Mechanical and thermal performance of magnesium carbon fiber sandwich composites with variable fiber orientations for aerospace structures

M E Annadorai; M Ramakrishna

doi:10.1038/s41598-026-38567-8

. 2026 Feb 7;16:7710. doi: 10.1038/s41598-026-38567-8

Mechanical and thermal performance of magnesium carbon fiber sandwich composites with variable fiber orientations for aerospace structures

M E Annadorai ^1,^✉, M Ramakrishna ¹

PMCID: PMC12946178 PMID: 41654620

Abstract

Sandwich composites are broadly recognised for their excellent stiffness, strength, and durability while minimising their weight. This study investigates the fabrication, mechanical testing, thermal analysis, and microstructural characterisation of magnesium-carbon fiber sandwich composites for aircraft applications. The composite panels were fabricated using traditional methods such as filament drum winding, followed by hand layup and compression moulding with six stacking sequences, including unidirectional ([0°], [45°], [90°]), bidirectional ([0°/90°], [45°/-45°]), and multidirectional ([0°/90°/45°/-45°]) fiber orientations. X-ray diffraction (XRD) investigated phase composition and crystalline structure. Thermogravimetric analysis (TGA) revealed two-stage degradation, with [0°/90°]₄ retaining the highest residue (73–75%), followed by the quasi-isotropic laminate (69–71%), while [90°]₈ showed the lowest stability (51–53%). Differential Scanning Calorimetry (DSC) confirmed decomposition onset between 250 °C and 380 °C, with balanced laminates exhibiting reduced endothermic peaks. Mechanical testing revealed strong orientation dependence, with the [0°]₈ laminate exhibiting the highest tensile (223.8 MPa) and flexural strength (700.9 MPa), while the [90°]₈ configuration showed the lowest values. Impact behaviour contrasted with tensile and flexural results, where the [45°/-45°]₄ laminate absorbed the highest impact energy (0.42 kg·m; 101.43 kJ/m²). The quasi-isotropic stacking sequence Mg/CF[0°/90°/45°/-45°]2/Mg exhibited a well-balanced multifunctional response, attaining the third-highest tensile strength, second-highest flexural strength, and second-highest impact energy absorption among the evaluated laminates, while also demonstrating enhanced thermal stability as evidenced by higher TGA residual mass and more stable DSC transition behaviour. Microstructural analysis (SEM/EDS) revealed elemental composition and microstructural features, including fiber-matrix bonding, delamination, and fracture morphology. Results showed substantial differences in mechanical and thermal performance across fiber orientations, aiding in the design of magnesium-carbon fiber sandwich composites for lightweight, high-performance aerospace applications.

Keywords: Sandwich composite, Magnesium alloy, Carbon fiber, Fiber orientation, Aerospace application

Subject terms: Engineering, Materials science

Introduction

The aerospace sector demands advanced materials capable of lightweight structures and also delivering high specific strength, structural efficiency, and long-term reliability under extreme thermo-mechanical environments. Sandwich composites, composed of stiff and strong face sheets bonded to lightweight cores, have therefore become essential structural systems due to their superior stiffness-to-weight ratios, enhanced energy absorption, and excellent damage tolerance¹. These characteristics make them suitable for critical aerospace components such as fuselage skins, wing structures, control surfaces, radomes, and spacecraft panels. Their overall performance is significantly influenced by the choice of face-sheet and core materials, with polymer foams, Nomex honeycomb, and metallic foams enhancing load distribution, bending stiffness, and crashworthiness in demanding aerospace applications².

Magnesium and its alloys are particularly attractive due to their extremely low density, high specific stiffness, favourable damping capacity, high thermal conductivity, and inherent electromagnetic shielding capability. These properties render magnesium alloys promising candidates for next-generation lightweight aircraft structures where reductions in mass directly enhance fuel efficiency and payload capability^3,4. However, monolithic magnesium has limitations such as modest strength, limited ductility, and sensitivity to corrosion. To address these constraints, several alloy systems, such as the AZ (Mg-Al-Zn), AM (Mg-Al-Mn), AS (Mg-Al-Si), WE (Mg-rare earth materials), and ZK (Mg-Zn-Zr) series, have been developed, with AZ alloys gaining prominence due to their balanced strength, corrosion resistance, and formability achieved through controlled alloying and thermomechanical processing⁵. Carbon fibers, known for their exceptional tensile strength, stiffness, low density, and superior resistance to environmental factors, are suitable as reinforcement in aerospace composites. Their excellent thermal stability further enables aerospace components to withstand extreme temperature fluctuations encountered during flight and space missions. When integrated with magnesium, carbon fibers provide a synergistic combination of weight reduction, mechanical robustness, and thermal stability^6,7. Fiber orientation plays a critical role in determining this mechanical behaviour, as variations in alignment significantly affect stiffness anisotropy, interlaminar shear strength, and damage evolution patterns⁸. Beyond mechanical performance, fiber orientation significantly influences thermal conduction pathways and residual stress evolution in multi-layered systems. These effects are particularly relevant in aerospace structures exposed to fluctuating thermal loads or multidirectional mechanical stresses⁹. Prior studies have demonstrated that orientation-dependent mechanical and thermal responses, making optimised stacking essential for structural applications such as aerospace and automotive.

Mohamed and Abdelbary evaluated unidirectional carbon-fibre laminates at 0°, 30°, 45°, 60°, and 90° and found that tensile performance is strongly orientation-dependent, with the 0° configuration providing the highest stiffness and failure resistance¹⁰. Olhan et al. reported that three-dimensional (3D) woven fabric composites exhibited significantly higher energy absorption and deformation resistance compared with unidirectional (UD) and bidirectional (2D) laminates across all evaluated impact energy levels¹¹. Extending to multidirectional systems, Parmiggiani et al. studied how variations in fiber alignment significantly influence mechanical properties such as tensile and flexural properties¹². Hao et al. reported that aligned fibres enhance heat conduction along the fiber axis, making them suitable for aerospace structures requiring efficient thermal management¹³. Similarly, Kim and Kim demonstrated that multidirectional fiber orientations provide balanced thermal expansion properties, reducing warping under temperature fluctuations¹⁴.

Recent advancements in composite material science further emphasise the importance of developing multifunctional, damage-resistant, and structurally repairable hybrid systems for both aerospace and automotive sectors¹⁵. Studies employing multiple reinforcements such as basalt, E-glass, and carbon fibers integrated into metallic substrates through techniques like friction stir processing have shown substantial improvements in tensile, flexural, and impact behaviour compared to monolithic metals¹⁶. Simultaneously, advancements in solid-state processing and additive manufacturing have enabled the fabrication of defect-free metal-composite hybrids with refined microstructures and superior mechanical stability, surpassing conventional fusion-based approaches¹⁷. In the context of energy-absorbing structures, modern crash box research demonstrates that metal-composite hybrid configurations substantially enhance impact resistance, energy absorption, and post-impact stability capabilities aligned with the requirements of automotive and aerospace sandwich structures designed to withstand crash and debris impacts¹⁸. Cryogenic treatment of metal-composite hybrids enhances microstructure and residual stress, improving hardness, wear resistance, fatigue life, and low-temperature performance for aerospace applications¹⁹. Additionally, recent aerospace materials selection studies highlight a continued shift toward lightweight metal-composite systems due to their superior strength-to-weight efficiency, thermal resistance, and durability, reinforcing their relevance in advanced aircraft design and next-generation sandwich architectures²⁰.

Despite this growing body of research, systematic studies on magnesium-carbon fiber sandwich composites remain scarce, particularly those examining how fiber orientation influences multifunctional behaviour for structural applications. Existing investigations predominantly focus on polymeric sandwich cores or monolithic magnesium alloys, leaving significant gaps in understanding the interplay between carbon fiber reinforcement and magnesium face sheets within a multi-layered sandwich architecture. Challenges related to optimising interfacial bonding between metal and composite layers, mitigating galvanic interactions between magnesium and carbon fibers, and correlating mechanical-thermal performance with underlying microstructural features also remain insufficiently addressed in the current literature.

