NPL 9660 AIRFOIL (npl9660-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NPL 9660 AIRFOIL (npl9660-il) Reynolds number: 200,000 Max Cl/Cd: 54.91 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-npl9660-il-200000.txt Download as CSV file: xf-npl9660-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NPL 9660 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6027 0.08888 0.08533 -0.0121 1.0000 0.0574
-9.250 -0.6165 0.08068 0.07713 -0.0206 1.0000 0.0580
-9.000 -0.6355 0.07444 0.07081 -0.0255 1.0000 0.0585
-8.750 -0.6520 0.06969 0.06593 -0.0272 1.0000 0.0593
-8.500 -0.6874 0.06627 0.06181 -0.0275 1.0000 0.0620
-8.000 -0.6762 0.05622 0.05162 -0.0272 1.0000 0.0635
-7.750 -0.6634 0.05315 0.04856 -0.0267 1.0000 0.0643
-7.500 -0.6527 0.05047 0.04581 -0.0257 1.0000 0.0654
-7.250 -0.6439 0.04798 0.04320 -0.0241 1.0000 0.0668
-7.000 -0.6655 0.03490 0.02827 -0.0156 1.0000 0.0392
-6.750 -0.6518 0.03224 0.02547 -0.0138 1.0000 0.0387
-6.500 -0.6376 0.02979 0.02277 -0.0117 1.0000 0.0384
-6.250 -0.6218 0.02745 0.02013 -0.0097 1.0000 0.0384
-6.000 -0.6042 0.02535 0.01765 -0.0077 1.0000 0.0389
-5.750 -0.5851 0.02344 0.01550 -0.0062 1.0000 0.0398
-5.500 -0.5650 0.02240 0.01449 -0.0051 1.0000 0.0412
-5.250 -0.5444 0.02135 0.01332 -0.0038 1.0000 0.0436
-5.000 -0.5232 0.02008 0.01189 -0.0025 1.0000 0.0469
-4.750 -0.5023 0.01946 0.01128 -0.0015 1.0000 0.0504
-4.500 -0.4809 0.01869 0.01041 -0.0004 1.0000 0.0544
-4.250 -0.4600 0.01827 0.01005 0.0004 1.0000 0.0584
-4.000 -0.4384 0.01801 0.00967 0.0014 1.0000 0.0619
-3.750 -0.4130 0.01705 0.00880 0.0014 0.9992 0.0650
-3.500 -0.3710 0.01646 0.00824 -0.0018 0.9946 0.0687
-3.250 -0.3295 0.01602 0.00777 -0.0048 0.9895 0.0722
-3.000 -0.2882 0.01527 0.00709 -0.0079 0.9848 0.0763
-2.750 -0.2458 0.01486 0.00674 -0.0112 0.9799 0.0823
-2.500 -0.2038 0.01437 0.00628 -0.0143 0.9743 0.0887
-2.250 -0.1586 0.01394 0.00587 -0.0181 0.9701 0.0980
-2.000 -0.1173 0.01342 0.00546 -0.0210 0.9632 0.1200
-1.750 -0.0946 0.01073 0.00523 -0.0211 0.9549 0.6628
-1.500 -0.0685 0.01056 0.00553 -0.0194 0.9446 0.7782
-1.250 -0.0427 0.01071 0.00580 -0.0175 0.9342 0.8337
-1.000 -0.0206 0.01092 0.00603 -0.0148 0.9222 0.8690
-0.750 -0.0002 0.01114 0.00625 -0.0117 0.9095 0.8930
-0.500 0.0210 0.01132 0.00641 -0.0086 0.8977 0.9146
-0.250 0.0457 0.01140 0.00645 -0.0061 0.8869 0.9363
0.000 0.0850 0.01140 0.00640 -0.0069 0.8732 0.9554
0.250 0.1302 0.01122 0.00617 -0.0097 0.8598 0.9655
0.500 0.1689 0.01096 0.00585 -0.0115 0.8437 0.9730
0.750 0.2088 0.01072 0.00553 -0.0135 0.8247 0.9794
1.000 0.2482 0.01051 0.00522 -0.0156 0.8022 0.9856
1.250 0.2872 0.01036 0.00494 -0.0175 0.7748 0.9921
1.500 0.3274 0.01027 0.00469 -0.0199 0.7447 0.9978
1.750 0.3570 0.01026 0.00454 -0.0203 0.7173 1.0000
