Fuselage Stiffener

Objective

Design a lightweight fuselage with internal stiffeners capable of supporting over 30 lbs in compression.

Utilize trade studies and FEA to optimize the overall strength to weight ratio

Requirements

Fuselage must hold atleast 30 lbs in compression
Each fuselage section must be 3D printed within a 6" x 6" x 6" print bed
The length of all fuselage sections together must equal 36"

Diagram of the 3D print bed helped determine the diameter of each fuselage

Design Considerations Prior to Modeling

Proper stringer alignment is critical—any mismatch can create stress concentrations or joint eccentricity, which would be the primary source of failure.
Stringer spacing must avoid large gaps; excessive spacing increases unsupported skin regions and raises the risk of local buckling.
Larger fuselage diameter increases moment of inertia, improving bending stiffness and load-carrying capability.
Skin distributes loads and maintains geometry, stringers prevent buckling, and joints remain the most critical elements for structural integrity.
Print orientation was selected to maximize load-bearing performance because the material is non-isotropic, with strength higher along layers than across them.

Modeling Preperation

The expression feature in Nx was utilized to build a robust model to control the key dimensions of the stiffener
This minimized the number of features in the design tree and reduced the potential for errors to occur

Design Considerations Prior to FEA

The CAD model was idealized and split in half since the fuselage is symmetric.
This resulted in reduced solving time to run multiple studies while maintaining accuracy of the FEA
The objective was to determine the highest areas of stress and lowest areas of stress to further optimize the fuselage

FEA

Boundary Conditions

Fixed bottom
Symmetry constraint
Load: 44lbs (to simulate 2 x 10kg dumbbells for physical testing)

Identifying the most important parameter:

6 adjustable parameters
5 tests for each parameter
Record the stress/displacement
Record the volume/mass

Key Design Parameters

As fuselage wall thickness increases, both stress and displacement decrease in a clear, nearly linear manner, indicating a strong inverse relationship where thicker walls significantly reduce structural response.
As the hat angle decreases, displacement increases almost linearly while stress shows moderate variability with a slight overall upward trend, peaking near 105°, indicating a stronger correlation between hat angle and displacement than stress.

As hat height increases, displacement decreases steadily and nearly linearly, while stress shows a mild overall decrease with some variability, indicating that increased hat height more effectively reduces displacement than stress.
As hat thickness increases, both stress and displacement decrease in a clear, nearly linear fashion, showing that increasing hat thickness effectively reduces structural loading and deformation.

As hat width increases, displacement decreases nearly linearly, while stress varies more irregularly with only a slight overall decreasing trend, indicating that hat width has a stronger and more consistent effect on displacement than on stress.
As leg width increases, displacement decreases steadily and nearly linearly, while stress varies with a mild overall decreasing trend, indicating leg width has a stronger and more consistent influence on displacement than on stress.

Final Design Decision

Most important design parameters:

Fuselage wall thickness
Stiffener hat thickness

Re-ran three additional tests using different combinations of these two parameters → Test Run 3 fit our criteria the best.

Reason:

lightweight + minimal changes in stress and displacement gave us confidence that the Test Run 3 parameters were the most effective

Final Design Parameters:

Fuselage Wall Thickness = 0.0875”
Hat Angle = 120 degrees
Hat Height = 0.25”
Hat Thickness = 0.0625”
Hat width = 0.3125”
Leg Width = 0.1875”
# of Stiffeners = 6
Outer Diameter = 1.9”

PLA - Material:
- Mass Density (RHO) = 1.21 g/cm^3
- Young’s Modulus (E) = 1860 MPa

Final Design

Segment-Joint Design Investigation

After finalizing the fuselage and stiffener design, we focused on creating an efficient joint to reach the full 36" length.
Ensuring perfect stiffener alignment was the key requirement for proper load transfer and buckling prevention.
We selected a simple transition-fit: small nubs on one segment’s stiffeners mate with the open ends of the next segment.

Results of Compression Load Testing: Trial 1

Displacement increases smoothly along the length of the top hat stiffener, with a maximum nodal displacement of approximately 0.0124 in, indicating global bending-dominated deformation with no abrupt local concentrations.
Von Mises stress (unaveraged) ranges from about 79 psi to 104 psi, showing localized peak stresses along element boundaries, typical of raw finite element results.
Von Mises stress (averaged) smooths these local variations, with stresses ranging from approximately 85 psi to 103 psi, providing a more representative view of the overall stress distribution.
Stress contours are relatively uniform along the stiffener, suggesting no critical stress concentrations and a well-distributed load path.
The close agreement between averaged and unaveraged peak stresses indicates mesh quality is adequate and results are numerically stable.

Results of Compression Load Testing: Trial 2

This trial was intended to replicate how the fuselage would behave when divided into seven segments, matching the way our physical model was assembled.

Used the Split function to divide the fuselage into seven equal-length segments.
Applied MMC to the face edges of the split fuselage by gluing and setting up face contacts, though some edges did not allow MMC to be applied.

Results of Compression Load Testing: Trial 2 (Continued)

The displacement contour shows a smooth, continuous deformation along the full length of the fuselage, with a maximum displacement of approximately 0.0071 in, indicating that the seven split segments behave as a single structural member despite being divided.
Unaveraged von Mises stress ranges from approximately 30 psi to 91 psi, with no sharp stress spikes at the segment interfaces, suggesting effective load transfer across most glued faces.
Averaged stress results further smooth local variations, with stresses ranging from approximately 38 psi to 77 psi, confirming a stable and physically consistent stress distribution.
The absence of visible stress concentrations at the split locations indicates that the applied MMC gluing successfully replicated bonded behavior between the segmented fuselage sections.
Minor discontinuities where MMC could not be applied did not significantly affect global displacement or stress trends, implying the segmented model reasonably represents the behavior of the physically assembled fuselage.
Overall, the results validate the use of the Split function combined with MMC face contacts as an effective method for modeling a multi-segment fuselage assembly in NX.

Results of a Mock Physical Compression Load Testing

For preliminary testing, we conducted an in-home load test: the parts were 3D-printed in PLA, assembled, and subjected to a 44-lb load using two dumbbells.

Summary

Successfully completed an end-to-end workflow including design, simulation, optimization, and physical validation of a lightweight fuselage capable of supporting ≥30 lb within size constraints.
Used parametric modeling, sensitivity studies, and FEA to develop an efficient top-hat–stiffened fuselage with strong structural performance and minimal material usage.
Identified fuselage wall thickness and hat thickness as the most influential parameters on stress and displacement, leading to selection of Test 3 as the optimal configuration.
Designed a modular, segmented fuselage with reliable load transfer enabled by alignment nubs and controlled manufacturing tolerances.
Simulation of both the continuous model and seven-segment assembly confirmed stresses well below PLA yield limits.
Physical testing closely matched simulation results, showing no visible deformation, cracking, or joint instability, validating modeling assumptions and analysis methods.
Overall, delivered a structurally sound, manufacturable, and weight-efficient design, demonstrating a robust methodology applicable to more complex aerospace structures.