NACA M24 AIRFOIL (m24-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M24 AIRFOIL (m24-il) Reynolds number: 1,000,000 Max Cl/Cd: 129.3 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m24-il-1000000.txt Download as CSV file: xf-m24-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M24 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3273 0.09376 0.09096 0.0000 0.5938 0.0186
-7.750 -0.3242 0.09028 0.08747 -0.0029 0.5903 0.0193
-7.500 -0.3264 0.08570 0.08286 -0.0088 0.5876 0.0201
-7.250 -0.3188 0.08117 0.07826 -0.0134 0.5845 0.0202
-7.000 -0.3074 0.07725 0.07431 -0.0159 0.5815 0.0202
-6.750 -0.3006 0.07220 0.06921 -0.0179 0.5784 0.0204
-6.500 -0.2879 0.06941 0.06639 -0.0184 0.5746 0.0206
-6.250 -0.2729 0.06689 0.06381 -0.0191 0.5707 0.0207
-6.000 -0.2568 0.06436 0.06124 -0.0201 0.5672 0.0209
-5.750 -0.2395 0.06179 0.05862 -0.0211 0.5637 0.0211
-5.500 -0.2211 0.05919 0.05597 -0.0222 0.5599 0.0215
-5.250 -0.2017 0.05653 0.05322 -0.0233 0.5563 0.0220
-5.000 -0.1809 0.05367 0.05026 -0.0243 0.5525 0.0231
-4.750 -0.1510 0.04884 0.04521 -0.0262 0.5500 0.0242
-4.500 -0.1294 0.04560 0.04185 -0.0265 0.5466 0.0242
-4.250 -0.1153 0.04135 0.03747 -0.0262 0.5432 0.0246
-4.000 -0.0938 0.03958 0.03562 -0.0262 0.5394 0.0248
-3.750 -0.0712 0.03791 0.03386 -0.0263 0.5357 0.0251
-3.500 -0.0479 0.03618 0.03206 -0.0262 0.5323 0.0255
-3.250 -0.0239 0.03445 0.03024 -0.0261 0.5285 0.0261
-3.000 0.0009 0.03237 0.02802 -0.0259 0.5248 0.0273
-2.750 0.0242 0.02566 0.02080 -0.0236 0.5222 0.0294
-2.500 0.0489 0.02471 0.01979 -0.0234 0.5187 0.0297
-2.250 0.0743 0.02365 0.01865 -0.0233 0.5151 0.0301
-2.000 0.1001 0.02256 0.01746 -0.0231 0.5112 0.0306
-1.750 0.1262 0.02124 0.01599 -0.0228 0.5073 0.0314
-1.500 0.1513 0.01759 0.01173 -0.0210 0.5042 0.0351
-1.250 0.1780 0.01663 0.01076 -0.0210 0.5011 0.0358
-1.000 0.2056 0.01602 0.01009 -0.0210 0.4972 0.0365
-0.750 0.2334 0.01537 0.00933 -0.0210 0.4931 0.0376
-0.500 0.2619 0.01483 0.00862 -0.0209 0.4889 0.0399
-0.250 0.2915 0.01145 0.00465 -0.0200 0.4864 0.0301
0.000 0.3202 0.01104 0.00420 -0.0201 0.4828 0.0301
0.250 0.3485 0.01071 0.00382 -0.0202 0.4789 0.0302
0.500 0.3767 0.01050 0.00358 -0.0203 0.4747 0.0309
0.750 0.4048 0.01026 0.00334 -0.0204 0.4712 0.0313
1.000 0.4328 0.01003 0.00310 -0.0204 0.4675 0.0317
1.250 0.4606 0.00985 0.00292 -0.0204 0.4637 0.0324
1.500 0.4884 0.00975 0.00278 -0.0204 0.4598 0.0331
1.750 0.5163 0.00965 0.00267 -0.0205 0.4563 0.0338
2.000 0.5445 0.00956 0.00259 -0.0206 0.4527 0.0343
2.250 0.5721 0.00944 0.00246 -0.0206 0.4487 0.0351
2.500 0.5998 0.00939 0.00238 -0.0207 0.4445 0.0366
2.750 0.6279 0.00936 0.00236 -0.0208 0.4408 0.0388
3.000 0.6562 0.00933 0.00234 -0.0210 0.4374 0.0413
3.250 0.6840 0.00927 0.00234 -0.0211 0.4336 0.0600
