CERES :: Browsing by Author "Guo, Shijun J."

Browsing by Author "Guo, Shijun J."

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Open Access
Adaptive control of a nonlinear aeroelastic system
(Elsevier Science B.V., Amsterdam., 2011-07-01T00:00:00Z) Li, Daochun; Xiang, Jinwu; Guo, Shijun J.
Aeroelastic two-dimensional wing section with both trailing-edge (TE) and leading-edge (LE) was investigated in this paper through numerical simulation in time domain. Structural stiffness and damping in pitch degree of freedom were represented by nonlinear polynomials. Open-loop limit cycle oscillation (LCO) characters of two examples were studied, and flutter boundaries with initial conditions were obtained. Parametric uncertainties in both pitch stiffness and damping were considered in the design of adaptive control laws to depress LCOs. Firstly an adaptive controller based on partial feedback linearization was derived for the wing section with a single TE control surface. Secondly a structured model reference adaptive control law was designed for the aeroelastic system with both TE and LE control surfaces. The results show that the designed control laws are effective for flutter suppression, and that considering damping uncertainty has positive effect on flutter control. It may reduce convergent time or increase flutter speed.
Open Access
Aerodynamic analysis and experiment of a micro flapping wing rotor
(Cranfield University, 2015-03) Li, Hao; Guo, Shijun J.
This project is aimed at developing a bio-inspired flyable micro/nano aerial vehicle (MAV) of high agility and performance capable of vertical take-off and landing and hovering (VTOLH). To achieve the aim, a novel flapping wing rotor (FWR) concept invented by Dr. Guo has been adopted, which is ideal for MAVs of sub 60 gm and especially for nano scale of sub 5 gm according to aerospace industry’s definition. The advantages and potential of the FWR concept for MAV development has been studied consistently by Dr. Guo’s research team in the last five years. However making a flyable micro FWR model especially in sub 5gm and demonstrate its VTOLH feasibility remains as a big challenge and has not been achieved in previous projects. To meet the above objective, the first achievement in the project is the successful design, build and test of a flyable micro FWR model (FWR-EX1) of only 3 gm based on off-the-shelf available micro motor. The key breakthrough is to achieve the necessary large aeroelastic twist of the flapping wing during the upstroke in an adaptive manner for structural and aerodynamic efficiency. To achieve the next objective for design and performance improvement, study has also been focused on deeper scientific understanding and analysis of the FWR mechanisms. Attention has therefore been paid to a systematic study on aerodynamic modelling and efficiency of the FWR. The method is based on a revised quasi-steady aerodynamic model that combines the theoretical method and experimental data. The numerical results of the revised quasi-steady aerodynamic model are in agreement with existing results obtained via CFD methods. Based on the model and analysis, the optimal kinematics for the FWR has been determined. Subsequently a comparison of the FWR aerodynamic efficiency was made with two other most studied configurations of MAVs, the insect flapping wing and rotorcraft ... [cont.].
Open Access
Aeroelastic dynamic response and control of an airfoil section with control surface nonlinearities
(Elsevier Science B.V., Amsterdam, 2010-10-31T00:00:00Z) Li, Daochun; Guo, Shijun J.; Xiang, Jinwu
Nonlinearities in aircraft mechanisms are inevitable, especially in the control system. It is necessary to investigate the effects of them on the dynamic response and control performance of aeroelastic system. In this paper, based on the state-dependent Riccati equation method, a state feedback suboptimal control law is derived for aeroelastic response and flutter suppression of a three degree-of-freedom typical airfoil section. With the control law designed, nonlinear effects of freeplay in the control surface and time delay between the control input and actuator are investigated by numerical approach. A cubic nonlinearity in pitch degree is adopted to prevent the aeroelastic responses from divergence when the flow velocity exceeds the critical flutter speed. For the system with a freeplay, the responses of both open- and closed-loop systems are determined with Runge-Kutta algorithm in conjunction with Henon's method. This method is used to locate the switching points accurately and efficiently as the system moves from one subdomain into another. The simulation results show that the freeplay leads to a forward phase response and a slight increase of flutter speed of the closed-loop system. The effect of freeplay on the aeroelastic response decreases as the flow velocity increases. The time delay between the control input and actuator may impair control performance and cause high-frequency motion and quasi-periodic vibration.
