Since 2009, research focused on damage formation of composite aircraft fuselage by High Energy Wide Area Blunt Impact (HEWABI) has been conducted at UCSD. As a major damage source to composite aircraft fuselage structures, HEWABI, caused by accidental contact by heavy ground service equipment (GSE), potentially leaves significant internal damage in multiple structural components with barely visible damage signs on the outside skin surface. Accounting for key structural components existing in real fuselage structures, specifically floor structures and continuous shear tie, a series of new 2nd generation large scale specimens were designed, fabricated, and tested. The objectives of the research described herein are to: (1) use of large-scale experiments to understand damage formation from HEWABI events near the floor structure location using the 2nd generation composite fuselage panel; of interest is examining how damage development is affected by the major changes from the 1st to 2nd generation panel design and boundary conditions, (2) investigation of C-frame failure both experimentally and analytically, and (3) evaluation of current FE modeling capability correlated with the 2nd generation HEWABI test results.

The increased used of complex geometries in aerospace and automotive applications has been possible due to recent advancements in material manufacturing methods. Materials such as woven polymer matrix composites and additively manufactured (AM) metals offer benefits in weight and part reduction compared to traditional aerospace materials. The need for the experimental and numerical characterization of the nonlinear behavior in advanced composites and defects in laser powder bed fusion (LPBF) metals under intermediate loading rate is needed to better use such materials in aerospace and automotive applications which are subjected to impact or crash events. The establishment of comprehensive material databases is both labor intensive and costly. The need for an efficient methodology to characterize the rate-dependent behavior is therefore needed.

To address this problem, experimental characterization of the in-plane shear response of two balanced plain weave carbon fiber composites is studied using test methods based on a v-notch shear specimen and a shear fixture designed using a combination of the ASTM D7078M and a modified Arcan fixture. Additionally, specimen preparation for digital image correlation (DIC) with Phantom high speed cameras is discussed. Results from 2D DIC and measured force values are used to obtain experimental stress-strain curves. Comparison between quasi-static and intermediate rates mechanical behavior show increase in strength and modulus for both plain weave carbon fabrics.

Additionally, a hybrid experimental and numerical characterization methodology is developed for woven polymer matrix composites that uses models from Micromechanics Analysis Code with Generalized Method of Cells (MAC/GMC) with experimental stress-strain curves. Validation of this methodology is performed through coupon verification in finite element analysis (FEA) software LS-DYNA for two woven materials using a tabular plasticity material model, MAT213. Additionally, validation studies of an adhesively bonded composite wingbox structure subjected low velocity impact is simulated and compared with experimental tests.

Lastly, experimental investigation of the combined effects of heat treatment on AM steel metals with fusion defects is performed through uniaxial tension experiments at various loading rates. High-speed camera images of mechanical tests seek to determine how heat treatment affects the material response and failure. Optical microscopy images are used to observe the failure mechanisms of specimens with internal defects.

Primary aerospace composite structures are commonly assembled with bolted joints due to their ability to transfer high loads and ease of assembly. However, when bolted joints are used beyond their originally intended design life, joint strength can be significantly reduced due to the accumulation of internal damage, necessitating frequent inspections. Furthermore, internal damage in composites (delamination, matrix cracks) can continue to propagate without visual indications, thus nondestructive testing is required. As a result, maintenance can become very costly, particularly for aircraft that are in-service beyond their designed life expectancy. By establishing a comprehensive understanding of damage propagation behavior, engineers can determine which damage modes to inspect for and reduce inspection frequency. This research aims to support and improve maintenance operations, fleet management, and aircraft design practice by investigating delamination initiations and propagations in the vicinity of fastener holes within fiber-reinforced composite materials.

Static and fatigue bearing were performed using novel test methods developed as part of this research for countersunk fastener joints: the modified countersunk double lap shear (DLS), single lap shear (SLS), and semi-circular notch (SCN) test configurations. DLS and SLS static and fatigue experimental test results were compared to study joint configuration, laminate stacking sequence, and loading condition effect on bearing damage initiation and growth under both static and fatigue loading. From static and fatigue tests, it was observed that major bearing damage accumulates in the straight shank region of the countersunk hole indicating most of the bearing load is carried by the straight shank region. Fatigue bearing test data showed that when the bolted hole elongates, stiffness decreases and internal delamination damage area growth becomes detectable through C-scan. Stated in reverse, if no measurable hole elongation is found, significant delamination is not expected. Complex damage morphology forms in this region, emanating from the loaded bearing face, and creating large wedge-shaped regions that drive delamination propagation with additional loading cycles. Additionally, optical microscopy observations indicated that pin bending might have affected bearing damage growth. In order to understand the effect of pin bending, a custom designed semi-circular notched experiment was performed on the countersunk hole geometry and compared to the DLS static experiments. Results indicated that the pin bending had no strong effect on the bearing failure morphology for the selected diameter to thickness ratio.

