EPPLER 1098 AIRFOIL (e1098-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 1098 AIRFOIL (e1098-il) Reynolds number: 200,000 Max Cl/Cd: 63.73 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e1098-il-200000-n5.txt Download as CSV file: xf-e1098-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 1098 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.000 -0.3737 0.07299 0.06798 -0.1193 0.8002 0.0131
-13.750 -0.3987 0.06631 0.06104 -0.1230 0.7930 0.0130
-13.500 -0.4163 0.06143 0.05593 -0.1253 0.7861 0.0130
-13.250 -0.4308 0.05726 0.05154 -0.1269 0.7794 0.0130
-13.000 -0.4428 0.05362 0.04766 -0.1279 0.7735 0.0130
-12.750 -0.4514 0.05039 0.04421 -0.1285 0.7674 0.0130
-12.500 -0.4578 0.04747 0.04107 -0.1288 0.7620 0.0130
-12.250 -0.4603 0.04495 0.03838 -0.1288 0.7570 0.0132
-12.000 -0.4606 0.04267 0.03594 -0.1286 0.7518 0.0133
-11.750 -0.4590 0.04054 0.03360 -0.1281 0.7473 0.0134
-11.500 -0.4545 0.03873 0.03161 -0.1276 0.7436 0.0135
-11.250 -0.4481 0.03708 0.02984 -0.1272 0.7395 0.0138
-11.000 -0.4395 0.03551 0.02811 -0.1265 0.7357 0.0140
-10.750 -0.4290 0.03407 0.02650 -0.1258 0.7323 0.0142
-10.500 -0.4170 0.03278 0.02504 -0.1251 0.7294 0.0146
-10.250 -0.4021 0.03153 0.02365 -0.1244 0.7265 0.0149
-10.000 -0.3864 0.03040 0.02239 -0.1237 0.7234 0.0152
-9.750 -0.3698 0.02937 0.02124 -0.1229 0.7205 0.0157
-9.500 -0.3533 0.02846 0.02019 -0.1221 0.7178 0.0164
-9.250 -0.3376 0.02755 0.01919 -0.1213 0.7154 0.0173
-9.000 -0.3229 0.02669 0.01828 -0.1204 0.7132 0.0179
-8.750 -0.3085 0.02586 0.01744 -0.1195 0.7105 0.0186
-8.500 -0.2947 0.02502 0.01657 -0.1184 0.7078 0.0193
-8.250 -0.2821 0.02419 0.01567 -0.1172 0.7052 0.0198
-8.000 -0.2715 0.02339 0.01480 -0.1157 0.7029 0.0205
-7.750 -0.2636 0.02265 0.01398 -0.1138 0.7008 0.0213
-7.500 -0.2551 0.02186 0.01312 -0.1121 0.6990 0.0222
-7.250 -0.2435 0.02110 0.01230 -0.1108 0.6973 0.0232
-7.000 -0.2288 0.02047 0.01165 -0.1098 0.6953 0.0248
-6.750 -0.2114 0.01995 0.01105 -0.1090 0.6933 0.0273
-6.500 -0.1948 0.01930 0.01039 -0.1082 0.6914 0.0307
-6.250 -0.1750 0.01880 0.00981 -0.1076 0.6895 0.0352
-6.000 -0.1555 0.01821 0.00922 -0.1071 0.6879 0.0429
-5.750 -0.1354 0.01759 0.00866 -0.1067 0.6863 0.0615
-5.500 -0.1173 0.01666 0.00801 -0.1064 0.6847 0.1220
-5.250 -0.1009 0.01531 0.00722 -0.1063 0.6832 0.2484
-5.000 -0.0830 0.01428 0.00706 -0.1061 0.6816 0.4265
-4.750 -0.0559 0.01444 0.00725 -0.1061 0.6800 0.4795