Open Access
Aeroelastic investigation of conventional fixed wings and bio-inspired flapping wings by analysis and experiment.
(2018-09) Li, Hao; Guo, Shijun J.
In this thesis, the structure and aeroelastic design, analysis and optimization of conventional fixed wing is firstly addressed. Based on the study results of conventional fixed wing, the study then focuses on the more complicated aerodynamics and aeroelasticity of flapping wing Micro Air Vehicles (MAV). A Finite Element (FE) model of a composite aircraft wing is firstly used as case study for the aeroelasticity of conventional fixed wing. A MATLAB-NASTRAN interfaced optimization platform is created to explore the optimal design of the wing. Optimizations using the developed platform show that 13% of weight reduction can be achieved when the optimization objective is set to minimize wing weight; and 18.5% of flutter speed increase can be achieved when aeroelastic tailoring of composite laminate layups is carried out. The study results further showed that the most sensitive part of the wing for aeroelastic tailoring is near the engine location, which contributes to the majority of flutter speed increment for optimization. In order to facilitate the structural design of non-circular cross section fuselage of Blended-Wing-Body (BWB) aircraft, an analytical model of 2D non-circular cross section is developed, which provides efficient design and optimization of the fuselage structure without referring to FE models. A case study based on a typical BWB fuselage using the developed model shows that by optimizing the fuselage structure, significant weight saving (17%) can be achieved. In comparison with the conventional fixed wing, insect flapping wings demonstrate more complicated aerodynamic and aeroelastic phenomena. A semi-empirical quasi-steady aerodynamic model is firstly developed to model the unsteady aerodynamic force of flapping wing. Based on this model, the aerodynamic efficiency of a Flapping Wing Rotor (FWR) MAV is investigated. The results show that the optimal wing kinematics of the FWR falls into a narrow range of design parameters governed by the dimensionless Strouhal number (St). Furthermore, the results show that the passive rotational of the FWR converges to an equilibrium state of high aerodynamic efficiency, which is a desirable feature for MAV applications. Next, the aerodynamic lift coefficient and efficiency of the FWR are calculated and compared with typical insect-like flapping wings and rotary wing. The results show that the aerodynamic efficiency of FWR in typical wing kinematics is higher than insect-like flapping wings, but slightly lower than the conventional rotary wing; the FWR aerodynamic lift coefficient (CL) surpassed the other wings significantly. Based on the numerical results, the study then continued to experimental investigations of the FWR. A prototype FWR model of weight 2.6g is mounted on a load cell to measure the instantaneous lift production. The kinematics of the wing is captured using high speed camera. Aeroelastic twist of the wing is measured using the resulting wing motion. Analyses by CFD and the quasi-steady aerodynamic model is then carried out and compared with experimental results. The study revealed that passive twist of the FWR wing due to aeroelastic effects forms desirable variations of wing Angle of Attack (AoA), which improves the aerodynamic performance of FWR. The results of the thesis provide guidance for structural, aerodynamic and aeroelastic design, analysis and optimization of conventional fixed wing, as well as bio-inspired flapping wing MAVs.
Open Access
Analysis and Experiment of an Ultra-light Flapping Wing Aircraft
(Cranfield University, 2013-08-21) Hu, Xiaowei; Guo, Shijun J.