To address these critical gaps, the present study fabricates magnesium-carbon fiber sandwich composites using a combined manufacturing strategy involving filament drum winding, hand layup, and compression moulding, incorporating six distinct stacking sequences, such as unidirectional, bidirectional, and quasi-isotropic, to assess their orientation-dependent mechanical, thermal, and microstructural behaviour. Comprehensive tensile, flexural, and impact testing is complemented by thermal analyses using TGA and DSC and detailed microstructural investigations through SEM, EDS, and XRD. This integrated approach enables the establishment of structure-property relationships that clarify how fiber orientation governs the multifunctional performance of magnesium-CFRP sandwich systems.

Materials & methodology

Materials

AZ31 magnesium alloy (face sheet)

AZ31 magnesium alloy exhibits good ductility, impact resistance, and thermal conductivity, making it well suited for lightweight composite structures. Its nominal composition of approximately 3 wt% Al, 1 wt% Zn, and the balance Mg provides an effective combination of low density, adequate mechanical strength, corrosion resistance, and weldability²¹. These attributes have led to its established use in lightweight structural and aerospace-related applications where mass reduction is critical. Compared with higher-alloyed magnesium systems such as the WE or ZK series, AZ31 offers superior manufacturability, wider commercial availability, and lower material and processing costs, while still providing sufficient stiffness and durability for sandwich face-sheet applications. Accordingly, commercially available AZ31B magnesium alloy was selected as the face-sheet material in the present study to ensure practical relevance, mechanical reliability, experimental reproducibility, and sustainability in lightweight sandwich composite structures²².

Carbon fiber roving 3 K yarn (core sheet & reinforcement)

Continuous carbon fiber roving with a 3 K filament count was chosen as the core reinforcement owing to its high tensile strength, low density, excellent thermal stability, and resistance to chemical and environmental degradation²³. The 3 K roving yarn, consisting of approximately 3,000 filaments per tow, is widely employed in aerospace-grade CFRP systems, as it provides an optimal balance between mechanical performance, flexibility, and efficient resin impregnation. Compared with higher tow-count fibers (e.g., 6–12 K), 3 K yarn enables improved resin wetting and reduced void content, which is critical for achieving reliable interfacial bonding in metal-composite sandwich architectures. In this study, carbon fiber core sheets were fabricated using 3 K roving in combination with an epoxy matrix, ensuring uniform fiber distribution, effective load transfer, and structural continuity²⁴. This selection represents a practical compromise between stiffness, damage tolerance, and manufacturability for aerospace-orientated sandwich composites.

Epoxy resin (matrix)

An epoxy resin system was employed as the matrix material due to its excellent mechanical properties, strong adhesion to carbon fibers, chemical resistance, and dimensional stability, which are essential for structural composite applications²⁵. In this study, LY 556 epoxy resin combined with Aradur 5200 hardener was selected for its low viscosity, well-defined curing behaviour, and compatibility with filament winding, hand layup, and compression moulding processes. This resin system is widely used in aerospace and high-performance composite research, offering a balanced combination of stiffness, toughness, damping capacity, and thermal stability supported by established long-term service data²⁶. Although alternative matrix systems such as bismaleimide or cyanate ester resins provide higher service-temperature capability, they involve increased processing complexity and cost. Therefore, the LY556/Aradur 5200 system was adopted as a practical and aerospace-relevant matrix for magnesium-carbon fiber sandwich composites.

Overall, the selected AZ31 sheet, 3 K carbon fiber, and LY556/Aradur 5200 material provide a pragmatic and aerospace-relevant balance of mechanical performance, damping behaviour, thermal stability, manufacturability, and availability of service data. The results of this study establish a robust baseline framework against which future investigations employing higher-performance magnesium alloys, surface-engineered face sheets, or advanced high-temperature resin systems can be benchmarked to further optimise stiffness and durability. Figure 1 illustrates the working materials used to fabricate the Mg-CF sandwich composites.

Fig. 1 — Materials used to fabricate magnesium-carbon fiber sandwich composite.

Preliminary characterization of base materials

To ensure the suitability of the raw materials used in fabricating the magnesium carbon fiber-based sandwich composite, initial mechanical and microstructural tests were conducted on the AZ31 magnesium alloy sheet and carbon fiber 3 K yarn. Tensile tests were carried out according to relevant ASTM standards. The AZ31 magnesium alloy sheet exhibited a peak tensile load of 1080.5 N, corresponding to a tensile strength of 172.88 MPa, confirming its ductility and suitability as a face sheet material. The carbon fiber roving 3 K yarn reached a peak tensile load of 396 N, resulting in a tensile strength of 3429.2 MPa, which highlights its excellent load-carrying capacity. The load vs. elongation behaviour for both materials is illustrated in Fig. 2, providing a visual comparison of their tensile responses.

Fig. 2 — Load vs. elongation curves for AZ31 sheet and carbon fiber under tensile loading.

Elemental analysis using Energy Dispersive X-ray Spectroscopy (EDS) was also conducted. The AZ31 sheet contained magnesium as the major element, along with minor quantities of aluminium, zinc, and other materials. The carbon fiber yarn showed a dominant carbon peak, with minor oxygen content likely due to surface oxidation or residual epoxy.

Figure 3 shows representative EDS spectra for the AZ31 magnesium sheet and carbon fiber roving 3 K yarn, respectively. And the properties of working materials are tabulated in Tables 1 and 2, and 3.

Fig. 3 — EDS spectra images of AZ31 sheet and carbon fiber yarn with elemental composition.

Table 1.

Properties of AZ31 magnesium alloy Sheet.

Material type	Material properties
Magnesium Alloy (AZ31B)	Tensile strength	172.88 MPa
	Elastic modulus	5472.97 MPa
	Yield Strength	140.08 MPa
	Peak Load	1080.5 N
	Peak Elongation	2.15 mm
	Density	1.77 g/cm³
	Poisson’s ratio	0.35
	Thermal conductivity	96.0 W/m-K
	Melting point	605–630 °C

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Table 2.

Properties of carbon fiber roving 3 K yarn.

Material type	Material properties
Carbon fiber roving (3 K Yarn)	Tensile strength	3429.2 MPa
	Tensile modulus	230 GPa
	Peak Load	396 N
	Elongation at break	1.74%
	Fiber diameter	7 μm
	Density	1.79 g/cm³
	Thermal conductivity	10–160 W/m K

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Table 3.

Properties of epoxy resin and hardener.

Material type	Material properties
Epoxy LY556 (Resin)	Density	1.15–1.20 g/cm³ at 25 °C
	Viscosity	10,000–12,500 mPa·s at 25 °C
	Gel time	80 °C for 2 h
	Curing temperature	120 °C for 4 h
Aradur 5200 (hardener)	Density	1.20 g/cm³ at 25 °C
	Viscosity	500-1,000 mPa·s at 25 °C
	Pot life	2–4 h (depends on temp)
	Mixture ratio (Resin: Hardener)	100:23

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Methodology

The present work fabricates magnesium-carbon fiber sandwich composites using a combined manufacturing strategy involving filament drum winding, hand layup, and compression moulding, selected to ensure uniform fiber alignment, controlled resin distribution, and high-quality laminate consolidation. This hybrid fabrication strategy enabled the production of high-quality, flat sandwich panels with consistent microstructural integrity across all stacking sequences, thereby ensuring that the observed differences in mechanical and thermal performance could be attributed primarily to fiber orientation rather than processing-induced variability.

