Flush repairs to military aircraft are expected to become more prevalent as more thick skin composites are used, particularly on the surface of the fuselage, wings and other external surfaces. The use of these repairs, whilst difficult to manufacture provide an aerodynamic, ???stealthy??? finish that is also more structurally efficient than overlap repairs. This research was undertaken to improve the design methodology of scarf repairs with reduced material removal and to investigate the damage tolerance of scarf repair to low velocity impact damage. Scarf repairs involve shallow bevel angles to ensure the shear stress in the adhesive does not exceed allowable strength. This is important when repairing structures that need to withstand hot and humid conditions, when the adhesive properties degrade. Therefore, considerable amounts of parent material must be machined away prior to repair. The tips of the repair patch and the parent laminate are very sharp, thus a scarf repair is susceptible to accidental damage. The original contributions include: ??? Developed analytic means of predicting the stresses within optimised scarf joints with dissimilar materials. New equations were developed and solved using numerical algorithms. ??? Verified using finite element modelling that a scarfed insert with dissimilar modulus subjected to uniaxial loading attracted the same amount of load as an insert without a scarf. As such, the simple analytic formula used to predict load attraction/diversion through a plate with an insert may be used to predict the load attraction/diversion into a scarf repair that contains a dissimilar adherend patch. ??? Developed a more efficient flush joint with a doubler insert placed near the mid line of the parent structure material. This joint configuration has a lower load eccentricity than external doubler joint. ??? Investigated the damage tolerance of scarf joints, with and without the external doubler. The results showed that scarf joints without external doublers exhibited a considerable strength reduction following low velocity impact. Based on the observations, the major damage mechanics in the scarf joint region following impact have been identified. These results demonstrated that it is important to incorporate damage tolerance in the design of scarf repairs.

This master´s thesis is focused on aerodynamic analysis of 3D configuration of aircraft TL – 4000 and designing shape modifications to improve flight performances. Modifications are focused on wing – fuselage junction and junction between tail units. In continuation are analyzed the shapes of air intakes to cooling engine. Based on aerodynamic analysis of aircraft TL – 4000 are chosen convenient positions for pitot-static tube and inlets for ventilation of cabin.

A influência da fuselagem sobre a posição do centro aerodinâmico da asa é complexa e deve ser considerada nos cálculos de equilíbrio e estabilidade estática longitudinal da aeronave. Este trabalho apresenta uma análise comparativa para indicar o mais preciso dentre sete métodos teóricos para prever esta influência, em condições de baixa velocidade, utilizando seis configurações de modelos de asa mais fuselagem em escala reduzida, com proporções dimensionais características da aviação leve. Mediram-se os coeficientes de momento e sustentação para cada configuração, através de ensaios em túnel de vento de baixa velocidade, circuito aberto e seção de testes fechada. Calcularam-se as posições experimentais do centro aerodinâmico através da distância do eixo de rotação da balança ao bordo de ataque da asa e derivadas do coeficiente de momento em relação ao coeficiente de sustentação. Aplicaram-se os métodos teóricos às configurações. Os resultados demonstram que a maioria dos métodos prevê comportamentos na variação da posição do centro aerodinâmico semelhantes aos obtidos experimentalmente e apontados na revisão da literatura. A análise dos resultados teóricos ante os experimentais aponta o método descrito em Engineering Sciences Data Unit (1996a) como o mais preciso. / The fuselage influence on the wing aerodynamic center is complex and must be considered within longitudinal static stability and equilibrium calculations of the airplane. This work presents a comparative analysis to indicate the most accurate between seven theoretical methods that predict this influence, at low speed conditions, using six configurations of wing-fuselage reduced scale models, with the dimensional proportions found in light aviation. The moment and lift coefficients have been measured by experiments in a low speed open circuit wind tunnel with a closed test section. The experimental aerodynamic center positions have been found by the distance of the balance trunnion to wing leading edge and the derivation of the moment coefficient relative to the lift coefficient. The theoretical methods have been applied to all configurations. The results show that most of the methods predict variations in aerodynamic center position in the same way as those obtained in experimental results and shown in the literature review. The analysis between theoretical and experimental results indicates the method from Engineering Sciences Data Unit (1996a) as the most accurate.

A finite element model of fibre metal laminate (FML) fuselage was developed in order to evaluate its impact response under survivable crash event. To create a reliable crash finite element (FE) model of FML fuselage, a ‘building block approach’ is adapted. It involves a series of validation and verification tasks in order to establish reliable material and damage models, verified impact model with structural instability and large displacement and verified individual fuselage structure under crash event. This novel development methodology successfully produced an FE model to simulate crash of both aluminium alloy and FML fuselage under survivable crash event using ABAQUS/Explicit. On the other hand, this allows the author to have privilege to evaluate crashworthiness of fuselage that implements FML fuselage skin for the whole fuselage section for the first time in aircraft research field and industry. The FE models consist of a two station fuselage section with one meter longitudinal length which is based on commercial Boeing 737 aircraft. For FML fuselage, the classical aluminium alloy skin was replaced by GLARE grade 5-2/1. The impact response of both fuselages was compared to each other and the results were discussed in terms of energy dissipation, crushing distance, failure modes, failure mechanisms and acceleration response at floor-level. Overall, it was observed that FML fuselage responded similarly to aluminium alloy fuselage with some minor differences which conclusively gives great confidence to aircraft designer to use FML as fuselage skin for the whole fuselage section. In terms of crushing distance, FML fuselage skin contributed to the failure mechanisms of the fuselage section that lead to higher crushing distance than in aluminium alloy fuselage. The existence of various failure modes within FML caused slight differences from the aluminium fuselage in terms of deformation process and energy dissipation. These complex failure modes could potentially be manipulated to produce future aircraft structure with better crashworthiness performance.