II Inspired by flying animals in nature especially birds, human has designed and attempted to achieve man-powered flapping wing aircraft in very early aviation history. Limited by the understanding of the aerodynamic theory and materials in practise, the bird-like aircraft remains as a dream and ambition for over a contrary. As the relevant knowledge and technology are fast developing in the last decade, the research topic becomes attractive again with encouraging results from a few full scale aircraft flight tests. Although it is suspected that a manned scale flapping wing may not be as efficient as fixed wing, the unique advantages of high manoeuvrability and short take-off and landing capability will keep flapping wing as one of the most potential type of personal and aerobatic aircraft in the future market. The aim of this project is to investigate into the feasibility and development of a bio-inspired bird-like man-powered ultra-light flapping wing aircraft (ULFWA). The project is based on analytical and experimental study of a scaled model taking an existing hang glider as the baseline airframe. Based on the characteristics of flying animals in nature and manmade hang glider properties, this thesis focuses its study on evaluating the feasibility and analysis of primarily a human powered aircraft. For this purpose, there are four main features as guidance in the ULFWA design. Firstly the flapping frequency was limited to below 2Hz. Secondly the hang glider airframe was adapted with a simple flapping mechanism design. Thirdly the flapping wing stroke and kinematics has been kept with the simplest and resonant movement to achieve high mechanical efficiency. Finally the wing structure has flexible rib of chord wise unsymmetrical bending stiffness to offset the aerodynamic lift loss in upstroke. An engine powered mechanism design was also studied as additional option of the ULFWA. The initial design and aerodynamic calculation of the ULFWA was based on the hang glider data including dimensions, MTOW (226 kg) and cruising speed. The unsteady aerodynamic lift and thrust forces were calculated based on Theodorsen’s theory and unsteady panel method in 2D and extended to 3D using strip theory. A set of optimal flapping kinematic parameters such as amplitude and combination of the heaving and pitching motion of the 2D wing section were determined by calculation and comparison in the limited range. Considering the maximum power and lag motion that human could achieve, the flapping frequency in the ULFWA design is limited to 1Hz. This slow motion leads to a much lower propulsive efficiency in terms of the optimum Strouhal Number (St=0.2-0.4), which was used as the design reference. Mechanism and structure design with inertia force calculation was then completed based on the kinematics. This led to the evaluation of power requirement, which was divided into two components, drag and inertia forces. The results show that the ULFWA needs minimum 2452.25W (equals to 3.29Bhp) to maintain sustainable cruise flight. In order to demonstrate the ULFWA flapping mechanism and structure design, a 1:10 scaled model with two pairs of wings of different stiffness were built for testing and measurement. Two servomotors were used as to simulate human power actuation. With this model, simplified structure and one of mechanism designs was shown. Four experiments were carried out to measure the model’s lift and thrust force. Because of the limited response of the servo motors, the maximum flapping frequency achieved is only 0.75 Hz in the specified flapping amplitude which is close to reality and has improvement margin. By reducing the flapping amplitude, the frequency can be increased to gain higher thrust. Although it is fund that the result from scaled model test is a little lower than theoretical result, it has demonstrated the feasibility and potential of human powered flapping wings aircraft.
Open Access
Analysis of composite wing structures with a morphing leading edge
(Cranfield University, 2011-01) Morishima, Ryoko; Guo, Shijun J.
One of the main challenges for the civil aviation industry is the reduction of its environmental impact. Over the past years, improvements in performance efficiency have been achieved by simplifying the design of the structural components and using composite materials to reduce the overall weight. These approaches however, are not sufficient to meet the current demanding requirements set for a „greener‟ aircraft. Significant changes in drag reduction and fuel consumption can be obtained by using new technologies, such as smart morphing structures. These concepts will in fact help flow laminarisation, which will increase the lift to drag ratio. Furthermore, the capability to adapt the wing shape will enable to optimise the aerodynamic performance not only for a single flight condition but during the entire mission. This will significantly improve the aircraft efficiency. The current research work has been carried out as part of the European Commission founded Seventh Framework Program called „Smart High Lift Device for the Next Generation Wing‟ (SADE), which main aim is to develop and study morphing high lift devices. The author‟s investigation focused on developing a design concept for the actuation mechanism of a morphing leading edge device. A detailed structural analysis has been carried out in order to demonstrate its feasibility.In the first phase of the research the attention was directed on the preliminary design and analysis of the composite wing box. The parameters of the key structural components, such as skin, spars, ribs and stringers were set to satisfy the static stress and buckling requirements. Moreover, numerical and experimental studies were conducted to analyse the static failure and buckling behaviour of two typical composite wing structural components: a spar section and a web and base joint assembly. In the second stage of the research, a design for the morphing leading edge actuation mechanism was developed. The actuation system was designed in such a way that the target shape was reached with minimum actuation force demand. A geometrical nonlinear FE analysis was conducted to simulate the leading edge morphing deflection and ensure that structural strength requirements were satisfied. Furthermore, the behaviour of the skin integrated with the internal actuation mechanism was modelled under the aerodynamic pressure, at different flight conditions and gust loads, in order to prove that the proposed actuation system can compete with the conventional rigid rib. This study demonstrated that a feasible morphing leading edge design for a next generation large aircraft wing can be achieved. Developing the readiness of this technology will have a significant impact on aircraft efficiency and considerable contribution towards a more environmental friendly aviation.