Filament drum winding process

Filament drum winding was employed as the initial step to produce continuous, well-tensioned carbon-fiber sheets with controlled fiber orientation. Although it is conventionally used for axisymmetric components, its drum-based configuration enables the fabrication of flat laminates with superior fiber alignment accuracy compared to manual placement techniques²⁷. In this process, continuous carbon-fiber yarns were impregnated with the epoxy resin system and wound under controlled tension onto a rotating cylindrical mandrel coated with a suitable release agent. After winding, the laminates were allowed to cure at room temperature to form a preliminary sheet structure. Following initial curing, the circular composite sheets were carefully cut into the required orientations (such as 0°, 90°, and ± 45°) and dimensions as per design requirements. These trimmed sheets were then subjected to compression moulding, where they were cured at an elevated temperature of 80 °C for 2 h to enhance consolidation, resin flow, and final strength²⁸. The thickness of the fabricated composite sheet was measured to be approximately 0.22 mm, ensuring precision and consistency for further processing in the sandwich composite structure. Figure 4 illustrates the preparation of carbon fiber sheets using the filament drum winding process, ensuring uniform fiber alignment and controlled orientation during laminate fabrication.

Fig. 4 — Preparation of carbon fiber sheet using filament drum winding process.

Hand layup process

Hand layup was employed as an intermediate processing step to integrate the magnesium face sheets with the fiber-reinforced core in a controlled and reproducible manner using a mould²⁹. AZ31 magnesium alloy sheets, each measuring 295 × 295 × 0.5 mm, were placed as the top and bottom face sheets, while a carbon fiber reinforced epoxy laminate served as the core. Six different types of stacking sequences were used, each with eight layers, resulting in an overall core sheet thickness of approximately 2.1 mm. The stacking configurations included unidirectional, bidirectional, and multidirectional orientations. An epoxy resin system was applied between layers as a binding agent. Brushes and rollers were used to uniformly spread the resin and remove any trapped air, ensuring proper adhesion and minimising the formation of voids or delamination between plies. After stacking the core and positioning the magnesium face sheets, the final thickness of the sandwich laminates was approximately 3.2 mm, resulting in lightweight and structurally robust composite laminates. Figure 5 shows the hand layup process used for fabricating the sandwich composite.

Fig. 5 — Hand layup process of Mg-CF sandwich composite using a mould.

Compression moulding process

Compression moulding was employed as the final consolidation step to achieve precise thickness control, enhanced fiber-matrix interfacial bonding, and reduced void content in the sandwich laminate³⁰. After the hand layup process, the mould assembly containing the magnesium face sheets and carbon fiber core was placed in a compression moulding machine, where controlled pressure and temperature were applied to complete consolidation and curing of the composite. To ensure uniform bonding and structural integrity, a constant pressure of 35 bar was applied, while the temperature was maintained at 150 °C for 3 h³¹. The elevated temperature and pressure ensured that the epoxy resin fully permeated the fiber interfaces, eliminating any voids and achieving superior interlaminar bonding. After heating, the mould was cooled to room temperature to prevent thermal shock, then opened to demould the magnesium-carbon fiber sandwich laminate. The resulting laminate exhibited a uniform dimension of approximately 295 × 295 × 3.2 mm, with excellent surface finish and structural consistency, ready for further testing and characterisation. Figure 6 shows the compression moulding setup and fabrication process of sandwich composites.

Fig. 6 — Compression moulding process for fabricating Mg-CF sandwich composite.

This study looks at various designs for composite materials based on classical laminate plate theory (CLPT), including unidirectional, bidirectional, and multidirectional layers. Figures 7 and 8 show the six combinations of fabricated Mg-CF sandwich laminates. And the list of orientations, stacking sequences, and configuration details is shown in Table 4.

Fig. 7 — Prepared Mg-CF sandwich laminates with the combination of six orientations.

Fig. 8 — Schematic image of Mg-CF sandwich laminates with six orientation configuration.

Table 4.

Stacking sequence and configuration details of prepared Mg-CF sandwich composite laminates.

s. no	Laminate configuration	Fiber orientations	Stacking sequences
1	Unidirectional	0°	Mg/ CF [0°]₈/ Mg
2		90°	Mg/ CF [90°]₈/ Mg
3		45°	Mg/ CF [45°]₈/ Mg
4	Bidirectional	0°/90°	Mg/ CF [0°/90°]₄/ Mg
5	Bidirectional	45°/-45°	Mg/ CF [45°/-45°]₄/ Mg
6	Multidirectional	0°/90°/45°/-45°	Mg/ CF [0°/90°/45°/-45°]₂/ Mg

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Six stacking sequences were deliberately selected to capture the spectrum of anisotropic and quasi-isotropic behaviours commonly encountered in aerospace structural panels. The unidirectional laminates ([0°]₈ and [90°]₈ represent load paths dominated by axial tension and bending, typical of wing skins and stiffened fuselage structures. The cross-ply ([0°/90°]₄) and angle-ply ([45/-45°]₄) configurations were chosen to simulate combined axial-transverse and shear-dominated responses relevant to control surfaces, ribs, and torsion-loaded components. The quasi-isotropic laminate ([0°/90°/45°/-45°]₂) was designed to approximate in-plane isotropy, representative of structural regions subjected to multi-axial stress states such as cut-outs, joints, and stiffener run-outs^32,33. Collectively, these six layups provide a realistic yet experimentally tractable framework for systematically evaluating orientation-dependent structure-property relationships in aerospace sandwich composites.

Sample Preparation

After fabrication, the sandwich composite laminates were machined into standard test specimens using abrasive water jet machining (AWJM) in accordance with ASTM guidelines for tensile (ASTM D3039), flexural (ASTM D790), and impact (ASTM D256) testing. The machining was carried out using an S3015 model system equipped with a cutting bed of 3 m × 1.5 m. Garnet abrasive of 80 mesh size was employed at a controlled flow rate of 400 g/min, with the cutting pressure maintained at a maximum of 4200 bar. A carbide nozzle and diamond orifice were used to ensure dimensional accuracy and stable jet formation during cutting. The traverse speed was fixed at 700 mm/min, and the stand-off distance (SOD) was maintained within the range of 0.5–3 mm to minimise kerf taper and edge irregularities.

Although AWJM involves high-pressure material removal, the selected cutting parameters were carefully optimised to preserve edge quality and prevent damage to the layered architecture of the Mg-CF sandwich composites³⁴. The combination of moderate abrasive size, controlled abrasive flow rate, and optimised traverse speed limited excessive erosion, fibre pull-out, and matrix microcracking at the cut edges. Maintaining a small stand-off distance ensured jet stability and minimised local fibre deflection or misalignment, which are known to adversely affect tensile and impact properties when edge defects are present. As a non-thermal process, AWJM eliminated the formation of heat-affected zones, residual thermal stresses, or matrix softening that could otherwise influence mechanical test results³⁵. Figure 9 illustrates the preparation of test specimens using abrasive waterjet machining (AWJM).

Fig. 9 — Testing sample preparation using abrasive water jet machining.

Visual inspection and optical microscopy of representative specimen edges confirmed the absence of macroscopic delamination, edge chipping, or severe fibre distortion. Any minor surface roughness induced during cutting was consistent across all specimens and laminate configurations. Since identical machining parameters were used for all samples, the influence of edge-related effects on tensile, flexural, and impact results remained systematic and did not bias the comparative evaluation of different stacking sequences.

Interface preparation and adhesion considerations

The integrity of the interface between the AZ31 magnesium face sheets and the carbon fiber-epoxy core is critical for the structural performance and durability of sandwich panels in aerospace service. In this study, surface preparation of the magnesium face sheets was carefully controlled to promote effective interfacial bonding. Prior to layup, the AZ31 surfaces underwent degreasing and light abrasion to increase surface roughness and enhance epoxy wettability. Although no primers or coupling agents were employed, the inherent adhesion characteristics of the LY556/Aradur 5200 epoxy system, combined with surface roughening, were sufficient to achieve consistent bonding, as evidenced by the absence of interfacial delamination during machining and mechanical testing.

The cure schedule was designed to ensure effective crosslinking while minimising residual stresses arising from thermal expansion mismatch between magnesium and CFRP. Controlled heating and cooling during compression moulding limited interfacial stress accumulation, contributing to stable adhesion under the investigated loading conditions³⁶. Dedicated interfacial tests such as peel or lap shear were not performed, as the focus of the present study was on global thermo-mechanical performance; however, the absence of premature face-sheet/core separation and consistent failure modes across all mechanical tests indicate that interfacial adhesion was not the limiting factor. Future work will address quantitative interfacial strength evaluation and durability under extended thermal and environmental cycling.

