Composite materials are increasingly used in rotorcraft structures to reduce weight and improve efficiency. The rotorcraft industry is constantly in need of higherperformance materials that offer improved mechanical strength and stiffness at a lower weight. In polymer-matrix composite structures, matrix-dominated failures impose severe limitations on structural performance. The objective of this work is to advance composite material technologies for rotorcraft through the use of nanoadditives to improve structural efficiency. Technical challenges and potential solutions for improving matrix-dominated performance of prepreg composites through nanoparticle reinforcement, are discussed. In particular, a promising technology for improving compression and interlaminar strength and fatigue performance, is identified. The advanced materials technology is based on high weight content loading of approximately 100-nm diameter nanosilica particles in low-viscosity resins. Such technology resulted in 2 MAKEEV compression strength improvement for intermediate-modulus carbon-fiber/epoxy-matrix 250° F curing prepreg composites as recently demonstrated by 3M. This work not only supports the initial findings of 3M regarding the improvement of compression strength performance but also demonstrates improved interlaminar material properties including fatigue performance, and expands the material design space. Fatigue performance is critical to rotorcraft dynamic components as they are subject to extreme oscillatory flight loads that can result in material fatigue failures.

The compression modulus and compression strength of a proprietary ultra-high modulus carbon/epoxy composite are experimentally determined using SACMA SRM 1R-94 and ASTM D6641 for laminates with stacking sequences of [(0/ AE 60) s ] 2 and [(þ60/0/À60) s ] 2 . The moduli of both laminates are experimentally shown to be statistically equivalent; the strength of the [(0/ AE 60) s ] 2 laminate is shown to be significantly less than the [(þ60/0/À60) s ] 2 laminate. A finite element model is developed for each compression test to predict the compression modulus and compression strength using the built-in damage progression algorithm in Abaqus. Premature end crushing is predicted in the modulus specimen of SACMA SRM 1R-94 for the [(þ60/0/60) s ] 2 laminate. End crushing in the modeled SACMA SRM 1R-94 strength specimen is predicted to occur immediately after ultimate laminate failure for the [(þ60/0/60) s ] 2 laminate. Fiber misalignment is found to significantly affect the predicted compression modulus and compression strength of the ultra-high modulus carbon/epoxy composite analyzed.

A B S T R A C T In this work we present a strain field mining methodology where the intensities and spatial distributions of strains in double lap strap epoxy (EA 9394) bonded joints are measured, analysed and correlated with bond strength. While the global behaviour of the bonded joints was linear elastic at room temperature, discrete regions of elevated (relative to the full field average) shear strains were observed. The size, spatial distribution and other features of these strain 'hot spots' were described with a metric called the homogeneity index. This index captures the uniformity of the strain field with respect to regions of positive and negative strain at a specific stress level. The linearity of the progression of the homogeneity index and its magnitude was correlated with a joint's fracture strength and failure mode. The index was a robust predictor of double-lap strap bonded joint performance between 30% and 90% of a specimen's strength and can be used to improve joint design, manufacturing, quality assurance, maintenance and inspection.A = pixel area of a hot spot A1 = fitting parameter A2 = fitting parameter B1 = fitting parameter B2 = fitting parameter C1 = fitting parameter C2 = fitting parameter τ ult = average failure shear stress of the adhesive for all specimens δ centre = nearest neighbour distance based on the centroids of hot spots δ edge = nearest neighbour distance based on the edges of hot spots δ min = smallest resolvable displacement for digital image correlation ε min = minimum detectable strain between two features for digital image correlation ε th = threshold strain used to define a hot spot σ ia x = shear stress at the adhesive and inner adherend interface σ oa x = shear stress at the adhesive and outer adherend interface τ ult = failure shear stress of the adhesivẽ δ þ À centre = median nearest neighbour distance based on the centroids of hot spots δ þ À edge = median nearest neighbour distance based on the edges of hot spots à = median pixel area of a hot spot A coh = percent of epoxy failure ANOVA = analysis of variance A ROI = pixel area of the field of view σ x = average bond line shear stress DICT = digital image correlation and tracking Correspondence: C. L. Muhlstein.

Fatigue of fiber-reinforced polymer matrix composites has been studied exhaustedly over the years. However, most fully reversed fatigue testing has concentrated on flexural fatigue. The resulting stress state through the laminate thickness is non-uniform and as such may give rise to unconservative life estimates. Fully reversed (R=−1) axial testing has not been heavily reported due to the difficulties associated with precluding local or global buckling of the axial specimen. The few reported fully reversed axial tests utilize either cylindrical specimens or fixturing that is very large, expensive to fabricate, and difficult to handle. Herein we report on a simple side support fixture that allows the use of standard ASTM tensile test specimens (D3039 or D638, depending on test material), which are readily prepared. The fixture allows fully reversed axial fatigue testing as well as static baseline compression and residual compression strength characterization. Also presented herein are tension-tension (R=0.1) and fully reversed tension-compression (R=−1) S-N diagrams for an E-glass/epoxy composite.