Finite element analysis using Virtual Crack Closure Technique (VCCT) and Hashin failure criteria in Abaqus was used to further understand the internal stress state of the specimen configurations and to investigate the rate of delamination growth and arrest in the SCN and DLS configurations. Results from FEA were used to more comprehensively understand the observations from static and fatigue experiments and to verify hypotheses formulated to explain these observations.

Sandwich structural elements consist of two metallic or non-metallic laminated skins separated by a lightweight thicker core layer, maximizing in this manner weight efficiency and structural performance. From the available core configurations, phenolic resin impregnated Nomex® paper honeycomb cores are highly preferred in aerospace industry because of their lower density, excellent thermal resistance, and curved shape adaptability.

Despite the inherent benefits of Nomex® paper sandwich structures, their impact response is still questionable. Towards this direction, a series of simulated hail ice impact tests was performed on flat carbon fiber sandwich panels via gas gun test apparatus to assess Nomex® paper core damage. Impact test results revealed significant core crushing in the form of cell collapse without any externally visible signs of damage on the impacted skin surface.

Since high-velocity impact tests on sandwich panels are by nature highly complex, the out-of-plane response of over-expanded Nomex® cores was phenomenologically characterized through quasi-static flatwise compression tests on stabilized sandwich core coupons. Localized detachment and crushing of the phenolic resin accumulation regions between adjacent cell walls, were the main damage modes that triggered honeycomb collapse.

X-ray Computed Tomography (CT) scans were also employed to assess damage modes on the interior cells of tested sandwich core coupons. Given several sets of X-ray CT scan 3D data, an automated B-spline image-based reconstruction scheme was developed to numerically capture scanned honeycomb core geometries. Different applications of the proposed reconstruction scheme are going to be presented. Finally, the contribution of the phenolic resin accumulation regions on the out-of-plane response was investigated through numerical modeling of the reconstructed B-spline geometry. To this end, constituent material properties were determined through uniaxial tensile testing on in-situ phenolic resin impregnated paper coupons as well as bare aramid paper coupons. Numerical simulations on sandwich cores without the phenolic resin accumulations yielded significant reduction of the collapse limit, validating in this way the experimental observations.

Blunt impact damage (BID) is known to potentially cause significant strength reduction on composite structures even when the damage is barely noticeable. Therefore, non-destructive evaluation (NDE) to assess blunt-impacted composite aircraft structures is a crucial procedure to determine existence and severity of the damage to examine airworthiness. Ultrasonic guided waves (UGWs) test method has been studied herein as an NDE technique to evaluate large composite structures, because the typically thin-walled aircraft structures effectively act as has guided waves propagation paths.

UGW scanning methods were performed only from the external skin-side using pitch-catch method: an actuator and a receiver located at a distance apart to scan the structure within. The actuator and the receiver transducers were placed on the adjacent skin-to-shear tie bolted joint regions for scanning of internal shear tie and C-frame components, and placed at opposite sides of the hat stringer on the skin bay for investigation of co-cured stringer components. UGWs transmission through bolted joint structure was examined to determine the test frequency that allowed wave transmission through bolted joints at low attenuation. And UGW tests from the skin-to-shear tie joint regions identified significant wave energy drop at the test frequency from the comparison of the pristine and the damaged (HEWABI) shear ties. Furthermore, damage characteristics were analyzed from the UGW measurements of the stringer scan followed by successful blunt impacted stringer delamination damage detection. Features extracted from the UGW test results showed exceptional match to the damage location, severity, and modes determined from the conventional NDE methods. Finally, UGW measurements-based residual strength estimation was studied from simple notched coupon tests and is an on-going study for more complex structures with realistic damage modes.

