Skip to content

Commit 394969f

Browse files
cbell14cbell14
committed
Merge pull request #12 from barbagroup/master
updating
2 parents 6fecc15 + 6c03754 commit 394969f

26 files changed

Lines changed: 1461 additions & 6 deletions

‎lessons/11_Lesson11_vortexSourcePanelMethod.ipynb‎

Lines changed: 6 additions & 6 deletions
Original file line numberDiff line numberDiff line change
@@ -1,7 +1,7 @@
11
{
22
"metadata": {
33
"name": "",
4-
"signature": "sha256:cf4c65f24e5cc8c0fb88e140e3a39bbaef85fa46059dd161bc48bdff73413944"
4+
"signature": "sha256:fc91852ef4248b15d605d5513f948bd5496d03055b6dea91be90e4b274359323"
55
},
66
"nbformat": 3,
77
"nbformat_minor": 0,
@@ -27,22 +27,22 @@
2727
"level": 5,
2828
"metadata": {},
2929
"source": [
30-
"Version 0.4 -- April 2015"
30+
"Version 0.5 -- April 2015"
3131
]
3232
},
3333
{
3434
"cell_type": "heading",
3535
"level": 1,
3636
"metadata": {},
3737
"source": [
38-
"Source-vortex panel method"
38+
"Vortex-source panel method"
3939
]
4040
},
4141
{
4242
"cell_type": "markdown",
4343
"metadata": {},
4444
"source": [
45-
"In [Lesson 9](http://nbviewer.ipython.org/urls/github.com/barbagroup/AeroPython/blob/master/lessons/09_Lesson09_flowOverCylinder.ipynb) of _AeroPython_, you learned to use a _source panel method_ to represent a circular cylinder, and in [Lesson 10](http://nbviewer.ipython.org/urls/github.com/barbagroup/AeroPython/blob/master/lessons/10_Lesson10_sourcePanelMethod.ipynb) we used it for a symmetric airfoil at zero angle of attack. But what if we want the airfoil to generate some lift? If we place the airfoil at a non-zero angle of attack, we _should_ get lift, but will a source-panel representation be able to give you lift? Remember the _Kutta-Joukowski theorem_?\n",
45+
"In [Lesson 9](http://nbviewer.ipython.org/urls/github.com/barbagroup/AeroPython/blob/master/lessons/09_Lesson09_flowOverCylinder.ipynb) of _AeroPython_, you learned to use a _source panel method_ to represent a circular cylinder, and in [Lesson 10](http://nbviewer.ipython.org/urls/github.com/barbagroup/AeroPython/blob/master/lessons/10_Lesson10_sourcePanelMethod.ipynb) we used it for a symmetric airfoil at zero angle of attack. But what if we want the airfoil to generate some lift? If we place the airfoil at a non-zero angle of attack, we _should_ get lift, but will a source-panel representation be able to give you lift? Remember the [_Kutta-Joukowski theorem_](http://en.wikipedia.org/wiki/Kutta%E2%80%93Joukowski_theorem)?\n",
4646
"\n",
4747
"\n",
4848
"Historically, the first panel method ever developed was a source-sheet method. At the time, Douglas Aircraft Company was concerned with calculating the flow around bodies of revolution, and it was only later that the method was extended to lifting surfaces. (See the reference below for a nice historical account.)\n",
@@ -75,7 +75,7 @@
7575
"\n",
7676
"For example, using the source-sheet panel method of [Lesson 10](http://nbviewer.ipython.org/github/barbagroup/AeroPython/blob/master/lessons/10_Lesson10_sourcePanelMethod.ipynb) with an angle of attack $\\alpha=4^\\circ$ (using 40 panels), and plotting the streamlines in an area close to the trailing edge, we get the following plot:\n",
7777
"\n",
78-
"<center><img src=\"./resources/StreamlinesTrailingEdge.png\"></center>"
78+
"<center><img src=\"./resources/StreamlinesTrailingEdge.png\" width=\"600\"></center>"
7979
]
8080
},
8181
{
@@ -459,7 +459,7 @@
459459
"cell_type": "markdown",
460460
"metadata": {},
461461
"source": [
462-
"To enforce the *Kutta-condition*, we state that the pressure coefficient on the fisrt panel must be equal to that on the last panel:\n",
462+
"To enforce the *Kutta-condition*, we state that the pressure coefficient on the first panel must be equal to that on the last panel:\n",
463463
"\n",
464464
"$$C_{p_1} = C_{p_{N}}$$\n",
465465
"\n",