Open Access
Buckling and post-buckling of a composite C-section with cutout and flange reinforcement
(Elsevier, 2013-12-31) Guo, Shijun J.; Li, Daochun; Zhang, Xiang; Xiang, Jinwu
This paper presents an investigation into the effect of cutout and flange reinforcement on the buckling and post-buckling behaviour of a carbon/epoxy composite C-section structure. The C-section having a cutout in the web is clamped at one end and subjected to a shear load at the other free end. Three different stiffener reinforcements were investigated in finite element analysis by using MSC Nastran. Buckling load was predicted by using both linear and nonlinear FE analysis. Experiments were carried out to validate the numerical model and results. Subsequently post-buckling analysis was carried out by predicting the load–deflection response of the C-section beam in nonlinear analysis. Tsai-Wu failure criterion was used to detect the first-play-failure load. The effect of circular and diamond cutout shape and effective flange reinforcements were investigated. The results show that the cutout and reinforcement have little effect on the buckling stability. However an L-shape stiffener to reinforce the C-section flange can improve the critical failure load by 20.9%.
Open Access
Coupled piezoelectric fans with two degree of freedom motion for the application of flapping wing micro aerial vehicles
(Elsevier Science B.V., Amsterdam., 2008-10-03T00:00:00Z) Chung, H. C.; Lal Kummari, K.; Croucher, S. J.; Lawson, Nicholas J.; Guo, Shijun J.; Huang, Zhaorong
Piezoelectric fans consisting of a piezoelectric layer and an elastic metal layer were prepared by epoxy bonding and a coupled flexible wing was formed by a pair of carbon fibre reinforced plastic wing spars and polymer skin attached to two piezoelectric fans. Two sinusoidal voltages with phase differences were then used to drive the coupled piezoelectric fans. High speed digital cameras were used to characterize the two degree of freedom (DOF) motion of the wing and these results were compared to finite element model of the wing and the coupled piezoelectric fans. It has been observed that the phase delay between the driving voltages applied to the coupled piezoelectric fans plays an important role in the control of the flapping and twisting motions of the wing and this set-up has the potential for application to the control of flapping wings for micro aerial vehicles. (c) 2008 Elsevier B.V. All rights reserved.
Open Access
Cutout reinforcements for shear loaded laminate and sandwich composite panels
(Springer Science Business Media, 2008-06-30T00:00:00Z) Guo, Shijun J.; Zhou, L.; Cheung, C. W.
This paper presents the numerical and experimental studies of shear loaded laminated and sandwich carbon/epoxy composite panels with cutouts and reinforcements aiming at reducing the cutout stress concentration and increasing the buckling stability of the panels. The effect of different cutout sizes and the design and materials of cutout reinforcements on the stress and buckling behaviour of the panels are evaluated. For the sandwich panels with a range of cutout size and a constant weight, an optimal ratio of the core to the face thickness has been studied for the maximum buckling stability. The finite element method and an analytical method are employed to perform parametric studies. In both constant stress and constant displacement shear loading conditions, the results are in very good agreement with those obtained from experiment for selected cutout reinforcement cases. Conclusions are drawn on the cutout reinforcement design and improvement of stress concentration and buckling behaviour of shear loaded laminated and sandwich composite panels with cutouts.
