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Aerospace Science and Technology
Journal Prestige (SJR): 0.796
Citation Impact (citeScore): 3
Number of Followers: 359  
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 1270-9638
Published by Elsevier Homepage  [3160 journals]
  • High-dynamic baseline determination for the Swarm constellation
    • Abstract: Publication date: Available online 20 March 2019Source: Aerospace Science and TechnologyAuthor(s): X. Mao, P.N.A.M. Visser, J. van den IJssel Baseline determination for the European Space Agency Swarm magnetic field mission is investigated. Swarm consists of three identical satellites -A, -B and -C. The Swarm-A and -C form a pendulum formation whose baseline length varies between about 30 and 180 km. Swarm-B flies in a higher orbit, causing its orbital plane to slowly rotate with respect to those of Swarm-A and -C. This special geometry results in short periods when the Swarm-B satellite is adjacent to the other Swarm satellites. Ten 24-hr periods around such close encounters have been selected, with baseline lengths varying between 50 and 3500 km. All Swarm satellites carry high-quality, dual-frequency and identical Global Positioning System receivers not only allowing precise orbit determination of the single Swarm satellites, but also allowing a rigorous assessment of the capability of precise baseline determination between the three satellites. These baselines include the high-dynamic baselines between Swarm-B and the other two Swarm satellites.For all orbit determinations, use was made of an Iterative Extended Kalman Filter approach, which could run in single-, dual-, and triple-satellite mode. Results showed that resolving the issue of half-cycle carrier phase ambiguities (present in original release of GPS RINEX data) and reducing the code observation noise by the German Space Operations Center converter improved the consistency of reduced-dynamic and kinematic baseline solutions for both the Swarm-A/C pendulum pair and other combinations of Swarm satellites. All modes led to comparable consistencies between the computed orbit solutions and satellite laser ranging observations at a level of 2 cm. In addition, the consistencies with single-satellite ambiguity fixed orbit solutions by the German Space Operations Center are at comparable levels for all the modes, with reduced-dynamic baseline consistency at a level of 1-3 mm for the pendulum Swarm-A/C formation and 3-5 mm for the high-dynamic Swarm-B/A and -B/C satellite pairs in different directions.
  • Accurate Higher Order Automated Unstructured Triangular Meshes for Airfoil
           Designs in Aerospace Applications using Parabolic Arcs
    • Abstract: Publication date: Available online 20 March 2019Source: Aerospace Science and TechnologyAuthor(s): Supriya Devi, K.V. Nagaraja, T.V. Smitha, Sarada Jayan This paper presents automatically generated higher order curved triangular meshes around airfoil design using MATLAB code. This work shows a valuable basis for the finite element procedures involved in evaluating aerodynamic performances. Finite element method (FEM) effectively solves all computational fluid dynamics problems around the airfoil and for that region around the airfoil that has been discretized with unstructured curved triangular elements. Meshes have been formed on the basis of subparametric transformation created for the curved triangular element obtained from the nodal relations of parabolic arcs. This scheme can be used to obtain the output data of node coordinates, element connectivity and boundary values for all discretized elements over the airfoil design. A spectacular work done on linear triangular element meshing over a domain by Persson and Gilbert Strang is the basis of present meshing scheme. The proposed meshing scheme presents a refined higher order (HO) curved triangular discretization of few airfoil designs namely NACA0012, NACA0015 and NACA0021 inscribed inside a circle. The approach of the suggested meshing scheme described in this paper can be applied to numerous aerospace applications such as computing pressure gradients, understanding atmospheric nature study, evaluating laminar viscous compressible flow around the airfoil shape, etc. The element and nodal information gained from this discretization is useful for the numerical solutions of FEM and for the aerodynamic portrayal. This paper is aimed at the innovative discretization scheme which can be extended to all kinds of NACA airfoil designs. We have provided the MATLAB code AirfoilHOmesh2d for HO curved meshing around an airfoil with a cubic order triangular element. The mathematical explanation of this along with the description and implementation of it on few airfoil designs is described. The flowchart of the MATLAB code for cubic order meshing over airfoil design has been provided. This implementation supports many applications in an aerodynamic performance that have been elaborated in this paper. Two applications for the analysis of potential flow around airfoil and computation of pressure coefficient (Cp)on the surface of the airfoil design have been performed. It has been verified and found that the present HO curved meshing technique efficiently gives converging solution.