‎lessons/tutorial_xflr5/airfoils/NACA0012.dat‎

Lines changed: 102 additions & 0 deletions
Original file line numberDiff line numberDiff line change
@@ -0,0 +1,102 @@
1+
NACA0012
2+
1.000000 -0.000000
3+
0.980000 0.002870
4+
0.960000 0.005667
5+
0.940000 0.008395
6+
0.920000 0.011055
7+
0.900000 0.013650
8+
0.880000 0.016183
9+
0.860000 0.018655
10+
0.840000 0.021066
11+
0.820000 0.023420
12+
0.800000 0.025715
13+
0.780000 0.027953
14+
0.760000 0.030135
15+
0.740000 0.032260
16+
0.720000 0.034327
17+
0.700000 0.036337
18+
0.680000 0.038287
19+
0.660000 0.040178
20+
0.640000 0.042007
21+
0.620000 0.043772
22+
0.600000 0.045470
23+
0.580000 0.047100
24+
0.560000 0.048658
25+
0.540000 0.050140
26+
0.520000 0.051543
27+
0.500000 0.052862
28+
0.480000 0.054091
29+
0.460000 0.055226
30+
0.440000 0.056260
31+
0.420000 0.057186
32+
0.400000 0.057998
33+
0.380000 0.058686
34+
0.360000 0.059242
35+
0.340000 0.059655
36+
0.320000 0.059915
37+
0.300000 0.060007
38+
0.280000 0.059918
39+
0.260000 0.059631
40+
0.240000 0.059127
41+
0.220000 0.058383
42+
0.200000 0.057373
43+
0.180000 0.056065
44+
0.160000 0.054417
45+
0.140000 0.052379
46+
0.120000 0.049882
47+
0.100000 0.046828
48+
0.080000 0.043072
49+
0.060000 0.038376
50+
0.040000 0.032277
51+
0.020000 0.023598
52+
0.000000 0.000000
53+
0.020000 -0.023598
54+
0.040000 -0.032277
55+
0.060000 -0.038376
56+
0.080000 -0.043072
57+
0.100000 -0.046828
58+
0.120000 -0.049882
59+
0.140000 -0.052379
60+
0.160000 -0.054417
61+
0.180000 -0.056065
62+
0.200000 -0.057373
63+
0.220000 -0.058383
64+
0.240000 -0.059127
65+
0.260000 -0.059631
66+
0.280000 -0.059918
67+
0.300000 -0.060007
68+
0.320000 -0.059915
69+
0.340000 -0.059655
70+
0.360000 -0.059242
71+
0.380000 -0.058686
72+
0.400000 -0.057998
73+
0.420000 -0.057186
74+
0.440000 -0.056260
75+
0.460000 -0.055226
76+
0.480000 -0.054091
77+
0.500000 -0.052862
78+
0.520000 -0.051543
79+
0.540000 -0.050140
80+
0.560000 -0.048658
81+
0.580000 -0.047100
82+
0.600000 -0.045470
83+
0.620000 -0.043772
84+
0.640000 -0.042007
85+
0.660000 -0.040178
86+
0.680000 -0.038287
87+
0.700000 -0.036337
88+
0.720000 -0.034327
89+
0.740000 -0.032260
90+
0.760000 -0.030135
91+
0.780000 -0.027953
92+
0.800000 -0.025715
93+
0.820000 -0.023420
94+
0.840000 -0.021066
95+
0.860000 -0.018655
96+
0.880000 -0.016183
97+
0.900000 -0.013650
98+
0.920000 -0.011055
99+
0.940000 -0.008395
100+
0.960000 -0.005667
101+
0.980000 -0.002870
102+
1.000000 0.000000

