The research team, led by Nikolay V. Turbin, set out to answer a question that has lingered in the composite fatigue literature for decades: once an impact has seeded a delamination inside a laminate, how does that defect evolve when the structure is compressed over and over again? The question matters because modern airworthiness regulations, including the European Union Aviation Safety Agency’s CS 25.571 damage tolerance requirements and the Federal Aviation Administration’s Advisory Circular 20-107B on composite aircraft structure, demand that manufacturers demonstrate that damaged structure can sustain load until the damage is detected. For metals, fracture mechanics offers mature tools for predicting crack growth under fatigue. For composites, the picture has remained stubbornly incomplete, and the field has been searching for reliable slow-growth approaches for post-impact fatigue for years.

The experimental core of the study involved two test specimens fabricated from carbon-fiber reinforced laminate, a material class that dominates modern primary aircraft structures. Each specimen was first subjected to impact, creating the characteristic internal damage zone of delaminations at multiple depths through the thickness. The researchers then mapped the dimensions and depth of these internal delaminations before and after cyclic compression testing, using non-destructive inspection to build a before-and-after picture of the hidden damage architecture. This baseline-and-post-test comparison is what allowed the team to quantify how much the delaminations grew, in which directions, and at which interfaces between plies, rather than simply recording whether the specimen eventually failed.

Under cyclic compression, an impacted composite panel does not behave like a pristine one. The delaminated region behaves as a local zone of separated plies that can buckle outward or inward under compressive stress, and this local instability drives the interlaminar crack faces apart, pumping energy into the delamination front with every load cycle. The result is a distinctive damage evolution pattern: the delamination front advances through the laminate, the buckling zone enlarges, and the residual compressive capacity of the element declines. Previous studies by groups including Butler and Almond, Melin and Schön, and Uda and colleagues have documented aspects of this behavior, but the field still lacks a validated, practical framework that connects measured damage growth to a computational prediction that engineers could use in certification-oriented analysis.

The Moscow Aviation Institute team approached the problem with a two-pronged strategy. On the theoretical side, they developed a framework for describing specific damage growth in the impacted laminate, drawing on the classical foundations of fracture mechanics established by Griffith in 1921 and the crack propagation laws formalized by Paris and Erdogan in 1963, adapted to the particular geometry and mechanics of a delamination embedded in a compressed anisotropic plate. On the computational side, they built numerical simulations of the approximated damage growth, using finite element tools of the kind employed in commercial platforms such as Siemens Simcenter 3D, to reproduce the stress state around the delamination and the progression of the interlaminar crack. The theoretical predictions and the numerical model were then tested against the experimental observations from the cyclic compression tests.

One of the most striking findings from the cyclic testing was the appearance of a stage of unstable delamination growth followed by arrest. Rather than advancing smoothly and continuously with every cycle, the delamination in the tested specimens went through a period of rapid, unstable extension and then stopped growing, remaining arrested until the prescribed number of cycles expired. This behavior has important implications for how engineers think about post-impact fatigue life. It suggests that damage growth in these laminates is not a monotonically accelerating march toward failure but a more nuanced process in which the local mechanics of the buckled delamination zone, the changing geometry of the crack front, and the laminate’s resistance to interlaminar fracture interact to produce periods of dormancy as well as bursts of growth.

The study’s central claim is that this damage initiation and growth can be predicted. The researchers report that both their theoretical analysis and their numerical modeling captured the detected damage initiation and the subsequent growth observed in the experiments. In other words, the computational tools did not merely describe the damage after the fact; they anticipated where and when the delaminations would begin to extend and how far they would travel under the cyclic compression loading. The authors conclude that the possibility of a computational estimate of the damage tolerance of a composite element under cyclic loads is confirmed, which is precisely the kind of validated predictive capability that the slow-growth damage tolerance philosophy requires.

The significance of this confirmation reaches well beyond a single laboratory experiment. Certification practice for composite aircraft structures has traditionally leaned on conservative assumptions, treating barely visible impact damage as a strength knockdown rather than tracking its growth over time, because reliable slow-growth prediction has been elusive. As Pascoe and co-workers argued in a widely cited 2021 review, current research approaches had not yet delivered a workable slow-growth damage tolerance method for fatigue after impact in fiber-reinforced polymers. If computational frameworks of the kind demonstrated in this study can be refined and generalized, they could allow designers to quantify how a specific impact scar will behave under a specific service load spectrum, potentially reducing conservatism while maintaining or improving safety margins.