  • A Statistical Energy Analysis (SEA) model of a fuselage section for the
           prediction of the internal Sound Pressure Level (SPL) at cruise flight
    • Abstract: Publication date: Available online 20 March 2019Source: Aerospace Science and TechnologyAuthor(s): Giuseppe Petrone, Giacomo Melillo, Aurelio Laudiero, Sergio De Rosa Comfort plays an increasingly important role in the interior design of airplanes. In general, comfort is defined as ‘freedom from pain, well-being’; in scientific literature, indeed, it is defined as a pleasant state of physiological, psychological and physical harmony between a human being and the environment or a sense of subjective well-being. Cabin noise in passenger aircraft is one of the comfort parameter, which creates straightaway discomfort when exceeding personal thresholds. In general the cabin noise varies by the seat position and changes with flight condition. It is driven by several source types, which are transmitted through different transfer paths into the cabin. In the forward area the noise is mainly dominated by the turbulent boundary layer described by pressure vortexes traveling along the fuselage surface.In this paper evaluation of the Sound Pressure level, for the medium-high frequency range, of an aircraft fuselage section at different stations and locations inside the cabin has been performed numerically by using Statistical Energy Analysis (SEA) method. Different configurations have been considered for the analysis: from the “naked” cabin (only primary structure) up to “fully furnished” (primary structure with interiors and noise control treatments) one. These results are essential to understand which are the main parameters affecting the noise insulation. Furthermore the Power Inputs evaluation has been determined to see the contribution of each considered aeronautic component on the acoustic insulation. Finally, the effect of a viscoelastic damping layer embedded in the glass window has been evaluated.
  • Rotorcraft blade-vortex interaction noise prediction using the
           Lattice-Boltzmann method
    • Abstract: Publication date: Available online 18 March 2019Source: Aerospace Science and TechnologyAuthor(s): Gianluca Romani, Damiano Casalino The aim of this paper is to assess the accuracy, capabilities and computational performances of the Lattice-Boltzmann/Very Large Eddy Simulation Method to predict the unsteady aerodynamic loads, the rotor wake development and the noise radiation of helicopter rotors in strong Blade-Vortex Interaction conditions. The numerical flow solution is obtained by solving the explicit, transient and compressible Lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW®. The acoustic far-field is computed by using the Ffwocs-Williams & Hawkings integral solution applied to a permeable surface encompassing the whole helicopter geometry. The employed benchmark configuration is the 40% geometrically and aeroelastically scaled model of a BO-105 4-bladed main rotor tested in the open-jet anechoic test section of the German-Dutch wind tunnel in the framework of the HART-II project. In the present study, only the baseline operating condition of the experimental campaign, without Higher-Harmonic Control enabled, is considered. All simulations are performed by assuming a rigid blade motion, but a computational strategy based on a combination of a rigid blade pitching motion and a transpiration velocity boundary condition applied on the blade surface is employed to take into account the blade elastic deformation motion measured during the experiments. As expected, modeling the blade elastic deformation leads to more accurate predictions of control settings, unsteady air-loads and noise footprint. The effects of the computational grid on the aerodynamic and aeroacoustic prediction is documented as well.
  • Structural design and verification of an innovative whole adaptive
           variable camber wing
    • Abstract: Publication date: Available online 18 March 2019Source: Aerospace Science and TechnologyAuthor(s): Anmin Zhao, Hui Zou, Haichuan Jin, Dongsheng Wen A whole adaptive variable camber wing (AVCW) equipped with an innovative double rib sheet (DRS) structure is experimentally and numerically studied in this work. The DRS structure adopts the surface contact mode for the force transmission of changeable camber wing instead of the conventional rigid hinge joint contact. The whole AVCW design allows to adjust the shape of wing in a real-time at various flight conditions, which is of great interest for Unmanned Aerial Vehicle (UAV) applications. The flight-test experiments demonstrate that the total AVCW carrying the developed adaptive control system (ACS) can enhance UAV flight efficiency by 14.1% comparing to a traditional fixed-wing of Talon UAV. In addition, it indicates that employing the whole AVCW structure can sustain a larger flight load, without increasing the weight of entire wing structure except for the actuator device and adhesives, which is promising for future engineering applications.
  • Disturbance observer-based gain adaptation high-order sliding mode control
           of hypersonic vehicles
    • Abstract: Publication date: Available online 15 March 2019Source: Aerospace Science and TechnologyAuthor(s): Xiaomeng Yin, Bo Wang, Lei Liu, Yongji Wang In this study, hypersonic vehicle (HV) tracking control in the presence of uncertainties and external disturbance is investigated. As the exact bounds of uncertainties and disturbances are usually unknown during flight, a disturbance observer (DOB)-based gain adaptation high-order sliding mode control (HOSMC) method is proposed for HVs. To mitigate the chattering effect while maintaining strong robustness, a method combining the advantages of the DOB and gain adaptation is introduced into the HOSMC. The DOB is employed to estimate and reject the uncertainties and external disturbance. Additionally, an adaptive control law is developed to compensate for estimation errors. By combining the DOB and the adaptive HOSMC, the unnecessarily large gain for maintaining robustness is reduced; thus, the chattering is mitigated. The effectiveness of the proposed control method is validated via simulation, in which a strong robustness, high tracking performance, and reduced chattering effect are achieved under uncertainties and external disturbance.
