80 lines
2.7 KiB
Python
80 lines
2.7 KiB
Python
import matplotlib.pyplot as plt
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import math
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def raymers_iterate(current_max_weight, payload_weight, fuel_weight, a_constant, c_constant, variable_sweep_constant):
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empty_max_weight_ratio = (a_constant * current_max_weight**c_constant * variable_sweep_constant)
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new_max_weight = payload_weight / (1 - (fuel_weight / current_max_weight) - (empty_max_weight_ratio))
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return new_max_weight
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payload_mass_metric = 125 # Grams
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payload_mass_imperial = payload_mass_metric * 0.002204623 # Converting grams to pounds
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a_constant = 0.86 # Sailplane - Unpowered
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c_constant = -0.05 # Sailplane - Unpowered
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max_weight_guess = 6 # Initial weight guess of 6 lbs
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current_estimate = max_weight_guess
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imperial_estimates = []
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for i in range(0,11):
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current_estimate = raymers_iterate(current_estimate, payload_mass_imperial, 0, a_constant, c_constant, 1)
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imperial_estimates.append(current_estimate)
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metric_estimates = []
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for imperial_estimate in imperial_estimates:
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metric_estimates.append(imperial_estimate / 0.002204623)
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plt.plot(metric_estimates, linestyle='--', marker='o')
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# This section has been commented out because jordan (senft) does not believe the customer wants to see the numbers
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# Add labels to each point
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#for index, weight in enumerate(metric_estimates):
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# plt.text(index, weight + 5, f'{index}: {weight:.2f}', fontsize=9, ha='right')
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plt.title("Aircraft Raymer's Method Weight Estimation")
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plt.xlabel("Iteration Number")
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plt.ylabel("Weight Estimation (g)")
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plt.grid(True)
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#plt.show()
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print("Raymer's Aircraft Calculations (IMPERIAL UNITS)\n")
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print("Weight Estimation: %0.4f lbs" % imperial_estimates[-1])
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# Wing loading
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typical_weight_area_ratio = 6 # Historical sailplane ratio
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print("Typical Weight Area Ratio: %0.4f lbs/ft^2" % typical_weight_area_ratio)
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# To find area of the wings we must work backwards (ratio = weight/area)
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# Area of wings
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wing_area = imperial_estimates[-1] / typical_weight_area_ratio
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print("Main Wing Area: %0.4f ft^2" % wing_area)
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# L/D Ratio
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ld_ratio = 7.5 / 1.5
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print("Lift to Drag Ratio: %0.4f" % ld_ratio)
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# Aspect Ratio (Wing span) - Need L/D
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aspect_ratio = 4.464 * (ld_ratio**0.69)
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print("Aspect Ratio: %0.4f" % aspect_ratio)
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# Wing span
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wing_span = math.sqrt(aspect_ratio*wing_area)
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print("Main Wing Span %0.4f ft" % wing_span)
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# Chord Length
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chord_length = wing_area / wing_span
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print("Main Wing Chord Length %0.4f ft" % chord_length)
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# Fuselage Length
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fuselage_length = 0.86 * imperial_estimates[-1]**0.48
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print("Fuselage Length %0.4f ft" % fuselage_length)
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# Tail Positioning (Moment arm)
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tail_poisitioning = fuselage_length * 0.65
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print("Tail Positioning (Moment arm location) %0.4f ft" % tail_poisitioning)
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# Tail Dimensions
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