*ARCHIVED* development moved to aircraft-studio.

bugfix: negative x-coord error for high-camber cases

Commit
3ce3509ea818c289d45de8264068791de91a1b99
Author
Marius Peter <blendoit@gmail.com>
Author date
Committer
Marius Peter <blendoit@gmail.com>
Committer date
Changed files
creator.py
index 0bfe5292..3749bf7a 100644..100644
@@ -122,7 +122,7 @@
122 122 self.naca_num = int()
123 123 # Mean camber line
124 124 self.x_c = []
125 Removed: self.y_c = []
125 Added: self.z_c = []
126 126
127 127 def add_naca(self, naca_num):
128 128 '''
@@ -148,22 +148,22 @@
148 148
149 149 def get_camber(x):
150 150 '''
151 Removed: Returns camber y-coordinate from 1 'x' along the airfoil chord.
151 Added: Returns camber z-coordinate from 1 'x' along the airfoil chord.
152 152 '''
153 Removed: y_c = float()
153 Added: z_c = float()
154 154 if 0 <= x < p_c:
155 Removed: y_c = (m / (p ** 2)) * (2 * p * (x / self.chord)
155 Added: z_c = (m / (p ** 2)) * (2 * p * (x / self.chord)
156 156 - (x / self.chord) ** 2)
157 157 elif p_c <= x <= self.chord:
158 Removed: y_c = (m / ((1 - p) ** 2)) * ((1 - 2 * p)
158 Added: z_c = (m / ((1 - p) ** 2)) * ((1 - 2 * p)
159 159 + 2 * p * (x / self.chord)
160 160 - (x / self.chord) ** 2)
161 Removed: return (y_c * self.chord)
161 Added: return (z_c * self.chord)
162 162
163 163 def get_thickness(x):
164 Removed: '''
165 Removed: Returns thickness from 1 'x' along the airfoil chord.
166 Removed: '''
164 Added: '''Returns thickness from 1 'x' along the airfoil chord.'''
165 Added:
166 Added: x = 0 if x < 0 else x
167 167 y_t = 5 * t * self.chord * (
168 168 + 0.2969 * sqrt(x / self.chord)
169 169 - 0.1260 * (x / self.chord)
@@ -173,12 +173,12 @@
173 173 return y_t
174 174
175 175 def get_theta(x):
176 Removed: dy_c = float()
176 Added: dz_c = float()
177 177 if 0 <= x < p_c:
178 Removed: dy_c = ((2 * m) / p ** 2) * (p - x / self.chord)
178 Added: dz_c = ((2 * m) / p ** 2) * (p - x / self.chord)
179 179 elif p_c <= x <= self.chord:
180 Removed: dy_c = (2 * m) / ((1 - p) ** 2) * (p - x / self.chord)
181 Removed: theta = atan(dy_c)
180 Added: dz_c = (2 * m) / ((1 - p) ** 2) * (p - x / self.chord)
181 Added: theta = atan(dz_c)
182 182 return theta
183 183
184 184 def get_upper_coord(x):
@@ -204,7 +204,7 @@
204 204 # Generate our airfoil geometry from previous sub-functions.
205 205 for x in x_chord:
206 206 self.x_c.append(x)
207 Removed: self.y_c.append(get_camber(x))
207 Added: self.z_c.append(get_camber(x))
208 208 self.x.append(get_upper_coord(x)[0])
209 209 self.z.append(get_upper_coord(x)[1])
210 210 for x in x_chord_rev:
@@ -361,7 +361,7 @@
361 361 plt.plot(airfoil.chord / 4, 0, '.', color='g',
362 362 markersize=24, label='Quarter-chord')
363 363 # Plot mean camber line
364 Removed: plt.plot(airfoil.x_c, airfoil.y_c, '-.', color='r', linewidth='2',
364 Added: plt.plot(airfoil.x_c, airfoil.z_c, '-.', color='r', linewidth='2',
365 365 label='Mean camber line')
366 366 # Plot airfoil surfaces
367 367 plt.fill(airfoil.x, airfoil.z, color='b', linewidth='1', fill=False)
@@ -374,7 +374,7 @@
374 374 plt.plot(x, y, '-', color='b')
375 375 except AttributeError:
376 376 print('No spars to plot.')
377 Removed: # Plot upper stringers
377 Added: # Plot stringers
378 378 try:
379 379 for _ in range(0, len(airfoil.stringer.x)):
380 380 x = airfoil.stringer.x[_]
@@ -382,11 +382,6 @@
382 382 plt.plot(x, y, '.', color='y', markersize=12)
383 383 except AttributeError:
384 384 print('No stringers to plot.')
