Frontiers in Aerospace Science

This study introduces an aeroelastic model for the linear/non-linear analysis of thick fins in supersonic or hypersonic regimes. In the first step a linear aeroelastic model for the analysis of thick fins is developed. To this aim, a thick fin with two degrees of freedom (2 DOF) as well as an elastic double-wedged fin in supersonic/hypersonic flight regimes are considered. An unsteady aerodynamic model is developed based on the shock/expansion theory by local aspplication of the piston theory over the flat surfaces of the fin. The structural model is also obtained based on the Lagrangian approach. Employing such developed model, the effects of initial angle of attack, thickness and some other geometrical parameters on the aeroelastic stability boundaries and unsteady aerodynamic loads are studied. In the next step, a more sophisticated model describing the non-linear aeroelastic behavior of fins with 3 DOF and free-plays in flapping, plunging and pitching motions is investigated. To this aim, governing equation are obtained by a modification of the linear model and effects of several geometrical parameters (e.g. thickness, initial angle of attack, hinge frictional torque etc.) on the aeroelastic behavior of fins are assessed1.

This paper focus on the assessment of aerodynamic forces applied on an aircraft. First, numerical analysis is performed on a simplified linear two-degree-of-freedom (2DOF) airfoil system model. Methods of aerodynamic force determination are established at different airspeed conditions and the calculated dynamic forces correlated well to the actual forces. Next, a finite element model of the airfoil is developed to represent its structural dynamics, and the established force determination methods are applied to determine the aerodynamic forces acting on such model. After that, a physical downsizing model of the airfoil is manufactured and its natural frequencies, damping ratios, and mode shapes are calculated and compared to those obtained from the analytical models. Vibration experiments are also conducted to measure the physical model’s structural response in the wind tunnel, and its aerodynamic forces are calculated based on the data recorded by the attached accelerometers.

This chapter provides a combined computational and analytical study to investigate the lateral impact behavior of pressurized pipelines and inspect effects of important parameters such as the outside diameter and internal pressure on such behavior. A total of more than 300 numerical simulations are carried out on mild steel pipe models with different internal pressure levels and were struck at the mid-span and at the one quarter span positions. These numerical simulations of the impact tests are performed using 3D dynamic nonlinear finite element analysis (FEA) through LS-DYNA, where both geometrical and material nonlinearities are considered. The computational results for the first time systematically reveal the effects of internal pressure, impact position, and outside diameter on the lateral impact behavior of the pipeline models. Quartic polynomial functions are applied to formulate the maximum crushing force (Fmax), maximum permanent displacement (Wf), and absorbed energy (Ep) of the pressurized pipelines during the impact problem. The effects of the diameter and pressure on F, W, and E are therefore illustrated through analyzing those functions. Response surfaces are also plotted based on the generated quartic polynomial functions and the quality (accuracy) of those functions are verified through several techniques. The outcomes of this study have potential benefits on research of safety and reliability of pressurized pipelines in hydraulic system of aerospace and development of advanced pipeline materials.

A number of existing and emerging industrial applications are dependent on layered substrates through adhesive bonding. The interfacial fracture of adhesively bonded structures is a critical issue for their extensive applications to a variety of modern industries. In the recent two decades, cohesive zone models (CZMs) have been receiving intensive attentions for fracture problems in adhesively bonded joints due to its fairly simple and accurate predictive ability. In CZMs the nonlinear interfacial fracture behaviors are described by the traction-separation laws (also referred to as cohesive laws). The cohesive laws represent the local constitutive behavior, instead of the global parameter, such as toughness. While numerous global tests have been conducted to measure the interfacial toughness of adhesive joints, limited local tests have been conducted to determine the interfacial tractionseparation laws or interfacial cohesive laws. Among the limited local tests in some recent experimental studies, very few studies have considered the effects of adhesive thickness on the local interfacial traction-separation laws. In the present work, within the framework of nonlinear fracture mechanics, comprehensive experimental studies are conducted to investigate the effect of adhesive layer thickness on the local nonlinear interfacial behaviors. The fracture tests of adhesive joints with various adhesive layer thicknesses were conducted under different fracture modes: pure Mode-I (peel fracture), pure Mode-II (shear fracture), and mixed Mode I/II. The experimentally determined interfacial traction-separation laws provide valuable baseline data for parameter calibrations in numerical models. The current experimental results may also facilitate the understanding of adhesive thickness dependent interface fracture of bonded joints.