3.500 0.7112 0.00920 0.00244 -0.0212 0.4297 0.1334
3.750 0.7385 0.00916 0.00254 -0.0213 0.4258 0.2118
4.000 0.7432 0.00746 0.00267 -0.0166 0.4232 0.9372
4.250 0.7749 0.00767 0.00291 -0.0173 0.4191 0.9661
4.500 0.8185 0.00800 0.00320 -0.0208 0.4145 0.9763
4.750 0.8577 0.00827 0.00343 -0.0235 0.4103 0.9811
5.000 0.9233 0.00860 0.00375 -0.0319 0.4061 0.9854
5.250 0.9635 0.00875 0.00388 -0.0349 0.4019 0.9870
5.500 1.0081 0.00897 0.00407 -0.0389 0.3972 0.9898
5.750 1.0608 0.00917 0.00427 -0.0446 0.3932 0.9946
6.000 1.0939 0.00924 0.00434 -0.0461 0.3892 0.9952
6.250 1.1238 0.00935 0.00443 -0.0470 0.3841 0.9956
6.500 1.1535 0.00947 0.00454 -0.0478 0.3789 0.9961
6.750 1.1833 0.00955 0.00463 -0.0486 0.3730 0.9966
7.000 1.2127 0.00972 0.00477 -0.0494 0.3670 0.9971
7.250 1.2426 0.00982 0.00488 -0.0503 0.3611 0.9976
7.500 1.2723 0.00997 0.00502 -0.0512 0.3542 0.9982
7.750 1.3020 0.01014 0.00519 -0.0521 0.3472 0.9988
8.000 1.3313 0.01032 0.00536 -0.0530 0.3394 0.9993
8.250 1.3602 0.01052 0.00555 -0.0538 0.3310 0.9997
8.750 1.4110 0.01109 0.00605 -0.0542 0.3077 1.0000
9.000 1.4335 0.01149 0.00638 -0.0538 0.2916 1.0000
9.250 1.4536 0.01210 0.00686 -0.0533 0.2649 1.0000
9.500 1.4577 0.01411 0.00836 -0.0510 0.1887 1.0000
9.750 1.4390 0.01702 0.01076 -0.0456 0.1078 1.0000
10.000 1.4163 0.01903 0.01262 -0.0390 0.0764 1.0000
10.250 1.3895 0.02042 0.01405 -0.0314 0.0698 1.0000
10.500 1.3646 0.02279 0.01637 -0.0258 0.0515 1.0000
10.750 1.3412 0.02588 0.01937 -0.0217 0.0291 1.0000
11.000 1.3341 0.02826 0.02176 -0.0195 0.0225 1.0000
11.250 1.3344 0.03023 0.02377 -0.0181 0.0203 1.0000
11.500 1.3348 0.03235 0.02595 -0.0169 0.0191 1.0000
11.750 1.3342 0.03469 0.02834 -0.0159 0.0179 1.0000
12.000 1.3350 0.03697 0.03068 -0.0151 0.0172 1.0000
12.250 1.3370 0.03919 0.03296 -0.0145 0.0167 1.0000
12.500 1.3378 0.04156 0.03540 -0.0139 0.0162 1.0000
12.750 1.3379 0.04409 0.03800 -0.0134 0.0157 1.0000
13.000 1.3369 0.04683 0.04080 -0.0131 0.0152 1.0000
13.250 1.3347 0.04980 0.04383 -0.0128 0.0148 1.0000
13.500 1.3291 0.05320 0.04731 -0.0126 0.0143 1.0000
13.750 1.3244 0.05658 0.05077 -0.0124 0.0139 1.0000
14.000 1.3257 0.05931 0.05357 -0.0124 0.0137 1.0000
14.250 1.3253 0.06228 0.05660 -0.0124 0.0135 1.0000
14.500 1.3237 0.06542 0.05982 -0.0125 0.0132 1.0000
14.750 1.3226 0.06854 0.06301 -0.0126 0.0129 1.0000
15.000 1.3205 0.07184 0.06637 -0.0129 0.0127 1.0000
15.250 1.3178 0.07524 0.06983 -0.0131 0.0124 1.0000
15.500 1.3157 0.07857 0.07322 -0.0134 0.0122 1.0000
15.750 1.3121 0.08218 0.07691 -0.0138 0.0120 1.0000
16.000 1.3080 0.08587 0.08066 -0.0142 0.0118 1.0000
16.250 1.3025 0.08976 0.08461 -0.0147 0.0116 1.0000
16.500 1.2942 0.09411 0.08904 -0.0154 0.0113 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M24 AIRFOIL (m24-il)