  • Continuum breakdown and surface catalysis effects in NASA arc jet testing
           at SCIROCCO
    • Abstract: Publication date: Available online 15 March 2019Source: Aerospace Science and TechnologyAuthor(s): Davide Cinquegrana, Raffaele Votta, Carlo Purpura, Eduardo Trifoni A facility characterization test campaign was performed for NASA in SCIROCCO Plasma Wind Tunnel, with increasing enthalpy and constant reservoir pressure. In the frame of numerical rebuilding of the experimental data, an important gap among measured and numerically predicted values of heat flux led to deep analyses the numerical methods and the models usually employed for test rebuilding. The Navier-Stokes model with chemical reacting non-equilibrium flows, denoted a lack of physical accuracy due to local rarefaction effects, as certified by means of the continuum breakdown parameter. Furthermore, the low stagnation pressure environment could influence the surface catalytic behaviour of the hemisphere copper probe, and the chemical contribution to the stagnation heat flux. At the end, a set of direct simulation Monte Carlo with partial catalytic behaviour of the probe was performed, in order to address both critical phenomena highlighted and close the gap with the measured heat fluxes, understanding the actual test chamber environment.
  • Evaluation of the damping capacity according to the geometric and the
           number of resonator with thermal environment using a Rijke tube
    • Abstract: Publication date: May 2019Source: Aerospace Science and Technology, Volume 88Author(s): Seonghwi Jo, Yunho Choi, Hong Jip Kim Combustion instabilities are often-observed phenomena which take place in acoustically closed spaces such as combustion chambers in propulsion systems. These phenomena should be studied consistently because they can make considerable damage on the combustion chambers and even entire engine systems. Acoustic cavity has been widely used as a passive stabilization device to suppress these combustion instabilities. To elucidate damping capacity of acoustic cavity according to various geometric shapes, a well-known Rijke tube facility has been used to simply include the interaction of heat and acoustic field. The damping capacity has been evaluated quantitatively in terms of bandwidth and amplitude ratio. Present results showed that the acoustic cavities having large orifice area showed better damping capacity. But, in case of orifice length, the length above a certain value makes cavity volume and resultant acoustic stiffness too small, so the damping capacity was decreased above a threshold-like specific length. As for the number of cavities, the damping capacity increases with the number, but, in a certain number or more, the damping capacity has been attenuated and the significant increase of the capacity was not observed. In case of the acoustic cavities which were thought to have sufficient damping capacity, decay time has been measured to quantify the damping capability in time domain. The effect of orifice area on decay time was much higher than those of orifice length. These results from heat and acoustic field interaction through Rijke tube showed that the pure acoustic approach would be insufficient for the fine tuning of acoustic cavity, and further combustion tests would also be necessary to optimize the shapes of the cavity.
  • Receding horizon guidance and control using sequential convex programming
           for spacecraft 6-DOF close proximity
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Ding Zhou, Yanquan Zhang, Shunli Li This paper presents a receding horizon implementation of sequential convex programming for the spacecraft six-degree-of-freedom close proximity to a non-cooperative target satellite. With linearized relative translational dynamics and newly derived discrete rotational equations in terms of modified Rodrigues parameters, nonlinear system dynamics of the original optimal guidance and control problem for proximity are converted into convex ones for sequentially planning. The nonconvex constraint on field-of-view of the visual sensor pointing with coupled attitude and relative position is then approximately relaxed as a convex standard second-order cone, and concave spherical and ellipsoidal obstacle regions respectively around the target's body and solar arrays are convexified by affine constraints in terms of tangent planes. The original nonlinear optimal guidance and control problem is accordingly transformed into a series of second-order cone programming sub-problems via iteratively successive convexification with the trust region constraints, and sequentially solved using disciplined convex programming method. A close-loop guidance and control using the proposed sequential convex programming scheme is then demonstrated by means of receding horizon to robustly drive the spacecraft maneuvering close to the target. Numerical simulations and results reveal that the proposed method provides rapid and reliable guidance and control performance for six-degree-of-freedom close proximity and shows the potential for on-board implementations in real-time applications.
  • Trajectory-following guidance based on a virtual target and an angle
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Qi Chen, Xugang Wang, Jing Yang, Zhongyuan Wang In this paper, a new trajectory-following guidance scheme based on the combination of the virtual target concept and a missile terminal guidance law is proposed. A line-of-sight (LOS) angle constraint is derived for the unmanned aerial vehicle (UAV) to track a virtual target moving along the reference trajectory, by which the trajectory-following problem is transformed into an angle-constrained terminal guidance problem such that well-developed terminal guidance laws can be incorporated into the design of a trajectory-following guidance scheme. The nonsingular sliding mode technique is employed to control the vehicle's LOS angle to converge to its desired value and thus to tightly track the reference trajectory. In addition, the finite-time trajectory-following position error convergence of the proposed guidance scheme is presented. The significant contribution of this paper lies in the fact that a LOS angle constraint has been presented for a virtual-target-based trajectory-following scheme, greatly enhancing the flexibility of the design of the trajectory-following guidance scheme. Numerical simulations are performed to demonstrate the performance of the proposed guidance law and its superiority over existing trajectory-following guidance laws.