‎lessons/tutorial_xflr5/airfoils/NACA0012.png‎

19.9 KB
Loading

‎lessons/tutorial_xflr5/airfoils/S1223.dat‎

Lines changed: 82 additions & 0 deletions
Original file line numberDiff line numberDiff line change
@@ -0,0 +1,82 @@
1+
S1223
2+
1.00000 0.00000
3+
0.99103 0.00784
4+
0.97802 0.01835
5+
0.96261 0.02680
6+
0.94323 0.03502
7+
0.91930 0.04356
8+
0.89079 0.05235
9+
0.85856 0.06113
10+
0.82381 0.06959
11+
0.78742 0.07760
12+
0.74999 0.08512
13+
0.71190 0.09213
14+
0.67331 0.09864
15+
0.63433 0.10466
16+
0.59499 0.11025
17+
0.55540 0.11542
18+
0.51556 0.12020
19+
0.47594 0.12455
20+
0.43691 0.12839
21+
0.39897 0.13154
22+
0.36239 0.13382
23+
0.32727 0.13507
24+
0.29380 0.13526
25+
0.26298 0.13413
26+
0.23369 0.13144
27+
0.20520 0.12719
28+
0.17753 0.12140
29+
0.15063 0.11416
30+
0.12508 0.10568
31+
0.10156 0.09630
32+
0.08059 0.08633
33+
0.06256 0.07625
34+
0.04762 0.06646
35+
0.03552 0.05724
36+
0.02596 0.04872
37+
0.01853 0.04078
38+
0.01277 0.03330
39+
0.00833 0.02630
40+
0.00500 0.01979
41+
0.00259 0.01372
42+
0.00098 0.00806
43+
0.00013 0.00283
44+
0.00003 -0.00207
45+
0.00068 -0.00682
46+
0.00246 -0.01098
47+
0.00619 -0.01366
48+
0.01123 -0.01502
49+
0.01719 -0.01550
50+
0.02435 -0.01576
51+
0.03302 -0.01582
52+
0.04390 -0.01544
53+
0.05829 -0.01458
54+
0.07754 -0.01302
55+
0.10295 -0.01040
56+
0.13352 -0.00658
57+
0.16679 -0.00133
58+
0.20387 0.00556
59+
0.24384 0.01316
60+
0.28373 0.02055
61+
0.32309 0.02747
62+
0.36202 0.03383
63+
0.40063 0.03957
64+
0.43901 0.04467
65+
0.47719 0.04910
66+
0.51512 0.05282
67+
0.55289 0.05580
68+
0.59047 0.05802
69+
0.62777 0.05944
70+
0.66474 0.06002
71+
0.70133 0.05974
72+
0.73744 0.05854
73+
0.77288 0.05640
74+
0.80748 0.05328
75+
0.84096 0.04915
76+
0.87283 0.04405
77+
0.90248 0.03804
78+
0.92915 0.03131
79+
0.95223 0.02400
80+
0.97196 0.01595
81+
0.98931 0.00672
82+
1.00000 0.00000

‎lessons/tutorial_xflr5/airfoils/S1223.png‎

19.7 KB
Loading

‎lessons/tutorial_xflr5/experimental_data/gregory_oreilly_1970_naca0012_AoA0.dat‎

