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Aerospace Science and Technology
Journal Prestige (SJR): 0.796
Citation Impact (citeScore): 3
Number of Followers: 341  
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 1270-9638
Published by Elsevier Homepage  [3161 journals]
  • Quasi-3D higher-order shear deformation theory for thermal buckling
           analysis of FGM plates based on a meshless method
    • Abstract: Publication date: Available online 21 September 2018Source: Aerospace Science and TechnologyAuthor(s): Vuong Nguyen Van Do, Chin-Hyung LeeAbstractThis study introduces a new quasi-3D (three-dimensional) higher-order shear deformation theory (HSDT) capable of accounting for the through-thickness deformations with only four unknowns, which is suited to the numerical method for buckling analysis of functionally graded material (FGM) plates in thermal environments. The refined quasi-3D HSDT is incorporated into the improved meshless radial point interpolation method (RPIM) in order to scrutinize the thermal buckling responses of FGM plates. In the improved RPIM, the radial basis function is presented in a compactly supported form to build the shape functions without any fitting parameters. Parametric studies on the buckling behavior of FGM plates under various types of through-thickness temperature changes are conducted and the effects of temperature distribution on the plate surface are investigated. Results illustrate the accuracy of the proposed meshless method based on the refined quasi-3D HSDT and the improved RPIM for predicting the thermal buckling behavior of FGM plates.
  • Experiment and analysis on radiation properties of SiO
    • Abstract: Publication date: Available online 18 September 2018Source: Aerospace Science and TechnologyAuthor(s): C. Sun, H.F. Sun, Z.Y. Wang, X.L. XiaAbstractThe thermal control coatings (TCC) on the surfaces of lunar rover/landers plays an important role for radiation heat transfer to allow the detectors survive the extremely critical thermal environment on the moon. Optical Solar Reflector (OSR) coating is usually considered as a good candidate TCC due to its excellent radiation properties such as low solar absorptance and high infrared emittance. In this study, SiO2/Ag coating, commonly used as OSR, was investigated to study its radiation properties, combining experiment measurement with numerical simulations. Based on the self-made setup with the comparison method, the temperature dependence of the normal infrared emittance of the OSR coating was obtained, covering the temperature range from 326 K to 404 K, to simulate the temperature condition on the Moon in the daytime. The direction hemisphere spectral reflectance at the typical wavelength of 0.6328 μm, 1.34 μm and 3.39 μm was also investigated in details. It was found that the spectral reflectance of the OSR coating at visible wavelengths is much higher than that at infrared wavelengths. In addition, numerical simulation result showed good agreement with the experimental data, and the relative error is 4.49% at 0.6328 μm.
  • Computational aerodynamics analysis of a light sport aircraft: compliance
           study for stall speed and longitudinal stability certification
    • Abstract: Publication date: Available online 18 September 2018Source: Aerospace Science and TechnologyAuthor(s): S. Piedra, E. Martinez, C.A. Escalante-Velazquez, S.M.A. JimenezAbstractThe present work reports an aerodynamics analysis conducted to analyze the compliance of flight requirements of a new light sport aircraft (LSA) design in accordance with the ASTM F2245 and CO AV-27/12. This new design is termed Halcon II and is property of the Horizontec aircraft company. At the conceptual design stage, vortex lattice method (VLM) calculations are implemented to compute the polar curves of the isolated wing and the complete aircraft. After the conceptual design of the aerodynamic surfaces is almost completed, the computational fluid dynamics simulations (CFD) are implemented in OpenFoam to characterize the flow around the aircraft. The CFD simulations consider steady state conditions and a RANS model is used to describe turbulent fluid flow. Next, detailed calculations are performed to quantify aerodynamic forces and the pitching moment. The results obtained from the CFD simulations strongly support the results yielded by the VLM calculations. Moreover, it is found that the aircraft is statically stable at the most critical velocities during stall and cruise flight conditions.
  • Adaptive fault-tolerant cooperative guidance law for simultaneous arrival
    • Abstract: Publication date: Available online 17 September 2018Source: Aerospace Science and TechnologyAuthor(s): Guofei Li, Yunjie Wu, Pengya XuAbstractThe problem of simultaneous arrival is investigated in this paper for multiple interceptors under partial actuator effectiveness. A fault-tolerant cooperative guidance is presented, in which an adaptive law is designed for disposing of the uncertainty. The convergence analysis demonstrates that the simultaneous arrival could be achieved in fixed-time interval under the actuation failures. The upper bound of convergence time is not dependent on the initial conditions of interceptors. The simulation results validate the effectiveness of proposed cooperative guidance law satisfactorily.
  • A Hybrid Control Scheme for Attitude and Vibration Suppression of a
           Flexible Spacecraft using Energy-Based Actuators Switching Mechanism
    • Abstract: Publication date: Available online 14 September 2018Source: Aerospace Science and TechnologyAuthor(s): Milad Azimi, Ghasem SharifiAbstractA novel energy-based switching logic scenario to design a hybrid thruster/reaction wheel (RW) system is investigated for active vibration suppression of a flexible spacecraft embedded with collocated and Non-collocated configuration of Piezoelectric patches. The system model is obtained using Lagrangian formulation and the finite element method. The control system includes the attitude and vibration controller, designed by Extended Lyapunov Design (ELD)/Strain Rate Feedback (SRF) control technique to take advantage of the globally stabilized feedback control strategy. An attractive feature of the proposed dual-stage system is the switching time, which is model-based and depends on the rigid-flexible body dynamics including PZT actions. Based on this approach, the Multi-objective Genetic Algorithm (MGA) determines the switching point concerning fuel consumption, settling time and vibration energy. The obtained results using a comparative study show the capabilities of the combination of the RWs and thrusters for cost-effective, high precise attitude control and residual vibration suppression of flexible spacecraft in future missions.
  • An analytical study of sound transmission through stiffened double
           laminated composite sandwich plates
    • Abstract: Publication date: Available online 13 September 2018Source: Aerospace Science and TechnologyAuthor(s): Tao Fu, Zhaobo Chen, Hongying Yu, Zhonglong Wang, Xiaoxiang LiuAbstractThe main objective of this research work is focused on sound transmission loss analysis of stiffened double laminated composite sandwich plate structures subjected to plane sound wave excitation, wherein the laminated composite plates are composed of perfectly bonded functionally graded carbon nanotubes (CNTs) reinforced composite layers, and in each layer, the carbon nanotubes is uniform or functionally graded along with the thickness direction of structures, and three types of the CNTs distributions are studied. The extended rule of mixture is employed to determine the properties of the composite material. The compatibility of displacements on the interface between the plate and the stiffeners is employed to derive the governing equation of each stiffener. Then fluid–structure coupling is considered by imposing velocity continuity condition at fluid–structure interfaces. By using the space harmonic approach and virtual work principle, the sound transmission loss is described analytically. Since no existing results of sound insulation can be found for such composite material plate structure, comparison studies can only be made with the isotropic and laminated case. Good agreement is found from these comparison studies. Based on the developed theoretical mode, the influences of the volume fractions of CNTs, distribution type of CNTs, structural damping, lamination angle and the number of layers on sound transmission loss are subsequently presented.
  • Equivalent and simplification of nickel-based single crystal plates with
           film cooling holes
    • Abstract: Publication date: Available online 13 September 2018Source: Aerospace Science and TechnologyAuthor(s): Zhixun Wen, Yamin Zhang, Zhenwei Li, Zhufeng YueAbstractThe mechanical properties of orthotropic material thin-walled plate with close-packed film cooling holes were studied based on the equivalent solid material concept. The equivalent principals of plane stress conditions, plane strain conditions, generalized plane strain conditions and general stress conditions were considered. A simplification method for square and triangular penetration patterns was presented. Extensive numerical simulation results covering different ligament efficiencies, penetration patterns and stress conditions were provided for the directionally solidified superalloy and nickel-based single crystal superalloy to verify the feasibility of equivalent principals and simplification method. Three crystal orientations [001], [011] and [111] of nickel-based single crystal superalloy were analyzed. The stress concentration factors were obtained for the mechanical behavior analysis of cooled blade. The values of equivalent error are all less than 10% when the ligament efficiency is larger than 0.4. The tensile deformation, Mises equivalent stress and stress distribution under the same stress level show the crystal orientation correlation. The tensile deformation of three crystal orientations is: [001]> [011]> [111]. The maximum Mises equivalent stress of [001] orientation and [011] orientation around the film cooling hole are basically the same, and they are all less than [111] orientation. The maximum and minimum values of stress concentration factors are 3.49 and 1.82 respectively.
  • Disturbance observer-based finite-time attitude maneuver control for micro
           satellite under actuator deviation fault
    • Abstract: Publication date: Available online 11 September 2018Source: Aerospace Science and TechnologyAuthor(s): Jianzhong Qiao, Dafa Zhang, Yukai Zhu, Peixi ZhangAbstractPrecise and fast attitude maneuvers are a requirement in many future space missions. However, the actuator deviation fault of micro satellite is inevitable, which affects the precision of spacecraft pointing. To overcome this difficulty, this paper presents a novel attitude control scheme to compensate the actuator deviation fault and achieve the finite-time attitude maneuver. Moreover, the proposed control scheme is constituted by an outer attitude tracking loop and an inner angular velocity tracking loop according to dynamic features of micro satellite. In the outer loop, a novel virtual control law, namely a desired angular velocity command, is developed, under which the finite-time attitude tracking can be achieved. Furthermore, in the inner loop, an effective composite controller which consists of a terminal sliding mode controller and a finite-time disturbance observer (FTDO) is proposed. The aim of composite controller is to ensure that the angular velocity can track the virtual control law in finite time, where the FTDO is designed to estimate and compensate the actuator deviation fault in the feed-forward channel. Finally, simulation results are conducted to demonstrate the effectiveness of the proposed control scheme.
  • Multi-phase smoothed particle hydrodynamics modeling of supercooled large
           droplet dynamics for in-flight icing conditions
    • Abstract: Publication date: Available online 11 September 2018Source: Aerospace Science and TechnologyAuthor(s): Vahid Abdollahi, Wagdi G. Habashi, Marco FossatiAbstractUnderstanding the dynamics of a single large water droplet is needed for accurate simulations of the in-flight icing phenomenon. Obtaining information on the ratio of ejected to deposited water and the post-impact droplet distribution should improve the numerical modeling of the bulk of impinging droplets. In this study, a weakly compressible multi-phase Smoothed Particle Hydrodynamics (SPH) method with shifting algorithm and surface tension model is presented to simulate the single droplet dynamics. The validity of the approach has been proved by modeling the classical problems of Rayleigh–Taylor instability, dam break, and droplet formation by comparing against other numerical and experimental data in the literature. Finally, droplet impingement on a liquid film and dry solid surface has been simulated and compared against the experimental data. The effect of impact angle and film thickness on the crown formation is studied to demonstrate the importance of modeling SLD impingement for in-flight icing conditions.
  • Investigation on self-pressurization and ignition performance of nitrous
           oxide fuel blend ethylene thruster
    • Abstract: Publication date: Available online 11 September 2018Source: Aerospace Science and TechnologyAuthor(s): Xuesen Yang, Xin Hong, Wei DongAbstractIn this paper, a novel method is developed to analyze the thermal parameters of Nitrous Oxide Fuel Blend Ethylene and establish the saturated vapor pressure relationship with temperature from −80∼30 °C based on Peng-Robinson equation of state, vapor/liquid phase equilibrium relation and Helmholtz equation. The experiments were conducted to validate the reliability of thermodynamic model. The self-pressurization flow model and sectionalized-centralized parameter gasification model of pipeline are established and validated. A chemical reaction mechanism with 129-species 900-reaction of Nitrous Oxide Fuel Blend Ethylene is established and the flame kernel parameters and laminar flame speed are obtained by steady-state, quasi-one-dimensional reacting flows simulations. Laminar flame speed of nitrous oxide fuel blend ethylene was measured and compared with calculated flame characteristics.
  • Cooperative aerial lift and manipulation (CALM)
    • Abstract: Publication date: Available online 10 September 2018Source: Aerospace Science and TechnologyAuthor(s): Hossein Rastgoftar, Ella M. AtkinsAbstractThis paper proposes a novel paradigm for aerial payload transport and object manipulation by an unmanned aerial vehicle (UAV) team. This new paradigm, called cooperative payload lift and manipulation (CALM), applies the continuum deformation agent coordination approach to transport and manipulate objects autonomously with collision avoidance guarantees. CALM treats UAVs as moving supports during transport and as stationary supports during object manipulation. Constraints are formulated to assure sufficient thrust forces are available to maintain stability and follow prescribed motion and force/torque profiles. CALM uses tensegrity muscles to carry a suspended payload or a manipulation object rather than cables. A tensegrity structure is lightweight and can carry both the tension and compression forces required during cooperative manipulation. During payload transport, UAVs are categorized as leaders and followers. Leaders define continuum deformation shape and motion profile while followers coordinate through local communication. Each UAV applies input–output (IO) feedback linearization control to track the trajectory defined by continuum deformation. For object manipulation, the paper proposes a new hybrid force controller to stabilize quadcopters when smooth or sudden (impulsive) forces and moments are exerted on the system.
  • Dispersion curves for a natural fibre composite panel: experimental and
           numerical investigation
    • Abstract: Publication date: Available online 10 September 2018Source: Aerospace Science and TechnologyAuthor(s): Giuseppe PetroneAbstractThe design of orthotropic panels, in which some mechanical properties are associated to high degrees of uncertainty, such as for the natural fibres, is very arduous and often techniques that give local information are needed. In this paper a unidirectional flax–polyethylene panel is experimentally and numerically investigated. Experimentally, for a given distance source–receiver, the group velocities are obtained by processing the guided waves time-transient signals via a time-frequency transform leading to an estimation of the dispersion curves for different fibre directions and frequencies. Test results are then compared and used to validate numerical ones, obtained by means of both Finite Element and Wave Finite Element models.
  • Aircraft Icing Safety Analysis Method in presence of Fuzzy Inputs and
           Fuzzy State
    • Abstract: Publication date: Available online 8 September 2018Source: Aerospace Science and TechnologyAuthor(s): Jiaqi Wang, Zhenzhou Lu, Yan ShiAbstractFor lack of safety measure under the widely existing fuzzy inputs and fuzzy state in the aircraft icing process, a safety analysis model is established to quantify the safety degree of aircraft icing under the fuzzy inputs and fuzzy state. Three-fold contribution is included in the established model. Firstly, by analyzing the recognition capability and properties, the failure credibility model possessing wide recognition capability and self-duality property is selected to measure the safety degree under the fuzzy inputs and binary state. Secondly, by a strictly mathematical derivation, an easily extended definition of the failure probability under random inputs and fuzzy state is proved, on which the failure credibility model under the fuzzy inputs and fuzzy state is established finally. After a numerical example is used to illustrate the feasibility of the established failure credibility model under the fuzzy inputs and fuzzy state in view of different perspectives, a real application of the established model is completed for the safety analysis of an aircraft icing problem.