-4.500 -0.0281 0.01468 0.00743 -0.1061 0.6784 0.5080
-4.250 -0.0005 0.01516 0.00788 -0.1059 0.6768 0.5338
-4.000 0.0270 0.01558 0.00828 -0.1056 0.6754 0.5497
-3.750 0.0551 0.01572 0.00827 -0.1059 0.6740 0.5589
-3.500 0.0831 0.01580 0.00832 -0.1060 0.6727 0.5629
-3.250 0.1114 0.01586 0.00829 -0.1062 0.6714 0.5663
-3.000 0.1400 0.01586 0.00820 -0.1066 0.6703 0.5698
-2.750 0.1690 0.01584 0.00804 -0.1072 0.6691 0.5734
-2.500 0.1981 0.01582 0.00790 -0.1077 0.6679 0.5761
-2.250 0.2258 0.01585 0.00790 -0.1080 0.6666 0.5779
-2.000 0.2528 0.01590 0.00793 -0.1081 0.6651 0.5798
-1.750 0.2802 0.01594 0.00795 -0.1083 0.6637 0.5818
-1.500 0.3076 0.01599 0.00796 -0.1086 0.6623 0.5840
-1.250 0.3354 0.01604 0.00797 -0.1090 0.6609 0.5863
-1.000 0.3633 0.01608 0.00796 -0.1094 0.6596 0.5884
-0.750 0.3916 0.01612 0.00793 -0.1099 0.6582 0.5905
-0.500 0.4196 0.01616 0.00795 -0.1102 0.6568 0.5922
-0.250 0.4478 0.01621 0.00799 -0.1105 0.6555 0.5939
0.000 0.4762 0.01628 0.00804 -0.1110 0.6545 0.5955
0.250 0.5049 0.01635 0.00809 -0.1114 0.6536 0.5972
0.500 0.5341 0.01642 0.00813 -0.1120 0.6526 0.5990
0.750 0.5601 0.01656 0.00830 -0.1121 0.6509 0.6010
1.000 0.5852 0.01673 0.00850 -0.1120 0.6490 0.6033
1.250 0.6112 0.01690 0.00869 -0.1121 0.6473 0.6058
1.500 0.6373 0.01704 0.00886 -0.1122 0.6454 0.6076
1.750 0.6638 0.01717 0.00902 -0.1123 0.6437 0.6092
2.000 0.6910 0.01728 0.00917 -0.1125 0.6421 0.6110
2.250 0.7189 0.01737 0.00929 -0.1128 0.6406 0.6130
2.500 0.7475 0.01745 0.00937 -0.1132 0.6392 0.6151
2.750 0.7771 0.01751 0.00945 -0.1139 0.6379 0.6174
3.000 0.8025 0.01771 0.00968 -0.1138 0.6358 0.6200
3.250 0.8234 0.01800 0.01007 -0.1131 0.6327 0.6226
3.500 0.8471 0.01818 0.01034 -0.1127 0.6300 0.6247
3.750 0.8733 0.01828 0.01052 -0.1127 0.6276 0.6271
4.000 0.9018 0.01830 0.01057 -0.1130 0.6254 0.6296
4.250 0.9325 0.01823 0.01052 -0.1137 0.6234 0.6323
4.500 0.9565 0.01837 0.01074 -0.1134 0.6200 0.6352
5.000 1.0020 0.01852 0.01107 -0.1121 0.6116 0.6405
5.250 1.0332 0.01833 0.01093 -0.1127 0.6087 0.6433
5.500 1.0556 0.01844 0.01113 -0.1121 0.6046 0.6468
5.750 1.0732 0.01866 0.01148 -0.1106 0.5996 0.6505
6.000 1.0992 0.01862 0.01150 -0.1104 0.5959 0.6544
6.250 1.1303 0.01843 0.01139 -0.1111 0.5929 0.6577
6.500 1.1417 0.01879 0.01192 -0.1086 0.5873 0.6614
6.750 1.1609 0.01888 0.01212 -0.1073 0.5822 0.6657
7.000 1.1905 0.01868 0.01197 -0.1077 0.5783 0.6700