  • Effective optimization on Bump inlet using meta-model multi-objective
           particle swarm assisted by expected hyper-volume improvement
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Shixin Cheng, Hao Zhan, Zhaoxin Shu, Huayu Fan, Ban Wang This paper presents an efficient multi-objective optimization method, focusing on aerodynamic optimization of a diverterless supersonic inlet (DSI) in transonic and supersonic flight conditions. The DSI inlet, through scrutinizing the Bump shape has potential to attain greater aerodynamic performance on exit plane of inlet. However, the high cost of computational fluid dynamic (CFD) simulations raises a significant challenge in the DSI optimization process. In order to obtain solution set in few numbers of objective function calls, a meta-model multi-objective particle swarm optimization (MOPSO) method is proposed based on a self-adaptive Kriging surrogate model, and applied to solve this kind of costly black-box optimization problem. The Kriging model is updated by using a dynamic expected hyper-volume improvement (EHVI) sample metric, which is developed by analyzing disadvantages of the original sample criterion. With the help of the dynamic sample metric, simulation results show that the surrogate-based MOPSO algorithm can obtain plenty enough non-dominated solutions and achieve high precision in the approximation of the Pareto front. In terms of DSI inlet optimization, the bump shape is parameterized by free form deformation (FFD) method, and the total pressure distortions of inlet exit plane are treated as two minimization objectives under transonic and supersonic flight conditions. A well distributed non-dominated solution set is generated by the proposed algorithm within the context of a small call number of cost evaluations, and optimized inlet configurated by the selected solution has better aerodynamic characteristics compared with the initial inlet.
  • Development of high fidelity reduced order hybrid stick model for aircraft
           dynamic aeroelasticity analysis
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Paul Vazhayil Thomas, Mostafa S.A. ElSayed, Denis Walch This paper presents a new high fidelity reduced order modeling methodology based on a hybrid stick model representation approach. Here, the traditional stick model developed by the unitary loading method is augmented by residual mass and stiffness matrices that account for the dynamic imparity between the stick model and the global finite element model, within a frequency range of interest, as well as the degrees of freedom coupling commonly ignored by the simplified stick model. The new method offers the handling flexibilities of the conventional stick model as well as the high dynamic accuracy of matrix based model order reduction methods such as the Guyan and the Craig-Bampton condensation techniques. Retaining the stick model in the proposed hybrid model representation intuitively enables aerospace development engineers to, accurately and efficiently, optimize the airframe mass and stiffness distribution for aircraft loads minimization and performance maximization without the need to engage an expensive global finite element model in such highly iterative analyses. Two hybrid stick models are presented in this paper that are developed based on the Guyan and the Craig-Bampton reduction methods. A case study is presented where the hybrid stick models developed along with their conventional stick model counterpart are employed in the dynamic aeroelasticity loads analyses of a Bombardier aircraft platform. Using monitor points method, the extracted aeroelastic loads using the reduced order models are compared against those generated employing the aircraft global finite element model. The dynamic characteristics of the reduced order models are also assessed based on their modal characteristics using modal assurance criteria along with their loads modal participation factors. Results obtained show that the developed hybrid stick models have superior dynamic characteristics compared to the conventional stick model.
  • / γ R e θ t +coupling+model+for+the+simulation+of+separated+transitional+flow&rft.title=Aerospace+Science+and+Technology&rft.issn=1270-9638&">Flow-dependent DDES / γ − R e ‾ θ t coupling model for the
           simulation of separated transitional flow
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Lin Zhou, Zhenghong Gao, Yiming Du A new DDES and γ−Re‾θt coupling model for separated transitional flow is proposed. The studies in this paper indicate that the activation of the LES branch in the free stream may lead to an inappropriate decrease in the inflow turbulence intensity and a delayed transition. To address this problem, a new flow-dependent RANS/LES switch function based on identification functions for the boundary layer and wake regions is designed. In this new model, the adoption of the RANS or LES branch is decided by both the local grid density and local flow properties, and the LES branch is only allowed to be active in the wake region. Therefore, the inflow turbulence intensity uncertainties caused by grid dependence of the LES branch in the free stream are largely eliminated. Separated transitional cases including flows over an A-Airfoil (small separation), a DBLN-526 airfoil (moderate separation), and a circular cylinder in subcritical and critical regimes (massive separation) are studied. The performance of two commonly employed boundary layer shielding functions is examined. Good agreements are achieved between the numerical and experimental results, and the accuracy and reliability of the new model are demonstrated.
