diff --git a/Raymers Code/raymers.py b/Raymers Code/raymers.py
index 7586892..154e011 100644
--- a/Raymers Code/raymers.py	
+++ b/Raymers Code/raymers.py	
@@ -41,16 +41,16 @@ plt.grid(True)
 #plt.show()
 
 print("Raymer's Aircraft Calculations (IMPERIAL UNITS)\n")
-print("Weight Estimation: %0.4f lbs" % imperial_estimates[-1])
+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/ft^2" % typical_weight_area_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 ft^2" % wing_area)
+print("Main Wing Area: %0.4f ft²" % wing_area)
 
 # L/D Ratio
 ld_ratio = 7.5 / 1.5
@@ -73,8 +73,58 @@ fuselage_length = 0.86 * imperial_estimates[-1]**0.48
 print("Fuselage Length %0.4f ft" % fuselage_length)
 
 # Tail Positioning (Moment arm)
-tail_poisitioning = fuselage_length * 0.65
-print("Tail Positioning (Moment arm location) %0.4f ft" % tail_poisitioning)
+tail_position = fuselage_length * 0.65
+print("Tail Positioning (Moment arm location) %0.4f ft" % tail_position)
 
-# Tail Dimensions
+# 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
+
+# 4mm Spars
+area_moment_inertia = ((0.004) * (0.004**3))/12
+print(f"Second moment of inertia for 4mm: {area_moment_inertia} m^4")
+
+max_stress = bending_moment_wing * ((2*(10**-3)) / area_moment_inertia)
+print(f"Max Stress for 4mm spar: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")
+
+# 6mm Spars
+area_moment_inertia = ((0.006) * (0.006**3))/12
+print(f"Second moment of inertia for 6mm: {area_moment_inertia} m^4")
+
+max_stress = bending_moment_wing * ((3*(10**-3)) / area_moment_inertia)
+print(f"Max Stress for 6mm spar: {max_stress:.4f} Pa or {max_stress/1000000:.4f} MPa")