0.0027 0.1720 0.9752
5.500 0.6354 0.01101 0.00466 -0.0038 0.1314 0.9821
5.750 0.6702 0.01170 0.00509 -0.0057 0.0835 0.9859
6.000 0.7042 0.01239 0.00558 -0.0074 0.0561 0.9905
6.250 0.7391 0.01290 0.00605 -0.0092 0.0499 0.9946
6.500 0.7759 0.01336 0.00654 -0.0114 0.0465 0.9981
6.750 0.8053 0.01377 0.00697 -0.0121 0.0445 1.0000
7.000 0.8191 0.01415 0.00738 -0.0096 0.0430 1.0000
7.250 0.8320 0.01468 0.00791 -0.0069 0.0413 1.0000
7.500 0.8444 0.01531 0.00857 -0.0041 0.0399 1.0000
7.750 0.8606 0.01572 0.00902 -0.0020 0.0392 1.0000
8.000 0.8763 0.01619 0.00955 0.0002 0.0382 1.0000
8.250 0.8921 0.01667 0.01006 0.0023 0.0371 1.0000
8.500 0.9073 0.01721 0.01062 0.0045 0.0361 1.0000
8.750 0.9215 0.01784 0.01126 0.0068 0.0350 1.0000
9.000 0.9321 0.01880 0.01223 0.0096 0.0340 1.0000
9.250 0.9441 0.01976 0.01326 0.0121 0.0332 1.0000
9.500 0.9601 0.02031 0.01387 0.0141 0.0326 1.0000
9.750 0.9757 0.02086 0.01448 0.0160 0.0317 1.0000
10.000 0.9904 0.02150 0.01517 0.0181 0.0308 1.0000
10.250 1.0044 0.02216 0.01587 0.0202 0.0299 1.0000
10.500 1.0171 0.02291 0.01665 0.0224 0.0292 1.0000
10.750 1.0274 0.02388 0.01762 0.0249 0.0285 1.0000
11.000 1.0401 0.02579 0.01958 0.0266 0.0276 1.0000
11.250 1.0479 0.02629 0.02020 0.0296 0.0271 1.0000
11.500 1.0565 0.02696 0.02098 0.0323 0.0265 1.0000
11.750 1.0658 0.02781 0.02193 0.0346 0.0258 1.0000
12.000 1.0751 0.02880 0.02300 0.0366 0.0253 1.0000
12.250 1.0834 0.02975 0.02402 0.0386 0.0247 1.0000
12.500 1.0916 0.03088 0.02522 0.0403 0.0241 1.0000
12.750 1.0999 0.03206 0.02642 0.0418 0.0236 1.0000
13.000 1.1084 0.03381 0.02820 0.0431 0.0231 1.0000
13.250 1.1095 0.03606 0.03063 0.0447 0.0226 1.0000
13.500 1.1105 0.03771 0.03245 0.0460 0.0223 1.0000
13.750 1.1104 0.03970 0.03461 0.0470 0.0219 1.0000
14.000 1.1108 0.04157 0.03661 0.0476 0.0213 1.0000
14.250 1.1087 0.04415 0.03936 0.0480 0.0211 1.0000
14.500 1.1102 0.04622 0.04151 0.0478 0.0205 1.0000
14.750 1.1084 0.04893 0.04434 0.0476 0.0203 1.0000
15.000 1.1036 0.05212 0.04766 0.0470 0.0201 1.0000
15.250 1.1033 0.05477 0.05038 0.0460 0.0197 1.0000
15.500 1.0931 0.05897 0.05475 0.0447 0.0197 1.0000
15.750 1.0872 0.06269 0.05857 0.0432 0.0195 1.0000
16.000 1.0834 0.06625 0.06218 0.0416 0.0192 1.0000
16.250 1.0685 0.07168 0.06779 0.0390 0.0192 1.0000
16.500 1.0593 0.07640 0.07260 0.0367 0.0190 1.0000
16.750 1.0415 0.08277 0.07911 0.0333 0.0188 1.0000
17.000 1.0212 0.09010 0.08661 0.0289 0.0189 1.0000
17.250 0.9964 0.09865 0.09534 0.0237 0.0189 1.0000
17.500 0.9577 0.11072 0.10765 0.0160 0.0190 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 30 AIRFOIL (raf30-il)