Open Access
Cutout shape and reinforcement design for composite C-section beams under shear load
(Elsevier Science B.V., Amsterdam., 2009-04-30T00:00:00Z) Guo, Shijun J.; Morishima, R.; Zhang, X.; Mills, A.
This paper reports a study of the performance of two forms of cutout and various edge reinforcements in a composite C-section beam under static shear load. Firstly, cutout shape effect on stress concentration was studied. This was followed by a comparative study of a range of reinforcement doublers, which were 20 mm wide rings made of a steel alloy or composite laminate, or by a novel fibre tow placement technique. The comparisons are made in terms of the stress and strain reductions at a hot spot at the cutout edge. Good agreement between the numerical and test results has been achieved for different cutout shapes and reinforcements, and this study has demonstrated that the cutout induced stress concentration can be reduced significantly by appropriate cutout shape and edge reinforcements. The stress reduction magnitude is found to be strongly related to the stiffness of the reinforcement rings. The diamond-shaped cutout and the fibre tows reinforcement show clear advantages over the widely adapted circular cutout and laminated ring reinforcement. These findings should contribute to future design improvement of composite aircraft structures in similar shape and loading conditions.
Open Access
Design & modelling of a composite rudderless aeroelastic fin structure
(Cranfield University, 2010-11) Trapani, Matteo; Guo, Shijun J.
This thesis presents the study of a gapless and rudderless aeroelastic fin (GRAF) to enhance the directional stability and controllability of an aircraft. The GRAF concept was proposed and developed in the wake of previous research, targeted to improve flight performance and manoeuvrability, and to reduce fuel consumption and airframe weight. The study involved the subjects of aerodynamics, structural design and analysis, and flight mechanics. The work includes conceptual design, structural modelling, aeroelastic analysis and flight performance evaluation of a GRAF variant designed for a small subsonic Unmanned Aerial Vehicle (UAV). The Eclipse UAV, a platform designed by part time students at the Department of Aerospace Engineering of Cranfield University, was chosen as a case study. A new approach to design a more effective fin with an unconventional structural layout and novel techniques which have not been investigated in previous research is proposed. Despite the GRAF planform being similar to classical fin-hinged rudder configurations, it is provided with a flexible gapless control surface, kept as one continuous piece and integrated with the fin primary structure. With its fixed root and rudderless feature, the GRAF adopts an original method of operation. Its way of working relies upon an unconventional technique of combining morphing technology and aeroelastic effect. The morphable configuration is twisted to gain an aeroelastically beneficial effect to enhance the efficiency and manoeuvrability of the aircraft. This warping capability of the fin is the key role player enabling the GRAF surface to seamlessly generate the required aerodynamic forces. Unlike the conventional structures designed to be as rigid as possible to withstand the external loads, the GRAF will exploit its structure‟s flexibility to use the aeroelastically induced twist deformations for a self-adaptive warping behaviour and improve flight dynamic response and performance. In order to ensure the above features are achievable in practice, further study on the structural configuration was conducted. To achieve performance improvement, together with the original structural layout and aeroelastic effect exploitation, another three novel key components are investigated, proposed and introduced in the GRAF model. A structurally integrated actuation system, termed L-shape stringers device (LSS), is designed to transform actuator axial forces in spanwise distributed bending moments, to create seamless deformations of the trailing edge (TE) section. An innovative trailing edge joint, namely the swivel edge closure, is specifically designed to enhance the mobility and degrees of freedom of the trailing edge box. It is a revolutionary concept which, by virtually interrupting the structural integrity of the closed TE section, allows relative translation and rotation of the TE panels. Finally, it is the novel concept of the slot-connection that, whilst appearing to clamp the GRAF structure inside the slot, actually enables the design to increase the twist angle at the tip of the fin without overstressing the materials. In order to enhance the GRAF efficiency, a tailored design of the fin structure was conducted. A novel internal structure configuration integrated with the key components has been designed to be connected to a flexible cladding skin, rotating ribs and a load-carrying tubular beam all of which constitute the primary parts of the GRAF model. With the ultimate goal of a lighter tail version, the entire design has been made by using composite, light frames, in an engineering trade-off of stiffness, elasticity, weight and cost of both glass and carbon fibre laminates. The analysis via 2-D aerodynamic codes and FEA was conducted to assess and validate the GRAF model and the obtained performance. Static linear elastic analysis has been carried out to verify the structural layout of the novel design subject to strength and stiffness criteria in addition to the fin warping and cambering capabilities. Also an investigation of aeroelastic stability related to steady and unsteady aerodynamic conditions has been carried out during the model analysis phase. The study has shown that although the GRAF divergence and flutter margins are slightly smaller than those of the conventional fin, the design and performance requirements are satisfied within the very challenging objective of a lighter vertical tail structure.The dynamic analysis study has also demonstrated the beneficial effect obtained by damping yawing oscillations when such a self-adaptive structure, compared to a rigid one, can be operated under cross wind circumstances. The manufacturing feasibility and assembly of the GRAF structure has been explored with the construction of a 1:1 scale model of the fin prototype. The model has been used as concept demonstrator to assess the functionality of the introduced technical novelties, the ease of manufacturing and the structural weight of the final assembly.