With the continual development of the aircraft industry, aircraft engines have provoked people’s attention more and more. The turbine blade plays a vital and critical component of aircraft engines. In order to conform to the dimensional tolerances of wax pattern die-profile for turbine blade in investment casting process, this chapter provides an optimization method of geometric parameter for turbine blades based on inverse adjustment. The geometric parameters for optimizing were extracted, and the bending and torsional deformation can be compensation. Therefore the nonlinear deformation compensation during solidification and cooling procedure can be efficiently realized. This method set the theoretical foundation on optimization method of die-cavity for turbine blade. The die-profile optimization system which was developed in this paper proves better effect for the die-cavity design. This chapter also offers a reverse design methodology for investment die casting using ProCAST. In industry, the performance of the engine depends not only on shape, but also on the dimensions of the components. This process is difficult as super-alloy blade material cannot be easily machined. However investment casting is an ideal process for such net - shape components, but it still requires an accurate determination of the casting-die profile. In order to investigate and analyze the methods that affect the shape and dimensions of the turbine blade most, similar simulations have been conducted by ProCAST. By combining the methods of simplifying grid files and quick sorting, the efficiency of sorting and matching can be largely improved. Furthermore, the mold die cavity anti-deformation system can be easily built by utilizing that reverse design methodology. The optimized die profile for investment casting can be established with ProCAST.

This chapter discusses the structural analysis aspects of aerospace structures focusing on vibration induced fatigue. Metal Fatigue of dynamically loaded structures is a very common phenomenon in engineering practice. Several aerospace structures are used in environments where they experience dynamic loading. Furthermore, when the loading is dynamic, the response of the structure is affected by the structural resonances. Thus, the structural response to the loading will amplify at the regions of resonance. As a case study, an antenna (12-59 V/UHF) integration process on helicopter is investigated where the importance of the dynamic response is highlighted. Flight tests and finite element analyses (FEA) are carried out to ensure the safety of the integration process. Furthermore, another important aspect of the structural analysis of aerospace structures is the testing strategy. In this study, the fatigue tests performed by uni-axial tests which is a procedure defined in military standards are compared to multi axial testing. The results are shown for a helicopter Chaff/Flare Dispenser Bracket. The military standards assume that cumulatively uniaxial testing has equal multi axial testing fatigue damage. In this study, the uni-axial fatigue tests were compared to multi axial fatigue tests which were performed by FEA simulations. Furthermore, the effects of various loading conditions and geometries were investigated. The study showed that the assumption of multi axial testing can be represented by uni-axial testing does not hold for various cases.

Surface treatments, such as laser peening, can increase the life of the component by generating compressive residual stresses on the surface. Laser peening of an already peened component, termed as re-peening, can further increase the fatigue life of the component. Re-peening has several applications in the aerospace industry. The huge population of ageing aircraft components is one such application, which can benefit significantly from the re-peening process. However, this process is not optimized for maximum fatigue life due to the presence of many design variables and the complex nature of the problem which requires a large number of experimental testing to reach conclusions. Therefore, a computationally efficient optimization strategy needs to be developed to conduct large-scale laser peening simulations for problems related to fatigue life, such as aircraft lug failure, a problem that requires consideration of component curvature and residual stress relaxation effects. Deciding the time to peen an already peened component (re-peening time) is another variable which makes the problem further complicated. The ultimate goal of this research is to construct the framework to predict the optimum parameters for maximum fatigue life on structural components. A two-step optimization strategy is adopted for the fatigue life optimization of an aircraft lug component. The strategy employs laser peening process parameters, residual stress relaxation, and re-peening schedule as design variables.