Lines changed: 31 additions & 0 deletions
Original file line numberDiff line numberDiff line change
@@ -0,0 +1,31 @@
1+
# Data from Gregory and O'Reilly, NASA R&M 3726 (1970)
2+
# Reynolds number Re=2.88E+06
3+
# Angle of attack alpha=0deg
4+
# Data was digitized from a photocopy, hence is only approximate
5+
# Data for the upper airfoil surface only
6+
# [x/c] [Cp]
7+
0 1
8+
0.0023497 0.847673
9+
0.00496048 0.456198
10+
0.00526903 0.173569
11+
0.0142406 -0.044407
12+
0.0209337 -0.175278
13+
0.0473501 -0.372653
14+
0.0779437 -0.396388
15+
0.0976194 -0.41941
16+
0.128166 -0.418874
17+
0.150001 -0.411087
18+
0.178387 -0.402938
19+
0.289702 -0.36672
20+
0.322431 -0.347115
21+
0.387891 -0.307906
22+
0.448983 -0.268412
23+
0.514442 -0.229203
24+
0.579902 -0.189994
25+
0.638834 -0.159098
26+
0.704317 -0.114629
27+
0.767593 -0.065278
28+
0.835236 -0.026211
29+
0.896305 0.03502
30+
0.959533 0.0978565
31+
1.0009 0.173854

‎lessons/tutorial_xflr5/experimental_data/gregory_oreilly_1970_naca0012_AoA10.dat‎

Lines changed: 30 additions & 0 deletions
Original file line numberDiff line numberDiff line change
@@ -0,0 +1,30 @@
1+
# Data from Gregory and O'Reilly, NASA R&M 3726 (1970)
2+
# Reynolds number Re=2.88E+06
3+
# Angle of attack alpha=10deg
4+
# Data was digitized from a photocopy, hence is only approximate
5+
# Data for the upper airfoil surface only
6+
# [x/c] [Cp]
7+
0 -3.66423
8+
0.00218341 -5.04375
9+
0.00873362 -5.24068
10+
0.0131004 -4.67125
11+
0.0174672 -4.32079
12+
0.0480349 -2.74347
13+
0.0742358 -2.26115
14+
0.0982533 -1.95405
15+
0.124454 -1.7345
16+
0.146288 -1.55884
17+
0.176856 -1.36109
18+
0.28821 -1.00829
19+
0.320961 -0.941877
20+
0.384279 -0.787206
21+
0.447598 -0.654432
22+
0.515284 -0.543461
23+
0.576419 -0.432633
24+
0.637555 -0.343703
25+
0.700873 -0.254725
26+
0.766376 -0.1657
27+
0.831878 -0.098572
28+
0.893013 -0.00964205
29+
0.958515 0.0793835
30+
1 0.124088

‎lessons/tutorial_xflr5/experimental_data/gregory_oreilly_1970_naca0012_AoA15.dat‎

Lines changed: 30 additions & 0 deletions
Original file line numberDiff line numberDiff line change
@@ -0,0 +1,30 @@
1+
# Data from Gregory and O'Reilly, NASA R&M 3726 (1970)
2+
# Reynolds number Re=2.88E+06
3+
# Angle of attack alpha=15deg
4+
# Data was digitized from a photocopy, hence is only approximate
5+
# Data for the upper airfoil surface only
6+
# [x/c] [Cp]
7+
-7.59438e-05 -8.65066
8+
0.0024302 -10.1789
9+
0.00450442 -9.72033
10+
0.00870506 -9.04329
11+
0.0129722 -8.67192
12+
0.0167741 -6.16084
13+
0.0467387 -3.99796
14+
0.0769928 -3.16694
15+
0.0964534 -2.68574
16+
0.146315 -2.05038
17+
0.174528 -1.83081
18+
0.287443 -1.23636
19+
0.317853 -1.12586
20+
0.380854 -0.9266
21+
0.443854 -0.727343
22+
0.509042 -0.593492
23+
0.576404 -0.459546
24+
0.635076 -0.347813
25+
0.698095 -0.235891
26+
0.761123 -0.167637
27+
0.8285 -0.0991921
28+
0.893707 -0.0526765
29+
0.954576 -0.0500185
30+
1.00022 -0.00435728

‎lessons/tutorial_xflr5/images/anderson_horseshoe_vortex.png‎

44.4 KB
Loading

‎lessons/tutorial_xflr5/images/anderson_horseshoe_vortices.png‎

44.7 KB
Loading

0 commit comments

Comments
 (0)