  • Neural network based online predictive guidance for high lifting vehicles
    • Abstract: Publication date: Available online 7 September 2018Source: Aerospace Science and TechnologyAuthor(s): Zhenhua Li, Xiangdong Sun, Chen Hu, Gang Liu, Bing HeAbstractIn this paper, a data-driven online entry guidance framework is proposed. Based on the proposed framework, a novel neural network based online predictive guidance algorithm for high lifting vehicles is developed, which combines the benefits of the existing predictive guidance and the neural network predictor. By introducing the neural network predictor, the proposed algorithm can effectively overcome the long-standing contradiction between guidance accuracy and real-time guidance of existing numerical predictive guidance methods. Take the augmented predictor-corrector guidance algorithm as guidance pattern, a large number of sample trajectory data can be generated by performing full envelop trajectory simulations with different perturbation terms. Based on the sample data, the mapping relationship between the real-time flight states of high lifting vehicles and guidance commands is approximated by multi-layer feedforward neural network. By substituting the off-line trained neural network predictor for the trajectory integrations of each guidance cycle in the augmented algorithm, the proposed algorithm can successfully realize the online precision guidance for high lifting vehicles. The simulation results for nominal and dispersed cases show that the proposed algorithm has better performance on real-time capability and robustness than the existing numerical predictive guidance methods, and it is suitable for engineering practice.
  • Experimental study on the accretion and release of ice in aviation jet
    • Abstract: Publication date: Available online 7 September 2018Source: Aerospace Science and TechnologyAuthor(s): Mathias Schmitz, Gerhard SchmitzAbstractIce formations in aircraft fuel systems pose a serious safety threat with potentially disastrous consequences, when restricting the fuel flow towards the engines. This is an ongoing challenge in the aerospace industry. In this work experimental studies have been performed to investigate the effects of temperature, flow rate and surface properties on the accretion and release of ice in flowing fuel. A test rig with a glass-windowed pipe has been employed to quantitatively measure the transient icing process under controlled conditions. The accreted ice exhibited soft and fluffy characteristics and was most likely the result of impinging solid ice particles that were entrained in the fuel flow. The ice particles were most sticky in a temperature range between −6 °C and −20 °C. The thickness of accreted ice decreased with roughness on aluminium surfaces and there was a significant reduction on polytetrafluoroethylene (PTFE) and polymethyl methacrylate (PMMA) in comparison to aluminium, copper or stainless steel surfaces. Comparison of the thickness of accreted ice with the ice adhesion strength reported in the literature showed a clear correlation. The experimental results will help to gain better understanding of the ice accretion process in flowing fuel and may serve as basis for design guidelines to minimize ice formation within an aircraft fuel system.
  • The efficiency of a pulsed detonation combustor–axial turbine
    • Abstract: Publication date: Available online 5 September 2018Source: Aerospace Science and TechnologyAuthor(s): Carlos Xisto, Olivier Petit, Tomas Grönstedt, Andrew Rolt, Anders Lundbladh, Guillermo PaniaguaAbstractThe paper presents a detailed numerical investigation of a pulsed detonation combustor (PDC) coupled with a transonic axial turbine stage. The time-resolved numerical analysis includes detailed chemistry to replicate detonation combustion in a stoichiometric hydrogen–air mixture, and it is fully coupled with the turbine stage flow simulation. The PDC–turbine performance and flow behaviors are analyzed for different power input conditions, by varying the system purge fraction. Such analysis allows for the establishment of cycle averaged performance data and also to identify key unsteady gas dynamic interactions occurring in the system. The results obtained allow for a better insight on the source and effect of different loss mechanisms occurring in the coupled PDC–turbine system. One key aspect arises from the interaction between the non-stationary PDC outflow and the constant rotor blade speed. Such interaction results in pronounced variations of rotor incidence angle, penalizing the turbine efficiency and capability of generating a quasi-steady shaft torque.
  • Trajectory optimization for accompanying satellite obstacle avoidance
    • Abstract: Publication date: Available online 5 September 2018Source: Aerospace Science and TechnologyAuthor(s): Qinglei Hu, Jingjie Xie, Xinfu LiuAbstractThis paper aims at the trajectory optimization problem of an accompanying satellite in the space station under multi-obstacles by model predictive control (MPC) with successive linearization. This optimization problem is reformulated as a mixed integer second-order cone programming (MISOCP) problem by considering the obstacles avoidance, and various physical constraints. More specifically, a novel variation weight cost function is established with the distance information of the accompanying satellite and obstacles, such that a multi-objective optimization performance is achieved by incorporating fuel consumption and time expenditure further. Through the successive linearization of the MPC scheme, the satisfaction of the multiple constraints is established. These results lead to a feasible solution with harsh constraints as well as the fast convergence to the optimal solution. The obstacles avoidance constraints are first implemented through the appropriate selection of initial values, and then solved by successive approximation with the MPC framework for the convergence of solution. Numerical simulation has been conducted to demonstrate the effectiveness and applicability of the proposed method.
  • Optimization strategy for a single-stage axisymmetric hub endwall in axial
           compressor by a modified transonic area rule
    • Abstract: Publication date: Available online 5 September 2018Source: Aerospace Science and TechnologyAuthor(s): Zhiping Li, Yafei Zhang, Tianyu Pan, Hanan Lu, Miao Wu, Jian ZhangAbstractThe secondary loss is the primary factor limiting the performance of highly loaded compressors. As inspired by the transonic area rule used in optimization of airplane fuselage, considering the similarities of the blockage effect induced by wing to fuselage and the blockage induced by blade to flow channel, the idea of using the transonic area rule method in compressor hub profile modification has been proposed and discussed. Some related analyses are investigated due to the failure of a direct application of the transonic area rule, which includes flow field analyses, the study of related parameters and a sensitivity evaluation of eleven control parameters on the adiabatic efficiency of the compressor. Then, two parameters (rotor maximum concave displacement ratio ΔZ‾R and magnification ratio of rotor hub leading edge l¯1R) which have great influence on the compressor efficiency are chosen to modify the transonic area rule and to establish a new optimization guideline in the optimization of compressor hub profile. After all, the developed optimization method is applied to the first stage of a four-stage embedded compressor with an achievement of 0.97% peak efficiency rise without any penalty of total pressure ratio. Based on the numerical results, the main reasons for the peak efficiency rise are also presented and discussed.
  • Aerodynamic characterization of a transonic axial flow compressor stage
           – with asymmetric tip clearance effects
    • Abstract: Publication date: Available online 5 September 2018Source: Aerospace Science and TechnologyAuthor(s): S. Satish Kumar, Dilipkumar Bhanudasji Alone, Shobhavathy M. Thimmaiah, Janaki Rami Reddy Mudipalli, Ranjan Ganguli, S.B. Kandagal, Soumendu JanaAbstractThis manuscript discusses the performance of a single-stage transonic axial flow compressor with non-uniform/asymmetric rotor tip clearance. Detailed steady and unsteady experimental measurements are carried out for a compressor with non-uniform clearance over the rotor. Measurements are focused at the peak clearance region of the compressor annulus where stall inception is likely to occur. A sector-based steady-state CFD simulation is performed on the compressor stage with uniform averaged rotor running clearance obtained from the blade growth estimates at design speed using structural analysis. The role of the tip leakage vortex on the stall dynamics of the compressor is elucidated. The stall events/disturbances occurring at every two rotor revolutions are shown using hotwire measurements. Varied flow features of the transonic compressor stage are discussed for different compressor speeds investigated. The level of clearance eccentricity/asymmetry existing in this compressor due to its complex spread does not degrade the overall compressor performance behavior.
  • Numerical study of pore-scale flow and noise of an open cell metal foam
    • Abstract: Publication date: Available online 5 September 2018Source: Aerospace Science and TechnologyAuthor(s): Chen Xu, Yijun Mao, Zhiwei HuAbstractThis paper studies numerically the three-dimensional pore-scale flow inside a single cell structure of an open cell metal foam and its aeroacoustic features. Since the Reynolds number based on the pore diameter is very low, the Navier-Stokes equations are solved directly to simulate the unsteady pore-scale flow. The permeability and pressure drop obtained from numerical simulations are compared with existing reference results. Numerical results reveal that the flow drag of the pore-scale structure is dominated by the pressure drag which is mainly caused by the flow separation, while the friction drag is much smaller than the pressure drag despite a very high surface area-volume ratio of the metal foam. The unsteady flow separation contributes primarily to the pressure drag and causes the self-noise of the metal foam, therefore suppressing the unsteady flow separation, e.g., by optimizing the cell structure of the metal foam, would reduce the drag and aerodynamic noise. The aeroacoustic features, such as noise sources, spectra and the directivity pattern, are investigated, and the results reveal a high correlation between the noise generation and the flow separation. The direction of the maximum sound pressure for the studied cell is parallel to the flow, which is different to flow separation from a cylinder where the direction of maximum radiation is perpendicular to the flow.
  • Multidisciplinary optimisation of bipropellant rocket engines using H2O2
           as oxidiser
    • Abstract: Publication date: Available online 3 September 2018Source: Aerospace Science and TechnologyAuthor(s): Adam Okninski, Jan Kindracki, Piotr WolanskiAbstractToday's rocket chemical propulsion systems for in-orbit use rely on toxic liquid rocket propellants. The need for environmentally-friendly spacecraft and upper-stage high-performance engines can be seen. This paper covers the topic of optimisation of Geostationary Transfer Orbit apogee bipropellant rocket engines, using highly concentrated hydrogen peroxide as oxidiser and kerosene as fuel. Performance, structural and heat transfer analyses are described. In particular a detailed mass model for hydrogen peroxide/kerosene rocket propulsion systems is discussed. Special care is given to bipropellant rocket systems using staged combustion configurations where a catalyst bed is utilised. The optimisation process was conducted using Matlab software. The ultimate goal of this work was the development of a tool for hydrogen peroxide/hydrocarbon bipropellant rocket propulsion system optimisation in terms of given requirements and design constraints. The presented software may enable the development of advanced, environmentally-friendly satellites and highly-efficient architectures utilizing storable green propellants, including small upper-stage propulsion systems.
  • Semi-analytical prediction of macroscopic characteristics of open-end
           pressure-swirl injector
    • Abstract: Publication date: Available online 3 September 2018Source: Aerospace Science and TechnologyAuthor(s): A. Kebriaee, Gh. OlyaeiAbstractAfter proposing a semi-analytical solution for swirl laminar flow, macroscopic characteristics of open-end pressure-swirl injector including discharge coefficient and spray cone angle are calculated. In the presence of air core of the axial region inside the injector, the laminar rotational flow equations are simplified, and with the assumption of the quasi-developed axial flow along the nozzle, the equations are iteratively solved employing separation of variables method. The accuracy of the proposed semi-analytical solution is compared by some numerical and experimental results on an open-end injector. The validity of quasi-developed flow defined in the present work is confirmed based on the results of numerical simulations in different axial cross-sections. Moreover, findings demonstrate that the spray cone angle is predicted with an accuracy of about 3.7% for different operating conditions of experimental tests. The calculation of pressure distribution along the liquid film illustrates that viscous effect is negligible in pressure drop in the injector and the discharge coefficient is dominantly dependent on the flow inlet condition in the nozzle. Calculation of discharge coefficient presented in this paper shows an acceptable agreement with observations for different tests. Deviation of the theoretical discharge coefficient is less than 4.37% for various case studies with empirical results.
  • Characteristics of the combustion chamber of a boron-based solid
           propellant ducted rocket with a chin-type inlet
    • Abstract: Publication date: Available online 3 September 2018Source: Aerospace Science and TechnologyAuthor(s): Binbin Chen, Zhixun Xia, Liya Huang, Likun MaAbstractThe scope of this study is to improve combustion performance of a boron-based solid propellant ducted rocket (SPDR). A numerical approach for the internal reaction flow in SPDR was built, and an improved ignition and combustion model of boron particles was used. Many integrated physical processes of boron ignition and combustion, such as the balance of (BO)n on particle surface at ignition stage, the evaporation and boiling processes of boron at combustion stage, were considered. A new configuration of SPDR with swirling flow was proposed. Moreover, numerical and experimental investigations were conducted, and the effects of equivalence ratio and swirling flow on combustion performance were analysed. Validation of the developed numerical approach showed good agreement between the approach's results and available experimental data in open literature and also those in this study. Experimental results indicated that combustion efficiencies increase from 88% to 95% with swirling flow.
  • Numerical investigation of shock train unsteady movement in a mixing duct
    • Abstract: Publication date: Available online 29 August 2018Source: Aerospace Science and TechnologyAuthor(s): Lei Lu, Yi Wang, Xiao-qiang Fan, Guo-wei Yan, Jing-wen PanAbstractSupersonic mixing layers and complex background waves exist in the mixing duct of the RBCC (Rocket-Based Combined Cycle) engine and supersonic-supersonic ejector. Different from the isolator with uniform incoming flow, the shock train in a mixing duct has special structures and motion characteristics. To study the characteristics of shock train forward movement in a mixing duct, a 2d unsteady numerical simulation was conducted. The results indicated that many unsteadiness factors existed in the background flow field such as the development of mixing layers, and multiple complex interactions. With backpressure linearly increasing over flow time, the shock train moved upstream in the process of which two forward movement patterns were found, namely slow pattern and jump pattern. Two patterns alternatively governed the whole forward movement, resulting in periodical changes of leading shock structures. The leading shocks were ‘stretched’ and ‘squeezed’ during the movement process, which was referred to as ‘spring’ characteristics in this research. Further analysis indicated that the characteristics of shock train movement were closely related to the background flow structure, or rather the surface pressure gradient. The shock train leading edges jumped forward in an adverse wall pressure gradient and crept forward in a favorable wall pressure gradient. The whole movement process consisted of 5 motion periods. In the 2nd motion period, speed of leading edges and wave-front Mach number varied, leading to the periodical changes of shock train structures. In the theoretical analysis, shock polar method was applied to clarify the intensity variation of reflection shocks in the shock train.
  • Plume flowfield and propulsive performance analysis of a rotating
           detonation engine
    • Abstract: Publication date: Available online 25 August 2018Source: Aerospace Science and TechnologyAuthor(s): Jian Sun, Jin Zhou, Shijie Liu, Zhiyong Lin, Wei LinAbstractRotating detonation engine is a promising detonation-based propulsion system. In this paper, a series of numerical simulations of H2/air RDE are performed to show the plume flowfield of a RDE in detail and to obtain the pressure distribution on the end faces of the inner cylinder and the outer wall. The thrust force of the RDE is calculated with/without considering the two end faces to show its effect on the propulsive performance. Then the effects of the total massflow rate and the environment pressure on the propulsive performance are discussed. Under high environment pressure condition, the inner cylinder end face brings a notable disadvantage of the propulsive performance due to the strong suction effects. This disadvantage disappears under low environment pressure condition. The fluctuation of the total thrust force is dominated by the outer wall end face. As the massflow rate increases, the engine's thrust force becomes more steady. Under low environment pressure condition, the engine's thrust force is very stable and its fluctuation can be neglected. Under high environment pressure conditions, the use of the conventional thrust calculation formula overestimates the average thrust force and underestimates the unsteadiness of the thrust force obviously.
  • Nonlinear robust control of tail-sitter aircrafts in flight mode
    • Abstract: Publication date: Available online 22 August 2018Source: Aerospace Science and TechnologyAuthor(s): Zhaoying Li, Lixin Zhang, Hao Liu, Zongyu Zuo, Cunjia LiuAbstractIn this paper, a nonlinear robust controller is proposed to deal with the flight mode transition control problem of tail-sitter aircrafts. During the mode transitions, the control problem is challenging due to the high nonlinearities and strong couplings. The tail-sitter aircraft model can be considered as a nominal part with uncertainties including nonlinear terms, parametric uncertainties, and external disturbances. The proposed controller consists of a nominal H∞ controller and a nonlinear disturbance observer. The nominal H∞ controller based on the nominal model is designed to achieve the desired trajectory tracking performance. The uncertainties are regarded as equivalent disturbances to restrain their influences by the nonlinear disturbance observer. Theoretical analysis and simulation results are given to show advantages of the proposed control method, compared with the standard H∞ control approach.