Open Access
Development of piezoelectric actuated mechanism for flapping wing micro-aerial vehicle applications
(Maney Publishing, 2010-03-31T00:00:00Z) Lal Kummari, K.; Li, Daochun; Guo, Shijun J.; Huang, Zhaorong
A piezoelectric actuated two-bar two-flexure motion amplification mechanism for flapping wing micro-aerial vehicle application has been investigated. f(r)*A as an optimisation criterion has been introduced where f(r) is its fundamental resonant frequency of the system and A the vibration amplitude at the wing tip, or the free tip deflection at quasi-static operation. This criterion can be used to obtain the best piezoelectric actuation mechanism with the best energy transmission coefficient for flapping wing micro-aerial vehicle applications, and is a measurable quantity therefore can be compared with experimental results. A simplified beam model has been developed to calculate the fundamental resonant frequency for the full system consisted of piezoelectric actuator, motion amplification mechanism and the attached wing and the calculated values were compared with the measured results. A clear trend of the criteria f(r)*A varying with the two-flexure dimension, stiffness and setting angle have been obtained from the measured data and also the predicted results as a guideline for optimal design of the system.
Open Access
The effect of laminate lay-up on the flutter speed of composite wings.
(Professional Engineering Publishing, 2003-06-01T00:00:00Z) Guo, Shijun J.; Bannerjee, J. R.; Cheung, C. W.
This paper presents an analytical study on optimization of a laminated composite wing structure for achieving a maximum flutter speed and a minimum weight without strength penalty. The investigation is carried out within the range of incompressible airflow and subsonic speed. In the first stage of the optimization, attention has been paid mainly to the effect on flutter speed of the bending, torsion and, more importantly, the bending-torsional coupling rigidity, which is usually associated with asymmetric laminate lay-up. The study has shown that the torsional rigidity plays a dominant role, while the coupling rigidity has also quite a significant effect on the flutter speed. In the second stage of the optimization, attention has been paid to the weight and laminate strength of the wing structure, which is affected by the variation in laminate lay-up in the first stage. Results from a thin-walled wing box made of laminated composite material show that up to 18 per cent increase in flutter speed and 13 per cent reduction in weight can be achieved without compromising the strength. The investigation has shown that a careful choice of initial lay-up and design variables leads to a desirable bending, torsional and coupling rigidities, with the provision of an efficient approach when achieving a maximum flutter speed with a minimum mass of a composite wing.
Open Access
Efficient method for aeroelastic tailoring of composite wing to minimize gust response
(Hindawi, 2017-01-26) Yu, Yang; Zhengjie, Wang; Guo, Shijun J.
Aeroelastic tailoring of laminated composite structure demands relatively high computational time especially for dynamic problem. This paper presents an efficient method for aeroelastic dynamic response analysis with significantly reduced computational time. In this method, a relationship is established between the maximum aeroelastic response and quasi-steady deflection of a wing subject to a dynamic loading. Based on this relationship, the time consuming dynamic response can be approximated by a quasi-steady deflection analysis in a large proportion of the optimization process. This method has been applied to the aeroelastic tailoring of a composite wing of a tailless aircraft for minimum gust response. The results have shown that 20%–36% gust response reduction has been achieved for this case. The computational time of the optimization process has been reduced by 90% at the cost of accuracy reduction of 2~4% comparing with the traditional dynamic response analysis.