Composites are frequently used in aerospace structural applications due to their high strength to weight performance, but due to their layered structure they are vulnerable to transverse impacts. Impact damage in composite laminates often consists of highly interactive damage modes composed of delamination, matrix cracking, and fiber breakage. In order to ensure the safety of composite structures, a variety of non-destructive evaluation (NDE) techniques are used to characterize impact damage. However, procedures for utilizing NDE to create and validate models of residual strength after impact are not yet established due to either limitations in the characterization of impact damage, as in the case of Ultrasonic pulse-echo scanning (UT), or due to the complexity of interpretation of the NDE technique, as in the case of X-ray computed tomography (CT). Improved quantification of damage from CT and UT characterization may lead to improved predictive capabilities for the prediction of structural performance after an impact event.

This work presents a novel automatic damage segmentation procedure for CT scans of impacted composites that converts the complex 3D dataset into simplified damage visualizations and 2D damage maps for each composite layer. The results of this procedure were utilized to create and validate a modeling procedure to improve UT characterization of impact damage, and to validate and generate finite element models of impact damage and residual strength performance. The generated residual strength models were created with varying levels of damage modeling fidelity and it was found that the level of damage modeling needed for accurate failure prediction depends greatly on the structural geometry and the presence of major damage features. This NDE and modeling effort was supported by a series of impact and residual strength experiments for flat and stringer stiffened composite panels. The developed techniques proved capable of characterizing impact damage in a variety of structural configurations and establishing models that incorporate this damage at different levels of complexity.

This dissertation focuses on investigating and developing techniques for measuring residual strength estimate of a damaged composites from purely nondestructive signals. The nondestructive evaluation (NDE) system relates an ultrasonic guided wave(s) (UGW) emanating from a structure and reasonably estimates the structures residual strength using the Average Stress Criterion.

The UGW signals are flexural and dependent upon the material properties and plate geometry, which change under impact or some other damaging mechanism and is thus destructively observable as a strength reduction. Correlating strength to signal requires three objectives to be accomplished: 1) Parameterize broadband mini-impactor excitation source, 2) develop algorithm for processing spatiotemporal data for phase velocity information, and 3) relate UGW degradation to physical strength all in the context of Average Stress Criteria.

Parameterization of the mini-impactor hammer, which is a broadband excitation source, was the starting point. Seventy-two design permutations of plate material, thickness, and mini-impactor define the experiment domain. Ten experiments per permutation results in seven hundred and twenty experiments filling the domain with temporally aligned raw data. Temporal and spectral analyses are conducted with the modus: “follow the energy” and found 75\% of the analyses agree in terms of significant and deployable permutations to real world applications.

Knowing future data will be spatiotemporal UGW, an algorithm (WAVSVD) uses singular value decomposition to reveal all spatial and temporal patterns, some which are muted in our physical experience. WAVSVD is also used to denoise data. WAVSVD captures more datapoints than traditional 2DFFT and captures the phase velocity and frequency attenuation because of damage. WAVSVD is the UGW processing backbone and helps identifies the max characteristic wavenumber at attenuation as the degrading metric to use in the residual strength estimate.

Impact residual strength can be modeled with Average Stress Criterion. The assumption being that the stress distribution from edge of damage to edge of plate is developed as an Average Stress Criterion mixture of the material degradation and cracking. The Average Stress Through-hole correlates residual strengths of impacts to those of holes, while the Average Stress Crack captures the characteristic shift that needs to occur to relate the damage size to an equivalent crack.

Low velocity impact damage to laminated composite structures can cause a range of damage modes. Of interest is small damage composed of delamination (i.e. separation of plies) that is less than 50 mm in size, with barely visible impact-side damage. Formation of actual impact-induced damage, mainly multiple planes of delamination that are interconnected via matrix cracks, is the main goal of this study, which supports a larger project where injection repair of the damage is being investigated. Three 25-ply composite panel types having varying percentages of 0° fiber content have been low velocity impacted at 10, 15, and 20 Joule energy levels. 81 total impact studies were performed using a pendulum impact system with the composite panels clamped on all edges. Multiplanar internal delamination regions with barely visible impact damage on the impacted surface were created. In addition to measuring the force history during the impacts event, the internal damage is investigated and characterized using ultrasonic pulse-echo scanning methods. The perimeter of damaged regions is determined using gel-coupled hand-mapped A-scan methods, while an extensive investigation of the multiple planes of delamination damage is investigated using C-scanning methods where the specimen is immersed in de-ionized water. Results show that the panel type with the least amount of 0° plies has the largest internal delamination damage, but this panel type also requires the highest impact force for damage to occur. In addition to impact-side C-scans, nine specimens are back side C-scanned to investigate the pulse-echo shadowing effect which commonly hides lower-depth delaminations.