  • A PCA–ANN-based inverse design model of stall lift robustness for
           high-lift device
    • Abstract: Publication date: Available online 20 August 2018Source: Aerospace Science and TechnologyAuthor(s): Xinyu Wang, Shuyue Wang, Jun Tao, Gang Sun, Jun MaoAbstractThe concept of stall lift robustness for high-lift device (HLD) measures the stability of lift values under a series of angles of attack (AoA) around stall, which plays a significant role in flight safety. In this article, a stall lift robustness design with the consideration of aerodynamic constraints (stall AoA, average lift, etc.) is carried out, where design targets are set as a series of lift values on Lift-AoA curve. A Principle Component Analysis (PCA)–Artificial Neutral Network (ANN)-based inverse design model is introduced. The design targets are transformed by PCA for data dimension reduction. Then, the new set of design targets are input into the surrogate model of ANN, and corresponding geometry of new HLD is predicted. The ANN is constructed through database and sample points are screened considering lift unsteadiness. The design procedure is iterated to meet the design accuracy. The process of stall lift robustness design with the proposed model is discussed in this article, and the design results are validated by Detached-Eddy Simulation (DES).
  • Sensitivity assessment of optimal solution in aerodynamic design
           optimisation using SU2
    • Abstract: Publication date: Available online 17 August 2018Source: Aerospace Science and TechnologyAuthor(s): Guangda Yang, Andrea Da Ronch, Jernej Drofelnik, Zheng-Tong XieAbstractComputational fluid dynamics has become the method of choice for aerodynamic shape optimisation of complex engineering problems. However, the sensitivity of the final aerodynamic shape to numerical parameters has been largely underestimated to date. The purpose of this work is to investigate the influence that numerical parameters have on the optimisation results for two aerofoil problems (NACA 0012 and RAE 2822) in transonic flow, and to provide compact guidelines for best practice. Numerical parameters include: a) two parameterisation methods, Hicks–Henne bump functions and free-form deformation; b) numerical settings related to the tuning of each parameterisation method; and c) closure coefficients of Spalart–Allmaras (SA) turbulence model. All optimisations were performed using the open-source software tool SU2, and gradients were computed using the continuous adjoint method. It was found that: a) the optimisation result of NACA 0012 aerofoil exhibits strong dependence on all numerical parameters investigated, whereas the optimal design of RAE 2822 aerofoil is insensitive to those numerical settings; b) the degree of sensitivity reflects the difference in the design space, particularly of the local curvature on the optimised shape; c) the closure coefficients of SA model affect the final optimisation performance, raising the need for a good calibration of the turbulence model.
  • Decentralized formation flight via PID and integral sliding mode control
    • Abstract: Publication date: Available online 16 August 2018Source: Aerospace Science and TechnologyAuthor(s): Rebbecca T.Y. Thien, Yoonsoo KimAbstractThis paper solves a formation control problem for a group of vehicles such as UAVs on a directed network subject to constant and time-varying disturbances or commands. A celebrated PID control is first proposed for constant disturbance rejection. A set of necessary and sufficient conditions for the vehicles to achieve a desired formation is proposed. Then, integral sliding mode control is introduced to tackle unknown but bounded time-varying disturbances or commands. In addition, the control sensitivity with respect to the network topology is analyzed. Finally, a quadcopter formation platform is introduced and used to verify all the presented theoretical results experimentally.
  • Effect of low dorsal fin on the breakdown of vortices over a slender delta
    • Abstract: Publication date: Available online 14 August 2018Source: Aerospace Science and TechnologyAuthor(s): Xuanshi Meng, Feng Liu, Shijun LuoAbstractPrevious theoretical and experimental works show that the originally steady symmetric vortices over a flat-plate delta wing at zero side slip but high angles of attack are destabilized by the addition of a low dorsal fin, which renders the separation vortices asymmetric, unsteady, or non-conical. This paper examines the effect of the dorsal fin on the onset of vortex breakdown. A sharp-edged flat-plate delta wing with a 7.5 deg semi-apex angle is tested in a low-speed wind tunnel. The unsteady velocity and vorticity fields are mapped out at the 60% wing root chord location to quantify the vortex burst process caused by the addition of the low dorsal fins by using the laser Particle-Image-Visualization (PIV) technique. Two fin heights with the ratio of the local fin height to the local wing semi-span 0.3 and 0.6 are tested. The results demonstrate that the loss of global stability of the vortex configuration due to the addition of the low dorsal fin accelerates the onset of vortex breakdown and causes one of the vortices in the pair to burst periodically. The frequency and size of the burst are related to the height ratio of the fin.
  • Airbrake controls of pitching moment and pressure fluctuation for an
           oblique tail fighter model
    • Abstract: Publication date: Available online 14 August 2018Source: Aerospace Science and TechnologyAuthor(s): Wenyao Cui, Jian Liu, Yuanhao Sun, Qibing Li, Zhixiang XiaoAbstractThe pitching moments of a fighter model at an incidence of 32° without and with nine airbrakes are simulated by solving the unsteady Reynolds-averaged Navier–Stokes equations. The maximum reduction (60.22%) of the total pitching moment is obtained by the airbrake with deflection angle of 60° and length of 0.115 times the wing span. Furthermore, the unsteady flowfields are predicted by the improved delayed detached-eddy-simulation model. The breakdowns of the forebody and strake vortices are advanced, and the flow is blocked by the airbrake, jointly leading to a reduction of the pitching moment. Surprisingly, the pressure fluctuations on the oblique vertical tail are also attenuated due to the existence of the airbrake. A maximum reduction of overall sound pressure level by 11.8 dB is found near the leading-edge on the outer surface of the vertical tail due to the use of the 60T1 airbrake. The bursting vortices from the forebody and strake move towards the symmetry plane owing to the low-pressure region after the airbrake, weakening the unsteady interactions between the bursting vortices and oblique vertical tail.
  • Numerical investigations on lifting and flow performance of finger seal
           with grooved pad
    • Abstract: Publication date: Available online 14 August 2018Source: Aerospace Science and TechnologyAuthor(s): Xingyun Jia, Qun Zheng, Zhitao Tian, Yuting Jiang, Hai ZhangAbstractWe investigate non-contact finger seal owing to its potential to reduce the specific fuel consumption of a gas turbine engine by 2–3%. The compliance features combined with the non-contact characteristic permits the position adjustment of the finger seal to the rotor excursions without destroying the stability of the rotor-seal system. The grooved structures on the lift pad of the three-layer finger seal are proposed with a view to improve the lifting and leakage capacities. Two-way fluid-structure interaction (FSI) methods are used, and the results show that the grooved structures positively affect the sealing performance. First, an uneven clearance occurs because of the seal deformation. Next, the sealing flow field is changed owing to the grooved structure. Finally, the grooved structure is proven to improve the seal's lifting capability, and the vortices in the grooves improve the sealing performance. The vortices in the groove cavities contribute significantly to the improvement in the lifting and leakage capacities. The grooved structure changes the pressure distribution under the lift pad, and a larger force is generated on the bottom surface of the lifting pad.
  • Operation matching model and analysis between an air inlet and a
           compressor in an Air Turbo Rocket
    • Abstract: Publication date: Available online 13 August 2018Source: Aerospace Science and TechnologyAuthor(s): Yang Liu, Yonggang Gao, Xiaohang Pu, Jiang Li, Guo-qiang HeAbstractOperation matching between an air inlet and a compressor in an Air Turbo Rocket (ATR) is crucial to the full development of the advantages of engines. In this study, operation characteristic models of an air inlet and a compressor were established and verified through numerical simulation to gain the matched operation characteristics of the air inlet and the compressor in an ATR. The air inlet–compressor matching model was constructed on the basis of coupling relations, such as “flow equilibrium,” “interface parameter equilibrium,” and “pressure ratio equilibrium,” between the air inlet and the compressor in the ATR. Finally, the integrated 3D numerical simulation of the air inlet–compressor was verified. Results demonstrate that (1) the operation characteristic model of the air inlet can accurately calculate the effects of backpressure in the air inlet (Pb) on the total pressure recovery coefficients of the throat, the supersonic diffuser, and the whole air inlet in the ATR. Model prediction results are highly accurate as confirmed by 3D numerical simulation results in the following ranges: 2.75 < Ma < 3.75 and 0.5 atm < Pb < 3 atm. The maximum absolute error is less than 0.05. (2) The centrifugal compressor model can indicate the airflow and performance parameters of different characteristic sectors of impeller entrance, compressor exit, and axial diffuser entrance and exit. The 3D numerical simulation reports that the maximum relative errors of static pressure, total pressure, and total temperature predictions are 13.25%, 11.30%, and 5.65%, respectively. (3) The constructed air inlet–compressor matching model can rapidly predict the airflow parameters and performance parameters of their characteristic sections under specific operation conditions. The “most stable operation curve” introduced in the pressure ratio characteristic curve of the compressor can quickly and accurately accomplish the initialization of matching iterative computations, thereby solving aerodynamic coupling problems on an air inlet–compressor interface. (4) The integrated 3D numerical simulation of the air inlet/compressor reveals that the relative errors of 17 in 18 airflow parameters of two components are lower than 8%, and the relative error of the other parameters is 10.91%. The model exhibits high accuracy and can satisfy the requirements for the performance prediction of air inlet systems in ATR.
  • An efficient setup for freestream turbulence on transition prediction over
           aerospace configurations
    • Abstract: Publication date: Available online 13 August 2018Source: Aerospace Science and TechnologyAuthor(s): Gustavo Luiz Olichevis Halila, Enda Dimitri Vieira Bigarella, Alexandre Pequeno Antunes, João Luiz F. AzevedoAbstractInflow turbulence variables play a key role in the performance of correlation-based transition models. In order to consider this important effect, a wing-body configuration in transonic flow is simulated using the γ−Reθ Langtry–Menter transition model. The influences of the freestream turbulence intensity, Tui, and freestream eddy viscosity ratio, μt/μ, over the simulation results are addressed, and an efficient setup for these variables is suggested. Two distinct turbulence conservation boxes, regions in which the turbulence source terms are turned off and in which no turbulence decay occurs, are addressed in the paper. This strategy is an option to the approach of specifying higher values for the freestream turbulence variables and allowing them to decay up to the geometry near field. Discussions on the decay of the turbulent kinetic energy (k) outside of the conservation boxes support the results presented in the paper and provide relevant guidelines regarding the specification of boundary values for turbulence variables.
  • Effect of arbitrary yaw/pitch angle in bird strike numerical simulation
           using SPH method
    • Abstract: Publication date: Available online 13 August 2018Source: Aerospace Science and TechnologyAuthor(s): Zhuo Zhang, Liang Li, Dingguo ZhangAbstractThe influences of arbitrary attitude angles on bird strike on a fixed rigid flat plate and a rotary jet-engine fan are studied. A verified real bird model and a hemispherical-ended cylinder substitute bird model are modeled by using the smoothed particle hydrodynamics (SPH) method. Since birds can strike aircraft engine from any orientations, simulations of bird models striking a rigid flat plate at random attitude angles are first done as a validation test. Results show that different attitude angles of bird model have distinct effect on the response of bird strike. As the attitude angle increases, the peak impact force becomes larger and the bird model loses more energy. By considering the rotation of the jet-engine fan ignored before, the impact behavior of real bird models striking on rotating engine blades from arbitrary attitude angles is investigated. Effects of the attitude angle on the most concerned impact force, kinetic energy and von Mises stress of blade roots are discussed. It is concluded that considering attitude angles of real bird and rotation of the jet-engine fan in bird strike simulation has practical significance on structural tolerant design.
  • Performance improvement of variable speed rotors by Gurney flaps
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Dong Han, Chen Dong, George N. BarakosAbstractGurney flaps are used for improving the performance of variable speed rotors. An analytical model able to predict helicopter rotor power is first presented, and the flight data of the UH-60A helicopter is used for validation. The predictions of the rotor power are in good agreement with the flight test data, justifying the use of this tool in analyzing helicopter performance. A fixed Gurney flap can enhance the performance of variable speed rotors and expand the corresponding flight envelope, especially near stall and high speed flight. A retractable Gurney flap at 1/rev yields more power savings than a fixed Gurney flap or a retractable one with a higher harmonic prescribed motion. At a speed of 200 km/h, the retractable Gurney flap, actuated at 1/rev, can obtain 3.22% more power reduction at a rotor speed of 85% nominal rotor speed, and this value is 8.37% at a speed of 220 km/h. The height corresponding to the minimum power increases slowly in low to medium speed flight, and increases dramatically in high speed flight. With increasing take-off weight (i.e. rotor thrust), the retractable Gurney flap at 1/rev can obtain more rotor power savings.
  • Automatic carrier landing control for unmanned aerial vehicles based on
           preview control and particle filtering
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Ziyang Zhen, Shuoying Jiang, Kun MaAbstractFor the carrier-based unmanned aerial vehicles (UAVs), one of the important problems is the design of an automatic carrier landing system (ACLS) that would enable autonomous landing of the UAVs on a moving aircraft carrier. However, the safe autolanding on a moving aircraft is a complex task, mainly because of the deck motion and airwake disturbances, and dimension limitation. In this paper, an innovative ACLS system for carrier-based UAVs is developed, which is composed of the flight deck motion prediction, reference glide slope generation and integrated guidance and control (IGC) modules. The particle filtering method is used to online predict the magnitudes and frequencies of the deck motion, which are used to correct the reference glide slope to achieve minimum dispersion around the ideal touchdown point. An optimal preview control (OPC) scheme is presented for the IGC subsystem design, which fuses the preview information of the reference glide slope, equality constraint of UAV dynamics and performance index function, and predicted information of the carrier deck motion. Simulation results of a nonlinear UAV model show the effectiveness of the ACLS system in carrier autolanding under the deck motion and airwake disturbances.
  • H +performance&rft.title=Aerospace+Science+and+Technology&rft.issn=1270-9638&">Application of ICC LPV control to a blended-wing-body airplane with
           guaranteed H ∞ performance
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Tianyi He, Ali K. Al-Jiboory, Guoming G. Zhu, Sean S.-M. Swei, Weihua SuAbstractThis paper addresses the Input Covariance Constraint (ICC) control problem with guaranteed H∞ performance for continuous-time Linear Parameter-Varying (LPV) systems. The upper bound of the output covariance is minimized subject to the constraints on input covariance and H∞ output performance. This problem is an extension of the mixed H2/H∞ LPV control problem, in that the resulting gain-scheduling controllers guarantee not only closed-loop system robustness in terms of H∞ norm bound but also output covariance performance over the entire scheduling parameter space. It can be shown that this problem can be efficiently solved by utilizing the convex optimization of Parameterized Linear Matrix Inequalities (PLMIs). The main contributions of this paper are to characterize the mixed ICC/H∞ LPV control problem using PLMIs and to develop the optimal state-feedback gain-scheduling controllers, while satisfying both input covariance and H∞ constraints. The effectiveness of the proposed control scheme is demonstrated through vibration suppression of a blended-wing-body airplane model.