Testing and characterization

Mechanical testing

Mechanical testing of the fabricated magnesium-carbon fiber sandwich composite samples was carried out to evaluate their tensile, flexural, and impact properties. Tensile and flexural tests were performed using a computerised electronic tensometer (Model PC-2000), while the impact test was carried out using Izod KI-1.4 impact testing equipment³⁷. Figure 10 shows the testing arrangements that are used for mechanical test characterisation.

Fig. 10 — Experimental equipment setup for tensile, flexural, and impact testing.

Thermal analysis

Thermal characterisation of the magnesium-carbon fiber sandwich composites was performed using the Thermal Analysis System STA7200 (HITACHI) for Thermogravimetric Analysis (TGA) and the DSC7200 (HITACHI) for Differential Scanning Calorimetry (DSC). TGA measured weight change with temperature, while DSC evaluated heat flow during thermal transitions, providing insights into the composites’ thermal stability and phase behaviour³⁸. Figure 11 shows the equipment setup for thermal studies.

Fig. 11 — Thermal analysis systems: (a) STA7200 for Thermogravimetric Analysis (TGA), (b) DSC7200 for Differential Scanning Calorimetry (DSC).

Microstructural characterization

Microstructural analysis of the fabricated composites was conducted using the Rigaku MiniFlex 600 X-ray Diffractometer for precise phase identification and crystallinity assessment and the TESCAN VEGA3 SBH Scanning Electron Microscope (SEM) with integrated Energy-Dispersive X-ray Spectroscopy (EDS) for detailed surface morphology and elemental composition analysis³⁹. Figure 12 shows the XRD and SEM-EDS equipment used for material characterisation.

Fig. 12 — Experimental Equipment Setup Showing (a) X-ray Diffractometer (XRD) and (b) Scanning Electron Microscope (SEM) with Energy-Dispersive X-ray Spectroscopy (EDS).

Results and discussion

X-ray diffraction (XRD) analysis

To investigate the phase composition and structural characteristics of the fabricated magnesium-carbon fiber sandwich composite laminates, X-ray diffraction (XRD) analysis was performed on both powdered and solid-state specimens. Two complementary approaches were adopted: powder XRD, conducted on finely ground composite material to represent the bulk phase composition, and solid-state XRD, performed directly on the laminate surface to assess the as-fabricated texture without destructive preparation. All XRD measurements were carried out using a Rigaku MiniFlex 300/600 diffractometer, operated at 40 kV and 15 mA with a Cu-Kα radiation source (λ = 1.5418 Å), in θ-2θ scan mode, over the range of 10° to 90°, with a step size of 0.02° and a scan speed of 10°/min.

Powder XRD analysis

Powder X-ray diffraction (XRD) was conducted on a finely ground magnesium-carbon fiber (Mg-CF) sandwich composite to analyse the bulk crystalline phases. The diffractogram (Fig. 13) displays sharp, well-defined peaks corresponding to the hexagonal close-packed (HCP) crystal structure of AZ31 magnesium, with dominant reflections observed at approximately 20°, 25.9°, 32.3°, 34.6°, 36.78°, 42.66°, 48.08°, 57.6°, 63.42°, and 68.86° (2θ). The most intense peak, located at 36.78° (2θ) with a measured intensity of 270 counts, corresponds to the (101) plane of Mg according to JCPDS. The absence of peaks related to MgO, intermetallic compounds, or crystalline carbon confirms that no secondary crystalline phases were formed during fabrication, indicating high phase purity, high crystallinity, and no undesirable phase transformations in the composite.

Fig. 13 — Representative powder XRD pattern of Mg-CF sandwich composite.

Solid-state XRD analysis

Solid-state X-ray diffraction (XRD) was performed directly on the intact laminate surface of the magnesium-carbon fiber (Mg-CF) sandwich composite to evaluate the crystallographic condition of the AZ31 Mg face sheets after fabrication. The diffractogram (Fig. 14) displays sharp peaks corresponding to the hexagonal close-packed (HCP) crystal structure of AZ31 magnesium, with the most intense reflection at 35.06° (2θ) with a measured intensity of 1718 counts, attributed to the (101) plane, in agreement with JCPDS. No secondary crystalline phases such as MgO, intermetallics, or crystalline carbon were detected. All six laminate configurations exhibited identical peak positions and relative intensities; therefore, only one representative pattern is presented for clarity. These findings confirm that the Mg-CF sandwich composite fabrication process preserved the original hexagonal close-packed (HCP) crystal structure of AZ31 magnesium without inducing secondary phases, thereby ensuring phase stability and structural integrity. The retention of the HCP structure is consistent with previous studies on AZ31 processed below recrystallisation temperatures, where compression moulding primarily introduces residual stresses rather than crystallographic phase transformations⁴⁰.

Fig. 14 — Representative solid XRD pattern of Mg-CF sandwich composite.

It is noted, however, that conventional θ-2θ XRD provides averaged crystallographic information and may not capture localised residual stresses or subtle texture evolution arising from the combined thermal and pressure history during moulding. Magnesium alloys are particularly sensitive to thermomechanical processing, and such grain-scale effects may occur without producing detectable peak shifts. Advanced techniques such as EBSD or microhardness mapping could offer further insight into local texture variations and stress gradients, particularly near the metal-composite interface⁴¹. While these analyses were beyond the scope of the present study, the absence of peak broadening, shifts, or secondary phases indicates that the fabrication route preserved the overall crystallographic integrity of the AZ31 face sheets.

Energy dispersive X-ray spectroscopy (EDS) analysis

To verify elemental composition and consistency across the fabricated magnesium-carbon fiber sandwich composite laminates, Energy Dispersive X-ray Spectroscopy (EDS) was performed. The EDS scans, executed under high-vacuum conditions, targeted areas representing both magnesium face sheets and carbon fiber core regions. All six samples showed consistent spectra, with magnesium (Mg) as the dominant element, confirming the AZ31 magnesium alloy in the face sheets. Prominent carbon (C) peaks corresponded to the carbon fiber core reinforcement. Aluminium (Al) and zinc (Zn), key alloying elements in AZ31, were present in minor proportions. Oxygen (O) appeared in all spectra, likely from the surface oxide layer or epoxy resin matrix. Trace elements such as iron (Fe), manganese (Mn), calcium (Ca), silicon (Si), nickel (Ni), and copper (Cu) were detected at lower intensities, possibly from impurities, surface contamination, tool interactions during fabrication, or the resin system. Figure 15 presents a representative EDS spectrum of an Mg-CF composite sample.

Fig. 15 — EDS Spectra of a magnesium-carbon fiber sandwich composite laminate.

Thermal studies

Thermogravimetric analysis (TGA)

TGA was performed on the Mg-CF sandwich composites to study thermal stability. Small rectangular pieces of the composites with an approximate weight of 15 mg were used, balancing sufficient signal with uniform heating. The AZ31 Mg sheet showed negligible weight loss up to 750 °C, confirming that the CFRP core primarily governs the degradation. All samples exhibited a two-stage degradation profile: minor weight loss below 150 °C, due to evaporation of moisture and volatiles, followed by major loss between 350 °C and 500 °C, corresponding to epoxy matrix decomposition and fiber-matrix interface breakdown. Beyond 600 °C, the weight stabilised due to char formation and intact magnesium sheets.

