import matplotlib.pyplot as plt
import math

def raymers_iterate(current_max_weight, payload_weight, fuel_weight, a_constant, c_constant, variable_sweep_constant):
    empty_max_weight_ratio =  (a_constant * current_max_weight**c_constant * variable_sweep_constant)

    new_max_weight = payload_weight / (1 - (fuel_weight / current_max_weight) - (empty_max_weight_ratio))
    return new_max_weight

payload_mass_metric = 125 # Grams
payload_mass_imperial = payload_mass_metric * 0.002204623 # Converting grams to pounds

a_constant = 0.86 # Sailplane - Unpowered
c_constant = -0.05 # Sailplane - Unpowered

max_weight_guess = 6 # Initial weight guess of 6 lbs
current_estimate = max_weight_guess

imperial_estimates = []

for i in range(0,11):
    current_estimate = raymers_iterate(current_estimate, payload_mass_imperial, 0, a_constant, c_constant, 1)
    imperial_estimates.append(current_estimate)

metric_estimates = []

for imperial_estimate in imperial_estimates:
    metric_estimates.append(imperial_estimate / 0.002204623)

plt.plot(metric_estimates, linestyle='--', marker='o')

# This section has been commented out because jordan (senft) does not believe the customer wants to see the numbers
# Add labels to each point
#for index, weight in enumerate(metric_estimates):
#    plt.text(index, weight + 5, f'{index}: {weight:.2f}', fontsize=9, ha='right')

plt.title("Aircraft Raymer's Method Weight Estimation")
plt.xlabel("Iteration Number")
plt.ylabel("Weight Estimation (g)")
plt.grid(True)
#plt.show()

print("Raymer's Aircraft Calculations (IMPERIAL UNITS)\n")
print("Weight Estimation (mass): %0.4f lbs" % imperial_estimates[-1])

# Wing loading 
typical_weight_area_ratio = 6 # Historical sailplane ratio
print("Typical Weight Area Ratio: %0.4f lbs/ft²" % typical_weight_area_ratio)

# To find area of the wings we must work backwards (ratio = weight/area)
# Area of wings
wing_area = imperial_estimates[-1] / typical_weight_area_ratio
print("Main Wing Area: %0.4f ft²" % wing_area)

# L/D Ratio
ld_ratio = 7.5 / 1.5
print("Lift to Drag Ratio: %0.4f" % ld_ratio)

# Aspect Ratio (Wing span) - Need L/D
aspect_ratio = 4.464 * (ld_ratio**0.69)
print("Aspect Ratio: %0.4f" % aspect_ratio)

# Wing span
wing_span = math.sqrt(aspect_ratio*wing_area)
print("Main Wing Span %0.4f ft" % wing_span)

# Chord Length
chord_length = wing_area / wing_span
print("Main Wing Chord Length %0.4f ft" % chord_length)

# Fuselage Length
fuselage_length = 0.86 * imperial_estimates[-1]**0.48
print("Fuselage Length %0.4f ft" % fuselage_length)

# Tail Positioning (Moment arm)
tail_position = fuselage_length * 0.65
print("Tail Positioning (Moment arm location) %0.4f ft" % tail_position)

# Vertical Tail Area
C_VT = 0.02  # Assumed constant (sail plane, raymers)
vertical_tail_area = C_VT * ((wing_span * wing_area) / tail_position)
print("Vertical Tail Area: {:.4f} ft²".format(vertical_tail_area))

# Horizontal Tail Area
C_HT = 0.5  # Assumed constant (sail plane, raymers)
horizontal_tail_area = C_HT * ((chord_length * wing_area) / tail_position)
print("Horizontal Tail Area: {:.4f} ft²\n".format(horizontal_tail_area))


# Convert everything to metric 

print("End of imperial numbers (THANK THE LORD!)\n")

print("Weight Estimation (mass): %0.4f kg" % (metric_estimates[-1]*(10**-3)))

# Total Lift Force
lift_force_total = (metric_estimates[-1]*(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))
#print("Lift Force per unit length: {:.4f} lbf/ft".format(lift_force_per_ft))
#print("Lift Force per unit length (1 wing): {:.4f} lbf/ft".format(lift_force_per_wing_per_ft))

# Bending Moments

number_of_wings = 2

print("BENDING MOMENTS\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 Stress Calcs

B = 4 * 0.001 # Width in m
D = 4 * 0.001 # Height or depth in m

# 4mm Spars
area_moment_inertia = ((B) * (D**3))/12
print(f"Second moment of inertia for 4mm: {area_moment_inertia} m^4")

max_stress = bending_moment_wing * ((D/2) / area_moment_inertia)
print(f"Max Stress for 4mm spar: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")

B = 6 * 0.001 # Width in m
D = 6 * 0.001 # Height or depth in m

# 6mm Spars
area_moment_inertia = ((B) * (D**3))/12
print(f"Second moment of inertia for 6mm: {area_moment_inertia} m^4")

max_stress = bending_moment_wing * ((D/2) / area_moment_inertia)
print(f"Max Stress for 6mm spar: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")


B = 12 * 0.001 # Width in m
D = 2 * 0.001 # Height or depth in m

# 6mm Spars
area_moment_inertia = ((B) * (D**3))/12
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")