  • Pressure wave damping in transonic airfoil flow by means of micro vortex
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): M. Gageik, J. Nies, I. Klioutchnikov, H. OlivierAbstractIn transonic airfoil flow, pressure waves are generated mainly at the trailing edge and in the case of a shock in the region of the shock/boundary layer interaction. Depending on the Mach number, these waves lead to oscillating shock waves and an unsteady pressure distribution. For a free steam Mach number of M=0.76 and a chord length based Reynolds number of Re=106, micro vortex generators (μVG) are applied to dampen pressure waves. This is studied experimentally in a shock tube and numerically by using a high-order finite difference scheme (under-resolved Direct Numerical Simulation). The agreement of the pressure distribution and Schlieren pictures between simulation and experiment is good. By means of numerical visualizations, instability waves are identified within the separated boundary layer above a marginal boundary layer separation bubble. The applicability of μVG for dampening the pressure waves and stabilizing the flow field is possible and is studied in this paper. By numerical Schlieren pictures and further visualizations, the flow around the VG is characterized. The spanwise oriented instability waves are partly disintegrated which is also confirmed by the analysis of the vorticity. Finally, the nonlinear wave propagation is investigated and an explanation for the typical 1 to 2 kHz pressure oscillation is given.
  • Assessment of dynamic instability of laminated composite-sandwich plates
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Rosalin Sahoo, B.N. SinghAbstractThe current work deals with the assessment of dynamic stability behavior of laminated composite and sandwich plates subjected to in-plane static and periodic compressive loads based on a recently developed zigzag theory by the authors. This theory satisfies the traction-free boundary conditions at top and bottom surfaces of the laminate as well as the inter-laminar stress continuity at layer interfaces. Also, it obviates the need of artificial shear correction factor. The theory is based upon shear strain shape function assuming non-linear distribution of transverse shear stresses. An efficient C0 continuous, eight-noded isoparametric element with seven field variables is employed for the dynamic stability analysis of laminated composite and sandwich plates. The boundaries of principal instability domains are obtained following Bolotin's approach and are represented either in the non-dimensional load amplitude-excitation frequency plane or load amplitude-load frequency plane. A series of numerical examples on the dynamic stability analysis of laminated composite and sandwich plates are studied to demonstrate the effects of modular ratio, span to thickness ratio, boundary conditions, thickness ratio, static load factor and various load parameters on the principal instability regions. The predicted results are compared with the available existing results in order to ensure the performance of the proposed model.
  • An optimal approach to the preliminary design of small hybrid-electric
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Carlo E.D. RiboldiAbstractHybrid-electric propulsion is an interesting alternative for the light aviation market, carrying the advantages of electric propulsion in terms of lower noise and pollutive emissions in terminal maneuvers, while not renouncing to the flight performance – especially range – typical to conventional propulsion, based on hydrocarbon fuel. Some difficulty in the spreading of this new technology in light aviation may be ascribed to the lack of consolidated techniques to preliminary design hybrid-electric aircraft, complicating the negotiation of specifications and making design choices difficult. This is also the effect of a notable increase in the number of design variables needed to describe the hybrid-electric power-train, which include characteristics of both its thermal and electric parts, with respect to conventionally powered aircraft. The present paper presents a methodology to efficiently cope with this design problem. The procedure is based on an optimal approach where take-off weight is minimized, and constraints are included to assure meeting the mission performance requirements while not exceeding any technological limit. The paper recalls at first some simple mathematical models, allowing to translate flight performance requirements into constraints on the power-train. Then the proposed optimal design approach is thoroughly presented at a theoretical level. Finally, an example design of a hybrid-electric motor-glider is shown, where the optimal design tool is used both to find a baseline solution and to investigate the sensitivity of that design point with respect to constraints due to performance requirements and technological specifications.
  • Fractional-order controllers optimized via heterogeneous comprehensive
           learning pigeon-inspired optimization for autonomous aerial refueling
           hose–drogue system
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Yongbin Sun, Haibin Duan, Ning XianAbstractDynamic modeling and control system design for the hose–drogue system (HDS) in the docking stage of autonomous aerial refueling (AAR) are investigated in this paper. The dynamics and kinematics of hose are modeled via a finite-segment multi-body method, which describes the hose–drogue assembly as a link-connected system. A controllable drogue is connected to the hose for automatically stabilizing the drogue's relative position under the influences of tanker trailing vortex, receiver bow wave, atmospheric turbulence, gust, and wind shear. Thus, a drogue position control law based on fractional-order method is designed to resist the multi-wind disturbances. Noting that it is difficult to tune the parameters of fractional-order controller (FOC), a modified pigeon-inspired optimization (PIO), the hybrid of heterogeneous comprehensive learning strategy and PIO (HCLPIO), is carried out to optimize the parameters of FOC. The simulation results show that the proposed optimized fractional-order feedback controllers effectively stabilize the controllable drogue to swing within an acceptable range.
  • Free vibrations of functionally graded polymer composite nanoplates
           reinforced with graphene nanoplatelets
    • Abstract: Publication date: October 2018Source: Aerospace Science and Technology, Volume 81Author(s): Mohammad Arefi, Elyas Mohammad-Rezaei Bidgoli, Rossana Dimitri, Francesco TornabeneAbstractA two-variable sinusoidal shear deformation theory (SSDT) and a nonlocal elasticity theory are applied in this paper to analyze the free vibration behavior of functionally graded (FG) polymer composite nanoplates reinforced with graphene nanoplatelets (GNPs), resting on a Pasternak foundation. Based on the proposed theory, the transverse deflection is assumed as summation of bending and shear transverse deformations. Four different FG reinforcement patterns are here employed, namely a uniform distribution UD, and non-uniform distributions FG-O, FG-X and FG-A. The effective elastic modulus, the Poisson's ratio and the density of composite nanoplates are computed using the Halpin–Tsai model and the rule of mixture, respectively. The numerical results are validated through a comparative assessment of the results with respect to predictions from literature, including nanoplates and FG polymer composite plates. A wide parametric investigation shows the influence of some significant parameters, such as nonlocal parameters, total number of layers, weight fraction, as well as parameters related to the Pasternak foundation and geometry, on the free vibration response of FG polymer composite nanoplates reinforced with GNPs.
  • Acoustic energy absorption and dissipation characteristic of Helmholtz
           resonator enhanced and broadened by acoustic black hole
    • Abstract: Publication date: Available online 10 August 2018Source: Aerospace Science and TechnologyAuthor(s): X.Q. Zhou, D.Y. YuAbstractIn the present analysis, a novelty Helmholtz resonator (HR) enhanced by the acoustic black hole (ABH) is established. Dynamic partial differential equations are formulated according to the equilibrium of the flexible plate under harmonic uniformly distributed load, and the load is generated by acoustic fluids in the coupled system. By applying the Wentzel-Kramer-Brillouin (WKB) method, two fundamental functions are respectively assumed to describe the characteristic in the fluid and solid fields, and then the functions are determined by the boundary conditions and continual properties of the coupled structures. Moreover, relationships between the two functions are identified. By utilizing the energy conservation theorem of the closed system, energy storage and absorption characteristic in HR and ABH are separately calculated, which is further depicted as the normalized energy absorbed according to the input and output of the system, then energy transmission loss of the coupled system is calculated. The factors influence on the energy transmission loss in the frequency domain are quantitatively and qualitatively analyzed, especially the ABH parameter m. Finally, comparison is presented, the analysis shows the advantage of the coupled system by comparing with traditionally HR.
  • Rotating detonation in a ramjet engine three-dimensional modeling
    • Abstract: Publication date: Available online 9 August 2018Source: Aerospace Science and TechnologyAuthor(s): N.N. Smirnov, V.F. Nikitin, L.I. Stamov, E.V. Mikhalchenko, V.V. TyurenkovaAbstractA rotating detonation engine (RDE) combustion chamber fed by hydrogen–air mixtures of different composition was modeled numerically using 3D geometry. The RDE is a new type of engines capable to create higher thrust than the traditional ones based on the combustible mixture deflagration process. The dynamical process of combustion in the RDE is more than 100 times faster than that for the classical slow deflagration combustion mode. This type of engine has a more efficient thermodynamic cycle. In numerical experiments, different combustible mixture compositions were tested, and different scenarios of the engine performance were obtained. The computational domain used a regular mesh of uniform cubic elements. The time-consuming parts of the numerical code were parallelized using the OpenMP technique. Our calculations were made at APK-5 with a peak performance of 5.5 Tera Flops.
  • Influence of thickness on performance characteristics of non-sinusoidal
           plunging motion of symmetric airfoil
    • Abstract: Publication date: Available online 9 August 2018Source: Aerospace Science and TechnologyAuthor(s): R. Sankarasubramanian, Adithya Sridhar, M.S. Prashanth, Akram Mohammad, Kishore V. Ratna, V. LaxmanAbstractFor the past few decades flapping wing aerodynamics has attracted a great deal of research interest from both the aeronautical and biological communities pertaining to the development of MAVs. The objective of this study is to examine and understand the effect of non-dimensional plunge amplitude and reduced frequency on propulsive performance of NACA 4-digit airfoil series and to examine the performance characteristics of square plunge motion and trapezoidal plunge motion. Two dimensional flow simulations around plunging symmetric aerofoils were performed using FLUENT. The simulations were carried out at Reynolds number of 20000 using incompressible laminar, NS solver. The reduced frequency (k) was varied from 0.5–5 and the plunging amplitude (h) was varied from 0.25–1.5. The plunging motions to the aerofoils were provided through UDFs. The effect of variation of k and h on the thrust coefficient (CT), power-input coefficient (CP) and propulsive efficiency (η) is studied. CT value is maximum for square plunge profile for all the airfoils. However, for a given value of h, with the increase in k, CT increases with increasing thickness of the airfoil and reaches a maximum value for airfoil thickness of NACA0018 and then starts decreasing. With varying h and k, it was observed that the propulsive efficiency reached a peak value and the peak shifts to higher h and k with increasing airfoil thickness. From the above study, it was concluded that airfoil thickness played a major part in influencing the thrust generation at low Strouhal number. However, at high Strouhal numbers airfoils showed diverse trends with respect to thrust generation. Sinusoidal plunging motion was more efficient but generated less thrust when compared to square and trapezoidal plunging motions.
  • Design of a transonic wing with an adaptive morphing trailing edge via
           aerostructural optimization
    • Abstract: Publication date: Available online 9 August 2018Source: Aerospace Science and TechnologyAuthor(s): David A. Burdette, Joaquim R.R.A. MartinsAbstractNovel aircraft configurations and technologies like adaptive morphing trailing edges offer the potential to improve the fuel efficiency of commercial transport aircraft. To accurately quantify the benefits of morphing wing technology for commercial transport aircraft, high-fidelity design optimization that considers both aerodynamic and structural design with a large number of design variables is required. To address this need, we use high-fidelity aerostructural that enables the detailed optimization of wing shape and sizing using hundreds of design variables. We perform a number of multipoint aerostructural optimizations to demonstrate the performance benefits offered by morphing technology and identify how those benefits are enabled. In a comparison of optimizations considering seven flight conditions, the addition of a morphing trailing edge device along the aft 40% of the wing can reduce cruise fuel burn by more than 5%. A large portion of fuel burn reduction due to morphing trailing edges results from a significant reduction in structural weight, enabled by adaptive maneuver load alleviation. We also show that a smaller morphing device along the aft 30% of the wing produces nearly as much fuel burn reduction as the larger morphing device, and that morphing technology is particularly effective for high aspect ratio wings.
  • The crack detection and evaluation by elastic wave propagation in open
           hole structures for aerospace application
    • Abstract: Publication date: Available online 31 July 2018Source: Aerospace Science and TechnologyAuthor(s): Marek Barski, Adam StawiarskiAbstractThe rectangular plate made of aluminum alloy with a relatively large circular hole in the geometrical center of the structure is subjected to uniform tension. This loading causes that in the vicinity of a hole a process of crack formation is triggered. The crack is detected and evaluated with the use of elastic wave propagation method. The piezoelectric activator is located directly on the edge of the hole and the rest of sensors are placed at an equal distance from the activator. The crack is detected by comparison of the reference, obtained for the intact plate, and the actual signal (pitch-catch measurement). The appropriate damage magnitude measurement is proposed. Moreover, the system, which works without knowledge of an intact structure, is also discussed. In order to estimate and visualize the length of a crack, the advanced diagnostic imaging is introduced. The applied method is based on a visualization of sensing paths with an assigned value of correlation coefficients, computed for reference and actual signals. The effectiveness of the described system is verified with the use of numerical simulation and experimental test. The theoretical analysis is carried out with the use of the finite element method. The computations are performed for several assumed lengths of cracks. The obtained theoretical results, as well as the experimental analysis, confirm the fact that the crack can be detected at an early stage of formation. Moreover, a relatively good agreement between the real crack lengths and the ones obtained with the use of the proposed method is observed in the numerical analysis as well as in the experiment.
  • Investigation of stall process flow field in transonic centrifugal
           compressor with volute
    • Abstract: Publication date: Available online 31 July 2018Source: Aerospace Science and TechnologyAuthor(s): Ce Yang, Wenli Wang, Hanzhi Zhang, Changmao Yang, Yanzhao LiAbstractOwing to the asymmetric structure of the volute, the internal flow field in a centrifugal compressor is circumferentially non-uniform. Previous studies have paid little attention to whether the circumferentially non-uniform flow field has an influence on the circumferential position of stall inception under transonic inlet conditions. In this study, the performance of the centrifugal compressor and static pressure distribution around the casing wall are obtained by means of an experimental method. Thereafter, the experimental results are compared with the time-averaged results of unsteady simulations, and the stall process with shock waves is obtained. The results demonstrate that, in the presence of shock waves, the volute tongue (VT) determines the stall inception circumferential position. Owing to the asymmetric structure of the volute, the positions, strengths, and shapes of shock waves exhibit significant differences in varying blade passages. The VT induces the high static pressure region in a blade passage; the shock wave in the corresponding blade passage is much closer to the impeller inlet than the other blade passages; and the forward velocity of the shock wave in the stall process is significantly larger. Furthermore, leading edge spillage, which represents the stall inception, will first occur in this passage. The reason for these phenomena is that the high static pressure region can increase the static pressure at the blade passage outlet position, which compels the shock wave to move closer to the impeller inlet; therefore, stall inception first occurs in this passage. In the stall inception process, the shock wave in the blade passage, located in the low static pressure region, is not constantly moving to the impeller inlet; the shock wave moving direction also changes, and the shock wave moves downstream during a certain period. The tip leakage flow (TLF) trajectory in different blade passages differs significantly. Tip leakage vortex (TLV) breakdown first occurs in the blade passage that is most influenced by the VT.
  • Effect of bio-inspired sinusoidal leading-edges on wings
    • Abstract: Publication date: Available online 31 July 2018Source: Aerospace Science and TechnologyAuthor(s): Martiqua L. Post, Robert Decker, Anthony R. Sapell, Jonathan S. HartAbstractObservations of maneuvering humpback whales have revealed unique hydrodynamic performance hypothesized to be a result of tubercles on the leading-edge of the whales' pectoral flippers. Inspired by this biological observation, it is shown sinusoidal leading-edge wings prevent the dramatic loss of lift caused by stall and instead generate a gradual decrease in lift with as much as 25% higher lift in the poststall regime. Six different wing geometries, smooth and sinusoidal leading-edge models, swept and unswept configurations, were tested at angles of attack of −2 to 24 degrees at Reynolds numbers between 100,000 and 500,000. Oil surface flow visualization and CFD results reveal variations in flow phenomena between the smooth and sinusoidal leading-edge configurations.