  • Optimal autonomous multirotor motion planning in an obstructed environment
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Emanuele L. de Angelis, Fabrizio Giulietti, Goele Pipeleers, Gianluca Rossetti, Ruben Van Parys This paper addresses a methodology for autonomous motion planning of multirotor aircraft in obstructed environments. The control strategy allows the vehicle to online generate quasi-optimal trajectories with limited computational load while performing collision avoidance tasks. The problem is formulated in a model-predictive control architecture in which motion planning and trajectory tracking processes are solved separately. The first process is based on a spline path planning approach to generate smooth and safe trajectories. The second process elaborates trajectory inputs in terms of commanded thrust magnitude and desired attitude rates in order to steer the vehicle during the mission task. Results of both numerical simulations and, for the first time, an experimental validation are provided in order to assess the performance of the approach in the presence of external disturbances and unmodeled dynamics, provided adequate time horizon and update frequency are selected for the numerical optimization algorithm.
  • The optimization design with minimum power for variable speed control
           moment gyroscopes with integrated power and attitude control
    • Abstract: Publication date: Available online 15 March 2019Source: Aerospace Science and TechnologyAuthor(s): Feng Liu The task of Integrated Power and Attitude Control (IPAC) of a spacecraft can be implemented by Variable Speed Control Moment Gyros (VSCMGs). The Integrated Power and Attitude Control System (IPACS) singularity problem is the key factor for the spacecraft to successfully perform IPAC task, which can be overcome by rational designing steering law. The singularity characteristic and the steering law should be considered during the process of design parameters for the VSCMG cluster with IPACS task. There is no research report in this area at present. The steering results of weighted pseudo-inverse and null motion of Weighted Pseudo-Inverse with Null Motion (WPINM) can be canceled by each other under some certain condition. So the flywheel torque requirement of the WPINM steering law can be greatly increased, which is contradictory to the original design intention of the weighted matrix. A steering law with minimum requirement of flywheel power and torque is introduced from existing research result. Then, the constraint of the IPACS singularity problem and the SGCMGs singularity problem cannot be encountered during the whole process of IPAC task is given. At last, the parameters design problem of VSCMGs for the IPAC task is cast as a multi-objective optimization problem with minimum whole system power and maximum utilization ratio of flywheel momentum under the condition of consideration of the steering. The intelligent algorithm of Non-dominated Sorting Genetic Algorithm (NSGA) is used to solve the nonlinear multi-objective problem. The flywheel power can be greatly reduced by the new parameters design method.
  • Electric sail trajectory design with Bezier curve-based shaping approach
    • Abstract: Publication date: Available online 14 March 2019Source: Aerospace Science and TechnologyAuthor(s): Mingying Huo, Giovanni Mengali, Alessandro A. Quarta, Naiming Qi The aim of this paper is to propose a shape-based method in which the concept of Bezier curve is used to efficiently design the three-dimensional interplanetary trajectory of a spacecraft whose primary propulsion system is an Electric Solar Wind Sail. The latter is a propellantless propulsion concept that consists of a spinning grid of tethers, kept at a high positive potential by a power source and maintained stretched by the centrifugal force. The proposed approach approximates the time variation of the components of the spacecraft position vector using a Bezier curve function, whose geometric coefficients are calculated by optimizing the total flight time with standard numerical methods and enforcing the boundary conditions of a typical interplanetary rendezvous mission. The paper also discusses a geometrical approach to include, in the optimization process, the propulsive acceleration vector constraints obtained with the latest Electric Solar Wind Sail thrust model.
  • Control of cylinder wake flow and noise through a downstream porous
    • Abstract: Publication date: Available online 14 March 2019Source: Aerospace Science and TechnologyAuthor(s): Chen Xu, Yijun Mao, Zhiwei Hu Various passive and active methods have been developed to control flow separation from bluff bodies. However, these methods require adjusting features of the solid surface, such as modifying its geometry or porosity, or applying external force or momentum. This paper develops an off-body-based method, without adjusting any features of the solid surface, for controlling the unsteady flow separation by fixing porous materials downstream of bluff bodies. Numerical study on flow past a circular cylinder at a subcritical Reynolds number is performed, and the result indicates that the added downstream porous material changes flow in the wake and re-laminarizes the turbulent flow around the curved cylinder surface, reducing the wall pressure fluctuation around the cylinder. therefore, the associated aerodynamic noise is reduced greatly.