Among the laminates, the [0°/90°]₄ showed the best stability with approximately 73–75% residue, owing to cross-ply reinforcement and reduced resin-rich zones. The [0°/90°/45°/-45°]₂ showed the second highest (69–71%), benefiting from quasi-isotropic reinforcement but with slightly more interfacial breakdown. The [45°/-45°]₄ laminate exhibited moderate stability, retaining approximately 63–66% of the residue. The [0°]₈ laminate also exhibited moderate stability, with around 61–64% residue. The [45°]₈ laminate degraded earlier, retaining only 57–59%, while the poorest performance was observed for [90°]₈, which retained just 51–53% due to transverse fiber orientation and the presence of resin-rich zones. These results indicate that balanced cross-ply and quasi-isotropic laminates are more resistant to resin decomposition and heat damage. Furthermore, the orientation of fibers plays a decisive role in delaying thermal degradation by minimising resin exposure and enhancing load transfer. The observed two-stage mass loss behaviour is consistent with reported TGA studies on carbon fiber-epoxy composites, where initial moisture loss is followed by epoxy matrix decomposition, and thermal stability is governed by fiber orientation and resin distribution^42,43. Figure 16 presents the comparative TGA profiles of Mg-CF sandwich composites with six fiber orientations.

Fig. 16 — TGA comparison of Mg-CF sandwich composites with six fiber orientations.

Differential scanning calorimetry (DSC)

DSC was performed on the Mg-CF sandwich composites to study thermal transitions and decomposition. Specimens of approximately 12 mg were used, ensuring uniform heating and stable pan contact. The AZ31 Mg sheet showed no thermal events up to 450 °C, confirming that the CFRP core dominated the response. All samples exhibited three main features: a slight baseline drift below 150 °C due to moisture loss and physical relaxations, minor endothermic changes between 150 °C and 250 °C related to post-curing and initial matrix degradation, and a pronounced endothermic slope between 250 °C and 380 °C corresponding to epoxy decomposition and interfacial breakdown. Beyond 380 °C, the curves stabilised without any melting of Mg (expected near 650 °C). The DSC thermal transitions observed in this study agree with prior reports on CFRP systems, showing that laminate architecture and fiber orientation influence heat-flow behaviour and epoxy decomposition characteristics^44,45. Figure 17 presents the comparative DSC thermograms of Mg-CF sandwich composites with six fiber orientations.

Fig. 17 — DSC comparison of Mg-CF sandwich composites with six fiber orientations.

Among the laminates, balanced and quasi-isotropic lay-ups ([0°/90°]₄ and [0°/90°/45°/-45°]₂) showed slightly reduced endothermic magnitudes, indicating improved in-plane heat conduction and more uniform resin distribution. Unidirectional laminates ([0°]₈, [90°]₈, [45°]₈) exhibited deeper endothermic dips due to resin-rich zones and anisotropic heat flow. These results confirm that fiber orientation not only influences thermal stability in TGA but also modulates the apparent heat absorption and decomposition profile in DSC.

Quantitative aerospace relevance and thermal margin analysis

Based on the TGA and DSC results, the thermal response of the Mg-CF sandwich composites can be quantitatively related to realistic aerospace thermal environments. All laminates remain thermally stable up to approximately 150 °C, with only minor moisture loss and physical relaxations observed in DSC. This range fully covers typical aerospace service conditions, including fuel tank surroundings, wing and fuselage skins, and internal structural regions (-55 °C to 120 °C). Minor endothermic features detected between 150 °C and 250 °C are attributed to post-curing and initial matrix softening rather than structural degradation. These temperatures overlap with localised hot-zone exposures such as engine bay peripheral components and hot bleed air ducts, which generally operate below 200 –250 °C for secondary structures. Consistently, TGA revealed no significant mass loss below 300 °C, confirming the absence of chemical degradation within this operational window^46,47.

Major mass loss associated with epoxy decomposition and fiber-matrix interfacial breakdown occurs between 350 °C and 500 °C, while DSC indicates pronounced decomposition only above 250–300 °C. This establishes a minimum thermal margin of approximately 100 –150 °C between the highest anticipated aerospace service temperatures and the onset of irreversible degradation. Balanced cross-ply and quasi-isotropic laminates ([0°/90°]₄ and [0°/90°/45°/-45°]₂) exhibited higher residual mass (≈ 70–75%) and reduced endothermic response, reflecting superior thermal robustness under transient overheating conditions. Although these composites are not intended for extreme environments such as atmospheric re-entry (> 600 °C), the retention of intact magnesium face sheets and substantial char residue beyond this temperature suggests limited short-duration thermal tolerance. Overall, the results demonstrate that the Mg-CF sandwich laminates provide adequate thermal margins for conventional and high-speed aerospace structural applications, with fiber orientation playing a critical role in enhancing thermal stability and damage tolerance.

Mechanical studies

Tensile testing

Tensile testing was conducted in accordance with ASTM D3039 to determine the tensile strength, modulus of elasticity, and elongation at break of the sandwich composite specimens. Samples were prepared with dimensions of 250 mm × 25 mm × 3.2 mm. A constant crosshead speed of 5 mm/min was maintained during testing. Load and displacement data were captured in real-time to generate stress-strain curves. The average tensile stress obtained from three replicate specimens was taken as the representative tensile strength for each laminate stacking configuration. Experimental variability was quantified by reporting the minimum and maximum tensile stress values observed among the replicates, and the corresponding positive and negative error bars were determined relative to the mean. This methodology enables a clear assessment of data scatter and facilitates comparison of variability across different laminate configurations.

Among all samples, the [0°]₈ laminate exhibited the highest tensile load and strength, owing to the alignment of fibers parallel to the loading direction, which enabled maximum stress transfer through continuous fiber paths. The second-highest performance was observed for the [0°/90°]₄ bidirectional laminate, as its orthogonal fiber structure offered moderate longitudinal strength. The [0°/90°/45°/-45°]₂ multidirectional laminate ranked third, which provided a balance of axial and shear load resistance. Lower values were recorded for [45°/-45°]₄ and [45°]₈, where fibers were orientated diagonally relative to the tensile axis, reducing effective axial load transfer. The [90°]₈ laminate exhibited the lowest tensile performance, as fibers-orientated perpendicular to the load direction contributed minimally to resisting axial stress. The strong dependence of tensile strength on fiber alignment is consistent with classical laminate theory and previous studies on CFRP laminates, which report maximum tensile load-carrying efficiency when fibers are orientated parallel to the applied loading direction, owing to direct axial stress transfer through continuous fibers^48,49. Figure 18 presents the comparative tensile stress-strain curves of all six laminate configurations, highlighting the influence of fiber orientation on the mechanical response of the Mg-CF sandwich composites.

Fig. 18 — Comparative stress-strain curves of magnesium-carbon fiber sandwich composites with six laminate configurations.

Where Inline graphic = Peak Load (N), A= Cross-sectional area, Change in tensile stress within the linear elastic region, Corresponding change in tensile strain.

Figure 19 presents a bar graph comparing the tensile strength of all six Mg-CF composite laminates, clearly highlighting the influence of fiber orientation and stacking sequence on tensile performance. The corresponding values for tensile strength, tensile modulus, and peak load for each stacking configuration are summarised in Table 5.

Fig. 19 — Comparative bar graph of tensile strength for six Mg-CF composite laminate configurations.

Table 5.

Measured tensile strength values of six laminate stacking configurations.

Stacking configurations	Peak load (N)	Tensile strength (MPa)		Standard deviation (± MPa)	Tensile strength (MPa)	Tensile modulus (GPa)
Stacking configurations	Peak load (N)	Min	Max	Standard deviation (± MPa)	Tensile strength (MPa)	Tensile modulus (GPa)
Mg/CF[0°]₈/Mg	16936.69	217.7	229.6	± 4.1	223.8	40.0
Mg/CF[90°]₈/Mg	7021.81	86.6	94.5	± 2.9	91.0	7.2
Mg/CF[45°]₈/Mg	9905.07	122.5	136.8	± 4.8	130.2	8.4
Mg/CF[0°/90°]₄/Mg	15034.13	192.8	200.1	± 3.2	196.8	21.1
Mg/CF[45°/-45°]₄/Mg	12690.26	156.7	171.9	± 3.9	164.8	13.9
Mg/CF[0°/90°/45°/-45°]₂/Mg	13308.10	170.0	182.4	± 4.5	176.5	16.4

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Flexural testing

Flexural behaviour of the composite laminates was evaluated according to ASTM D790 using the three-point bending method. Standard specimen dimensions were 127 mm × 12.7 mm × 3.2 mm, with a span length set to 16 times the thickness (~ 51.2 mm). A crosshead speed of 5 mm/min was used. The maximum load at failure and corresponding deflection were recorded to calculate flexural strength and flexural modulus. The average flexural stress obtained from the three replicate specimens was considered as the representative flexural strength for each laminate configuration. To account for experimental scatter, the minimum and maximum flexural stress values were reported, and the corresponding positive and negative error bars were calculated relative to the mean value.