The diploma thesis deal with CFD based a to the aerodynamic optimalization of a fuselage and a wing fuselage junction of LSA category aircraft EV 007. Software Fluent is used.

The diploma thesis deals with the project of composite constuction of aeroplane VUT 001 Marabu and its FEM stress-deformation analysis. The project of pilt´s cockpit is solved in the light of basic ergonomic requirements. Mass analysis, calculation and resulting selection of conclusive events of load are parts of the work. The results are applied on FEM model of fuselage and the state of stress in individual sections of composite structure are determined by the MSC Patran/Nastran system.

This master’s thesis deals with the analysis of geometric shapes fuselage amfibie SEAGLE and propose optimal shapes for the improvement of hydrodynamic and aerodynamic properties.

Målet med detta examensarbete var att undersöka och ta fram ett konstruktionsförslag av en bärgningsmodul anpassad för Saabs robotsystem, RBS15, vilket skulle möjliggöra återanvändning av komponenter och underlätta utvärderingar av flygkroppen. Liknande system används av andra företag inom försvarsindustrin för att kunna utvärdera sina produkter. I examensarbetet har olika aspekter av bärgningssystemet undersökts, inklusive analys av fallskärmsinitiering, dimensionering av fallskärmar och infästningspunkter, samt packningsstudier av fallskärmsystemet. Litteraturstudier, intervjuer och simuleringar har använts som lösningsmetoder för att analysera och modellera systemet och dess komponenter. Beräknade krafter och flygbana visade att de befintliga spanten och infästningspunkterna på roboten klarade belastningen från fallskärmssystemet. Slutsatsen är att en bärgningsmodul till RBS15 har stor potential. Med detta system skulle roboten kunna återanvändas och utvärderas effektivt, vilket kan leda till ekonomiska fördelar och underlätta framtida utveckling av systemet. / The objective of this thesis was to investigate and develop a design proposal for a recovery module for Saab's missile system, RBS15, which would enable component reusability and facilitate evaluations of the missile. Similar systems are utilized by other companies in the defense industry to assess their products. The thesis project has explored various aspects of the recovery system, including parachute initiation analysis, dimensioning of parachutes with their affiliated attachment points, and packing studies of the parachute system. Literature reviews, interviews, and simulations have been employed as methods to analyze and model the system and its components. Calculated forces and flight trajectory demonstrate that the existing frames and attachment points on the missile can withstand the load from the parachute system. The conclusion is that a recovery module for RBS15 holds significant potential. With this system, the missile could be effectively reused and evaluated, which can lead to economic benefits and facilitating future system development.

The details of an experimental investigation focusing on obtaining the static and vibration characteristics of a full-scale carbon composite wing and fuselage structural assemblies of an ultralight unmanned aerial vehicle (UAV) are presented. The UAV has a total empty weight of 155-lb and an overall length of approximately 20.6t. A three-tier whiffletree system and the tail fixture were designed and used to load the wing and the fuselage in a manner consistent with a high-g flight condition. A shaker-table approach was used for the wing vibration testing, whereas the modal characteristics of the fuselage structure were determined for a freeree configuration. The static responses of the both structures under simulated loading conditions as well as their dynamic properties such as the natural frequency, damping coefficient and associated mode shapes were obtained. The design and implementation of the static and vibration tests along with the experimental results are presented in this thesis.

This study looks at the feasibility of using structural optimization techniques to address the problem of designing composite fuselage frames for crashworthiness. A key feature of any optimization strategy for increasing structural crashworthiness is a progressive failure analysis. Currently, the most widely used analysis methods for progressive failure of composite structures are considered too expensive computationally for practical optimization in today's computing environment. Developing an efficient analysis method for progressive failure of composite frames is a first step in the optimization for crashworthiness. In the current work a progressive failure analysis for thin-walled open cross-section curved composite frames is developed using a Vlasov type beam theory. A curved thin-walled composite beam theory is developed and a finite element implementation of the beam theory is used for progressive failure analysis. The accuracy and limitations of this analysis method are discussed. A model for progressive failure of the composite fuselage frame is developed from an extension of the laminate progressive failure analysis of Tsai-Wu. Comparisons based on a limited amount of available experimental data are encouraging. The first major failure event is captured by the theory, and the prediction of total energy absorbed follows the trend of the experimental data. It is believed that this accuracy is sufficient for preliminary design and optimization for crashworthiness. This progressive failure analysis is then incorporated into a frame optimization for crashworthiness based on the genetic algorithm method. The optimization methodology is demonstrated analytically to obtain frame designs with substantially increased crashworthlness. Laminate stacking sequence and cross-section shape are design variables for optimization / Ph. D.