Advanced Systems Laboratory

This experimental study investigates the impact of carbon black nanoparticles (CBNP) on the ablative properties of two commonly used composite materials in aerospace applications: Carbon/Phenolic and Silica/Phenolic composites. Neat and Nanofillers added laminates of these two composite types were prepared using a hand layup process followed by curing in an autoclave at 170 ℃ temperature and 5 bar pressure. Physical properties such as density, fibre content, degree of cure, and glass transition temperatures of all four types were found. Mechanical tests were conducted to observe the change in Interlaminar shear strength (ILSS), Flexural strength (FS), and Flexural Modulus (FM) of the laminates due to the addition of fillers. Oxy-acetylene torch tests were carried out on all classes of laminates to study the influence of these new class of fillers on the ablative properties of composites such as Liner Ablation Rate (LAR) and Mass ablation rate (MAR). A heat flux of 835 W/cm² was applied for the ablation test in this experiment. The results indicated that the addition of CBNP (5 wt %) significantly enhances the ablation properties of Carbon/Phenolic composites (LAR & MAR reduce by 60 % & 62 % respectively), while it degrades for Silica/Phenolic composites (LAR & MAR increase by 17 % & 27 % respectively.

ABSTRACT Acousto-Ultrasonic Testing (AUT) was performed on Hydroxyl-Terminated Polybutadiene (HTPB)-based solid rocket propellant samples to study the dispersion behaviour of mechanical waves within the material. Low-frequency mechanical waves, ranging from 10 kHz to 300 kHz, were used for testing. The results demonstrated that the stress wave velocity increased non-linearly with increasing frequency and sample thickness. Additionally, propellant samples subjected to natural ageing were analysed for mechanical properties, such as modulus, and these properties were correlated with stress wave velocity. This study aims to monitor variations in mechanical properties over time, revealing correlation between changes in mechanical properties and computed stress wave velocity. An average increase in wave velocity of approximately 25% is observed over a duration of 14 months. A corresponding enhancement in the elastic modulus, estimated at ~ 1.25 MPa (approximately 30%), is evident during the early stages of natural ageing. Experimental investigations on the variations in velocity with stress wave frequency and sample thickness are detailed.

Abstract The effect of 200 days of cyclic weathering test (CWT) and 1,000 h of salt spray test (SST) on the performance of polyurethane (PU) coating was investigated. A 200 μm coating was applied to a composite and cured before CWT and SST ageing studies. CWT was conducted in ten cycles (each cycle: −40°C for 5 days, 70°C for 5 days, 40°C and 90% RH for 5 days, and salt fog for 5 days). Initial water immersion tests concluded that the coated specimens have 37% less water absorption than uncoated specimens. The degradation behaviour of the coating was evaluated by measuring pull-off adhesion strength, critical load from the scratch test, and wear depth from the pin-on-disk test. Similarly, the surface characteristics of the coating were studied by SEM, gloss value, and contact angle. CWT ageing caused a decrease in gloss value (93–78 GU), pigment content (18–8 wt%), contact angle (92–70°), and an increase in wear depth at 20 N of applied load (45–200 µm). Similarly ageing also decreased the pull-off adhesion strength to 6.77 MPa (CWT) and 7.93 MPa (SST). CWT and SST also decreased the thermal stability by 11–13°C, considering 5 wt% loss and tensile strength of the coating from 42 to 29 MPa. The damage to the composite substrate was not significant due to prolonged exposure periods under CWT and SST, though the damage was noted in PU coating, like increased surface roughness and wear depth.

In this study, maximum back wall temperature and erosion rate of both the carbon phenolic composites (C-Ph) and ZrO 2 fabric incorporated carbon phenolic composites (Z-C-Ph) are evaluated after oxy-acetylene flame test to assess the effect of ZrO 2 fabric incorporation on the thermal and erosion properties of the composite. Theoretical estimation of back wall temperature is also carried out using ANSYS 15.0. Theoretical as well as experimental estimation shows that the heat affected zone of composite during oxy-acetylene flame test is wider and deeper for the C-Ph than those of the Z-C-Ph composite owing to the higher thermal conductivity and lower density of the C-Ph than those of the Z-C-Ph composite. Theoretically estimated maximum back wall temperatures (∼500°C for C-Ph and ∼450°C for Z-C-Ph) are found to be similar with the experimentally determined maximum back wall temperatures (516 ± 106°C for C-Ph and 428 ± 70°C for Z-C-Ph). The effects of several other parameters such as surface roughness of the sample, presence of blind drilled holes on the composite sample, char formation thickness on the maximum back wall temperature and erosion rate of both the composites are also studied. It is, in general, observed that the maximum back wall temperature and erosion rate is relatively higher for C-Ph (516 ± 106°C, 0.020 ± 0.007 mm/s, respectively) than those of Z-C-Ph (428 ± 70°C, 0.016 ± 0.003 mm/s, respectively). The percentage reduction of experimentally measured thermal conductivity and erosion rate due to ZrO 2 fabric incorporation in C-Ph composite is found to be ∼40% and ∼28%, respectively. Post-test examination with SEM reveals that more severe crack formation in the matrix and more matrix/carbon fibre de-bonding are observed in case of C-Ph than those in case of Z-C-Ph. Presence of blind drilled holes mitigates delamination thus increases thermal conductivity and renders the increase of back wall temperature. The higher char formation thickness causes decrease in the maximum back wall temperature since char reduces thermal conductivity. Smoother front surface in general causes lower heat absorption from the oxy-acetylene flame and thus decrease both the front/back wall temperature and erosion rate in case of C-Ph. However, smoothening of the surface by machining causes exposure of ZrO 2 fabric and subsequent reaction of ZrO 2 with char (C) which causes formation of highly porous char, ZrC and gaseous CO. This causes increase in the erosion rate due to greater char removal and consequently increase in back wall temperature for Z-C-Ph.

