Finite Element Analysis of Inter Spar Ribs of Composite Wing of Light Transport Aircraft against Brazier Load

Inter spar ribs of wing of a transport aircraft is ubjected to various types of loads. One of the loads that poses stability problem to the interspar ribs of a wing is brazier load, which arises due to flexure of the wing. This paper describes about the finite element analysis of inter spar ribs of a wing at local level against brazier load. This study has been taken place while converting metal wing in to composite wing. The objective of this study is to reduce the weight penalty to the maximum possible extent by removing material wherever feasible. This paper is limited to discuss about the linear buckling analysis of ribs against brazier load. The buckling factor of ribs under consideration are reported in terms of square root times the eigenvalue obtained from finite element analysis, which represent the nonlinear effect of bending moment on brazier load. This study has helped to reconfigure/redesign the interspar ribs of wing. This has led to substantial weight saving of 2.85 Kg which accpunts 15.77% reductions of total mass of inter spar ribs

Realisation of Shear Flow at Crucial Spar Splice Joints of Composite Wing in Idealised Wing Test Box

CSIR -National Aerospace Laboratories (CSIR-NAL) has developed a patented processing technique called “Vacuum Enabled Resin Infusion Technology (VERITy” for the manufacturing of composite wing for its transport aircraft programme. The building block approach was adopted during the design stage of the composite wing to understand the structural response in increasingly complex structural levels. In the subcomponent level, a wing test box was designed, fabricated and was subjected to structural static testing for critical loads of the wing. The wing test box which was idealized into a rectangular box included the most critical joints such as top & bottom skin splice joints and front & rear spar splice joints. The stacking sequence and thickness of laminates used in the respective positions in wing test box replicated the wing structure. The challenge was to maintain the shear flow despite differences in the geometry of wing and test box without compromising on the magnitude of shear force at the critical spar splice joint. This paper discusses the methodology adopted for transforming shear flow from wing structure on to the idealized wing test bo

Interface Toughness Measurements in Blade Stiffened Composite Skin Specimens

Separation of skin and stringer is likely to be a failure mode in structures where there is considerable out-of-plane deformation. Such deformations are possible when a stiffened skin structure is loaded in compression/shear beyond buckling or in structures which contain a flaw/defect at the skin-stringer interface. Prediction of such a failure in numerical simulations requires parameters such as interface fracture toughness which have to be obtained through specimen tests. Since, interface toughness is generally mode dependent, this study deals with testing of blade stiffened composite skin specimens under pure Mode-I and Mode-II loading. Specimens are fabricated with Carbon/Epoxy unidirectional prepreg (IMA/M21) material using autoclave molding process. Mode-I & Mode-II interface fracture toughness tests are conducted on blade stiffened composite skin specimens. All specimens contain an initial disbond of known size at the interface between the skin and stringer flange. This disbond is created by placing a thin non-adhesive insert of appropriate size at the interface during layup. Test fixtures are also designed and fabricated. Tests are conducted in displacement control mode in a UTM. Disbond length at various loads is recorded. Compliance calibration method is used to determine the interface fracture toughness in both modes. The interface toughness of blade stiffened skin specimens are found to be different from interlaminar fracture toughness obtained from conventional double cantilever beam and end-notched flexure tests

Damage Tolerance of Stiffened Skin Composite Panels

This work discusses design, fabrication and damage tolerance evaluation of co-cured skin stringer carbon fiber composite panels. Such stiffened panels are typically found in aircraft wing skins. Composite panels representing a portion of an aircraft wing box are designed and fabricated using Carbon fiber and epoxy matrix using resin infusion process. Fixtures to support the panels during low velocity impact tests are also designed and fabricated. A drop tower is used to conduct impact tests. Panels are subjected to various impacts to study the effect of impact energy on damage visibility and damage size. Impacts are also categorized according to their location: (a) Impact exactly above stringer, (b) Impact above skin, and (c) Impact above stringer flange. The extent of damages is studied based on non-destructive inspection techniques such as ultrasonic inspection. Further, one of the panels containing impact damages is subjected to residual strength test. Displacements and strains are measured using digital image correlation technique and resistance strain gages. Finite element model of the panel is also developed. Deformations and strains obtained from FE simulations are compared with test data. Results show that impact damages did not alter the load path significantly in the composite panel

Buckling & postbuckling of cocured composite stiffened panel under axial compression load- Computation & Testing

A cocured carbon fibre composite stiffened panel was designed to have an early onset of skin buckling and larger postbuckled strength. IMA/M21 prepreg was used for fabricating the panel using autoclave moulding process. Special end fixtures were developed using a metallic casing and the panel was encased in it using a mixture of epoxy resin and aluminium powder. The panel was tested under axial compression to its full load bearing capacity in a universal testing machine. Full field displacement measurement was carried out using advanced techniques like Digital Image Correlation (DIC) apart from acoustic emission technique and strain gauges for monitoring the structural response. After the test, the panel was subjected to ultrasonic scan to check for any delamination or disbonding that might have occurred due to post buckled response. Concurrently, a finite element model was developed to predict the buckling and postbuckling response of panel. DIC captured the onset of skin buckling and deformations/mode shapes in postbuckling regime. These experimental observations were correlated with numerical simulations. In the postbuckled regime, severe bending and twisting of skin and stringers was observed, resulting in complete loss of global axial stiffness of the panel. It is suspected that such a state of stress in the panel could lead to delamination, debonding or fiber failures. However, acoustic emission sensors bonded to the panel did not record any significant events during tests, even in the deep postbuckled regime. This was further corroborated through ultrasonic inspection conducted after unloading the panel