The [0°]₈ laminate again demonstrated the highest flexural strength, as fibers aligned along the span direction effectively resisted bending-induced tension and compression. The [0°/90°/45°/-45°]₂ laminate showed the second-best flexural performance, due to its multidirectional architecture enabling good resistance to complex bending stress states. The [45°/-45°]₄ laminate ranked third, with diagonal fibers partially contributing to flexural rigidity. The [0°/90°]₄ laminate ranked fourth, slightly below expectation, possibly due to the early onset of delamination or interfacial shear under transverse stress. The [45°]₈ and [90°]₈ laminates showed lower flexural strength, with the latter again recording the lowest value, consistent with its poor alignment to the primary load path during bending. The flexural performance trends observed in this study also align with the reported behaviour of laminated composites, where fiber orientation governs bending stiffness and strength through its influence on tensile-compressive stress distribution and interlaminar shear resistance under three-point bending conditions^50,51. Figure 20 displays the comparative flexural stress-strain curves of the six laminate stacking sequences, emphasising the effect of fiber orientation on the bending performance of the Mg-CF sandwich composites.

Fig. 20 — Flexural Stress-Strain Curves of Mg-CF Composite Laminates with Different Stacking Sequences.

Where Inline graphic = Peak Load (N),

L = Support span (mm).

b = Specimen width (mm).

d = Thickness (mm).

Inline graphic = Slope of initial linear region of load- deflection curve.

Figure 21 shows a bar graph comparing the flexural strength across all configurations, emphasising the directional dependence of flexural performance. The corresponding values for flexural strength, flexural modulus, and peak load for each stacking configuration are provided in Table 6.

Fig. 21 — comparative bar graph of flexural strength for Six Mg-CF composite laminate configurations.

Table 6.

Measured flexural strength values of six laminate stacking configurations.

Stacking configurations	Peak load (N)	Flexural strength (MPa)		Standard deviation (± MPa)	Flexural strength (MPa)	Flexural modulus (GPa)
Stacking configurations	Peak load (N)	Min	Max	Standard deviation (± MPa)	Flexural strength (MPa)	Flexural modulus (GPa)
Mg/CF[0°]₈/Mg	1186.65	691.6	711.8	± 10.1	700.9	33.3
Mg/CF[90°]₈/Mg	666.88	390.4	398.7	± 4.2	393.7	11.0
Mg/CF[45°]₈/Mg	715.91	411.5	429.3	± 8.9	422.6	10.4
Mg/CF[0°/90°]₄/Mg	794.37	454.2	478.9	± 12.4	469.7	15.3
Mg/CF[45°/-45°]₄/Mg	863.02	505.9	520.6	± 7.4	509.5	16.6
Mg/CF[0°/90°/45°/-45°]₂/Mg	1010.12	583.2	604.0	± 10.4	596.4	20.8

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Impact testing

Impact performance was evaluated in accordance with ASTM D256 using the Izod impact test method. For each stacking configuration, three replicate rectangular specimens with dimensions of 64 mm × 12.7 mm × 3.2 mm were prepared. A standard notch was introduced at the centre of each specimen, which was then mounted vertically with the notched side facing the striking edge of the pendulum hammer. The energy absorbed during fracture was measured in joules and employed to determine the impact strength of the laminates.

In contrast to the trends observed in tensile and flexural loading, the [45°/-45°]₄ laminate exhibited the highest impact strength. This superior performance is attributed to the angled fiber layers, which enhanced energy dissipation through fiber-matrix shearing, crack deflection, and delamination during fracture. The [0°/90°/45°/-45°]₂ laminate ranked second, as its multidirectional nature allowed for multi-axial energy absorption. Surprisingly, the [0°]₈ laminate ranked third, suggesting that while unidirectional fibers resist axial loads efficiently, their ability to absorb impact is limited to the fiber direction alone. Moderate impact strength was observed for [45°]₈, followed by [0°/90°]₄, which had restricted energy dissipation due to orthogonal layering. The [90°]₈ laminate recorded the lowest impact strength, consistent with its inability to resist force propagation along the direction of impact. Enhanced impact resistance in ± 45° and multidirectional laminates has also been reported in aerospace-grade composites, where off-axis fiber orientations promote shear deformation, crack deflection, and energy dissipation mechanisms^52,53. Figure 22 presents a bar chart comparing the impact strength of the six Mg-CF laminate stacking sequences, while the corresponding absorbed energy and impact strength values are detailed in Table 7.

Fig. 22 — Bar chart comparison of impact strength across six Mg-CF laminate stacking sequences.

Table 7.

Measured impact strength values of six laminate stacking configurations.

Stacking configurations	Minimum energy (kg·m)	Maximum energy (kg·m)	Average energy (kg·m)	Standard deviation (± kg·m)	Impact strength (kJ/m²)
Mg/CF[0°]₈/Mg	0.165	0.215	0.188	± 0.025	45.44
Mg/CF[90°]₈/Mg	0.095	0.130	0.110	± 0.018	26.54
Mg/CF[45°]₈/Mg	0.155	0.190	0.168	± 0.018	40.59
Mg/CF[0°/90°]₄/Mg	0.135	0.165	0.147	± 0.015	35.45
Mg/CF[45°/-45°]₄/Mg	0.375	0.460	0.420	± 0.043	101.43
Mg/CF[0°/90°/45°/-45°]₂/Mg	0.270	0.310	0.293	± 0.020	70.70

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Conversion of Energy from kg m into Joules

Figure 23 shows the fractured specimens of Mg-CF composite laminates subjected to tensile, flexural, and impact tests, visually illustrating the distinct failure patterns associated with different loading modes and fiber orientations.

Fig. 23 — Fractured Mg-CF composite samples after tensile, flexural, and impact testing.

Fracture surface morphology and semi-quantitative SEM analysis

Scanning Electron Microscopy (SEM) was employed to elucidate fracture mechanisms and damage evolution in magnesium-carbon fiber (Mg-CF) sandwich composites subjected to tensile, flexural, and impact loading. Fractured specimens were sputter-coated with gold to ensure electrical conductivity and examined using a TESCAN VEGA3 SBH SEM operated under high-vacuum conditions. The SEM observations revealed distinct, orientation-dependent failure features including fiber breakage, fiber pull-out, matrix cracking, interfacial debonding, voids, and localized delamination. To extend the qualitative interpretation of fracture mechanisms, a semi-quantitative analysis was performed on representative SEM micrographs of tensile, flexural, and impact tested specimens.

Tensile fracture morphology

Tensile fracture behavior was strongly governed by fiber orientation relative to the loading direction. In the [0°] laminates, SEM images revealed dominant fiber breakage with limited pull-out, reflected by a short average pull-out length (≈ 12 μm) and low crack density (≈ 4 cracks/mm²). The minimal fiber extraction and low matrix cracking indicate efficient stress transfer and strong interfacial bonding, consistent with the highest tensile strength observed experimentally. Representative SEM micrographs of these tensile-fractured surfaces for different laminate configurations are presented in Fig. 24.

Fig. 24 — SEM micrographs of tensile-fractured surfaces in Mg-CF sandwich composites with different fiber orientations.