The work also illustrates a broader trend in composite structural mechanics: the convergence of high-fidelity simulation, fracture-mechanics theory, and advanced non-destructive inspection. Phased array ultrasonic scanning of impact damage in carbon-fiber composites, as documented in the inspection literature, provides the detailed three-dimensional damage maps that models need as input and validation data. Finite element analysis of composite materials, supported by dedicated modeling techniques for delamination and progressive damage, provides the predictive engine. And carefully designed cyclic compression-after-impact tests, following standards such as ASTM D7137 for compressive residual strength of damaged laminates and ASTM D7136 for damage resistance to drop-weight impact, provide the ground truth. The new study weaves all three strands together in a single, self-consistent methodology.

There remain, of course, limits to what two specimens can establish. The authors are explicit that the aim of the work is to enhance the understanding of post-impact damage growth in order to refine the model used for composite material damage tolerance estimation, not to deliver a finished certification tool. Extending the approach across a wider range of layups, impact energies, load levels, and environmental conditions will be the necessary next step, and the statistical variability inherent in both impact damage and fatigue behavior will demand larger datasets. The study was carried out under a grant from the Russian Ministry of Science and Higher Education for large scientific projects in priority areas of scientific and technological development, reflecting the strategic weight that aviation nations place on composite damage tolerance. But the core message is clear and consequential: the hidden damage left by an impact need not remain an unquantified unknown. With the right combination of theory, simulation, and experiment, the life story of a delamination under cyclic compression can be read in advance, and that ability may reshape how the next generation of composite aircraft structures is designed, inspected, and certified.

Subject of Research: Post-impact delamination growth in carbon-fiber composite laminates under cyclic compression loading

Article Title: Computational and experimental analysis of the post-impact delamination growth in a composite structural element subjected to cyclic compression

Article References: Computational and experimental analysis of the post-impact delamination growth in a composite structural element subjected to cyclic compression. (n.d.). https://doi.org/10.1007/s42401-026-00547-1

Image Credits: AI Generated

DOI: 10.1007/s42401-026-00547-1

Keywords: composite structures, damage tolerance, delamination, compression fatigue after impact, carbon-fiber reinforced laminate, cyclic loading, fracture mechanics, finite element analysis, non-destructive inspection, aerospace structures, impact damage, fatigue life prediction

The research, published as an open-access paper in the International Journal of Aeronautical and Space Sciences, simulated a 50.8 millimeter hailstone, the classic two-inch benchmark of aviation hail threat models, slamming into a radome modeled on the Airbus A320neo at a combined velocity of roughly 240 meters per second, the speed of a typical cruise condition. The team, led by Anand Pai with co-authors including P. K. Sowmya, Khushhali Chaudhary, Suhas Yeshwant Nayak, Chandrakant R. Kini and B. Satish Shenoy, compared two industry-standard fiber-reinforced polymer systems: glass fiber-reinforced polymer, or GFRP, and Kevlar, or p-aramid, fiber-reinforced polymer, known as KFRP. The verdict was unambiguous. The KFRP radome limited peak deformation to 8.1 millimeters, an 18 percent reduction compared with the 9.88 millimeters suffered by the GFRP variant, while weighing about 19 percent less.

The stakes are far from academic. The researchers compiled a survey of documented hailstone strikes on commercial aircraft over the past decade, and the record shows a rising frequency of encounters, a trend the authors attribute to the growing incidence of severe convective weather and the continued expansion of global air traffic. Hailstones routinely strike the nose cone, fuselage, wing skins, leading edges, control surfaces, engine nacelles and fan blades. The radome is particularly exposed because of its forward-facing geometry, and the minimum skin thickness requirements imposed to protect it carry a real weight penalty, one that matters acutely for smaller jets where every kilogram affects performance and fuel burn.

Modeling a hailstone is deceptively hard. Natural hail is polycrystalline ice, a layered, anisotropic material riddled with air and water pockets, and its mechanical behavior swings dramatically with strain rate. At the high strain rates of 100 to 300 per second typical of these impacts, ice transitions from ductile to brittle failure, yet high-speed imaging shows it retains residual strength through fragmentation and axial splitting even after damage initiates. Its compressive strength and hardness can climb by more than 300 percent as temperature falls from minus 15 to minus 50 degrees Celsius. The team represented the projectile using Smoothed Particle Hydrodynamics, a mesh-free method that sidesteps the severe mesh distortion that cripples traditional Lagrangian finite element approaches during extreme deformation. The hailstone was discretized into 67,731 particles of one millimeter each, governed by a bilinear isotropic constitutive model that captures the critical moment when the ice’s deviatoric stress vanishes at the failure threshold and the projectile begins to behave like a fluid.