Abstract Mechanical properties of carbon‐epoxy composites are highly affected by the operating temperature. In this study, the effect of temperature on the compressive strength and in‐plane shear strength of composites has been analyzed. The unidirectional composites are processed by filament winding using T700 carbon fiber rovings and two different epoxy resin systems. The compressive strength measured from combined loading compression tests showed a decreasing trend with an increase in temperature. The compressive strength decreased from 700 to 400 MPa when the temperature increased from –20 to 150°C, which represents 54% reduction in strength. The in‐plane shear tests were performed using longitudinal and transverse extensometers and found that shear strength decreased by nearly 50%. Fractography of compression samples revealed that the primary compressive failure mode is micro buckling followed by kinking, at all temperatures. The existing micro mechanical model based on microbuckling and kinking for estimating the compressive strength has been modified to include temperature effects. The key model input parameters are temperature‐dependent resin tensile modulus and in‐plane shear strength. It is found that the modified model will predict the temperature‐dependent compressive strength in the temperature range of −20 to 150°C.Highlights Compressive strength decreased from 700 MPa at −20°C to 400 MPa at 150°C. Microbuckling followed by kinking failure was observed at all temperatures. In‐plane shear strength decreased from 44 MPa at −20°C to 27 MPa at 150°C. Resin tensile modulus decreased from 2.5 GPa at −20°C to 0.75 GPa at 150°C. Analytical model for compressive strength was extended for all temperatures.

Polymeric composites find extensive usage in aerospace applications, and their performance is influenced by environmental conditions throughout their life cycle. This study focuses on assessing the performance of composite laminates under different environmental conditions to evaluate the load carrying capacity LCC due to delamination. The laminates were specifically designed to withstand high pressure and temperature, ensuring satisfactory performance throughout their service life. The specimens, prepared according to ASTM standards with a thickness of 3 mm, featured different fibre orientations between the upper and lower laminates, including 0/0°, 0/30°, 0/45°, and 0/60°. The change in the delamination growth behavior for specimens subjected to different initial delamination lengths (a0) was studied using pre and post-radiographic tests RT. The investigation encompassed a range of initial delamination lengths, from 70 mm to 110 mm, incremented by 10 mm. Notably, failure was observed in specimens with a 0/30º angle when the initial crack length a0 reached 110 mm, while specimens with a 0/60º angle failed at an initial crack length of 80 mm. Additionally, it was noted that the maximum force required for the 0/30º angle laminate was observed when the initial crack length was 70 mm.

Polymeric composites find extensive usage in aerospace applications, and their performance is influenced by environmental conditions throughout their life cycle. This study focuses on assessing the performance of composite laminates under different environmental conditions to evaluate the load carrying capacity LCC due to delamination. The laminates were specifically designed to withstand high pressure and temperature, ensuring satisfactory performance throughout their service life. The specimens, prepared according to ASTM standards with a thickness of 3 mm, featured different fibre orientations between the upper and lower laminates, including 0/0°, 0/30°, 0/45°, and 0/60°. The change in the delamination growth behavior for specimens subjected to different initial delamination lengths (a0) was studied using pre and post-radiographic tests RT. The investigation encompassed a range of initial delamination lengths, from 70 mm to 110 mm, incremented by 10 mm. Notably, failure was observed in specimens with a 0/30º angle when the initial crack length a0 reached 110 mm, while specimens with a 0/60º angle failed at an initial crack length of 80 mm. Additionally, it was noted that the maximum force required for the 0/30º angle laminate was observed when the initial crack length was 70 mm.

Many military weapon systems used for the battlefield are to be stored longer before they launch from aircraft, ships, or land. These weapon systems are subjected to extreme stresses from long-term exposure in harsh environments, which affects their total life cycle and reliability. A thorough reliability analysis of weapon-integrated pressure sensors exposed to environmental stresses during long-term storage and transportation, emphasizing performance drift and ageing, is presented in this article. Laboratory experiments were conducted under controlled conditions simulating climatic and dynamic stresses as per MIL-STD-810H. Furthermore, it has been demonstrated how to use interval-censored field failure data to estimate the life distribution of a population in service, where pressor sensors encounter varying stress levels over time, and to determine key reliability metrics, including failure rate, mean time to failure, and probability over a 10-year storage period. The study offers quantitative insights into stress-related deterioration, supporting the development of condition-based calibration strategies, predictive maintenance planning, and robust sensor designs suitable for extended military storage applications.