P-51D TIP (BL215) AIRFOIL (p51dtip-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: P-51D TIP (BL215) AIRFOIL (p51dtip-il) Reynolds number: 200,000 Max Cl/Cd: 54.8 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-p51dtip-il-200000.txt Download as CSV file: xf-p51dtip-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: P-51D TIP (BL215) 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
------ -------- --------- --------- -------- -------- --------
-10.500 -0.3942 0.09390 0.09080 -0.0383 0.9087 0.0468
-10.250 -0.3960 0.08926 0.08616 -0.0403 0.9050 0.0481
-10.000 -0.4933 0.08632 0.08321 -0.0488 0.9364 0.0438
-9.750 -0.4911 0.08153 0.07838 -0.0526 0.9252 0.0446
-9.500 -0.4969 0.07649 0.07328 -0.0565 0.9155 0.0453
-9.250 -0.5097 0.07168 0.06840 -0.0598 0.9068 0.0454
-9.000 -0.5242 0.06802 0.06464 -0.0607 0.8992 0.0457
-8.750 -0.5366 0.06494 0.06147 -0.0603 0.8925 0.0464
-8.500 -0.5432 0.06172 0.05810 -0.0599 0.8872 0.0474
-8.250 -0.5470 0.05843 0.05463 -0.0594 0.8824 0.0494
-8.000 -0.5546 0.05770 0.05327 -0.0574 0.8766 0.0538
-7.750 -0.5608 0.05243 0.04762 -0.0562 0.8729 0.0550
-7.500 -0.5457 0.04778 0.04310 -0.0566 0.8705 0.0568
-7.250 -0.5304 0.04516 0.04044 -0.0565 0.8674 0.0599
-7.000 -0.5251 0.04380 0.03830 -0.0544 0.8637 0.0684
-6.750 -0.5053 0.03985 0.03458 -0.0550 0.8613 0.0725
-6.500 -0.4887 0.03198 0.02528 -0.0505 0.8586 0.0426
-6.250 -0.4648 0.02947 0.02245 -0.0495 0.8563 0.0408
-6.000 -0.4409 0.02777 0.02035 -0.0489 0.8535 0.0415
-5.750 -0.4184 0.02419 0.01659 -0.0489 0.8510 0.0458
-5.500 -0.3921 0.02271 0.01493 -0.0487 0.8488 0.0484
-5.250 -0.3656 0.02168 0.01369 -0.0483 0.8468 0.0520
-5.000 -0.3393 0.02079 0.01257 -0.0479 0.8450 0.0547
-4.750 -0.3144 0.01926 0.01106 -0.0475 0.8434 0.0576
-4.500 -0.2886 0.01864 0.01045 -0.0476 0.8414 0.0625
-4.250 -0.2632 0.01801 0.00979 -0.0474 0.8393 0.0656
-4.000 -0.2394 0.01737 0.00912 -0.0470 0.8370 0.0685
-3.750 -0.2171 0.01673 0.00850 -0.0465 0.8348 0.0739
-3.500 -0.1934 0.01632 0.00807 -0.0461 0.8329 0.0832
-3.250 -0.1710 0.01573 0.00757 -0.0455 0.8313 0.1083
-3.000 -0.1737 0.01346 0.00777 -0.0408 0.8295 0.6745
-2.750 -0.1503 0.01422 0.00851 -0.0392 0.8278 0.7305
-2.500 -0.1265 0.01488 0.00916 -0.0377 0.8261 0.7558
-2.250 -0.1048 0.01563 0.00996 -0.0354 0.8243 0.7796
-2.000 -0.0866 0.01645 0.01082 -0.0317 0.8225 0.8058
-1.750 -0.0693 0.01708 0.01146 -0.0281 0.8206 0.8296
-1.500 -0.0499 0.01747 0.01184 -0.0252 0.8193 0.8453
-1.250 -0.0272 0.01766 0.01198 -0.0240 0.8181 0.8549
-1.000 -0.0023 0.01771 0.01197 -0.0238 0.8170 0.8588
-0.750 0.0236 0.01778 0.01198 -0.0239 0.8161 0.8612
-0.500 0.0490 0.01790 0.01206 -0.0241 0.8151 0.8639
-0.250 0.0722 0.01811 0.01226 -0.0246 0.8130 0.8665
0.000 0.0958 0.01831 0.01244 -0.0251 0.8109 0.8686
0.250 0.1203 0.01853 0.01265 -0.0258 0.8095 0.8700
0.500 0.1444 0.01875 0.01287 -0.0262 0.8082 0.8711
0.750 0.1682 0.01900 0.01312 -0.0265 0.8070 0.8723
1.000 0.1925 0.01923 0.01336 -0.0267 0.8057 0.8737
1.250 0.2174 0.01945 0.01360 -0.0269 0.8045 0.8754
1.500 0.2418 0.01975 0.01393 -0.0272 0.8036 0.8772
1.750 0.2689 0.01997 0.01418 -0.0277 0.8025 0.8787
2.000 0.2859 0.02048 0.01473 -0.0272 0.7991 0.8810
2.250 0.3029 0.02091 0.01520 -0.0266 0.7945 0.8834
2.500 0.3287 0.02112 0.01546 -0.0268 0.7919 0.8851
2.750 0.3599 0.02114 0.01557 -0.0274 0.7897 0.8866
3.000 0.3795 0.02131 0.01582 -0.0263 0.7822 0.8892
3.250 0.4310 0.01985 0.01440 -0.0274 0.7723 0.8893
3.500 0.4668 0.01838 0.01299 -0.0262 0.7563 0.8905
3.750 0.5005 0.01696 0.01160 -0.0249 0.7419 0.8920
4.000 0.5320 0.01566 0.01034 -0.0236 0.7274 0.8938
4.250 0.5619 0.01448 0.00923 -0.0225 0.7109 0.8961
4.500 0.5911 0.01355 0.00838 -0.0217 0.6946 0.8984
4.750 0.6176 0.01262 0.00757 -0.0204 0.6694 0.9009
5.000 0.6368 0.01162 0.00605 -0.0171 0.4879 0.9044
5.250 0.6230 0.01417 0.00703 -0.0114 0.2169 0.9116
5.500 0.6227 0.01598 0.00803 -0.0077 0.0910 0.9177
5.750 0.6384 0.01677 0.00882 -0.0061 0.0727 0.9238
6.000 0.6544 0.01759 0.00964 -0.0046 0.0624 0.9303
6.250 0.6713 0.01849 0.01052 -0.0034 0.0532 0.9384
6.500 0.6934 0.01951 0.01164 -0.0029 0.0468 0.9472
6.750 0.7198 0.02128 0.01336 -0.0033 0.0397 0.9562
7.000 0.7550 0.02267 0.01490 -0.0047 0.0354 0.9661
7.250 0.7926 0.02459 0.01696 -0.0066 0.0325 0.9787
7.500 0.8287 0.02852 0.02105 -0.0085 0.0290 1.0000
7.750 0.8533 0.03006 0.02290 -0.0082 0.0279 1.0000
8.000 0.8775 0.03308 0.02630 -0.0080 0.0278 1.0000
8.250 0.8966 0.03732 0.03094 -0.0073 0.0284 1.0000
11.750 0.6990 0.10114 0.09814 0.0067 0.0522 1.0000
12.000 0.6780 0.10781 0.10487 0.0029 0.0522 1.0000
12.250 0.6586 0.11509 0.11220 -0.0021 0.0522 1.0000
|
Polar data table (+)
Polar graphs
<< Back to P-51D TIP (BL215) AIRFOIL (p51dtip-il)