385 Removed: # # Plot lower stringers
386 Removed: # for _ in range(0, len(airfoil.stringer.x)):
387 Removed: # x = airfoil.stringer.x[_]
388 Removed: # y = airfoil.stringer.z[_]
389 Removed: # plt.plot(x, y, '.', color='y', markersize=12)
390 385
391 386 # Graph formatting
392 387 plt.xlabel('X axis')
evaluator.py
index 867ede93..9512a4b2 100644..100644
@@ -49,6 +49,8 @@
49 49 self.lift_total = []
50 50 # Drag
51 51 self.drag = []
52 Added: # centroid
53 Added: self.centroid = []
52 54 # Inertia terms:
53 55 # I_x = self.I_[0]
54 56 # I_z = self.I_[1]
@@ -141,19 +143,23 @@
141 143 stringer_area = self.stringer.area
142 144 caps_area = self.spar.cap_area
143 145
144 Removed: spar_x = self.spar.x + self.spar.x
146 Added: caps_x = [value for spar in self.spar.x for value in spar]
147 Added: caps_z = [value for spar in self.spar.z for value in spar]
145 148 stringers_x = self.stringer.x
146 149 stringers_z = self.stringer.z
147 150
148 Removed: denom = float(len(spar_x) * caps_area
149 Removed: + len(stringers_x) * stringer_area)
151 Added: denominator = float(len(caps_x) * caps_area
152 Added: + len(stringers_x) * stringer_area)
150 153
151 Removed: x_ctr = (sum([i * caps_area for i in spar_x[:][0]])
152 Removed: + sum([i * stringer_area for i in stringers_x])) / denom
153 Removed: z_ctr = (sum([i * caps_area for i in spar_x[:][0]])
154 Removed: + sum([i * stringer_area for i in stringers_z])) / denom
155 Removed: return(x_ctr, z_ctr)
154 Added: centroid_x = float(sum([x * caps_area for x in caps_x])
155 Added: + sum([x * stringer_area for x in stringers_x]))
156 Added: centroid_x = centroid_x / denominator
156 157
158 Added: centroid_z = float(sum([z * caps_area for z in caps_z])
159 Added: + sum([z * stringer_area for z in stringers_z]))
160 Added: centroid_z = centroid_z / denominator
161 Added: return(centroid_x, centroid_z)
162 Added:
157 163 def get_inertia_terms(self):
158 164 '''Obtain all inertia terms.'''
159 165
@@ -165,7 +171,6 @@
165 171 z_stringers = self.stringer.z
166 172 x_spars = self.spar.x[:][0] + self.spar.x[:][1]
167 173 z_spars = self.spar.z[:][0] + self.spar.z[:][1]
168 Removed: print(x_spars)
169 174 stringer_count = range(len(x_stringers))
170 175 spar_count = range(len(self.spar.x))
171 176
main.py
index 97cd4927..f82ca89c 100644..100644
@@ -23,9 +23,9 @@
23 23 start_time = time.time()
24 24
25 25 # Airfoil dimensions
26 Removed: NACA_NUM = 4412
27 Removed: CHORD_LENGTH = 133
28 Removed: SEMI_SPAN = 140
26 Added: NACA_NUM = 2412
27 Added: CHORD_LENGTH = 101
28 Added: SEMI_SPAN = 40
29 29
30 30 # Airfoil thickness
31 31 T_UPPER = 0.1
@@ -37,11 +37,11 @@
37 37 STRINGER_MASS = 5 # lbs
38 38
39 39 # Area
40 Removed: SPAR_CAP_AREA = 0.3 # sqin
40 Added: SPAR_CAP_AREA = 0.0 # sqin
41 41 STRINGER_AREA = 0.1 # sqin
42 42
43 43 # Amount of stringers
44 Removed: TOP_STRINGERS = 5
44 Added: TOP_STRINGERS = 3
45 45 BOTTOM_STRINGERS = 4
46 46 NOSE_TOP_STRINGERS = 3
47 47 NOSE_BOTTOM_STRINGERS = 6
@@ -71,7 +71,7 @@
71 71 af.add_naca(NACA_NUM)
72 72 af.add_mass(AIRFOIL_MASS)
73 73 # af.info_print(2)
74 Removed: # af.info_save(SAVE_PATH, _)
74 Added: af.info_save(SAVE_PATH, _)
75 75
76 76 # Create spar instance
77 77 af.spar = creator.Spar()
@@ -82,7 +82,7 @@
82 82 af.spar.add_spar_caps(SPAR_CAP_AREA)
83 83 af.spar.add_mass(SPAR_MASS)
84 84 # af.spar.info_print(2)
85 Removed: # af.spar.info_save(SAVE_PATH, _)
85 Added: af.spar.info_save(SAVE_PATH, _)
86 86
87 87 # Create stringer instance
88 88 af.stringer = creator.Stringer()
@@ -95,7 +95,7 @@
95 95 af.stringer.add_area(STRINGER_AREA)
96 96 af.stringer.add_mass(STRINGER_MASS)
97 97 # af.stringer.info_print(2)
98 Removed: # af.stringer.info_save(SAVE_PATH, _)
98 Added: af.stringer.info_save(SAVE_PATH, _)
99 99
100 100 # Plot components with matplotlib
101 101 # creator.plot_geom(af)
@@ -106,7 +106,7 @@
106 106 eval.analysis()
107 107 # eval.info_print(2)
108 108 eval.info_save(SAVE_PATH, _)
109 Removed: evaluator.plot_geom(eval)
109 Added: # evaluator.plot_geom(eval)
110 110 # evaluator.plot_lift(eval)
111 111
112 112 # Print final execution time