NPL 9660 AIRFOIL (npl9660-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NPL 9660 AIRFOIL (npl9660-il) Reynolds number: 100,000 Max Cl/Cd: 41.86 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-npl9660-il-100000.txt Download as CSV file: xf-npl9660-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5732 0.10974 0.10466 -0.0002 1.0000 0.1114
-9.750 -0.6090 0.10300 0.09802 -0.0126 1.0000 0.1157
-9.500 -0.6516 0.09701 0.09199 -0.0215 1.0000 0.1162
-9.250 -0.5841 0.09539 0.09044 -0.0088 1.0000 0.1210
-9.000 -0.5784 0.09192 0.08698 -0.0095 1.0000 0.1252
-8.750 -0.5971 0.08540 0.08052 -0.0171 1.0000 0.1299
-8.500 -0.6598 0.08069 0.07551 -0.0246 1.0000 0.1331
-8.250 -0.6337 0.07479 0.06984 -0.0234 1.0000 0.1369
-8.000 -0.5485 0.06397 0.05937 -0.0270 1.0000 0.1546
-7.750 -0.5352 0.06114 0.05658 -0.0256 1.0000 0.1616
-7.500 -0.6252 0.06433 0.05928 -0.0233 1.0000 0.1574
-7.250 -0.6370 0.06119 0.05587 -0.0232 1.0000 0.1712
-7.000 -0.6169 0.05830 0.05315 -0.0218 1.0000 0.1784
-6.750 -0.6141 0.05552 0.05033 -0.0203 1.0000 0.1938
-6.500 -0.6095 0.05315 0.04795 -0.0183 1.0000 0.2114
-6.250 -0.6037 0.04017 0.03263 -0.0164 1.0000 0.0825
-6.000 -0.5906 0.03689 0.02902 -0.0143 1.0000 0.0793
-5.750 -0.5751 0.03415 0.02589 -0.0121 1.0000 0.0781
-5.500 -0.5577 0.03178 0.02317 -0.0102 1.0000 0.0784
-5.250 -0.5386 0.02958 0.02062 -0.0084 1.0000 0.0795
-5.000 -0.5180 0.02770 0.01840 -0.0068 1.0000 0.0821
-4.750 -0.4971 0.02652 0.01681 -0.0051 1.0000 0.0862
-4.500 -0.4758 0.02454 0.01492 -0.0044 1.0000 0.0918
-4.250 -0.4533 0.02343 0.01369 -0.0033 1.0000 0.0957
-4.000 -0.4310 0.02272 0.01279 -0.0021 1.0000 0.0998
-3.750 -0.4086 0.02137 0.01158 -0.0014 1.0000 0.1052
-3.500 -0.3865 0.02061 0.01084 -0.0004 1.0000 0.1096
-3.250 -0.3644 0.02000 0.01020 0.0005 1.0000 0.1139
-3.000 -0.3428 0.01925 0.00952 0.0014 1.0000 0.1191
-2.750 -0.3216 0.01861 0.00902 0.0022 1.0000 0.1255
-2.500 -0.2998 0.01825 0.00864 0.0030 1.0000 0.1333
-2.250 -0.2782 0.01766 0.00816 0.0035 1.0000 0.1448
-2.000 -0.2556 0.01719 0.00779 0.0038 1.0000 0.1616
-1.750 -0.2349 0.01458 0.00758 0.0043 1.0000 0.6082
-1.500 -0.2367 0.01478 0.00871 0.0134 1.0000 0.8384
-1.250 -0.2174 0.01573 0.00963 0.0178 0.9965 0.9289
-1.000 -0.1075 0.01663 0.01022 0.0045 1.0000 0.9851
-0.750 -0.0220 0.01683 0.01023 -0.0072 0.9987 1.0000
-0.500 0.0392 0.01683 0.01010 -0.0149 0.9869 1.0000
-0.250 0.1005 0.01683 0.01000 -0.0224 0.9760 1.0000
0.000 0.1633 0.01678 0.00990 -0.0300 0.9653 1.0000
0.250 0.2237 0.01665 0.00974 -0.0368 0.9515 1.0000
0.500 0.2709 0.01646 0.00954 -0.0407 0.9320 1.0000
0.750 0.3065 0.01621 0.00930 -0.0420 0.9105 1.0000