  • The impact of the length-to-depth ratio on aerodynamic surface quantities
           of a rarefied hypersonic cavity flow
    • Abstract: Publication date: Available online 14 March 2019Source: Aerospace Science and TechnologyAuthor(s): Rodrigo C. Palharini, Wilson F.N. Santos A computational investigation has been carried out to examine a non-reacting rarefied hypersonic flow over cavities by employing the Direct Simulation Monte Carlo (DSMC) method. The work focuses on the effects on the aerodynamic surface quantities due to variations in the cavity length-to-depth (L/H) ratio. The results highlight the sensitivity of the heat transfer, pressure and skin friction coefficients due to changes to the cavity L/H ratio. The L/H ratio ranged from 1 to 4, which corresponds to the transition flow regime based on an overall Knudsen number KnL. The analysis showed that the aerodynamic quantities acting on the cavity surface rely on the L/H ratio. It was found that pressure load and heating load to the cavity surfaces presented peak values along the forward face, more precisely in the vicinity of the cavity shoulder. Moreover, these loads are much higher than those found in a smooth surface, for the conditions investigated.
  • Rapid prototyping for hypervelocity impulse facility test models
    • Abstract: Publication date: Available online 14 March 2019Source: Aerospace Science and TechnologyAuthor(s): Steven F.T. Apirana, Christopher M. James, Robert Eldridge, Richard G. Morgan, Steven W. Lewis The X2 expansion tube facility at The University of Queensland is capable of simulating entry into most of the planetary bodies in our solar system, producing test conditions with stagnation enthalpies in excess of 100 MJ/kg. Models used in X2 are typically made from steel and, consequently, manufacture is often constrained to conventional methods, with associated long lead times, and high cost. Through an experimental campaign, the survivability and applicability of Rapid Prototype models was investigated. Three prototyping methods were investigated; Selective Laser Sintering, Stereolithography, and Fused Deposition Modelling, and these were compared to a steel baseline. A computational stress analysis was used to design an internally hollow test model geometry. All models survived the experimental test-time, one was destroyed by the post-experiment flow. It was thought that test-models might ablate during the experimental test time and this was investigated by using filtered imaging to capture Cyanogen radiation occurring in the model's boundary layer. Ablation was seen in all Rapid Prototype models, most strongly observed in the Selective Laser Sintered and Acrilonitrile Butadiene Styrene models. These models may be suitable for the study of non-equilibrium radiative emission just behind the shock wave, or ablation phenomena in a model's boundary layer.
  • Geometrically nonlinear static aeroelastic analysis of composite morphing
           wing with corrugated structures
    • Abstract: Publication date: Available online 14 March 2019Source: Aerospace Science and TechnologyAuthor(s): Natsuki Tsushima, Tomohiro Yokozeki, Weihua Su, Hitoshi Arizono In this paper, an integrated geometrically nonlinear aeroelastic framework to analyze the static nonlinear aeroelastic response of morphing composite wing with orthotropic materials has been developed. A flat plate/shell finite element, which can model plate-like wings, has been accommodated to model composite/corrugated panels to investigate effects of different laminate orientations and corrugations. A corotational approach is used to consider the geometrical nonlinearity due to large deformation produced by wing morphing. An unsteady vortex-lattice method is implemented to couple with the structural model subject to the large deformations. A homogenization method is also implemented to model corrugated panels as equivalent orthotropic plates. Individual structural, aerodynamic, and corrugated panel models, as well as the complete nonlinear aeroelastic framework, are verified. Numerical studies explore the static aeroelastic responses of a flat wing with composite/corrugated panels. This work helps to understand the nonlinear aeroelastic characteristics of composite/corrugated wings and demonstrates the capability of the framework to analyze the nonlinear aeroelasticity of such morphing wings.
  • Adaptive fault tolerant control for a small coaxial rotor unmanned aerial
           vehicles with partial loss of actuator effectiveness
    • Abstract: Publication date: Available online 13 March 2019Source: Aerospace Science and TechnologyAuthor(s): Zhankui Song, Kaibiao Sun This paper investigates the trajectory tracking problem for a small coaxial-rotor unmanned aerial vehicle (CRUAV) with partial loss of actuator effectiveness. The CRUAV model is decomposed into a dual loop structure based on back-stepping design idea. First, certain performance function specified a priori by the designer is introduced into the position loop such that the original position tracking error is transformed into an equivalent constrained variable providing for the performance judgment of the position loop. Then, a fault-tolerant control scheme is proposed based on adaptive strategies for compensating the effect caused by various adverse factors. Subsequently, an attitude control loop is derived by incorporating adaptive compensation method. It is proved that the proposed dual-loop structure is able to guarantee the satisfaction of the pre-specified constraint on the transformed errors. Finally, the effectiveness and benefits of the design dual-loop control system are validated via computer simulation.