Open Access
Feasibility of an electrostatic energy harvesting device for CFCs aircraft
(Elsevier, 2015-02-14) Xie, Huiling; Huang, Zhaorong; Guo, Shijun J.; Torru, Ekiyor
A novel energy harvesting concept is proposed for treating local electrostatic energy produced on flying composite aircrafts. This work focuses on the feasibility research on collecting static charges with capacitive collectors. The existing energy harvesting system and the electrification of the typical carbon fibre composites (CFCs) aircraft has been reviewed. The detailed model experiments were then designed to characterize different configurations for electrostatic energy harvesting on aeroplane. In the lab, the static charge was produced by a corona discharging device, and a capacitor or a metal sheet was put in the electric field to collect the charges under four different configurations. After that, the rest results for these configurations were analysed, which is followed by the discussion about the results application on the aircraft. This work has proved that it is feasible to collect the local static electricity on flying aircraft, and it could provide a new direction of energy harvesting system in aviation field.
Open Access
Gust alleviation of a large aircraft with a passive Twist Wingtip
(Wiley, 2015-04-03) Guo, Shijun J.; Espinosa De Los Monteros, Jaime; Liu, Ying
This paper presents an investigation into the gust response and wing structure load alleviation of a 200-seater aircraft by employing a passive twist wingtip (PTWT). The research was divided into three stages. The first stage was the design and analysis of the baseline aircraft, including aerodynamic analysis, structural design using the finite element (FE) method and flutter analysis to meet the design requirements. Dynamic response analysis of the aircraft to discrete (one-cosin) gust was also performed in a range of gust radiances specified in the airworthiness standards. In the second stage, a PTWT of a length of 1.13 m was designed with the key parameters determined based on design constraints and, in particular, the aeroelastic stability and gust response. Subsequent gust response analysis was performed to evaluate the effectiveness of the PTWT for gust alleviation. The results show that the PTWT produced a significant reduction of gust-induced wingtip deflection by 21% and the bending moment at the wing root by 14% in the most critical flight case. In the third stage, effort was made to study the interaction and influence of the PTWT on the symmetric and unsymmetrical manoeuvring of the aircraft when ailerons were in operation. The results show the that PTWT influence with a reduction of the aircraft normal velocity and heave motion by 1.7% and 3%, respectively, is negligible. However, the PTWT influence on the aircraft roll moment with a 20.5% reduction is significant. A locking system is therefore required in such a manoeuvring condition. The investigation has shown that the PTWT is an effective means for gust alleviation and, therefore, has potential for large aircraft application.
Open Access
Harvesting energy from the dynamic deformation of an aircraft wing under gust loading
(2012-07-06) Pozzi, Michele; Guo, Shijun J.; Zhu, Meiling; Tribikram, Kundu
Weight reduction and maintenance simplification are high in the agenda of companies and researchers active in the aerospace sector. Energy harvesters are being investigated because they enable the installation of wireless sensor nodes, providing structural health monitoring of the aircraft without additional cabling. This paper presents both a weight-optimized composite wing structure and a piezoelectric harvester for the conversion of mechanical strain energy into electrical energy. Finite elements modelling was used for the minimum- weight optimisation within a multi-constraints framework (strength, damage tolerance, flutter speed and gust response). The resulting structure is 29% more compliant than the original one, but is also 45% lighter. A strain map was elaborated, which details the distribution of strain on the wing skin in response to gust loading, indicating the optimal locations for the harvesters. To assess the potential for energy generation, a piezoelectric harvester fixed to a portion of the wing was modelled with a multi-physics finite elements model developed in ANSYS. The time-domain waveforms of the strain expected when the aircraft encounters a gust (gust frequencies of 1, 2, 5 and 10 Hz were considered) are fed into the model. The effects of harvester thickness and size, as well as adhesive thickness, were investigated. Energy generation exceeding 10 J/m2in the first few second from the beginning of the gust is predicted for 100μ-thick harvesters. The high energy density, low profile and weight of the piezoelectric film are greatly advantageous for the envisaged application
Open Access
Improvement of a tail-plane strucural model using vibration test data.