1.000 0.3349 0.01591 0.00899 -0.0417 0.8898 1.0000
1.250 0.3580 0.01559 0.00865 -0.0401 0.8693 1.0000
1.500 0.3767 0.01533 0.00837 -0.0377 0.8455 1.0000
1.750 0.3954 0.01502 0.00802 -0.0350 0.8222 1.0000
2.000 0.4139 0.01474 0.00768 -0.0323 0.7994 1.0000
2.250 0.4332 0.01455 0.00742 -0.0298 0.7763 1.0000
2.500 0.4533 0.01448 0.00730 -0.0278 0.7514 1.0000
2.750 0.4740 0.01443 0.00716 -0.0257 0.7276 1.0000
3.000 0.4951 0.01445 0.00706 -0.0238 0.7042 1.0000
3.250 0.5166 0.01455 0.00712 -0.0223 0.6775 1.0000
3.500 0.5383 0.01466 0.00716 -0.0207 0.6521 1.0000
3.750 0.5602 0.01480 0.00721 -0.0193 0.6273 1.0000
4.000 0.5819 0.01498 0.00735 -0.0179 0.6000 1.0000
4.250 0.6036 0.01517 0.00743 -0.0163 0.5733 1.0000
4.500 0.6248 0.01539 0.00761 -0.0149 0.5414 1.0000
4.750 0.6456 0.01565 0.00777 -0.0134 0.5077 1.0000
5.000 0.6659 0.01597 0.00797 -0.0118 0.4709 1.0000
5.250 0.6856 0.01638 0.00824 -0.0102 0.4310 1.0000
5.500 0.7046 0.01691 0.00855 -0.0085 0.3905 1.0000
5.750 0.7237 0.01753 0.00894 -0.0070 0.3533 1.0000
6.000 0.7433 0.01821 0.00948 -0.0056 0.3212 1.0000
6.250 0.7636 0.01896 0.01006 -0.0044 0.2961 1.0000
6.500 0.7846 0.01969 0.01074 -0.0033 0.2737 1.0000
6.750 0.8057 0.02045 0.01141 -0.0023 0.2538 1.0000
7.000 0.8266 0.02124 0.01212 -0.0014 0.2343 1.0000
7.250 0.8471 0.02200 0.01279 -0.0005 0.2150 1.0000
7.500 0.8673 0.02278 0.01352 0.0004 0.1959 1.0000
7.750 0.8869 0.02354 0.01434 0.0014 0.1763 1.0000
8.000 0.9060 0.02442 0.01523 0.0024 0.1567 1.0000
8.250 0.9247 0.02560 0.01621 0.0034 0.1378 1.0000
8.500 0.9423 0.02674 0.01737 0.0047 0.1189 1.0000
8.750 0.9607 0.02802 0.01860 0.0057 0.1046 1.0000
9.000 0.9811 0.02953 0.02009 0.0066 0.0947 1.0000
9.250 1.0027 0.03106 0.02160 0.0073 0.0877 1.0000
9.500 1.0228 0.03254 0.02320 0.0082 0.0817 1.0000
9.750 1.0429 0.03401 0.02473 0.0089 0.0767 1.0000
10.000 1.0630 0.03572 0.02654 0.0096 0.0728 1.0000
10.250 1.0797 0.03782 0.02892 0.0106 0.0694 1.0000
10.500 1.0998 0.03987 0.03102 0.0112 0.0669 1.0000
10.750 1.1091 0.04285 0.03452 0.0129 0.0653 1.0000
11.000 1.1183 0.04577 0.03780 0.0143 0.0641 1.0000
11.250 1.1282 0.04862 0.04090 0.0155 0.0631 1.0000
11.500 1.1416 0.05151 0.04388 0.0163 0.0621 1.0000
11.750 1.1314 0.05530 0.04817 0.0187 0.0619 1.0000
12.000 1.1163 0.05909 0.05236 0.0210 0.0618 1.0000
12.250 1.0967 0.06288 0.05645 0.0231 0.0618 1.0000
12.500 1.0754 0.06713 0.06097 0.0240 0.0619 1.0000
12.750 1.0545 0.07188 0.06593 0.0236 0.0621 1.0000
13.000 0.7574 0.14817 0.14287 -0.0247 0.1375 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NPL 9660 AIRFOIL (npl9660-il)