2.000 0.3812 0.01028 0.00443 -0.0197 0.6916 1.0000
2.250 0.4056 0.01033 0.00435 -0.0191 0.6672 1.0000
2.500 0.4305 0.01039 0.00433 -0.0187 0.6429 1.0000
2.750 0.4554 0.01046 0.00432 -0.0183 0.6195 1.0000
3.000 0.4802 0.01056 0.00432 -0.0178 0.5968 1.0000
3.250 0.5046 0.01069 0.00435 -0.0173 0.5746 1.0000
3.500 0.5289 0.01082 0.00442 -0.0167 0.5518 1.0000
3.750 0.5530 0.01097 0.00451 -0.0161 0.5281 1.0000
4.000 0.5768 0.01114 0.00461 -0.0155 0.5037 1.0000
4.250 0.6002 0.01133 0.00473 -0.0147 0.4781 1.0000
4.500 0.6232 0.01155 0.00487 -0.0140 0.4482 1.0000
4.750 0.6456 0.01181 0.00503 -0.0131 0.4142 1.0000
5.000 0.6671 0.01215 0.00522 -0.0121 0.3747 1.0000
5.250 0.6877 0.01260 0.00548 -0.0110 0.3328 1.0000
5.500 0.7079 0.01310 0.00580 -0.0099 0.2933 1.0000
5.750 0.7276 0.01366 0.00618 -0.0086 0.2621 1.0000
6.000 0.7473 0.01420 0.00659 -0.0074 0.2379 1.0000
6.250 0.7671 0.01470 0.00700 -0.0063 0.2166 1.0000
6.500 0.7872 0.01519 0.00739 -0.0052 0.1984 1.0000
6.750 0.8084 0.01564 0.00782 -0.0042 0.1836 1.0000
7.000 0.8303 0.01609 0.00828 -0.0034 0.1709 1.0000
7.250 0.8524 0.01658 0.00876 -0.0026 0.1582 1.0000
7.500 0.8746 0.01707 0.00925 -0.0020 0.1450 1.0000
7.750 0.8970 0.01757 0.00974 -0.0014 0.1305 1.0000
8.000 0.9191 0.01814 0.01028 -0.0008 0.1136 1.0000
8.250 0.9393 0.01895 0.01098 0.0000 0.0931 1.0000
8.500 0.9595 0.01983 0.01176 0.0008 0.0741 1.0000
8.750 0.9773 0.02095 0.01281 0.0019 0.0639 1.0000
9.000 0.9956 0.02199 0.01382 0.0029 0.0575 1.0000
9.250 1.0140 0.02305 0.01494 0.0039 0.0532 1.0000
9.500 1.0313 0.02425 0.01612 0.0050 0.0502 1.0000
9.750 1.0497 0.02534 0.01729 0.0060 0.0475 1.0000
10.000 1.0668 0.02674 0.01860 0.0070 0.0454 1.0000
10.250 1.0852 0.02788 0.01993 0.0080 0.0433 1.0000
10.500 1.1030 0.02909 0.02116 0.0089 0.0414 1.0000
10.750 1.1215 0.03073 0.02276 0.0096 0.0398 1.0000
11.000 1.1383 0.03219 0.02444 0.0107 0.0387 1.0000
11.250 1.1549 0.03379 0.02621 0.0117 0.0376 1.0000
11.500 1.1706 0.03546 0.02801 0.0126 0.0367 1.0000
11.750 1.1856 0.03722 0.02988 0.0136 0.0360 1.0000
12.000 1.2016 0.03927 0.03201 0.0143 0.0354 1.0000
12.250 1.2168 0.04214 0.03500 0.0149 0.0349 1.0000
12.500 1.2170 0.04435 0.03752 0.0172 0.0347 1.0000
12.750 1.2142 0.04685 0.04032 0.0193 0.0346 1.0000
13.000 1.2084 0.04967 0.04342 0.0212 0.0345 1.0000
13.250 1.1995 0.05283 0.04686 0.0226 0.0344 1.0000
13.500 1.1883 0.05639 0.05068 0.0234 0.0344 1.0000
13.750 1.1747 0.06040 0.05494 0.0235 0.0345 1.0000
14.000 1.1592 0.06494 0.05972 0.0227 0.0345 1.0000
14.250 1.1418 0.07006 0.06505 0.0211 0.0346 1.0000
14.500 1.1240 0.07567 0.07086 0.0188 0.0347 1.0000
14.750 1.1061 0.08175 0.07711 0.0159 0.0349 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NPL 9660 AIRFOIL (npl9660-il)