  • Performance analysis and path planning for UAVs swarms based on RSS
    • Abstract: Publication date: Available online 17 July 2018Source: Aerospace Science and TechnologyAuthor(s): Weijia Wang, Peng Bai, Xiaolong Liang, Jiaqiang Zhang, Lvlong HeAbstractIn the localization estimation system, it is well known that the sensor-emitter geometry can seriously impact the accuracy of the location estimate. In this paper, we analyze the optimal deployment for received signal strength (RSS) localization with the measurement noise is set to be distance dependent. First, the Cramer–Rao low bound (CRLB) with distance-dependent noise in RSS localization is calculated and chosen to be the optimality criterion. The optimal deployment is analyzed via angle and distance criterion, respectively. Then, the analytic solutions to the optimal deployment are derived in both static and movable target scenarios. Finally, we extend our work to the path planning problem with constraints and interior penalty method is applied to settle the constrained nonlinear optimization problem. The simulation results show that the path optimization verifies the accuracy of the analytical findings.
  • Interval analysis of the standard of adaptive cycle engine component
           performance deviation
    • Abstract: Publication date: Available online 10 July 2018Source: Aerospace Science and TechnologyAuthor(s): Min Chen, Jiyuan Zhang, Hailong TangAbstractThe deviation of engine component performance has great impact on the Adaptive Cycle Engine (ACE) overall performance and task adaptability. To make ACE performance reach the mission requirement, proper standard of component performance deviation (CPD) should be given in advance. In this paper, an approach based on the first order Taylor series expansion has been proposed and applied to set the standard of CPD without large amounts of calculation. This approach is an inversion process of the normal interval analysis and can take the distinction between the impacts of CPD indexes into consideration. Fifteen component performance parameters and four important operating conditions are investigated. The results show that the distinction between the impacts of CPD indexes and different operating conditions is obvious. Compared with setting uniform standard for all CPD indexes, this method can better utilize component characteristics and make the standard reasonable and economical. The standard of CPD can be derived within 10 times of off-design point calculation through this method, which is much less than 32768 times of calculation for the vertex method. This method is universal and can be applied to setting standard for the CPD of other type of gas turbine engines.
  • Experimental characterization of the transonic test section flow in a
           Ludwieg tube
    • Abstract: Publication date: Available online 24 May 2018Source: Aerospace Science and TechnologyAuthor(s): B. Hammer, H. OlivierAbstractA new Ludwieg tube for transonic testing has been put into operation. For this, an existing shock tube has been modified to a Ludwieg tube. The ideal operating principle is described by theory and corresponding theoretical values are compared to experimental ones. Pressure histories as well as test section flow visualization demonstrate a better performance concerning flow quality and testing time compared to the former shock tube. Static as well as total pressure measurements have been performed within the test section. Pitot rake measurements show the homogeneity of the test section flow field. Additional boundary layer profile measurements give evidence of the temporal growth of the test section wall boundary layer. These results are compared with theoretical ones deduced from a theoretical approach for a boundary layer behind an expansion wave. Both, boundary layer and Pitot rake measurements allow to discuss the influence of a turbulent boundary layer on these measurements. Cookie cutters at the entrance to the test section act as boundary layer bleed which partially avoid the entering of the tube wall boundary layer into the test section. Their influence on the test section flow is discussed as well.
  • Effect of design parameters on the mass of a variable-span morphing wing
           based on finite element structural analysis and optimization
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Pedro D.R. Santos, Diogo B. Sousa, Pedro V. Gamboa, Yifan ZhaoAbstractIn the past years, the development of morphing wing technologies has received a great deal of interest from the scientific community. These technologies potentially enable an increase in aircraft efficiency by changing the wing shape, thus allowing the aircraft to fly near its optimal performance point at different flight conditions. However, these technologies often present an undesired mass increase due to their inherent complexity. Therefore, the aim of the current work is to ascertain the influence of geometrical and inertial parameters on the structural mass of a Variable-span Wing (VSW). The structural mass prediction is based on a parametric study. A minimum mass optimization problem with stiffness and strength constraints is implemented and solved, being the design variables structural thicknesses and widths, using a parametric Finite Element Model (FEM) of the wing. The study is done for a conventional fixed wing and the VSW, which are then combined to ascertain the VSW mass increment, i.e., the mass penalization of the adopted morphing concept. Polynomials are found to produce good approximations of the wing mass. The effects of the various VSW design parameters in the structural mass are discussed. On one hand, it was found that the span and chord have the highest impact in the wing mass. On the other hand, the VSW to fixed wing mass ratio proved that the influence of span variation ratio in the wing mass is not trivial. It is found that the mass increase does not grow proportionally with span variation ratio increase and that for each combination of span and chord, exists a span variation ratio that minimizes the mass penalty. In the future, the developed polynomials could be used to create a mass prediction model to aid the design of morphing wings during the conceptual design phase.
  • Inverse design and Mach 6 experimental investigation of a pressure
           controllable bump
    • Abstract: Publication date: Available online 8 August 2018Source: Aerospace Science and TechnologyAuthor(s): Zonghan Yu, Guoping Huang, Chen XiaAbstractThe low kinetic flow starting from the airframe leading edge is a neglected aspect in designing hypersonic air-breathing flight vehicles. Compared with other boundary layer treating technologies, the bump concept can obtain a good balance of boundary layer removal, external drag control, shock system simplification, and integration design flexibility capabilities. On the basis of conventional conical-flow theory and the new 3D inverse design method, this study proposes a pressure controllable bump concept that can generate the bump configuration inversely by the prescribed pressure distribution. The bump/inlet integration pattern is analyzed, and the basic design methodology is presented. To validate whether the pressure distribution can be used in diverting the boundary layer, experimental study of the bump and numerical simulation are conducted. Results show that the bump has generated identical pressure distribution to the design. The bump can also divert approximately 50% of the boundary layer from the incoming low kinetic flow in Mach 6. Compared with the conventional cone-derived bump in Mach 6, the new bump is 25.8% shorter in height. The flow structure is adjusted nearly parallel to the x-direction, thereby promoting the flow quality of the inlet entrance. Hence, the new inverse design method of pressure distribution expands the applicable Mach range of the hypersonic airframe forebody.
  • Position tracking control of a helicopter in ground effect using nonlinear
           disturbance observer-based incremental backstepping approach
    • Abstract: Publication date: Available online 7 August 2018Source: Aerospace Science and TechnologyAuthor(s): Jinshuo Hu, Jianzhe Huang, Zhenxing Gao, Hongbin GuHelicopters are highly nonlinear and internally unstable air vehicles, the dynamic response of which can be strongly influenced by flight conditions (wind gust, ground effect, etc.). Therefore, it is a challenging task to design a reliable flight control system (FCS) with all safety and performance requirements satisfied. This paper investigates the robust position tracking problem of a helicopter in ground effect (IGE). Based on a highly reliable and computational efficient finite state representation of rotor flow field IGE, a novel nonlinear disturbance observer-based incremental backstepping (N
      DOI BS) controller is designed to track position commands (all derivatives of the reference trajectory are known) under the influence of system uncertainties and external disturbances. Without requiring the exact knowledge of helicopter dynamics, the N
      DOI BS approach guarantees that instant control increments are derived in terms of Lyapunov theory and ensures robustness in the presence of mismatched disturbances whose first derivatives are bounded. It is shown that all state variables of the closed-loop system are semi-globally uniformly ultimately bounded (SGUUB). In addition, to further improve the horizontal position tracking performance, rotor state feedback (RSF) technique and a disturbance feedback strategy are applied to develop a pitch stability augmentation system (SAS). Finally, controller performance is demonstrated through numerical simulations using the Bo-105 utility helicopter. With the efficiency and robustness properties verified, the suggested N
      DOI BS control scheme shows great potential for implementing advanced FCS designs in existing helicopters.
  • Overview and application of FEM methods for shock analysis in space
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Andrés García-Pérez, Félix Sorribes-Palmer, Gustavo Alonso, Ali RavanbakhshAbstractSpacecraft are subjected to severe mechanical loads, especially during the ascent phase of the launch. Among several vibrational environments, shock loads, which are caused mainly by the activation of pyrotechnic devices used for the separation of the payloads and the different stages of the launcher, are transmitted throughout the entire structure and reach the scientific instruments of the spacecraft. Therefore, it is important to verify if the space instruments can withstand this environment, considering the nature of the shock, which generally consists in an intensive and short load. In recent years, the demand of numerical analyses to predict the responses of the structures against shocks is increasing and, for this reason, it is necessary to establish adequate numerical methods, taking into account the complex mathematical treatment and the uncertainty in the load characterization. The purpose of this paper is to present the application of different methods to calculate the required structural results for a space instrument subjected to the shock environments using a finite element model (FEM). The procedures for each method, the type of the results that can be calculated and the comparison of the results are described in this paper. The objective is to select the most suitable analysis method for shock loads based on the precision of the results and the capability of obtaining all the variety of data for a complete evaluation of the structure.
  • Sequential multidisciplinary design optimization and reliability analysis
           under interval uncertainty
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Lei Wang, Chuang Xiong, Juxi Hu, Xiaojun Wang, Zhiping QiuAbstractWith the rapid development of modern technology and the rising demand for high reliability in complex multidisciplinary engineering systems, more and more attention has been paid to reliability based multidisciplinary design optimization (RBMDO). The regular RBMDO is a triple-level nested optimization loop which is computationally expensive. In this paper, a sequential multidisciplinary design optimization and reliability analysis method under non-probabilistic theory is developed to decouple the reliability analysis from the optimization. The multidisciplinary design optimization (MDO), interval uncertainty analysis and the reliability analysis are conducted in a sequential manner. Furthermore, a dimension-by-dimension method (DDM) is proposed to conduct the interval uncertainty analysis. The calculation of the reliability under interval uncertainty is deduced based on the volume ratio theory. The shifting distance of the limit state function is also deduced. Both numerical and engineering examples are employed to demonstrate the validity of the proposed method.
  • Verification and application of a mean flow perturbation method for jet
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Swagata Bhaumik, Datta V. Gaitonde, S. Unnikrishnan, Aniruddha Sinha, Hao ShenAbstractThe stability properties of basic states are often elucidated by examining the evolution of small disturbances. Such studies have recently been successfully applied to mean turbulent states, obtained through averaging of experimental measurements or Large-Eddy Simulations (LES), for both wall-bounded as well as free shear flows. Typically, the equations are employed using the disturbance form of the equations. To circumvent the necessity to linearize the governing equations, an especially tedious task for viscous and turbulent closure terms, Touber and Sandham (2009) [21], proposed an approach that achieves the same purpose by solving the full Navier–Stokes (NS) equations, with a forcing term to maintain mean flow invariance. The method places no restrictions, such as slow streamwise variations, on the underlying basic state. The goals of the current work are to first verify this mean flow perturbation (NS-MFP) technique and then apply it to the problem of jet noise. For the first thrust, we show that when the basic state is appropriately constrained, the technique reverts to Linear, Parabolized and Global stability methods. The method is then verified by reproducing the growth of unstable modes in an inviscid Mach 6 entropy layer. The application to jet noise considers subsonic Mach 0.9 and perfectly expanded supersonic Mach 1.3 round jets. The results are compared with those from Parabolized Stability Equations (PSE) and LES solutions, respectively, considering monochromatic and multi-frequency perturbations. The NS-MFP method successfully reproduces key features of the modal response, including Strouhal number dependent directivity of noise radiation. Aspects related to the manner in which the mean basic state is obtained, whether from LES or Reynolds-averaged Navier–Stokes (RANS) equation are also explored. In particular, the sensitivity of the perturbation to whether the eddy viscosity is included or not, is examined in reference to maximum intensity of pressure fluctuation, directivity of noise radiation and the rate of fall-off of the spectra at higher Strouhal numbers. The results indicate that a closer match on the noise-radiation characteristics is obtained when effects of eddy-viscosity on the disturbances are neglected.
  • Novel HW mixed zig-zag theory accounting for transverse normal
           deformability and lower-order counterparts assessed by old and new
           elastostatic benchmarks
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Ugo Icardi, Andrea UrraciAbstractMixed zig-zag plate theories are derived from a recently developed 3-D five d.o.f. zig-zag “adaptive” theory ZZA (a priori fulfillment of interfacial stress constraints) under steadily growing limiting assumptions on displacement, strain and stress fields. The intended aim is trying to save computational costs simultaneously preserving accuracy. Lower-order theories assume a uniform or a polynomial transverse displacement, out-of-plane stresses are derived from local equilibrium equations or retaken from higher-order theories. The focus of this study is twofold: (i) to assess whether and for which cases a higher-order through-thickness zig-zag transverse displacement representation is essential, or vice versa a simpler kinematics can be assumed; (ii) to compare accuracy of theories based on Murakami's and Di Sciuva's zig-zag functions with the same expansion order across the thickness. A number of challenging benchmarks are retaken from literature and new benchmarks with a strong variation of material properties (damaged layers), distributed or localized step loading and some different boundary conditions are considered to assess accuracy of theories and of FEA 3-D (constituting the reference solution in lack of exact results). For these benchmarks, closed-form solutions are obtained assuming the same expansion order across the thickness and the same in-plane trial functions for all theories. The numerical illustrations show that lower order and Murakami's based theories fail for cases having the strongest layerwise and transverse anisotropy effects, or a marked through-thickness asymmetry. The Hu–Washizu highest-order theory HWZZ is shown to be always the only as accurate as ZZA, despite it reduces the computational effort.
  • Effect of axisymmetric endwall contouring on the high-load low-reaction
           transonic compressor rotor with a substantial meridian contraction
    • Abstract: Publication date: Available online 3 August 2018Source: Aerospace Science and TechnologyAuthor(s): Shijun Sun, Shaowen Chen, Wei Liu, Yun Gong, Songtao WangAbstractFor a high-load low-reaction compressor transonic rotor based on an increase in outlet axial velocity, the stage reaction is decreased by a substantial meridian contraction of the endwall. The axisymmetric endwall contouring has a deep impact on the aerodynamic performance of the rotor. Therefore, nine combinations of endwall contours with three different shapes including a linear wall, a sinusoidal wall and a symmetrical wall of the sinusoidal wall about the linear wall were firstly studied through 3D steady Reynolds Average Navier-Stokes simulations. The results show that the cases with a linear hub have the highest peak pressure ratio and the largest peak efficiency, whereas the cases, featured with a sinusoidal hub, have the widest stall margin, the largest mass flow and the lowest pressure ratio. Endwall contouring not only influences the shock structure across the flow passage but also dominates the location and intensity of separation. In addition, the effect of endwall contouring on loading distribution is also presented and the case with a sinusoidal hub is characterized by aft-loading and the other two cases have a feature of fore-loading.