Laser peening is an advanced surface enhanced method which induces compressive residual stress on the critical regions of components prone to fatigue failure. However, the residual stresses relax under the fatigue loading conditions. Constitutive models have to be robust enough to predict the residual stress relaxation mechanism. Although tensile cold working increases the tensile yield strength, the compressive yield strength is reduced. As a result of this, a lower compressive load can relax the initial compressive residual stress. This phenomenon, termed as Bauschinger Effect, can be represented by an analytical stress-strain model to predict the relaxation effects based on the cold working of the material. Three dimensional finite element (FE) models are created to represent residual stress relaxation in a low cycle fatigue regime for Ti-6Al-4V material. The creation of the numerical model for simulating stress relaxation model involves two phases. The first phase is modeling the stress gradient effect which relates to the effect of cold working. This process utilizes a laser peening simulation model. The second phase is modeling the stress-strain response of the material by creating a mixed hardening model.

The aerospace industry has utilized high strength aluminum alloys to propel the production and manufacturing of advanced aerospace technology; however, welding of Aluminum and its constituent alloys introduce challenges which affect the structural integrity of the welded area if conducted indecorously. For this reason, Non-destructive Evaluation (NDE) of welded areas is conducted to ascertain defective regions to ensure structural integrity of the aerospace structure. NDE techniques are noninvasive and can determine whether the object contains irregularities, discontinuities, or flaws. Inspecting weld areas allows for cost reduction by detecting discontinuities in the early stages of manufacturing; consequently, reducing the time and money to rework the error and allows for the validation of sound welds. A variety of NDE techniques are available depending on the applications, each with its own advantages and disadvantages. Among the number of NDE techniques, radiography and ultrasonics are the most widely utilized for inspection of weld defects. In this study, a detailed analysis was thereby conducted to ascertain the critical phased array ultrasonic testing (PAUT) parameters for the detection of weld defects, more specifically with Friction-Stir-Welding (FSW). Consequently a comparison to X-ray radiography is also included. It was observed that both techniques produced similar detection results for defects in the range of 1.0mm; however, it was found that PAUT was the only technique able to discover defects in the range of 0.15 mm.

An applicable procedure for design, optimization, and manufacturing of a remotely piloted helicopter (RPH) is studied analytically and experimentally. The procedure is presented in four main phases of conceptual design, detailed design, manufacturing and assembly, and test and reliability analysis. Using this procedure, an RPH case study, called Parvan, for an arbitrary traffic monitoring designed and manufactured. Different subsystems of an RPH such as rotor assembly, engine and fuel systems, power transmission system, and control system for the proposed design are studied at each phase accordingly. Final ly, 3-D multiplatform software of CATIA is used to simulate the RPH and making the ful l-scaled prototype. The successful performance of Parvan at hover, climb, and forward flight modes showed the effectiveness of the proposed procedure in developing an RPH.

This chapter reviews the advances in the materials for applications in structures of both subsonicas well as supersonic aircrafts. An account of the operating and ambient environmental conditions during flight is first given and the resulting material requirements have been discussed. Design relationships have been established taking into consideration the loading conditions and the strength requirements. In particular, the aircraft skin temperatures at various mach numbers have been taken into account for selecting appropriate structural materials for both subsonic and supersonic aircrafts; and consequently various aerospace aluminum alloys, titanium alloys, superalloys, and composites have been suggested. Finally, a new materials-selection chart is presented which would help aerospace designers to select appropriate materials for structural application in subsonic and supersonic aircrafts.

Advanced materials with optimized properties are essential in addressing the stringent requirements imposed by future aerospace vehicles. The discovery of carbon nanotubes and their desirable properties, as recognized in diverse scientific disciplines, have therefore identified these materials as expedient candidates for usage in aerospace applications. Given the limited number of overarching review articles encapsulating the usefulness of carbon nanotubes in aerospace sciences, this chapter explores the prospective applications of these materials in aerospace applications with their possible implementation on future aircraft, unmanned aerial vehicles and rotorcraft. The prospects of carbon nanotube usage intended for fuselage/satellite weight reduction, aircraft icing mitigation, lightning protection for aircraft, and future space launch are further explored. Conclusively, present challenges associated with successful implementations of these materials and existing obstacles preventing their safe integration in the aerospace industry are revisited, outlined and discussed.