In contrast, [0°/90°] and quasi-isotropic [0°/90°/45°/-45°] laminates exhibited mixed failure modes. Fiber breakage in 0° plies coexisted with pull-out and interfacial debonding in off-axis layers, resulting in moderate pull-out lengths (≈ 22 μm) and intermediate crack densities (≈ 6 cracks/mm²). These features indicate non-uniform stress distribution and explain the reduction in tensile strength compared to unidirectional laminates. Laminates with [45°/-45°], 45°, and 90° orientations showed pronounced fiber-matrix debonding, matrix cracking, and fiber pull-out. The higher crack density (≈ 13 cracks/mm²) and short fiber engagement lengths (≈ 5 μm in 90° laminates) reflect inefficient load transfer, correlating directly with their lower tensile strength and brittle failure behavior. The observed transition from fiber-dominated fracture in 0° laminates to matrix cracking, fiber pull-out, and interfacial debonding in off-axis and transverse configurations is consistent with reported tensile failure mechanisms in CFRP laminates, where fiber alignment governs load transfer efficiency and fracture morphology under axial loading⁵⁴.

Flexural fracture morphology

Under flexural loading, the fracture surfaces revealed tension–compression asymmetry across all laminates. In [0°] laminates, the tensile side exhibited fiber rupture and limited matrix cracking, while the compression side showed localized fiber buckling. The low void fraction (≈ 0.7%) and low crack density contributed to delayed crack propagation and superior flexural strength. Representative SEM micrographs of the flexural-fractured surfaces for the different laminate configurations are presented in Fig. 25.

Fig. 25 — SEM micrographs of flexural-fractured surfaces in Mg-CF sandwich composites with different fiber orientations.

Multidirectional laminates, particularly [0°/90°/45°/-45°], displayed complex damage patterns including shear cracks, fiber bridging, and matrix deformation. These laminates showed moderate pull-out lengths (≈ 24 μm) and reduced crack density compared to transverse laminates, suggesting effective stress redistribution among plies. This distributed damage mechanism explains their balanced flexural response and improved damage tolerance. In [45°/-45°] laminates, diagonal shear cracking, fiber kinking, and interfacial debonding were dominant. Although crack density was moderate, the presence of shear-driven damage limited flexural strength. The 90° laminates exhibited extensive matrix cracking, high void fraction (≈ 2.2%), and poor fiber engagement, leading to premature flexural failure. Similar tension–compression asymmetry, fiber buckling on the compressive side, and shear-driven damage in multidirectional laminates have been widely reported in flexural studies of sandwich and laminated composites, where ply orientation and interfacial integrity control crack initiation and propagation under bending loads⁵⁵.

Impact fracture morphology

Impact fracture surfaces were characterized by rough, irregular morphologies indicative of multiple energy-dissipating mechanisms. Among all configurations, the [45°/-45°] laminates exhibited the longest average fiber pull-out length (≈ 36 μm), accompanied by matrix shear yielding and fiber bridging. These features significantly increase fracture surface area and energy dissipation, explaining the highest impact strength measured. Representative SEM micrographs of the impact-fractured surfaces for the different laminate configurations are presented in Fig. 26.

Fig. 26 — SEM micrographs of impact-fractured surfaces in Mg-CF sandwich composites with different fiber orientations.

Quasi-isotropic laminates showed crack deflection across multiple ply interfaces, resin-rich zones, and moderate fiber pull-out, resulting in balanced impact performance. In contrast, [0°] laminates failed primarily through fiber breakage with limited plastic deformation, reflected by shorter pull-out lengths and lower impact energy absorption. Transverse and off-axis laminates (45° and 90°) exhibited brittle matrix fragmentation, interfacial debonding, and higher void-assisted crack initiation. The limited fiber contribution and higher crack density in these laminates correlate with poorer impact resistance. The dominance of fiber pull-out, matrix shear yielding, crack deflection, and interfacial debonding under impact loading closely aligns with fracture mechanisms reported for CFRP and metal–polymer sandwich composites, in which energy absorption is governed by fiber orientation and interfacial damage evolution during dynamic loading⁵⁶.

Interfacial delamination between magnesium face sheet and composite core

Delamination at the interface between the Mg face sheet and the carbon fiber-epoxy composite core was observed in a few specimens, mainly under flexural and impact loading. SEM analysis revealed localised separation along the metal-composite boundary, likely due to interfacial stress concentration, poor surface preparation, or mismatched thermal and mechanical properties. In these regions, the discontinuity in fiber-matrix continuity and the presence of voids at the bonding layer contributed to crack initiation. Such delamination can weaken the load transfer efficiency across the interface and may compromise structural integrity under complex loading conditions. Representative SEM images illustrating well-bonded regions (a, b,c) and delaminated zones (d, e,f) for different fiber orientations are presented in Fig. 27. Similar interfacial delamination phenomena at metal-composite interfaces have been widely reported in magnesium-polymer and aluminium-CFRP sandwich systems, where elastic modulus mismatch, residual thermal stresses, and insufficient surface activation promote interfacial crack initiation and propagation under bending and impact loads⁵⁷.

Fig. 27 — SEM Images of Mg-CF composite interfaces with and without delamination in different fiber orientations.

Benchmarking against conventional aerospace sandwich systems

From an aerospace design perspective, it is instructive to benchmark the present Mg-carbon fiber sandwich laminates against established systems such as aluminium face sheets with Nomex honeycomb or polymer foam cores, using normalised performance metrics reported in the literature. Conventional Al-Nomex sandwiches offer excellent bending stiffness at low weight but are often limited by low transverse shear strength, reduced damage tolerance, and poor thermal conductivity. When normalised by areal density, the Mg-CF sandwich panels investigated in this study exhibit competitive specific stiffness and strength, while additionally benefiting from the toughness and thermal conductivity of magnesium face sheets and the load-bearing efficiency of continuous carbon fiber cores. Compared with aluminium-foam sandwich systems, the Mg-CF laminates demonstrate improved stiffness, energy absorption, and thermal robustness, particularly for balanced and quasi-isotropic layups. Although not intended as a universal replacement for conventional sandwich architectures, the Mg-CF system occupies a complementary design space where enhanced damage tolerance, thermal stability, and multifunctional performance are required.

Conclusions

In this study, magnesium-carbon fiber-based sandwich composites were successfully fabricated and characterised using six different fiber orientations: unidirectional ([0°]₈, [45°]₈, [90°]₈), bidirectional ([0°/90°]₄, [45°/-45°]₄), and multidirectional ([0°/90°/45°/-45°]₂) configurations. The influence of stacking sequence and fiber alignment on the structural and mechanical behaviour was systematically investigated through mechanical testing, thermal analysis, and microstructural studies.

Tensile testing revealed a strong orientation dependence, with the unidirectional [0°]₈ laminate exhibiting the highest tensile strength of 223.8 MPa, while the transverse [90°]₈ configuration showed the lowest strength due to ineffective load transfer. Flexural testing followed a similar trend, where the [0°]₈ laminate achieved the maximum flexural strength of 700.9 MPa, whereas multidirectional laminates provided improved resistance to complex bending stresses. Impact testing demonstrated a contrasting response, with the [45°/-45°]₄ laminate absorbing the highest impact energy (0.42 kg·m, corresponding to ≈ 101.43 kJ/m²), highlighting the role of off-axis fibers in enhancing energy dissipation through shear deformation and crack deflection. The quasi-isotropic [0°/90°/45°/-45°]₂ laminate exhibited a well-balanced multifunctional performance, ranking among the top three in tensile strength, second in flexural strength, and second in impact energy absorption.

Thermal analysis confirmed that fiber orientation also governs thermal stability. TGA results showed that cross-ply and quasi-isotropic laminates retained the highest residual mass, with [0°/90°]₄ maintaining 73–75% and [0°/90°/45°/-45°]₂ retaining 69–71% residue after heating to 750 °C, while the [90°]₈ laminate retained only 51–53%. DSC analysis indicated that all laminates remained thermally stable up to ≈ 150 °C, with major decomposition occurring only above 250–300 °C, establishing a thermal margin of approximately 100–150 °C relative to typical aerospace service environments.