Despite a high strength to weight ratio, composites are susceptible to impact damage which can be barely visible while still adversely affecting their strength, therefore detecting and characterizing damage is important. Non-destructive evaluation, specifically single-sided pulse-echo ultrasonic C-scans, can be used to detect damage. The goal is to develop a code that aids the quantitative assessment of damage within laminates by segmenting C-scans of damaged regions into 3D reconstructions. The code is open source to create a free alternative to commercial ultrasonic C-scan software and allow users to adapt the code for their own applications. The code processes C-scans and calculates time-of-flight (TOF), then uses unique characteristics in C-scans to create a map of boundaries where the damage changes from one layer to another. Using this map, the code processes the TOF to create a 3D reconstruction of damaged layers. When both front and back side C-scans of a sample are available, the code combines them to produce a hybrid C-scan that merges information from both. The code is developed using a dataset of 25-ply composite panels with varying layups and levels of impact damage. A 24-ply composite sample impacted at 48 Joules was also processed and compared with an X-ray CT dataset. The code's 3D reconstruction output is useful for an enhanced understanding of the impact damage state (e.g., area of damage at specific depth), provides the capability to extract quantitative metrics of the damage, and can be used to create actual-geometry structural analysis model representations of the damage state.

A common damage mechanism that occurs on polymer matrix composite (PMC) aircraft structures are delaminations which are interlaminar defects between composite plies. A current practice for repair of delaminations include injection using a repair resin. Injection repairs have traditionally been a faster, less expensive way to repair composite delaminations. However, the aerospace industry and safety regulating authorities consider injection repair cosmetic and have not accepted it as a strength restorative process. Questions about quantifying the removal of, internal contamination, how to characterize restored strength, and ensuring desired percentage of repair resin fill within the delaminations have been primary reasons why injection repairs have not been credited with strength restoration. Therefore, a new quantitative internal surface cleaning, preparation and repair process was developed for delaminated composite materials including the development and utilization of a new test coupon and configuration.

First, laboratory manufactured fractures were successfully compared to delaminations taken from actual aircraft structures using advanced surface metrology. Laboratory manufactured delaminations methods include a novel modified end notch flexure coupon (ENF) loaded in Mode II for fracture propagation and an out of plane impact specimen. Multiple lay-up configurations were characterized and surface metrology results indicate that the ENF fracture surfaces was most closely related to the in-service damage due to statistical comparison of arithmetic mean (Pa) and root mean square (Pq) surface profilometry and arithmetic mean roughness (Ra) results.

Materials and process development was also conducted to create a novel injection repair procedure. First, intentional contamination procedures were successfully developed, verified, and introduced into simulated delaminations. Contamination removal and internal fracture surface preparation using solvent and atmospheric plasma cleaning procedures were remarkably successful in removing hydrocarbon contaminant archetypes from contaminated fractures. Real-time mass spectrometry providing quantifiable cleaning verification of hydrocarbon contamination removal was successfully implemented into the novel procedure. Destructive post-processing verification of surface cleaning by infrared spectrometry, and goniometry was successfully completed. In addition, a modified low-viscosity injection resin was also developed for delamination infiltration and was found to exhibit shear strength characteristics comparable to un-modified structural adhesive while achieving a >84% reduction in viscosity. Following the development of a simulated enclosed delamination modified ENF coupon, application of the surface cleaning and modified repair resin to the previously fractured and contaminated coupons were tested. After re-pair of the modified ENF coupons, Mode II interlaminar fracture toughness (GIIc) and calculated stiffness was quantitatively restored or increased for all coupon configurations. Fracture analysis was completed to verify desired cohesive repair resin or adherend matrix brittle fracture failure modes of repaired test specimens. Therefore application of the novel injection repair procedure provided quantifiable local mechanical property restoration that contributes to the overall repair of aerospace PMC structures to carry designed loads.