The current use of fuel efficient and environmentally friendly aircraft is only possible by the development of innovative lightweight constructions and the use of lightweight materials, such as carbon fiber reinforced plastics. With the rising demand on fiber reinforced components in the aerospace industry new production processes have been built up. However, current production technologies for composites cause higher costs and obtain longer process cycle times in comparison to the manufacturing processes of metals. Moreover raw materials, such as carbon fibers and resin, and semi-finished products are very expensive. In contrast to this and compared with other manufacturing technologies for fiber reinforced plastics Sheet Molding Compound compression processes are characterized by cost efficiency, high productivity, the option of full automation and the possibility for the realization of complex shapes and integrated functions. However there are also some disadvantages like a low level of stiffness and strength in comparison to continuous fiber reinforced plastics. The reasons for these facts are the short fibre length, a lower fibre-volume fraction and an isotropic fibre distribution. Consequently, the combination of sheet moulding compound and pre-impregnated, tailored carbon fibre reinforcements in an one-shot compression moulding and curing process merges the advantages of both groups of composite materials. Therefore the creation of load-bearing, complex, functional and autoclave-quality parts without an autoclave can be realised. In this chapter, this innovative technology and its potentials are presented. This paper will also deal with the resulting material characteristics.

Automated fiber placement (AFP) machines can steer the fibers/tows to make the so-called variable stiffness (VS) composites. They allow the designers to fully exploit the directional properties of composite materials to tailor the internal load distribution and improve the structural performance. VS composites have been shown to be very promising in the design optimization of composite panels and shells for buckling and post-buckling performance and consequently for further reducing the mass of future aerospace structures. In this chapter, the buckling performance improvement of VS composite cylinders with circular and elliptical cross sections is investigated. A metamodeling based design optimization (MBDO) method is presented to maximize the buckling performance of VS composite cylinders compared with their constant stiffness (CS) designs. The structural improvement mechanism via stiffness tailoring in a VS composite cylinder is also presented and discussed. The effects of different parameters including the cylinders’ aspect ratio and size as well as the percentage of the steered plies in the laminate are also investigated.

Metal Matrix Composites (MMCs) are materials which consist of a metal alloy reinforced with ceramic particles or fibers. These materials possess a very high strength to weight ratio, good resistance to impact and wear, and a number of other properties which make them attractive for use in aerospace and defense applications. For example, MMCs have being extensively used for structural tubing in the space shuttle orbiter, the antenna mast of the Hubble Space Telescope, control surfaces and propulsion systems for aircraft, and tank armors. However, difficulties arise when joining those materials with fusion welding and impose limitations on the size of MMC components. Melting of the material leads to formation of an undesirable phase when molten Aluminum (Al) comes into contact and reacts with the reinforcement. This phase forms a strength depleted zone along the jointline. Friction Stir Welding (FSW) is a relatively joining technique, developed at The Welding Institute (TWI) in 1991. Because FSW occurs below the melting temperature of many metal alloys, it precludes formation of deleterious phases and results in a more favorable welded microstructure that is closer to that of the parent material. At NASA, this process was first applied to weld the super lightweight external tank for the space shuttles program. Today FSW is employed to join structural components in Delta IV, Atlas V, and Falcon IX rockets as well as NASA’s Orion Crew Exploration Vehicle and Space Launch System. Currently, FSW researchers are interested in extending the application of the process to new materials which are difficult to weld using conventional fusion techniques, such as MMCs. Rapid wear of the welding tool in FSW of MMCs is a consequence of the large discrepancy in hardness between the steel tool and the reinforcement material. This chapter summarizes the challenges encountered when joining MMCs to themselves or to other materials in structures. Specific attention is paid to the influence of the process variables for FSW on the wear process. A phenomenological model of the wear process was established based on the rotating plug model of FSW. The effectiveness of tool materials with high hardness (e.g. Tungsten Carbide, high speed steel, and tools with diamond coatings) in resisting abrasive wear is also considered. In-process force, torque, and vibration signals are analyzed to determine the feasibility of in situ monitoring of tool shape changes as a result of wear. One advantage of this model is that its successful implementation would eliminate the need for off-line evaluation of tool condition during joining. Monitoring, controlling, and reducing tool wear in FSW of MMCs are critical to full application of these materials in aerospace structures where they would be of most benefit. The work presented in this chapter can be further extended for machining of MMCs, where the wear of the tool materials is also a limiting factor.