The radome itself was built with equal care. The 7.5 millimeter skin was assembled from thirty discrete plies of 0.25 millimeters each in a symmetric [0/90]15 stacking sequence of 2 by 2 twill weave fabric. Rather than plugging in generic material cards, the researchers derived the effective orthotropic properties through a multi-scale homogenization workflow: representative volume elements of the unidirectional yarn architecture and the bidirectional woven architecture were simulated in ANSYS Material Designer, using manufacturer data for Kevlar 2200 aramid fibers with a Young’s modulus of 108 gigapascals and EC11 204 E-glass fibers at 72.4 gigapascals, both embedded in an epoxy matrix. The resulting laminate definitions were then exported into ANSYS Explicit Dynamics, where the coupled SPH-finite element framework ran the impact scenarios over a 100 microsecond window.

The simulation reproduced a two-stage impact response well known from experimental literature on flat composite panels: a rapid loading stage in which stress and deformation climb to a peak, followed by a rebound stage as the projectile fragments and the stress decays. On the curved radome geometry, the two materials diverged after about 40 microseconds. Up to that point their deformation histories were nearly indistinguishable, but the GFRP shell then deformed more steeply toward its larger final value, while the KFRP laminate held a more stable late-stage profile. Contact stress told a complementary story. The KFRP radome reached a higher peak stress of roughly 1581 megapascals at about 60 microseconds, versus about 890 megapascals at 70 microseconds for GFRP, indicating a stiffer, more localized response that accelerates hailstone fragmentation and spreads the debris tangentially along the radome’s curvature, a phenomenon the authors call geometric shedding.

Energy accounting reinforced the picture. Both simulations began with an identical projectile kinetic energy of 1925 joules, and total energy was conserved throughout, with hourglass energy held below 2 percent of peak internal energy, satisfying standard explicit dynamics stability criteria. But the KFRP radome accumulated more internal energy and showed a sharper decay of the projectile’s kinetic energy, signatures of enhanced momentum dissipation through multi-axial matrix shearing and plastic work. In other words, the aramid architecture does not merely deflect the blow; it absorbs and dissipates the impact energy more effectively across the convex surface of the nose cone.

Perhaps the most technically interesting contribution is the ply-by-ply failure analysis. Because global mechanical metrics poorly capture the directional nature of composite damage, the team implemented a sequential explicit-to-implicit mapping: deformation states extracted at 40, 60, 80 and 100 microseconds from the explicit solver were transferred into a static implicit environment, where the anisotropic Tsai-Wu failure criterion yielded an Inverse Reserve Factor for individual plies. An IRF at or above 1.0 signals the onset of material degradation. The results exposed a fundamental difference in damage character. In the GFRP radome, elevated failure indices spread diffusely, with the rear ply reaching IRF values up to 3.30 across wide, unconfined contours, a pattern that raises the risk of widespread micro-cracking, shear failure and interlaminar delamination. The KFRP radome, by contrast, confined its most severe indices, which climbed as high as 11.79, to a narrow zone at the impact apex, while the mid-plane plies stayed comparatively protected. Damage containment, not merely damage resistance, emerged as the Kevlar laminate’s decisive advantage, preserving the load-bearing capacity of the surrounding shell and the radar instruments behind it.

To bridge the gap between idealized single impacts and operational reality, the researchers also ran a cluster scenario of five concurrent 50.8 millimeter hailstones spaced 60 millimeters apart, extending the simulation to 5 milliseconds to capture staggered strikes and shockwave interactions. The result was a wider concave indentation with surface cratering across the central radome region, a damage envelope that closely resembles photographic evidence of real hail damage on an operational Interjet A320 nose cone. That qualitative agreement, the authors argue, validates the framework’s ability to capture the momentum transfer required to produce macroscopic cratering. The team is candid about limitations: the bilinear ice model omits thermal effects and microstructural heterogeneity, and future work will add stochastic multi-impact scenarios and cohesive zone modeling to probe the interplay between intralaminar softening and delamination. Still, the message for aircraft designers is striking. A radome built from Kevlar rather than glass fiber could shed roughly 19 percent of its mass, about 114 kilograms in the modeled geometry, while deforming less, absorbing more energy and keeping damage tightly contained where it matters least to the aircraft’s survival.

Subject of Research: Computational simulation of hailstone impact response of composite aircraft radomes

Article Title: Dynamic Impact Response of Composite Aircraft Radomes Subjected to Hailstone Impact: A Computational Study

Article References: Pai, A., Sowmya, P. K., Chaudhary, K., Nayak, S. Y., Kini, C. R., & Satish Shenoy, B. (2026). Dynamic Impact Response of Composite Aircraft Radomes Subjected to Hailstone Impact: A Computational Study. International Journal of Aeronautical and Space Sciences. https://doi.org/10.1007/s42405-026-01308-2

Image Credits: AI Generated

DOI: 10.1007/s42405-026-01308-2

Keywords: aircraft radome, hailstone impact, Kevlar, GFRP, composite laminates, SPH, finite element analysis, Tsai-Wu failure criterion, aviation safety, explicit dynamics, impact damage, A320neo