  • Performance modeling of pulse detonation engines using the method of
    • Abstract: Publication date: Available online 13 March 2019Source: Aerospace Science and TechnologyAuthor(s): James T. Peace, Frank K. Lu A quasi-one-dimensional method of characteristics (MOC) model is developed to evaluate the single-cycle gasdynamic flow field and associated propulsive performance of a fully- and partially-filled pulse detonation engine (PDE), and PDEs equipped with diverging nozzles. A detailed description of the PDE thrust chamber flow field is provided to highlight the dominant gasdynamic wave processes encountered in a general single-cycle operation. The MOC model is developed using a simplified unit process approach with an explicit inverse time marching algorithm in order to readily construct the complex thrust chamber flow field along a predefined grid. A grid dependency study is carried out to determine the appropriate grid resolution for the purposes of minimizing computational expense and maximizing numerical accuracy. Lastly, the MOC model is validated with existing numerical and experimental performance data for PDEs operating with oxyhydrogen and hydrocarbon mixtures in both atmospheric and sub-atmospheric environments. It is shown that the current MOC model has good agreement with existing numerical and experimental data for a variety of PDE configurations and operating conditions.
  • Application of flow control strategy of blowing, synthetic and plasma jet
           actuators in vertical axis wind turbines
    • Abstract: Publication date: Available online 13 March 2019Source: Aerospace Science and TechnologyAuthor(s): Haitian Zhu, Wenxing Hao, Chun Li, Qinwei Ding, Baihui Wu The vertical axis wind turbines (VAWTs), as common wind turbines for harvesting wind energy, have wide prospects of development. However, the VAWTs are periodically influenced by dynamic stall which can cause the aerodynamic losses and load fluctuation. Therefore, the VAWTs urgently require flow control technique to improve the aerodynamic characteristics. The jet actuators, as active flow control (AFC) techniques, are reasonable implementations for VAWTs. The current paper presents the review of jet flow control techniques which have been used or are worth being used in VAWTs, including the blowing, synthetic and plasma jet actuators. However, the jet flow control strategies to reduce the energy or matter consumption of jet actuators for VAWTs should be developed. Based on the validation of computational model, the VAWTs with upward-parabola blowing jet flow control strategy which can suppress the flow separation in advance was numerically investigated at different tip-speed ratios. The results show that the upward-parabola blowing control strategy can dramatically enhance the aerodynamic performance by using significantly low energy or matter consumption. Obviously, this novel control strategy customized for VAWT can be applied in other jet actuators, improving the aerodynamic performance of VAWT.
  • A theoretical and computational study of the high-temperature effects on
           the transition criteria of shock wave reflections
    • Abstract: Publication date: Available online 12 March 2019Source: Aerospace Science and TechnologyAuthor(s): Peng Jun, Zhang Zijian, Hu Zongmin, Jiang Zonglin In this paper, we study the high-temperature effects on the reflection of shock waves in hypersonic flows by using analytical and computational approaches. First, a theoretical approach is established to solve the shock relations which are further applied to develop the shock polar analytical method for high-temperature air. Then, a comparative investigation using ideal gas model and real gas model considering vibration excitation indicates that the high-temperature effects cause an obvious change to the overall profile of the shock polar. The post-shock pressure increases within the strong branch of the shock polar while decreases within the weak branch due to vibration excitation of air molecules. A more notable phenomenon is the increase in the maximum deflection angle of the shock polar which can significantly influence the detachment criterion of shock reflection transition in high-temperature air flows. The shock polar analysis of shock reflection shows that the high-temperature effects result in an obvious increase to the detachment criterion while a slight increase to the von Neumann criterion. A series of computations are conducted to confirm the above analytical findings on the shock reflection considering high-temperature effects. A slight difference of transition criterion between the theory and computations is found to be caused by the existence of the expansion fan which is an inherent flow structure. The proposed shock polar analytical method is proved to be an effective but simple approach for the study of shock wave reflections in hypersonic flows.
  • Lateral cutoff analysis of sonic boom using full-field simulation
    • Abstract: Publication date: Available online 12 March 2019Source: Aerospace Science and TechnologyAuthor(s): Rei Yamashita, Kojiro Suzuki This paper describes the world's first successful simulation for lateral cutoff phenomena of sonic boom far from the flight path due to variation in atmospheric temperature with altitude. A flow field around an axi-symmetric paraboloid has been analyzed by the full-field simulation method that solves the three-dimensional Euler equations with a gravity term to create a horizontally stratified atmosphere. A solution-adapted structured grid is constructed to align the grid lines with the front and rear shock-wave surfaces in the entire domain, including the near field around a supersonic body and far field reaching the ground beyond lateral cutoff. The flight is assumed to have a speed of Mach 1.2 at an altitude of 10 km, and the computational domain ranges over a distance of 30 km from the axis of symmetry. The computational results show that the evanescent wave in the shadow zone beyond lateral cutoff decays exponentially and changes into a progressive rounding waveform. The characteristics of the waveform transition are in good agreement with those observed in the flight tests. Therefore, the full-field simulation is recognized as a promising approach for investigating sonic boom strength in the full extent of sonic boom noise, including lateral cutoff and evanescent waves. Moreover, the computational results clarify that sonic boom focusing occurs above the ground, except for the vicinity of the ground, and the focusing strength along the lateral cutoff curve detected from the three-dimensional shock-wave surface increases with altitude. The results of ray tracing analysis collaborate the reasonability of the simulation results, and the caustic of downward convex agrees well with the lateral cutoff curve. In the shadow zone, the magnitude of exponential decay increases with altitude, and the lateral distance where the pressure rise decreases rapidly shortens with altitude.