7.250 1.1947 0.01911 0.01261 -0.1040 0.5712 0.6739
7.500 1.2122 0.01908 0.01267 -0.1023 0.5652 0.6787
7.750 1.2144 0.01948 0.01320 -0.0982 0.5574 0.6840
8.000 1.2301 0.01948 0.01329 -0.0962 0.5491 0.6891
8.250 1.2240 0.02033 0.01428 -0.0914 0.5373 0.6950
8.500 1.2225 0.02118 0.01523 -0.0876 0.5243 0.7017
8.750 1.2276 0.02183 0.01597 -0.0847 0.5083 0.7084
9.000 1.2420 0.02213 0.01628 -0.0831 0.4900 0.7164
9.250 1.2581 0.02232 0.01642 -0.0816 0.4665 0.7245
9.500 1.2687 0.02292 0.01692 -0.0796 0.4421 0.7336
9.750 1.2722 0.02402 0.01791 -0.0769 0.4174 0.7437
10.000 1.2718 0.02547 0.01930 -0.0742 0.3934 0.7548
10.250 1.2707 0.02707 0.02086 -0.0716 0.3720 0.7677
10.500 1.2695 0.02877 0.02254 -0.0691 0.3511 0.7833
10.750 1.2678 0.03054 0.02430 -0.0667 0.3317 0.8033
11.000 1.2664 0.03227 0.02606 -0.0644 0.3143 0.8323
11.250 1.2635 0.03384 0.02776 -0.0618 0.2970 0.9361
11.500 1.2637 0.03564 0.02951 -0.0600 0.2806 1.0000
11.750 1.2661 0.03759 0.03141 -0.0588 0.2641 1.0000
12.000 1.2679 0.03963 0.03339 -0.0575 0.2470 1.0000
12.250 1.2699 0.04172 0.03543 -0.0564 0.2303 1.0000
12.500 1.2717 0.04391 0.03757 -0.0554 0.2137 1.0000
12.750 1.2730 0.04618 0.03978 -0.0544 0.1968 1.0000
13.000 1.2749 0.04848 0.04201 -0.0535 0.1795 1.0000
13.250 1.2768 0.05084 0.04431 -0.0528 0.1620 1.0000
13.500 1.2781 0.05331 0.04670 -0.0521 0.1448 1.0000
13.750 1.2771 0.05607 0.04936 -0.0515 0.1269 1.0000
14.000 1.2781 0.05870 0.05190 -0.0510 0.1102 1.0000
14.250 1.2788 0.06142 0.05455 -0.0506 0.0949 1.0000
14.500 1.2811 0.06404 0.05712 -0.0503 0.0824 1.0000
14.750 1.2837 0.06668 0.05974 -0.0501 0.0723 1.0000
15.000 1.2867 0.06931 0.06238 -0.0499 0.0641 1.0000
15.250 1.2903 0.07193 0.06504 -0.0499 0.0556 1.0000
15.500 1.2932 0.07468 0.06781 -0.0499 0.0480 1.0000
15.750 1.2946 0.07769 0.07083 -0.0500 0.0407 1.0000
16.000 1.2950 0.08085 0.07400 -0.0502 0.0338 1.0000
16.500 1.2934 0.08769 0.08090 -0.0509 0.0229 1.0000
16.750 1.2922 0.09127 0.08454 -0.0514 0.0194 1.0000
17.000 1.2888 0.09524 0.08856 -0.0522 0.0162 1.0000
17.250 1.2879 0.09891 0.09233 -0.0529 0.0144 1.0000
17.500 1.2846 0.10300 0.09650 -0.0539 0.0130 1.0000
17.750 1.2829 0.10692 0.10055 -0.0550 0.0119 1.0000
18.000 1.2809 0.11095 0.10470 -0.0562 0.0110 1.0000
18.250 1.2774 0.11526 0.10912 -0.0577 0.0104 1.0000
18.500 1.2736 0.11968 0.11366 -0.0593 0.0100 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 1098 AIRFOIL (e1098-il)