Tufting thread and density controls the mode-I fracture toughness in carbon/epoxy composite

Herein, interlaminar crack initiation and its growth in tufted carbon/epoxy composite under Mode-I loading are investigated for different thread materials and tuft densities. Two configurations of Double Cantilever Beam (DCB) test specimens are fabricated – one with 2 rows of tuft (hereafter referred to as lower tuft density) and another with 3 rows of tuft (will be referred to as higher tuft density). The crack front is arrested and delamination growth is delayed by tufting, which increases interlaminar fracture toughness . Higher enhancement in fracture toughness is observed for carbon thread tufted specimens followed by Kevlar and then glass thread tufted specimens. Fracture toughness of tufted specimens is about 4.5 to 10 times of untufted specimen depending on the thread material and tuft density. An increase in tuft density increases fracture toughness from 175% for Kevlar to 272% for glass threads. Fiber bridging from the parent laminate layer is observed in the untufted specimen, whereas in the tufted specimen, this phenomenon is insignificant except for bridging due to tufting thread. Fracture analysis shows that the failure is mainly due to the rupture of thread at the interface. Thread pull-out or slippage is absent exhibiting good adhesion with epoxy matrix

Strain sensor's network for low-velocity impact location estimation on carbon reinforced fiber plastic structures: Part-II.

Identification of low velocity impact (LVI) location in composite aircraft structures is seamless need for safe, reliable operation and maintenance of aerospace industry. To locate the LVI’s an optimized sensor network has designed using the strain response from fiber Bragg grating (FBG) & resistance strain gauge (RSG) sensor bonded to the composite structure. Strain scan (SS) algorithm has been developed to locate such events reported as Part-I. In this work, we have developed a novel algorithm based on weighted energy (WE) of the sensor response. The LVI’s has been carried out on composite structures & the locations of LVI’s have estimated using SS, WE & previously developed machine learning base support vector machine (SVM) algorithms. The WE and SS algorithms are based on proximity of events (closer to the sensor, higher the response), whereas LS-SVR is a data-driven approach. Further, we have compared the performance of the developed algorithms and algorithms cited in the literature using the performance index (PI), a measure of estimation efficiency as a function of the number of sensors, dimension/area of the structure, error & number of test cases. It is established that WE algorithm shown suprema performance over the other algorithm with 34 mm mean Euclidian distance error & PI value of 5.5

Unfolding the effects of tuft density on compression after impact properties in unidirectional carbon/epoxy composite laminates

Post impact compressive residual strength of epoxy-based composite laminates is studied in the presence of through-thickness reinforcement (TTR). The tufting technique is used to introduce TTR in compact carbon unidirectional (UD) composites. Composite laminates are tufted using Kevlar thread with different tuft density by varying tuft pitch and spacing. The specimens are impacted with low-velocity impact to achieve barely visible impact damage (BVID). The damaged area is quantified using the ultrasonic C-scan method. Tufted specimens exhibited a significant reduction in the damaged area. The reduction in damage area is about 26% to 51% depending upon the tuft density. Residual strength post impact is determined by compression after impact (CAI) test. Upon increasing the tuft density up to 0.56%, the damaged area decreases by 51%, and CAI strength increases by 43%. Above this tuft density, damage area slightly decreases and CAI strength remains almost the same. The improvement in CAI strength is due to an increase in apparent interlaminar strength attributed to the enhancement in bridging effect due to TTR. Different failure modes such as delamination, fiber crushing, kink-band formation, etc. are observed in both the untufted and tufted specimens

Damage detection in composite aircraft wing-like test-box using distributed fiber optic sensors.

Structural Health Monitoring Systems for the aircraft structures are intended to detect the location and size of the damage with minimum downtime of the aircraft. As the damage produces localized strain changes, a sensor network of many sensors bonded/embedded throughout the structure is required for damage diagnosis. This paper describes a distributed fiber optic sensor's demonstration to detect the disbond damage in a typical aircraft wing-like structure where a single optical fiber can cover a large area. Distributed fiber optic sensor is being bonded along the bolt line of the test-box. Controlled disbond are created by the removal of the bolts. In the pristine state, the strain signature/map is measured under applied load, which act as a reference in this case. In this approach, localized strain signature difference between the reference and disbond-structure is used to detect the damage. A non-dimensional quantity damage index (DI) is calculated based on the normalized sum of the square magnitude of strain difference of all sensors to indicate the damage severity. We have found that the DI value greater than or equal to 0.2 indicates the damage region through the iterative approach. Further, the algorithm is validated by loading the structure with disbond at different load levels. The maximum error in the disbond estimation is found to be ∼11 mm. This system and methodology hold immense promise for a ground-based method for detecting damage (disbond) in the aircraft/unmanned aerial vehicle structures (UAV)

Strain sensor's network for low-velocity impact location estimation on carbon reinforced fiber plastic structures: Part-I

In this work, we have investigated the strain response (angular/spatial) from fiber Bragg grating (FBG) sensor & resistance strain gauge (RSG) sensors bonded to the composite structure due to the projectile low velocity impact (LVI). The number of sensor & its orientating has been optimized based on such experimental data and designed an optimum sensor network for faithful LVI detection. In order to study the efficacy of the sensor network, an impact localization algorithm based on peak strain amplitude from the sensor bonded to the structure was used in this study. Further the detection efficiency of the algorithm has been improved using weighted average value around the peak amplitude of strain experienced by the sensor. We found that for the high energy (~35 J) LVI the maximum distance error (Euclidian distance) was 50 mm for 80% of total trail case. Furthermore, we have developed and compared the relative performance of the algorithm cited in the literature, will be presented in PART-II of the same Journal