  • Investigation of performance and mode transition in a variable divergence
           ratio dual-mode combustor
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Chenlin Zhang, Juntao Chang, Shuo Feng, Wen Bao, Daren YuAbstractTo operate in a wide range of incoming flow Mach number, a variable divergence ratio dual-mode combustor is designed. Experimental and numerical investigation of the dual-mode combustor has been conducted in this paper. By regulating the divergence ratio in combustor, in a high incoming flow Mach number or low equivalence ratio, a small divergence ratio could increase pressure peak and boost combustor performance. Under a low flight Mach number or a high equivalence ratio, a large divergence ratio has a benefit to accommodate more heat release and prevent the inlet unstart. The results of experiments and numerical investigations have been testified that the variable divergence ratio dual-mode combustor could operate in a wide range of incoming flow Mach number. Another aspect, the combustor wall ramp makes an additional effect on the combustion zone. The combustion zone distribution is close to the upper wall because of the pressure differentials of the main flow aroused by lower wall ramp compressing acting on the supersonic airflow. Under the effecting of incoming flow Mach number and heat release, Mach number near the leading edge of the ramp decides the location of the combustion zone. Comparing with a common combustor, it is harder to form a thermal throat in the combustor because of the big divergence angle in the variable divergence ratio combustor. Based on the flow field analysis, a conclusion is drawn that the combustion mode transition is dominated by the combustion heat release in the variable divergence ratio dual-mode combustor configuration.
  • The influence of boundaries on sound insulation of the multilayered
           aerospace poroelastic composite structure
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): R. Talebitooti, M.R. Zarastvand, H.D. GohariAbstractAcoustic analysis of the four-sides simply supported doubly curved composite shell interlayered with porous material used in aerospace applications is considered based on Third order Shear Deformation Theory (TSDT). The focus is specifically placed on presenting the effect of boundaries on Sound Transmission Loss (STL) of the poroelastic structure. Then, the results are compared with those of infinite shell. In order to calculate the STL, the displacement and pressure terms are derived in the form of double Fourier series for a simply supported boundary condition doubly curved shell. A new approach is made to consider the number of modes for the finite composite structure treated with porous material. In addition, the simultaneous solution of the equations of porous and composite layers along with acoustic wave equations is performed. This process leads to obtain the unknown constants in all parts of the structure including upper and bottom composite shells as well as porous core which results in presenting the STL of the structure in the logarithmic scale. Since, the solution is provided in the form of modal infinite series which should be terminated at the sufficient modes, the convergence checking is prepared in 3D configurations with respect to various frequencies and shell dimensions to present the number of the appropriate modes for this process. Moreover, the results are validated with the aid of other researches including experimental as well as analytical models available in literature. Finally, some new results including Sound Pressure Level (SPL) in both of global view and contour map, transverse displacement configurations and the effects of various shell dimensions on STL, are presented for this finite structure.
  • Extendable chord for improved helicopter rotor performance
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Dong Han, Kelong Yang, George N. BarakosAbstractExtendable blade sections are investigated as a method for reducing rotor power and improving helicopter performance. A validated helicopter power prediction method, based on an elastic beam model is utilized. The static extendable chord can deliver a rather small power reduction in hover, and significant power savings at high speed flight, however, the cruise power is increased. In hover, the active chord is best deployed in the middle part of the blade, and just inboard of the tip at high speed flight. The increase in chord length can lead to power savings at high speed flight but the benefits decrease in other speeds. The 1/rev dynamically extendable chord can lead to an overall power reduction over the speed range of a helicopter. The best deployment location is at the blade tip, which is different from the statically extendable chord. It is best extended out in the retreating side, and retracted back in the advancing. The power reduction by the 1/rev dynamically extendable chord increases with the increase in the length of the chord extension and take-off weight of the helicopter. Generally, a lower harmonic extendable chord can save more power than one actuated at higher harmonics. The dynamic chord can reduce more power than the corresponding static chord.
  • Combustion characteristics of hydrogen and cracked kerosene in a DLR
           scramjet combustor using hybrid RANS/LES method
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Junsu Shin, Hong-Gye SungAbstractThis paper applies a zonal hybrid RANS/LES framework to analyze supersonic combustion in a model scramjet combustor. The geometries and boundary conditions of model scramjet combustor are based on an experiment conducted at DLR, German Aerospace Center. This model scramjet combustor was designed to achieve free flight Mach number of 5.5 and total air temperature of 1500 K. Hydrogen at subcritical conditions and thermal/catalytic cracked kerosene at supercritical conditions are injected as fuel. A surrogate of thermal/catalytic cracked kerosene is composed of ethylene and methane in supercritical conditions. To remain consistent with the hydrogen-fueled case, the total equivalence ratio is set to 0.034 for both cases. The total equivalence ratio is quite small, so it is not induced flow separation in the combustor duct. The thermodynamic and transport properties of the supercritical thermal/catalytic cracked kerosene are calculated using the Redlich–Kwong Peng–Robinson cubic equation of state and Chung's model for viscosity and conductivity. This paper focuses on comparisons of the subcritical hydrogen-fueled and supercritical cracked kerosene-fueled scramjet combustors in terms of intrinsic flow and combustion features. The analysis is demonstrated via a reacting regime diagram in nonpremixed turbulent combustion, flame index contours and scatter plots of the flamelet structure. It is found that the cracked kerosene surrogate flame is more vulnerable to quenching than the hydrogen flame, and flame quenching occurs in the immediate vicinity of the injector.
  • LPV gain-scheduled attitude control for satellite with time-varying
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Rongyu Jin, Xueqin Chen, Yunhai Geng, Zhili HouAbstractThe performance of an attitude control system is impacted by the inertia change, especially for small satellites. Although there are various control methods for satellites with inertia uncertainties, very few controllers are designed with consideration of time-varying inertia. In this paper, a linear parameter varying (LPV) model is established for a satellite with time-varying inertia. Moreover, because of their common occurrences, the actuator faults and saturation are considered during the controller design. A gain-scheduled controller is developed to guarantee the steady-state and transient performance of the system by limiting the steady-state variance and regional pole constraints. The simulations indicate that the proposed controller improves the stability performance by adjusting the gain with the inertia variations.
  • Ethylene flame-holding in double ramp flows
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Eric Won Keun Chang, Sungmo Yang, Gisu Park, Hojin ChoiAbstractIn this work, ethylene flame-holding in supersonic flows was investigated using a shock tunnel. The experiments were conducted at flow stagnation temperatures ranging from 1270 to 1810 K. The two-dimensional test model consisted of a double-ramp inlet and a constant-area combustor with a recessed wall cavity. The two fuel injectors were located at the inlet and the other upstream of the cavity. Shadowgraph and flame chemiluminescence images were captured for optical visualization. The inlet injection images showed various flame-holding patterns. At 1270 K, the flame was not maintained. At 1540 K, the flame was maintained inside the cavity, and the condition provided continuous combustion during the steady flow. At 1810 K, strong flame signals were observed from the inlet to the cavity and downstream. At 1540 K, the inlet injection with a low fuel pressure showed a gradual flame quenching in the cavity during flow establishment. On the other hand, the same injection in the combustor showed flame-holding in the shear layer above the cavity. The results showed that the flame patterns are strongly influenced by the flow stagnation temperature and the location of fuel-injection.
  • A composite impact-time-control guidance law and simultaneous arrival
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Zhongtao Cheng, Bo Wang, Lei Liu, Yongji WangAbstractA new composite impact time control guidance law is proposed for simultaneous attack against a ground target. The guidance strategy is composed of two phases. Specifically, the first phase is to guide the missile to proper switching relative range for the second phase. After which, the second phase will lead the missile to the target with a desired impact time by using the proposed Lyapunov based guidance scheme. The impact time is given in a simple analytical form of initial states and switching relative range, the value of which is determined to control the impact time. The proper range of the switching state is analyzed considering the line of sight angle constraint. And the permissible range for impact time can be calculated corresponding to the proper switching state. Two classes of simultaneous attack are considered to demonstrate the effectiveness of the impact time control. One is the sequential launch, and the other is a salvo launch of multiple missiles from different positions. Numerical simulations are carried out to verify the performance of the proposed methodology.
  • Adaptive RBF observer-sliding mode controller design for a two dimensional
           aeroelastic system with unsteady aerodynamics
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Jiaxin Yuan, Na Qi, Zhan Qiu, Fuxin WangAbstractAn adaptive RBF observer-sliding mode controller is designed for the vibration suppression of a two-dimensional aeroelastic system, using a single trailing-edge control surface. The prototypical aeroelastic model describes the plunge and pitch motion of a wing, including cubic nonlinear structural stiffness and unsteady aerodynamics. The unsteady aerodynamics are modeled with an approximation to Theodorsen's theory. It's assumed that only the pitch angle is measured and the remaining state variables needed for full state feedback are estimated by the designed observer. A neural-network is employed to approximate the nonlinear dynamics of the observer system, and then a sliding surface is put forward on the estimation space. The finite-time reachability of the predesigned sliding surface is guaranteed by the designed sliding mode control law. The effectiveness of the proposed strategy is finally demonstrated by simulation results.
  • Reset observer design for time-varying dynamics: Application to WIG crafts
    • Abstract: Publication date: Available online 2 August 2018Source: Aerospace Science and TechnologyAuthor(s): Arash Aminzadeh, Alireza KhayatianAbstractIn this article, a novel reset state estimator for time-varying dynamics is proposed. A first order reset law is used in the structure of observer to improve the performance of the state estimation response which decreases the settling time and overshoot of estimation. The sufficient conditions for stability of this observer based on systems with impulsive effect is addressed which is the main contribution of this study. This observer is applied to a WIG (Wing-In-Ground) craft in the presence of wave and gust which can be modeled by a linear time-varying dynamics. Finally, the five states for longitudinal dynamics of the WIG craft based on a full-order reset observer with time-varying parameters are estimated. The sufficient stability conditions of WIG state observer are stated and simulation results show the superior performance of the reset law.
  • Deep learning based trajectory optimization for UAV aerial refueling
           docking under bow wave
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Yiheng Liu, Honglun Wang, Zikang Su, Jiaxuan FanAbstractIn the autonomous aerial refueling (AAR) docking process, the bow wave generated by the receiver has a strong effect on the drogue, which affects the docking success rate greatly. Thus, a deep learning based trajectory optimization method which aims to decrease the bow wave effect on the drogue is proposed in this paper. There are mainly three parts in the proposed trajectory optimization method. Firstly, a precise bow wave model based on deep learning is presented to estimate the bow wave effect on the drogue. Furthermore, due to the dynamic characteristic of the drogue, a simple and practical drogue motion prediction model under multiple disturbances is carried out to provide a precise prediction of the drogue position at the next time. Moreover, considering the strict attitude constraints requirements in the AAR docking process, a novel reference observer is designed to estimate the receiver attitude from the optimized trajectory under wind perturbations. Then, the proposed trajectory optimization method could not only diminish the bow wave effect on the drogue largely but also satisfy the attitude constraints of the receiver. Finally, the effectiveness of the proposed method is demonstrated by the simulations.
  • Intelligent GNSS/INS integrated navigation system for a commercial UAV
           flight control system
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Guohao Zhang, Li-Ta HsuAbstractOwing to the increase in civil applications using quadcopters, commercial flight control systems such as Pixhawk are a popular solution to provide the sensing and control functions of an unmanned aerial vehicle (UAV). A low-cost global navigation satellite system (GNSS) receiver is crucial for the low-cost flight control system. However, the accuracy of GNSS positioning is severely degraded by the notorious multipath effect in mega-urbanized cities. The multipath effect cannot be eliminated but can be mitigated; hence, the GNSS/inertial navigation system (INS) integrated navigation is a popular approach to reduce this error. This study proposes an adaptive Kalman filter for adjusting the noise covariance of GNSS measurements under different positioning accuracies. The adaptive tuning is based on a proposed accuracy classification model trained by a supervised machine-learning method. First, principal component analysis is employed to identify the significant GNSS accuracy related features. Subsequently, the positioning accuracy model is trained based on a random forest learning algorithm with the labeled real GNSS dataset encompassing most scenarios concerning modern urban areas. To reduce the cases of misclassifying the GNSS accuracy, a fuzzy logic algorithm is employed to consider the GNSS accuracy propagation. Additionally, the process noise covariance of the INS is determined using the Allan variance analysis. The positioning performance of the proposed adaptive Kalman filter is compared with both a conventional Kalman filter and the positioning solution provided by the commercial flight control system, Pixhawk 2. The results show that the proposed adaptive Kalman filter using random forest with fuzzy logic can achieve a better classification of GNSS accuracy compared to the others. The overall positioning result improved by approximately 50% compared with the onboard solution.
  • Aeroelastic analysis of functionally graded rotating blades reinforced
           with graphene nanoplatelets in supersonic flow
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Reza Bahaadini, Ali Reza SaidiAbstractIn this study, the aeroelastic analysis of functionally graded (FG) multilayer graphene platelet reinforced polymer composite (GPLRPC) rotating blades under supersonic flow is investigated. It is considered that the graphene platelet (GPL) nanofillers are distributed in the matrix either uniformly or non-uniformly along the thickness direction. Four GPL distribution patterns namely, U-GPLRPC, Λ-GPLRPC, X-GPLRPC and O-GPLRPC are considered. The effective material properties of GPLRPC layers are obtained via the modified Halpin–Tsai micromechanics model and the rule of mixture. Based on the first-order shear deformation theory, the dynamic equations of thin-walled blades reinforced with GPL are obtained using extended Hamilton's principle. The aerodynamic pressure is assumed in accordance with the quasi-steady supersonic piston theory. The extended Galerkin method (EGM) is employed to transform the coupled equations of motion to a general eigenvalue problem. The influences of rotating speed, GPL distribution, GPL weight fraction, geometry of GPL nanofillers, geometric parameters and Mach number on the natural frequencies of the system are studied. The results indicate that the Λ-GPLRPC distribution pattern predicts the highest natural frequencies for the composite blade. Also, the natural frequencies of the composite blade significantly increase by adding a small amount of GPL to the polymer matrix.
  • A comprehensive high-order solver verification methodology for free fluid
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Farshad Navah, Siva NadarajahAbstractThe aim of this article is to present a comprehensive methodology for the verification of computational fluid dynamics (CFD) solvers with a special attention to aspects pertinent to discretizations with orders of accuracy (OOAs) higher than two. The method of manufactured solutions (MMS) is adopted and a series of manufactured solutions (MSs) is introduced that examines various components of CFD solvers for free flows (not bounded by walls), including inviscid, laminar and turbulent problems when the latter are modeled by the Reynolds-averaged Navier–Stokes (RANS) equations. The treatment of curved elements is also examined. These MSs are furthermore conceived with demonstrated suitability for the verification of OOAs up to the sixth. Each MS is as well utilized to discuss salient aspects useful to the code verification methodology such as the relative qualities of the most useful norms in measuring the discretization error, the sensitivity analysis of the verification process to forcing function terms, the relation between residual minimization and discretization error convergence in iterative solutions and finally the sensitivity of high-order discretizations to grid stretching and self-similarity. Furthermore, scripts and code are provided as accompanying material to assist the interested reader in reproducing the verification results of each manufactured solution (MS).