(Elsevier Science B.V., Amsterdam, 2002-09-26T00:00:00Z) Guo, Shijun J.
To achieve the best structural model improvement using vibration test data, a major effort has been made to identify poor modelling regions as a guideline for subsequent model updating. The method presented and used in this paper is the energy error estimation method. In the method the difference between analytical and test data based energies at element scale is estimated to indicate any poor structural mass and stiffness modelling. As a result, poor modelling regions can be distinguished from those modelled correctly and the improvement of the original structural model can be carried out effectively and accurately. To demonstrate the application of this method, a full-scale tail-plane structure has been studied by using simulated "test" modes as a simulated case and using measured modes as a practical case. In both cases poor modelling regions of the original structural model have been accurately located. Subsequently, a significant improvement of the structural model with a reduction of average frequency error from original 2·2% down to 0·1% for the simulated case and from 4·6 to 1·8% for the practical case has been achie
Open Access
Modelling and analysis of thin-walled structures for optimal design of composite wing
(2017-05) Mohd Saleh, Siti Juita Mastura; Guo, Shijun J.
At present, the option for composite usage in aircraft components and the associated manufacturing process is largely based on experience, knowledge, benchmarking, and partly market driven. Consequently, a late realisation involving the design and manufacture, and an inevitable iterative design and validation process has led to high costs. The aim of this research is to develop a Knowledge-Based Optimisation Analysis Tool (K-BOAT) for optimal design of composite structures, subject to multi design constraints. Extensive study has been carried out on composite structure design, modelling, testing and analysis method to optimise a design of a composite wing panel during the preliminary design stage. This approach will allow the maximum knowledge input and interface between users (design engineers) with the design tool, rather than be left to the optimiser to provide a solution. The K-BOAT will build a set of parameters in the initial design, including the ratio of component dimensions, layers of different fibre angles, and bending-torsion coupling of a panel and a wing box. This framework offers a guideline for the design engineers to understand and expect the optimal solution of composite structures at the early design stage. This research focused on the optimal design of aircraft composite wing skin. The first level involved the initial analysis of the composite wing by using a low fidelity model based on thin-walled structural analysis method. The second level focused on the optimal design of the wing skin using the analytical method and validation using the high fidelity finite element (FE) method. In-house computing programs and commercial software are used for this level of study. In the third level, the FE model has been used to present a baseline structure to perform further detailed analysis and optimisation. The study is related to an industrially funded project. A case study of a practical wing structure in the project has indicated an improvement in aircraft aeroelastic stability by 30.5% from the initial design. Validation of the real industrial application proved that K-BOAT is applicable to the conceptual and preliminary phases in aircraft design.
Open Access
Modelling, simulation and optimal control for an aircraft of aileron-less folding wing
(Wseas, 2008-10-31T00:00:00Z) Jiewang, Zheng; Guo, Shijun J.; Li, Wei
The purpose of this paper is to discuss the method of modeling and control system design for a loitering aircraft of aileron-less folding wing. A nonlinear model of the aircraft was established, and then linearized by small disturbance method. The lateral-directional stability augmentation options were analyzed through the root locus plots. The pole placement method based on linear quadratic regulator (LQR) technology was used to achieve desirable dynamic characteristics. In the analysis, the state parameters which represent rapid oscillation states of the aircraft such as roll rate and yaw rate were set as primary control parameters in the inner loop. The states oscillated slowly such as rolling angle and yaw angle were set as main control parameters in the outer loop. Based on the self-organizing fuzzy control algorithm, the aircraft can be controlled to fly in a desired path. Two types of course control plan were investigated and verified. The results show that the control plans are feasible and the control system is adequately robust to meet the requirements of the course control.