BOEING-VERTOL VR-8 AIRFOIL WITH TAB (vr8b-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-8 AIRFOIL WITH TAB (vr8b-il) Reynolds number: 500,000 Max Cl/Cd: 50.94 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr8b-il-500000-n5.txt Download as CSV file: xf-vr8b-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-8 AIRFOIL WITH TAB
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.7029 0.09454 0.09264 0.0341 1.0000 0.0042
-8.500 -0.7049 0.08874 0.08686 0.0297 1.0000 0.0042
-8.250 -0.7086 0.08225 0.08040 0.0233 1.0000 0.0042
-7.750 -0.7046 0.06542 0.06340 0.0048 1.0000 0.0042
-7.500 -0.6998 0.05719 0.05496 -0.0005 1.0000 0.0042
-7.250 -0.6883 0.05091 0.04845 -0.0036 1.0000 0.0044
-7.000 -0.6721 0.04566 0.04295 -0.0055 1.0000 0.0046
-6.750 -0.6532 0.04071 0.03772 -0.0069 1.0000 0.0048
-6.250 -0.6091 0.03059 0.02681 -0.0085 0.9805 0.0055
-6.000 -0.5879 0.02342 0.01885 -0.0077 0.9719 0.0065
-5.750 -0.5658 0.01871 0.01338 -0.0065 0.9642 0.0071
-5.500 -0.5407 0.01819 0.01276 -0.0062 0.9558 0.0074
-5.250 -0.5148 0.01751 0.01196 -0.0060 0.9461 0.0080
-5.000 -0.4881 0.01640 0.01062 -0.0057 0.9319 0.0091
-4.750 -0.4646 0.01534 0.00919 -0.0043 0.9149 0.0105
-4.500 -0.4424 0.01503 0.00886 -0.0032 0.8968 0.0113
-4.250 -0.4181 0.01455 0.00821 -0.0023 0.8739 0.0128
-4.000 -0.3932 0.01453 0.00779 -0.0012 0.8428 0.0140
-3.750 -0.3700 0.01354 0.00655 -0.0002 0.7652 0.0150
-3.500 -0.3414 0.01327 0.00580 -0.0006 0.6539 0.0166
-3.250 -0.3110 0.01332 0.00537 -0.0014 0.5653 0.0178
-3.000 -0.2812 0.01296 0.00477 -0.0020 0.5176 0.0182
-2.750 -0.2515 0.01218 0.00384 -0.0029 0.4658 0.0193
-2.500 -0.2168 0.01300 0.00346 -0.0057 0.0395 0.0196
-2.250 -0.1873 0.01274 0.00310 -0.0061 0.0330 0.0196
-2.000 -0.1578 0.01250 0.00279 -0.0066 0.0318 0.0197
-1.750 -0.1282 0.01228 0.00251 -0.0071 0.0243 0.0199
-1.500 -0.0984 0.01214 0.00226 -0.0076 0.0229 0.0203
-1.250 -0.0689 0.01200 0.00206 -0.0081 0.0225 0.0207
-1.000 -0.0394 0.01190 0.00191 -0.0086 0.0224 0.0212
-0.750 -0.0099 0.01183 0.00180 -0.0090 0.0223 0.0216
-0.500 0.0195 0.01179 0.00171 -0.0095 0.0221 0.0219
-0.250 0.0490 0.01178 0.00165 -0.0099 0.0222 0.0228
0.000 0.0784 0.01178 0.00162 -0.0103 0.0221 0.0237
0.250 0.1078 0.01181 0.00161 -0.0108 0.0222 0.0239
0.500 0.1371 0.01185 0.00162 -0.0112 0.0222 0.0241
0.750 0.1663 0.01191 0.00161 -0.0116 0.0223 0.0254
1.000 0.1958 0.01188 0.00168 -0.0121 0.0223 0.0260
1.250 0.2251 0.01191 0.00180 -0.0126 0.0221 0.0316
1.500 0.2544 0.01199 0.00192 -0.0130 0.0216 0.0320
1.750 0.2838 0.01208 0.00205 -0.0135 0.0213 0.0329
2.000 0.3131 0.01217 0.00218 -0.0140 0.0211 0.0378
2.500 0.3751 0.01113 0.00257 -0.0167 0.0208 0.5488
2.750 0.4049 0.01080 0.00285 -0.0176 0.0206 0.7154
3.000 0.4329 0.01077 0.00311 -0.0178 0.0205 0.7894
3.250 0.4571 0.01079 0.00340 -0.0169 0.0204 0.8643
3.500 0.4743 0.01084 0.00363 -0.0142 0.0203 0.9348
3.750 0.5002 0.01097 0.00380 -0.0137 0.0204 1.0000
4.000 0.5296 0.01129 0.00411 -0.0143 0.0205 1.0000
4.250 0.5588 0.01162 0.00445 -0.0149 0.0206 1.0000
4.500 0.5878 0.01197 0.00481 -0.0154 0.0208 1.0000
4.750 0.6168 0.01234 0.00522 -0.0159 0.0210 1.0000
5.000 0.6455 0.01275 0.00565 -0.0164 0.0212 1.0000
5.250 0.6740 0.01323 0.00615 -0.0169 0.0214 1.0000
5.500 0.7022 0.01379 0.00674 -0.0174 0.0216 1.0000
5.750 0.7299 0.01444 0.00744 -0.0178 0.0209 1.0000
6.000 0.7532 0.01674 0.00940 -0.0180 0.0194 1.0000
6.250 0.7820 0.01676 0.00974 -0.0184 0.0177 1.0000
6.500 0.8074 0.01812 0.01101 -0.0185 0.0159 1.0000
6.750 0.8350 0.01847 0.01158 -0.0188 0.0145 1.0000
7.000 0.8604 0.01992 0.01282 -0.0190 0.0135 1.0000
7.250 0.8886 0.01965 0.01307 -0.0193 0.0116 1.0000
7.750 0.9427 0.02032 0.01413 -0.0198 0.0078 1.0000
8.000 0.9687 0.02097 0.01488 -0.0199 0.0070 1.0000
8.250 0.9937 0.02190 0.01595 -0.0200 0.0063 1.0000
8.500 1.0191 0.02255 0.01672 -0.0201 0.0058 1.0000
8.750 1.0445 0.02301 0.01725 -0.0203 0.0054 1.0000
9.000 1.0687 0.02386 0.01822 -0.0204 0.0049 1.0000
9.250 1.0886 0.02606 0.02071 -0.0200 0.0046 1.0000
9.500 1.1061 0.02885 0.02390 -0.0194 0.0043 1.0000
9.750 1.1217 0.03166 0.02711 -0.0188 0.0039 1.0000
10.000 1.1361 0.03426 0.03002 -0.0183 0.0036 1.0000
10.250 1.1462 0.03733 0.03344 -0.0176 0.0034 1.0000
10.500 1.1498 0.04105 0.03753 -0.0168 0.0033 1.0000
10.750 1.1463 0.04521 0.04204 -0.0159 0.0032 1.0000
11.000 1.1299 0.04975 0.04689 -0.0149 0.0032 1.0000
11.250 1.1070 0.05542 0.05282 -0.0163 0.0032 1.0000
11.500 1.0760 0.06394 0.06162 -0.0214 0.0032 1.0000
11.750 1.0424 0.07476 0.07267 -0.0293 0.0033 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-8 AIRFOIL WITH TAB (vr8b-il)