X-ray diffraction analysis of both powder and solid-state specimens confirmed that the AZ31 magnesium face sheets retained their original hexagonal close-packed (HCP) crystal structure after compression moulding, with identical peak positions and no detectable secondary phases (e.g., MgO or intermetallics), indicating that the applied thermal and pressure history did not induce phase transformation or crystallographic degradation. The most intense Mg (101) reflection was consistently observed at ~ 35–37° (2θ) across all laminates, with no measurable peak shift, further confirming the absence of residual phase changes.

Microstructural investigations using SEM corroborated the mechanical findings, revealing orientation-dependent failure mechanisms such as fiber breakage in aligned laminates, extensive fiber pull-out and matrix shear in off-axis configurations, and limited interfacial delamination at the magnesium-composite interface. Overall, the results demonstrate that the Mg/CF[0°/90°/45°/-45°]₂/Mg sandwich configuration offers the most balanced combination of strength, stiffness, impact resistance, and thermal stability. These findings position magnesium-carbon fiber sandwich composites as a promising lightweight structural solution for aerospace applications where multifunctional performance and thermal robustness are required.

Future scope

Based on the present findings, two targeted research pathways are identified to directly extend the current Mg-CF sandwich system toward aerospace certification. First, building on the superior multifunctional performance of the quasi-isotropic [0°/90°/45°/-45°]₂ laminate, future work will focus on incorporating a thin interleaf or hybrid reinforcement (e.g., carbon nanotube- or graphene-modified epoxy layers) at the magnesium-composite interface. This approach is expected to enhance interfacial shear strength, damage tolerance, and thermal stability without altering the established laminate architecture. A systematic experimental plan involving lap-shear or peel testing, followed by tensile, flexural, and impact validation, would quantitatively establish gains in interfacial performance relative to the baseline system reported here.

Given magnesium’s inherent susceptibility to corrosion, future studies will focus on aerospace-relevant surface protection strategies for the AZ31 face sheets, such as plasma electrolytic oxidation (PEO) or sol-gel-based primer coatings compatible with epoxy bonding. These coatings would be evaluated through accelerated corrosion testing, thermal cycling, and residual mechanical performance assessment to verify long-term durability. Together, these two interfacing pathways, strengthening and corrosion protection, represent realistic, incremental developments that build directly on the present mechanical, thermal, and microstructural dataset and move the Mg-CF sandwich composite closer to qualification as a lightweight, damage-tolerant aerospace structural material.

Author contributions

A: M.E. Annadorai- Manuscript Drafting, Experimentation, Methodology. B: M. Ramakrishna- Reviewing and Editing, Result Anlaysis.

Data availability

The data supporting the findings of this study are available from the corresponding author upon reasonable request.

Declarations

Competing interests

The authors declare no competing interests.

Footnotes

Publisher’s note

Springer Nature remains neutral with regard to jurisdictional claims in published maps and institutional affiliations.

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32.Noman, A. A., Shohel, S. M., Riyad, S. H. & Gupta, S. S. Investigate the mechanical strength of laminated composite carbon fiber with different fiber orientations by numerically using finite element analysis. Mater. Today: Proc.10.1016/j.matpr.2023.02.132 (2023). [Google Scholar]
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36.Kausar, A., Ahmad, I., Rakha, S. A., Eisa, M. H. & Diallo, A. State-of-the-art of sandwich composite structures: manufacturing—to—high performance applications. J. Compos. Sci.7(3), 102. 10.3390/jcs7030102 (2023). [Google Scholar]
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57.de Castro Saiki, L. E. & Gomes, G. F. Understanding and mitigating delamination in composite materials: a comprehensive review. Mech. Adv. Mater. Struct.31(30), 13147–13167. 10.1080/15376494.2024.2333490 (2024). [Google Scholar]

Associated Data

This section collects any data citations, data availability statements, or supplementary materials included in this article.

Data Availability Statement

The data supporting the findings of this study are available from the corresponding author upon reasonable request.

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[CR30] 30.Alsubari, S. et al. Potential of natural fiber reinforced polymer composites in sandwich structures: A review on its mechanical properties. Polymers13(3), 423. 10.3390/polym13030423 (2021). [DOI] [PMC free article] [PubMed] [Google Scholar]

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[CR48] 48.Yan, B., Zhu, S., Tong, M. & Pan, S. Experimental study on the mechanical properties of laminates made of thin carbon fiber plies. Compos. Struct.245, 112336. 10.1016/j.compstruct.2020.112336 (2020). [Google Scholar]

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[CR50] 50.Tefera, G., Adali, S. & Bright, G. Flexural failure properties of fiber-reinforced hybrid laminated beam subject to three-point bending. Sci. Rep.14(1), 9792. 10.1038/s41598-024-60078-7 (2024). [DOI] [PMC free article] [PubMed] [Google Scholar]

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[CR52] 52.Sheikh, R., Moolrajani, N., Nadig, N. & Prasad, G. Mechanical characterization and fracture analysis of Aerospace-Grade fiber (Nanocomposites): A study on structural integrity and damage tolerance. In Fracture Behavior of Nanocomposites and Reinforced Laminate Structures (261–293). Cham: Springer Nature Switzerland. 10.1007/978-3-031-68694-8_11 (2024). [Google Scholar]

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[CR55] 55.Herath, H. U., Cai, D. A., Inusha, L., Macias, L., Wang, X. & P., & Quasi-Static indentation and compression behaviors of hybrid woven composite laminates. Coatings15(7), 791. 10.3390/coatings15070791 (2025). [Google Scholar]

[CR56] 56.Yang, Y., Miao, Z., Liu, Y., Tu, H. & Wei, Y. A comprehensive review of fiber-reinforced polymer-matrix composites under low-velocity impact. Mech. Adv. Mater. Struct. 1–39. 10.1080/15376494.2025.2458772 (2025).

[CR57] 57.de Castro Saiki, L. E. & Gomes, G. F. Understanding and mitigating delamination in composite materials: a comprehensive review. Mech. Adv. Mater. Struct.31(30), 13147–13167. 10.1080/15376494.2024.2333490 (2024). [Google Scholar]

PERMALINK

Mechanical and thermal performance of magnesium carbon fiber sandwich composites with variable fiber orientations for aerospace structures

M E Annadorai

M Ramakrishna

Abstract

Introduction

Materials & methodology

Materials

AZ31 magnesium alloy (face sheet)

Carbon fiber roving 3 K yarn (core sheet & reinforcement)

Epoxy resin (matrix)

Fig. 1.

Preliminary characterization of base materials

Fig. 2.

Fig. 3.

Table 1.

Table 2.

Table 3.

Methodology

Filament drum winding process

Fig. 4.

Hand layup process

Fig. 5.

Compression moulding process

Fig. 6.

Fig. 7.

Fig. 8.

Table 4.

Sample Preparation

Fig. 9.

Interface preparation and adhesion considerations

Testing and characterization

Mechanical testing

Fig. 10.

Thermal analysis

Fig. 11.

Microstructural characterization

Fig. 12.

Results and discussion

X-ray diffraction (XRD) analysis

Powder XRD analysis

Fig. 13.

Solid-state XRD analysis

Fig. 14.

Energy dispersive X-ray spectroscopy (EDS) analysis

Fig. 15.

Thermal studies

Thermogravimetric analysis (TGA)

Fig. 16.

Differential scanning calorimetry (DSC)

Fig. 17.

Quantitative aerospace relevance and thermal margin analysis

Mechanical studies

Tensile testing

Fig. 18.

Fig. 19.

Table 5.

Flexural testing

Fig. 20.

Fig. 21.

Table 6.

Impact testing

Fig. 22.

Table 7.

Fig. 23.

Fracture surface morphology and semi-quantitative SEM analysis

Tensile fracture morphology

Fig. 24.

Flexural fracture morphology

Fig. 25.

Impact fracture morphology

Fig. 26.

Interfacial delamination between magnesium face sheet and composite core

Fig. 27.

Benchmarking against conventional aerospace sandwich systems

Conclusions

Future scope

Author contributions

Data availability

Declarations