Final changes to code, sorry it's messy lol

.gitignore

@@ -164,3 +164,5 @@ cython_debug/
 .DS_Store
 
 Aircraft\ Design\ and\ Performance\ \-\ Assignment\ 1\ \-\ William\ Johnstone.docx
+
+CAD/

README.md

@@ -1,3 +1,5 @@
 # uni-raymers
 
 Code for estimating aircraft characteristics for the aircraft design module.
+
+Output of calculations can be found in output.txt

Raymers Code/raymers.py

@@ -1,3 +1,6 @@
+# William Johnstone
+# Apologies to anyone reading this, it became a mess. I wanted to tidy it but we don't always get what we want.
+
 import matplotlib.pyplot as plt
 import math
 
@@ -146,4 +149,83 @@ print(f"Second moment of inertia for new test: {area_moment_inertia} m^4")
 max_stress = bending_moment_wing * ((D/2) / area_moment_inertia)
 print(f"Max Stress for new test spar: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")
 
+# Aerofoil Spars
 
+B = 18 * 0.001 # Width in m
+D = 3.5 * 0.001 # Height or depth in m
+
+# 6mm Spars
+area_moment_inertia = ((B) * (D**3))/12
+print(f"Second moment of inertia for 6x3mm: {area_moment_inertia} m^4")
+
+max_stress = bending_moment_wing * ((D/2) / area_moment_inertia)
+print(f"Max Stress for 6x3mm: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")
+
+B = 6 * 0.001 # Width in m
+D = 0.5 * 0.001 # Height or depth in m
+
+# 6mm Spars
+area_moment_inertia = ((B) * (D**3))/12
+print(f"Second moment of inertia for 6x3.5mm: {area_moment_inertia} m^4")
+
+max_stress = bending_moment_wing * ((D/2) / area_moment_inertia)
+print(f"Max Stress for 6x3.5mm: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")
+
+
+# Bending Moment for tail (Rohan's numbers )
+bending_moment_wing = (lift_force_total / number_of_wings) * (wing_span_metric / 4)
+print(f"Bending Moment at of 1 Wing from root: {bending_moment_wing:.4f} Nm\n")
+
+
+
+######################################################################
+# 27/11/24
+# Updated stress calcs with new weight (from detailed CAD model, including payload mass)
+
+print("#############################################################")
+print("Everything after this is based on numbers from final CAD model. Thus more relevant")
+
+print("Weight (Final, from CAD Model) (mass): %0.4f kg" % (262*(10**-3)))
+
+# Total Lift Force
+lift_force_total = (262*(10**-3)) * 9.81
+#lift_force_per_ft = lift_force_total / wing_span
+#lift_force_per_wing_per_ft = lift_force_per_ft / 2
+print("Total Lift Force: {:.4f} N\n".format(lift_force_total))
+
+# Bending Moments
+
+number_of_wings = 2
+
+print("BENDING MOMENTS (LATEST, THESE ARE THE ONLY ONES RELEVANT)\n")
+wing_span_metric = wing_span * 0.3048
+print("Wing Span: {:.4f} m".format(wing_span_metric))
+
+# Bending Moment at Wing Root
+bending_moment_wing = (lift_force_total / number_of_wings) * (wing_span_metric / 4)
+print(f"Bending Moment at of 1 Wing from root: {bending_moment_wing:.4f} Nm\n")
+
+# Max bending stress for Balsa (https://woodworkly.com/is-balsa-wood-strong/)
+bending_strength_psi_lower = 2550
+bending_strength_psi_upper = 3170
+bending_strength_megapascal_lower = bending_strength_psi_lower * 0.006894757 # (1 psi = 0.006894757 MPa)
+bending_strength_megapascal_upper = bending_strength_psi_upper * 0.006894757 # (1 psi = 0.006894757 MPa)
+
+# Max Stress Calcs
+
+B = 18 * 0.001 # Width in m
+D = 2.5 * 0.001 # Height or depth in m
+
+# 4mm Spars
+area_moment_inertia = ((B) * (D**3))/12
+print(f"Second moment of inertia for final model: {area_moment_inertia} m^4")
+
+max_stress = bending_moment_wing * ((D/2) / area_moment_inertia)
+
+max_stress_megapascal = max_stress/1000000
+
+print(f"Max Stress for final model spar: {max_stress:.4f} Pa or {max_stress_megapascal:.4f} MPa")
+
+print(f"Balsa Wood Bending Stress Upper Limit: {bending_strength_megapascal_upper:.4f}MPa, Lower Limit: {bending_strength_megapascal_lower:.4f} MPa")
+
+print(f"Safety Factor: {bending_strength_megapascal_lower/max_stress_megapascal:.1f} - {bending_strength_megapascal_upper/max_stress_megapascal:.1f}")

output.txt

@@ -0,0 +1,50 @@
+Raymer's Aircraft Calculations (IMPERIAL UNITS)
+
+Weight Estimation (mass): 1.6966 lbs
+Typical Weight Area Ratio: 6.0000 lbs/ft²
+Main Wing Area: 0.2828 ft²
+Lift to Drag Ratio: 5.0000
+Aspect Ratio: 13.5523
+Main Wing Span 1.9576 ft
+Main Wing Chord Length 0.1444 ft
+Fuselage Length 1.1084 ft
+Tail Positioning (Moment arm location) 0.7205 ft
+Vertical Tail Area: 0.0154 ft²
+Horizontal Tail Area: 0.0283 ft²
+
+End of imperial numbers (THANK THE LORD!)
+
+Weight Estimation (mass): 0.7696 kg
+Total Lift Force: 7.5493 N
+
+BENDING MOMENTS
+
+Wing Span: 0.5967 m
+Bending Moment at of 1 Wing from root: 0.5631 Nm
+
+Second moment of inertia for 4mm: 2.1333333333333337e-11 m^4
+Max Stress for 4mm spar: 52786068.9659 Pa or 52.7861 MPa
+Second moment of inertia for 6mm: 1.08e-10 m^4
+Max Stress for 6mm spar: 15640316.7306 Pa or 15.6403 MPa
+Second moment of inertia for new test: 8e-12 m^4
+Max Stress for new test spar: 70381425.2878 Pa or 70.3814 MPa
+Second moment of inertia for 6x3mm: 6.431250000000002e-11 m^4
+Max Stress for 6x3mm: 15321126.5933 Pa or 15.3211 MPa
+Second moment of inertia for 6x3.5mm: 6.25e-14 m^4
+Max Stress for 6x3.5mm: 2252205609.2100 Pa or 2252.2056 MPa
+Bending Moment at of 1 Wing from root: 0.5631 Nm
+
+#############################################################
+Everything after this is based on numbers from final CAD model. Thus more relevant
+Weight (Final, from CAD Model) (mass): 0.2620 kg
+Total Lift Force: 2.5702 N
+
+BENDING MOMENTS (LATEST, THESE ARE THE ONLY ONES RELEVANT)
+
+Wing Span: 0.5967 m
+Bending Moment at of 1 Wing from root: 0.1917 Nm
+
+Second moment of inertia for final model: 2.3437500000000006e-11 m^4
+Max Stress for final model spar: 10223765.5151 Pa or 10.2238 MPa
+Balsa Wood Bending Stress Upper Limit: 21.8564MPa, Lower Limit: 17.5816 MPa
+Safety Factor: 1.7 - 2.1