  • Heavy rain effects on aircraft lateral/directional stability and control
           determined from numerical simulation data
    • Abstract: Publication date: Available online 25 July 2018Source: Aerospace Science and TechnologyAuthor(s): Zhenlong Wu, Benyin Lv, Yihua CaoAbstractIn this paper, effects of heavy rain on the lateral/directional stability and control performance of a DHC-6 Twin Otter aircraft are determined based on from the numerical simulation data. A two-way momentum coupled Eulerian–Lagrangian approach developed for two-phase flow simulation is adopted to obtain the fundamental aerodynamic coefficients in the rain condition. Then the lateral/directional aerodynamic derivatives of the aircraft in rain are evaluated based on the linear fitting processing and the strip theory. Finally, the lateral/directional stability of the aircraft in rain is obtained and the controllability is analyzed by simulation of the responses to unit step changes in the control surface deflections. The present results indicate that heavy rain can cause penalties in both the aerodynamic and flight mechanical performance of aircraft.
  • Numerical investigations on stator hub initiated stall in a single-stage
           transonic axial compressor
    • Abstract: Publication date: Available online 18 July 2018Source: Aerospace Science and TechnologyAuthor(s): Jiaguo Hu, Qiushi Li, Tianyu Pan, Yifang GongAbstractA new kind of stall inception initiated from the stator hub region was recently identified by experiment in a transonic compressor. To further explore the mechanisms of the new stall inception, the detailed stall evolution is studied in this paper using full-annulus unsteady simulations. The simulations correctly predict several key characteristics observed in the previous experiments: (1) the stall precursor is initiated in the stator hub region; (2) the initial disturbance is axisymmetric; (3) asymmetric rotating disturbance is developed afterwards. The numerical results also illustrate that the stall evolution has two distinct phases: the part span stall and the global stall, which are associated with the axisymmetric and the asymmetric disturbances respectively. The axisymmetric disturbance is caused by the ring-shaped flow separation in the stator hub region, while the asymmetric disturbance is initiated by the breakdown of the symmetry of the ring-shaped separation. For both disturbances, the axial velocity waveforms are in anti-phase at the stator tip and hub region, so they have the 1st-order mode in the span-wise direction. Further discussions on the radial distribution of load indicate that the localized critical load is the key factor leading to the earlier occurrence of flow breakdown in the stator hub region.
  • Autonomous orbit and attitude determination for Earth satellites using
           images of regular-shaped ground objects
    • Abstract: Publication date: Available online 17 July 2018Source: Aerospace Science and TechnologyAuthor(s): Muzi Li, Bo XuAbstractOrbit and attitude determination (OAD) is a crucial problem in spacecraft missions. For Earth satellites, especially Earth-observing satellites, images of ground landmarks are feasible to be used for autonomous OAD. Previous studies have demonstrated the possibility of image-based OAD using the light-of-sight (LOS) vector to ground point feature as measurement. However, valid ground point features might not always be available, for example, when the ground landmarks have smooth edges. Besides, when ground landmarks are regular-shaped, this method would discard useful shape information of the objects and affect OAD precision. Aimed at the problem, we present a new OAD scheme using onboard images of regular-shaped ground objects. In the scheme, the LOS vector to the center of the ground objects is used as measurement and is obtained based on the widely-used 3D reconstruction algorithms in machine vision. With the help of an extended Kalman filter (EKF), orbit and attitude parameters can be estimated. The OAD performance of the scheme is assessed using Monte-Carlo simulations. Results demonstrate the feasibility of the scheme and the substantial influence from both lighting constraints and image sampling frequency on the OAD performance. The scheme is also compared with a deterministic scheme, which directly derives position and attitude parameters purely based on the 3D reconstruction algorithm. For the current sensor precision, the proposed scheme is found to have a better performance.
  • Autonomous entry guidance using Linear Pseudospectral Model Predictive
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Liang Yang, Xiaoming Liu, Wanchun Chen, Hao ZhouAbstractThis paper aims at developing an autonomous entry guidance method that requires no mission-dependent adjustments and will be applicable to a wide range of entry scenarios with worldwide destinations. Firstly, a nonlinear reduced-order entry dynamical system with coupled lateral and longitudinal motions as well as a spherical and rotating Earth is proposed. This system relates the vertical and lateral lift-to-drag ratio profiles to the position and azimuth angle variables with the energy as the independent variable. It has a high computational accuracy even only three differential equations are involved. Secondly, a trajectory planning problem with only two bank reversals is formulated based on this reduced-order system with parameterized control. Trajectory integration prediction and linearization method are applied to transfer the original planning problem into iteratively solving a group of linear dynamical equations. Gauss Pseudospectral Method and Calculus of variations are employed to discrete them so as to derive a series of analytical correction formulas to eliminate the final errors, mathematically, which achieves high accuracy with a small number of points. Moreover, these control parameters include not only the magnitude of bank angle, but also bank reversal points, which will significantly increase its ability to shape the entry trajectory. After the last bank reversal, lateral and longitudinal guidance laws are designed to ensure multiple final constraints. Nominal testing and Monte Carlo simulations on the proposed method and the comparison with the typical predictor–corrector method are carried out. Results demonstrate that, even in highly dispersed environments, this method has wide applicability, strong robustness, and excellent performance. Moreover, its computational efficiency is so high that it sufficiently satisfies the requirement on onboard application.
  • Two controller designs of hypersonic flight vehicle under actuator
           dynamics and AOA constraint
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Yuyan Guo, Bin Xu, Xiaoxiang Hu, Xiangwei Bu, Yu ZhangAbstractThis paper studies the controller design for hypersonic flight vehicle (HFV) considering the actuator hysteresis and the angle of attack (AOA) constraint respectively. Firstly, the effect of input hysteresis dynamics are regarded as the combination of the partial loss of effectiveness (PLOE) and a bounded disturbance-like item, while the controller is modified by constructing adaptive compensation laws. Secondly, to ensure the constraint on AOA, a prescribed performance control (PPC) method is designed while the AOA tracking error is restricted by the pre-defined performance function. The two controllers are designed based on dynamic surface control, where a nonlinear adaptive filter is utilized to eliminate the bound layer error. The effectiveness of the designed controllers are demonstrated via the simulation results.
  • A new Gaussian mixture method with exactly exploiting the negative
           information for GMTI radar tracking in a low-observable environment
    • Abstract: Publication date: September 2018Source: Aerospace Science and Technology, Volume 80Author(s): Liyun Gong, Miao YuAbstractThis paper investigates the problem of ground vehicle tracking with a Ground Moving Target Indicator (GMTI) radar. In practice, the movement of ground vehicles may involve several different maneuvering types (acceleration, deceleration, standstill, etc.). Consequently, the GMTI radar may lose measurements when the radial velocity of the ground vehicle is below a threshold when it stops, i.e. falling into the Doppler blind region. Besides, there will be false alarms in low-observable environments where there exist high noises interferences. In this paper, we develop a novel algorithm for the GMTI tracking in a low-observable environment with false alarms while exactly incorporating the ‘negative information’ (i.e., the target is likely to stop when no measurements are recorded) based on the Bayesian inference framework. For the Bayesian inference implementation, the Gaussian mixture approximation method is adopted to approximate related distributions, while different filtering algorithms (including both extended Kalman filter and its generalization for interval-censored measurements) are applied for updating the Gaussian mixture components. Target state estimation can be directly obtained through the Gaussian mixture model for the GMTI tracking at every time instance. We have compared the developed method with other state-of-the-art ones and the simulation results show that the proposed method substantially outperforms the existing methods for the GMTI tracking problem.
  • Conceptual design and CFD analysis of a new prototype of agricultural
    • Abstract: Publication date: Available online 14 July 2018Source: Aerospace Science and TechnologyAuthor(s): Pedro David Bravo-Mosquera, Hernán Darío Cerón-Muñoz, Guillermo Díaz-Vázquez, Fernando Martini CatalanoAbstractIn the agricultural aviation, there are several aerodynamic factors that must be optimized in order to contribute to the successful application of agricultural products, such as the high aerodynamic efficiency (L/D) required at the working phase and the influence of aircraft wingtip phenomena on the spray deposition and movement. For these reasons, in this research is presented the conceptual design of an advanced prototype of agricultural aircraft, whose main characteristic is an adaptive multi-winglet device installed on the wingtips, which optimized the main aerodynamic issues presented in its mission. Traditional aircraft design methods were used to develop and assess the suitability of the aircraft, focusing on its design requirements and tackling studies of weight sizing, pilot ergonomics, aerodynamics, stability, and performance. Subsequently, analytical and computational methods were used to design the adaptive multi-winglet device, which is composed by three winglets with its own geometry fitted on a tip-tank. Six configurations were created by modifying only the cant angle of each winglet in order to determine the arrangement that provides the best aerodynamic characteristics through a study of computational fluid dynamics (CFD), using the Reynolds–Averaged–Navier–Stokes (RANS) equations coupled with the Shear Stress Transport (SST) turbulence model. First, the flow around the wing and the multi-winglet section of the aircraft was investigated exclusively. Afterward, the airflow around the entire aircraft was studied at the product application condition, in order to compare the overall aerodynamic performance of the baseline concept along with the optimal multi-winglet configuration installed on the aircraft. Lift, drag and pitching moment coefficients were assessed, as well as the wingtip vortex structure of the most relevant configurations. Results of this study showed that adaptive multi-winglet devices are a promising alternative to improve the overall performance of an agricultural aircraft, because they provide control over the size and strength of the wake-spray interaction on the sprayed product, reduce the induced drag, reduce the bending moment and improve the aerodynamic efficiency of the aircraft.
  • Three-dimensional cooperative guidance and control law for multiple
           reentry missiles with time-varying velocities
    • Abstract: Publication date: Available online 14 July 2018Source: Aerospace Science and TechnologyAuthor(s): Xiaofang Wang, Yiwei Zhang, Dongze Liu, Min HeAbstractA three-dimensional partial integrated guidance and control law is proposed for cooperative flight of multiple hypersonic reentry missiles with uncontrollable and time-varying velocities. As for the problem of multiply reentry missiles with varying velocities attacking the target synchronously, a cooperative scheme is presented by adjusting the lateral pre-setting angle of velocity. In addition, a partial integrated guidance and control method with a two-loop controller structure is designed to realize the proposed cooperative scheme. Considering the saturation of the velocity slope angle and the unknown uncertainty, the two-loop three-channel controller of each reentry missile is designed based on dynamic inverse theory, dynamic surface control theory and extended state observer. The stability of the closed-loop system is demonstrated by Lyapunov theory. Simulation results verify the effectiveness and superiority of the proposed guidance and control law.
  • Robust two-stage rank filter for Mars entry navigation under parameter
    • Abstract: Publication date: Available online 28 June 2018Source: Aerospace Science and TechnologyAuthor(s): Mengli Xiao, Yongbo Zhang, Huimin Fu, Qiang XiaoAbstractFuture Mars landing missions require high-precision entry navigation capability, but the uncertainties lie in the Mars atmosphere density and aerodynamic coefficients may cause serious performance degeneration of conventional Kalman filters. This paper proposed a robust two-stage rank filter (RTSRF) out of augmented state rank filter by decoupling the covariance matrices and modifying the measurement update equations. Therefore, this filter could obtain an accurate state estimation by treating the parameter uncertainties as the unknown inputs even the prior knowledge of the inputs are completely unknown, a novel method to dealing with the parameter uncertainties. Then, it was applied for Mars entry navigation simulation, and the results show that the navigation accuracy is greatly improved, fulfilling the requirement of future Mars pinpoint landing missions.
  • Criteria for designing low-loss and wide operation range variable inlet
           guide vanes
    • Abstract: Publication date: Available online 12 July 2018Source: Aerospace Science and TechnologyAuthor(s): Hengtao Shi, Baojie Liu, Xianjun YuAbstractCritical factors influencing the variable inlet guide vane (VIGV) profile loss at high incidence condition were studied by using numerical methods and a practical design criterion for designing wide low-loss operation range VIGVs in axial-flow compressor was proposed. At first, to acquire research samples, a series of airfoils with different shapes were generated for two selected representative VIGV cascade cases. Steady simulations based on Reynolds-Averaged Navier-Stokes method, carried out by commercial software CFX and validated with experimental data after grid independent study, were first conducted to predict the aerodynamic performances, the surface velocity distributions and the boundary-layer behaviors of the generated airfoils. Based on the simulated results, the influences of geometric parameters on airfoil performances were analyzed and the geometric features of low-loss VIGV airfoil were revealed. Further analysis indicated that the magnitude of airfoil loss at high incidence condition were mainly influenced by the scales of two boundary-layer separation regions: one at the leading edge caused by the high adverse pressure gradient induced by the suction spike and the other one caused by the adverse pressure gradient induced by the re-acceleration flow. To reveal the influence of the suction spike and the re-acceleration flow on the scales of separation regions, two practical parameters Dspike and Are were defined. It was found that there exists an optimized range of the Dspike and Are which could keep the separation flow to a small scale at high incidence condition and can be used as a surface velocity design criterion for designing wide low-loss operation range VIGVs. Moreover, the methods for choosing the airfoil geometric parameters to achieve the preferred surface velocity distribution were discussed. Finally, the developed design criterion was used to guide the airfoil modification of an axial-flow compressor VIGV and achieved an average of 19%, 52% and 73% loss coefficient reduction at three high stagger angle operating points, which confirms the applicability and effectiveness of the design criterion in three-dimensional environment.
  • Efficient uncertainty quantification for a hypersonic trailing-edge flap,
           using gradient-enhanced kriging
    • Abstract: Publication date: Available online 11 July 2018Source: Aerospace Science and TechnologyAuthor(s): Sudip Bhattrai, Jouke H.S. de Baar, Andrew J. NeelyAbstractWe present a numerical study on the uncertainty quantification (UQ) of aerodynamic forces acting on a hypersonic trailing-edge flap model, as a result of input uncertainties in the experimental boundary conditions. The complex fluid–thermal-structural interaction on aerodynamic surfaces of a hypersonic flight vehicle and fluctuations in flow conditions result in uncertainties in their aerodynamic characteristics. We run the numerical simulations in US3D to quantify these uncertainties. Altogether four input uncertain parameters—inlet flow velocity, density, temperature, and the model wall temperature—are considered. We obtain the aerodynamic forces from the primal solve, as well as gradient information from a dedicated sensitivity solver. We compare the surrogate-based UQ analysis using kriging as well as gradient-enhanced kriging (GEK), accounting for significant observation errors in the gradients, and quantify the accuracy of the output probability density function (PDF). The accuracy of the predicted output PDF converges faster for GEK than for kriging, implying the importance of the gradient information in order to reduce the computational cost—in the present case, the computational cost is reduced by a median speed-up of roughly 3.0 by exploiting the gradient information available from the sensitivity solver.
  • Dynamic surface control design of post-stall maneuver under unsteady
    • Abstract: Publication date: Available online 11 July 2018Source: Aerospace Science and TechnologyAuthor(s): Yongxi Lyu, Yuyan Cao, Weiguo Zhang, Jingping Shi, Xiaobo QuAbstractThis paper presents an efficient method that overcomes the problem of the control design of the post-stall maneuver under unsteady aerodynamics. On the basis of adequate data of the large amplitude oscillation experiment device in wind tunnel test, the unsteady aerodynamics model with nonlinearity, coupling and hysteresis is established by the improved Extreme Learning Machine (ELM) method. Considering the nonlinearity of the longitudinal model of the advanced fighter and the aerodynamics characteristics of the post-stall maneuver, the control law under large attack angle is designed combining the backstepping method and the daisy chain allocation method. The first order filter is adopted to prevent the “differential explosion” problem. The designed control allocation law guarantees that the conventional surfaces and the vector nozzle deflect coordinately within the position limits and the rate limits. The Radial Basis Function (RBF) network is applied to model the uncertainty, and the stability of the proposed control law which considering the uncertainty is also proved. Digital simulations of the typical “Cobra” maneuver under the unsteady aerodynamics are completed with comparisons under different conditions. Simulations results verify the validity of the proposed control law under unsteady aerodynamics and the aerodynamics uncertainty.
  • Rotorcraft comprehensive code assessment for blade-vortex interaction
    • Abstract: Publication date: Available online 11 July 2018Source: Aerospace Science and TechnologyAuthor(s): Massimo Gennaretti, Giovanni Bernardini, Jacopo Serafini, Gianluca RomaniAbstractThe scope of this paper is the presentation of the computational methodologies applied in the comprehensive code for rotorcraft developed in the last years at Roma Tre University, along with the assessment of its prediction capabilities focused on flight conditions characterized by strong blade-vortex interactions. Boundary element method approaches are applied for both potential aerodynamics and aeroacoustics solutions, whereas a harmonic-balance/modal approach is used to integrate the rotor aeroelastic equations. The validation campaign of the comprehensive code has been carried out against the well-known HART II database, which is the outcome of a joint multi-national effort aimed at performing wind tunnel measurements of loads, blade deflection, wake shape and noise concerning a four-bladed model rotor in low-speed descent flight. Comparisons with numerical simulations available in the literature for the same test cases are also presented. It is shown that, with limited computational cost, the results provided by the Roma Tre aero-acousto-elastic solver are in good agreement with the experimental data, with a level of accuracy that is in line with the state-of-the-art predictions. The influence of the vortex core modeling on aerodynamic predictions and the influence of the inclusion of the fuselage shielding effect on aeroacoustic predictions are discussed.
  • Study of boundary layer transition on supercritical natural laminar flow
           wing at high Reynolds number through wind tunnel experiment
    • Abstract: Publication date: Available online 11 July 2018Source: Aerospace Science and TechnologyAuthor(s): Jiakuan Xu, Ziyuan Fu, Junqiang Bai, Yang Zhang, Zhuoyi Duan, Yanjun ZhangAbstractIn order to achieve the goal of green aviation, energy conservation and emission reduction, laminar flow design technology has become a hot research topic. For transonic airliners, supercritical natural laminar flow wing design technology will significantly improve the aerodynamic performance(reduce flight drag, decrease fuel consumption and pollutant emissions). In this paper, airfoil optimization design system is applied to design the supercritical natural laminar flow airfoils based on high-precision boundary layer transition prediction technique. Then, three-dimensional layout of supercritical natural laminar flow wing is formed. Numerical simulations have been conducted to verify the laminar flow properties. In addition, the aerodynamic model with ratio of 1: 10.4 is processed to measure boundary layer transition phenomena in the high speed and low turbulence wind tunnel in Netherland. Temperature sensitive paint (TSP) technique is used to photograph laminar-turbulent distribution at different Mach numbers, Reynolds numbers and angels of attack. In the following content, boundary layer transition properties of the supercritical natural laminar flow wing are analyzed using TSP results and CFD simulations. Finally, key factors of supercritical natural laminar wing design and corresponding boundary layer transition properties are summarized. In addition, the research about transition properties of supercritical natural laminar flow wing at high Reynolds numbers have guiding significance for aircraft designers and transition researchers.
  • Numerical evaluation of station-keeping strategies for stratospheric
    • Abstract: Publication date: Available online 11 July 2018Source: Aerospace Science and TechnologyAuthor(s): Sai Sudha Ramesh, Juanli Ma, Kian Meng Lim, Heow Pueh Lee, Boo Cheong KhooAbstractThe trajectory control of stratospheric balloons poses a great challenge, given their importance in scientific explorations and military applications. This has led to the investigation of several trajectory control methods, which aim to retain the balloon system within a specific region. In particular, the use of a control device tethered to the main balloon is required to provide sufficient lateral forces to counteract the air drag on the main balloon. The present study evaluates the performance of two kinds of balloon systems namely, the dual-balloon and balloon–stratosail systems that have a control device tethered to the main balloon. The study compares their performances in the context of realistic wind conditions, and presents the best working range for each of the system, which may be useful in improvising the design of control devices for future applications.
  • The optimization and flow diagnoses for a transonic fan with stage flow
    • Abstract: Publication date: Available online 10 July 2018Source: Aerospace Science and TechnologyAuthor(s): Huanlong Chen, Yong Qin, Ruoyu WangAbstractThe redesign and optimization of a low-aspect ratio transonic fan is implemented in this study. An advanced 3D aerodynamic optimization design system is adopted, while flow diagnostic methods are employed to discuss the transonic flow in the blade passages. On the basis of maintaining high aerodynamic efficiency, the study seeks to improve the pressure ratio and through-flow capability of the redesigned fan stage. Furthermore, the dynamic principle for the redistribution of passage flow due to geometry change is revealed. In comparison with the prototype, the total pressure ratio of the redesigned fan is increased by 7.54% at the design point, while its mass flow rate and adiabatic efficiency are raised by 6.30% and 1.25%, respectively. Additionally, a wider high-efficiency operation range is also achieved by the optimization. Under stage flow condition, the control of shock wave at the rotor tip and the removal of low momentum fluid in the stator corner are the two keys in improving the aerodynamic performance of the redesigned fan. Moreover, tangential lean of the stator blade has also succeeded in delaying corner flow separations. Further research for these design techniques would give potential to expand the design system for transonic fan/compressor with low-aspect ratios.
  • Efficient numerical algorithm of profust reliability analysis: An
           application to wing box structure
    • Abstract: Publication date: Available online 10 July 2018Source: Aerospace Science and TechnologyAuthor(s): Kaixuan Feng, Zhenzhou Lu, Chao Pang, Wanying YunAbstractIn aerospace engineering, the reliability analysis technique attracts increasing attention from many structure designers. Compared with the conventional reliability analysis, the probability and fuzzy state assumption (profust) reliability analysis is proposed as a more comprehensive and objective theory to evaluate the structural safety. Because of great computational burden of the existing methods in estimating profust failure probability, an efficient numerical algorithm is developed in this paper. The proposed method is based on an equivalent transformation of the profust failure probability, then the profust failure probability can be rewritten as the form of a series of conventional failure probabilities which have similar constructions. The subset simulation method is employed to compute all the conventional failure probabilities by using a set of samples repeatedly. Next, this method is applied to estimate the profust failure probability of a wing box structure. The calculation results indicate that the proposed method can reduce the computational cost dramatically with acceptable precision.
  • Technology and opportunities of photon sieve CubeSat with deployable
           optical membrane
    • Abstract: Publication date: Available online 10 July 2018Source: Aerospace Science and TechnologyAuthor(s): Hyun Jung Kim, Shravan Hariharan, Matthew Julian, David MacdonnellAbstractA photon sieve (PS) is a revolutionary optical instrument that provides high resolution imaging at a fraction of the weight of typical telescopes, with an areal density of 0.3 kg/m2 compared to 25 kg/m2 for the James Webb Space Telescope. The photon sieve is a variation of a Fresnel Zone Plate consisting of many small holes arranged in a series of concentric rings. The sieve works by diffracting light of a certain wavelength to a specified focal point for imaging, so that only a specific wavelength can be imaged. Moreover, the better image contrast and higher signal-to-noise ratios come from suppressing higher diffracted orders by apodizing the number of pinholes in the outer rings. Finally, a photon sieve requires less supports and can withstand more deformation without a reduction in the imaging qualities.Due to these properties, various groups have created PS CubeSats for Earth and Sun imaging at a low cost and weight specifically using deployable technology. A team at the Air Force Research Laboratory created a design and prototype of a mechanism that deploys a 20 cm diameter photon sieve. The United States Air Force Laboratory used a similar design to create a CubeSat-based deployable photon sieve. The team at NASA Langley Research Center has researched photon sieves for conducting an Earth-observing experiment using LIDAR (Light Detection and Ranging) with a higher signal-to-noise ratio benefit from the PS. This paper provides a state of the art overview on existing PS CubeSat technology with deployable structures and applications. Especially, the paper introduces PS for LIDAR applications and discusses the CubeSat-based PS challenge being worked at the NASA Langley Research Center.
  • Switched propulsion force libration control for the low-thrust space tug
    • Abstract: Publication date: Available online 10 July 2018Source: Aerospace Science and TechnologyAuthor(s): Xin Sun, Rui ZhongAbstractThe number of micro-satellites launched by commercial rockets and designed by universities and institutions is experiencing rapid growth, inevitably increasing the amount of the space debris of the same size. This paper adopts a low-thrust tethered space tug system to achieve debris deorbit. A switched low-propulsion force control method using only two constant-thrust modes to achieve both deorbit and libration control is presented. This kind of control method benefits from its robustness and low demand on the output accuracy of the thruster. The harmonic-like libration dynamics of the tethered space tug system around the local horizontal configuration is discussed and the stability of the control method is proved. Moreover, the sufficient condition for the switching sustainability is presented. Afterwards, the control effects of such a system are illustrated using numerical simulations. A modified control law to adapt practical demands shows the flexibility of the switching control methodology.
  • Comparison of studies on flow and flame structures in different swirl
    • Abstract: Publication date: Available online 3 July 2018Source: Aerospace Science and TechnologyAuthor(s): L.X. ZhouAbstractSwirling gas combustion and two-phase combustion (spray-air or particle-air combustion) are widely encountered in gas-turbine combustors and internal combustion engines. In experimental studies of practical combustors frequently it is difficult to get the detailed information inside the combustors. Most of numerical simulation is Reynolds-averaged (RANS) modeling, which cannot give detailed flow and flame structures. Some investigators reported their results using large-eddy simulation (LES) with different combustion models. The present authors did systematic LES studies on swirling gas combustion and two-phase combustion using second-order moment (SOM) and EBU combustion models. The statistic results are assessed by the measurement results. The instantaneous results show the flow and combustion behavior in different swirl combustors.
  • Design of mistuning patterns to control the vibration amplitude of
           unstable rotor blades
    • Abstract: Publication date: Available online 2 July 2018Source: Aerospace Science and TechnologyAuthor(s): Roque Corral, Oualid Khemiri, Carlos MartelAbstractFlutter onset is one of the major causes for increased vibration levels in low pressure turbine (LPT) rotor blades. This paper describes the design process and the experimental testing of intentional mistuning patterns specifically chosen to show the possibility to control the flutter characteristics of an LPT rotor. The Asymptotic Mistuning Model (AMM) methodology is used to select the intentional mistuning patterns. The AMM formulation incorporates elastic and aerodynamic data from detailed FEM and CFD computations, and measured values of the rotor blades intrinsic mistuning. The intentional mistuning patterns are implemented in the rotor by mounting small masses at the tip-shroud of the blades, and the effect of these small masses is also introduced in the AMM description. Two intentional mistuning patterns are selected. The classical alternate mistuning pattern, designed to fully suppress flutter, and a second intentional mistuning pattern that is designed with the idea of halving the vibration amplitude of the tuned unstable rotor. This second mistuning pattern demonstrates that, through the implementation of the appropriate intentional mistuning pattern, flutter cannot only be suppressed but also modulated. The two mistuned LPT rotors were tested in a free flutter experiment at a high speed rotating wind-tunnel, and the experimental results showed a good agreement with the AMM stability predictions. This is the first time, to our knowledge, that the possibility to control flutter through intentional mistuning has been experimentally validated in a rotating rig. The AMM is also applied to evaluate the effect of the aerodynamic coupling in the stability calculations of the mistuned rotors, and the AMM results are compared with high fidelity numerical calculations. It is shown that, despite its very reduced formulation, the AMM produces quite accurate stability predictions, and that it is essential to take into account the aerodynamic coupling; if it is not considered then the instability level of the mistuned rotors can be substantially underestimated.
  • Fuzzy adaptive non-affine attitude tracking control for a generic
           hypersonic flight vehicle
    • Abstract: Publication date: Available online 2 July 2018Source: Aerospace Science and TechnologyAuthor(s): Yuhui Wang, Mou Chen, Qingxian Wu, Jun ZhangAbstractA fuzzy adaptive non-affine attitude tracking control method is proposed for a generic hypersonic flight vehicle (HFV). Due to the complexities of the hypersonic flows and nonlinear dynamics, the aerodynamic coefficients of the HFV are dependent not only on the attack of angle, sideslip angle, angular rates, and Mach number, but also on the deflection angles of the control surfaces. This cause the attitude tracking control problem becomes a non-affine multi-input–multi-output (MIMO) one. By analyzing the characteristics of the aerodynamic coefficients, it can be found that the non-affine terms have a great influence on the attitude dynamics and should not be ignored. Then, a non-affine MIMO attitude tracking controller is designed using fuzzy sliding mode adaptive techniques with consideration of the non-affine nonlinear aerodynamic coefficients. The proposed controller has good practicability because it does not need to know the exact bound values of the uncertainties, unmodeled dynamics, and external disturbances. Finally, simulation results show that the attitude angles can track the desired signals robustly and asymptotically under the control.
  • Synthesis of attitude control for statically unstable hypersonic vehicle
           with low-frequency aero-servo-elastic effect
    • Abstract: Publication date: Available online 2 July 2018Source: Aerospace Science and TechnologyAuthor(s): Minnan Piao, Zhihong Yang, Mingwei Sun, Jian Huang, Zenghui Wang, Zengqiang ChenAbstractIn this paper, an integrated attitude control scheme is proposed for a statically unstable hypersonic vehicle (HSV) with low-frequency aero-servo-elastic effect. Linear active disturbance rejection control (LADRC) is designed for the pitch angle to address the strong uncertainties and coupling effects. For the particular aero-servo-elastic effect with low frequency, a hybrid phase and gain stabilization technique is embedded into the LADRC framework to suppress the structural modes, which can reduce the phase loss around the crossover frequency. To achieve satisfactory tracking performance and robustness, a non-smooth H∞ optimization approach is applied to tune the controller gains, which simplifies the tuning and gain scheduling process. Monte Carlo simulation results demonstrate the effectiveness of the proposed control scheme.
  • Cooperative interception strategy for multiple inferior missiles against
           one highly maneuvering target
    • Abstract: Publication date: Available online 2 July 2018Source: Aerospace Science and TechnologyAuthor(s): Wenshan Su, Hyo-Sang Shin, Lei Chen, Antonios TsourdosAbstractThis paper proposes a novel cooperative guidance strategy, which aims to intercept one highly maneuvering target with multiple inferior missiles. In the scenario of interest, both the missiles and target are assumed to have bounded maneuverability, and the guidance goal is to make the joint reachable set of interceptors cooperatively cover the target maneuvering range. Under this guidance scheme, a preprogrammed covering strategy and an adaptive covering strategy are designed for the missile teams without and with communication capability respectively. The former attempts to specify different subsets of the target maneuvering range to different missiles as the expected reachable sets, while the latter aims to coordinate the expected reachable sets of different missiles dynamically according to the changing engagement situation. Considering the disadvantage of inferior maneuverability, the inherent limitations of the proposed guidance scheme are discussed. Numerical simulations with different target maneuvering modes demonstrate the prominent performance improvements of cooperative strategy over the traditional guidance laws.
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Heriot-Watt University
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