  • Design and aeromechanics investigation of compound helicopters
    • Abstract: Publication date: Available online 12 March 2019Source: Aerospace Science and TechnologyAuthor(s): Hyeonsoo Yeo This paper reviews lessons learned from the compound helicopter studies performed by NASA and the US Army Aviation Development Directorate to support the NASA Heavy Lift Rotorcraft Systems Investigation and the US Army's Joint Heavy Lift (JHL), Joint Multi-Role Technology Demonstrator (JMR-TD), and Future Vertical Lift (FVL) programs. These studies explored performance potential of advanced rotorcraft and investigated the impact of key modern-technologies in performance, weight, and aerodynamics on rotorcraft. The compound helicopter configurations considered in this paper represent a wide range of sizes, gross weight, rotor systems, and operating conditions. A brief description of design and aeromechanics analysis tools and methodologies is provided. Rotor performance correlation results at high advance ratio, which are critical for the accurate design and analysis of high-speed rotorcraft, are shown. Detailed aeromechanics analysis results, such as the effects of various compounding methods, lift share between rotor and wing, rotor rotational speed, blade twist, aircraft drag, and rotor/wing interference on aircraft performance, are presented.
  • A droplet/wall impact model and simulation of a bipropellant rocket engine
    • Abstract: Publication date: Available online 12 March 2019Source: Aerospace Science and TechnologyAuthor(s): Pengfei Fu, Lingyun Hou, Zhuyin Ren, Zhen Zhang, Xiaofang Mao, Yusong Yu Droplet impact on a solid wall contributes to droplet vaporization and cooling on the inner wall. A new droplet/wall impact model is developed and fitted from the experimental data, which takes both kinematic parameters of the impinging droplets and thermal parameters of the solid wall into consideration. The model describes the behavior of droplets after impact and identifies five representative regimes, i.e., stick/spread, suspend, rebound, boiling including breakup, and splash, through the critical Weber number of droplets and wall temperature. To assess the new model, numerical simulations were performed of propellant droplets impacting the wall of a bipropellant rocket engine chamber. A comparison with experimental data shows better predictions of the wall temperature of the chamber than obtained from previous models. A larger number of large-sized monomethylhydrazine and nitrogen tetroxide droplets in the boiling induced breakup and splash regimes break into smaller droplets when using the new model. There are also temperature peaks near the impact points and near the throat. Especially in the throat of the combustion chamber, the new model predicts the wall temperature distribution accurately, offering improved prediction of the combustion chamber and assessing thermal protection of the throat.
  • Experimental investigation on the energy absorption characteristics of
           honeycomb sandwich panels under quasi-static punch loading
    • Abstract: Publication date: Available online 28 February 2019Source: Aerospace Science and TechnologyAuthor(s): M. Zarei Mahmoudabadi, M. Sadighi The energy absorption characteristics of sandwich panels with aluminum plate as facesheet and metal hexagonal honeycomb as the core are investigated under quasi-static punch loading using two flat nose and spherical projectiles, experimentally. Failure modes are classified as plastic hinges, facesheet wrinkling, debonding of the adhesive layer between the facesheet and core, facesheet tearing, out of plane core crushing, in-plane core folding, core tearing and detachment from the support. Furthermore, the article examines the influences of six parameters including honeycomb wall thickness, sandwich core thickness, facesheet thickness, aspect ratio, adhesive layer between facesheet and core and existence of bottom facesheet. The results show that the increase in core thickness improves the energy absorption parameters of sandwich panel better than the increase in the facesheet thickness. Specific absorbed energy is increased linearly by increasing the honeycomb core thickness while it seems that the mentioned parameter has a meaningless dependence on the facesheet thickness. In addition, a 12 percent quota of adhesive layer between top facesheet and core is indicated in the energy absorption capacity of a sandwich panel for both flat nose and spherical projectiles, while its effect on the value of maximum force is 17% using flat nose projectile and 25% using the other one. Despite the major influence of the existence or non-existence of the bottom facesheet on the sandwich failure modes, its absorbed energy changes less than 3.5 percent; yet other parameters such as specific energy absorption and peak load are more dependent on the existence or non-existence of the bottom facesheet. Finally, keeping the honeycomb wall less thick improves the energy absorption characteristics of the sandwich panel.
  • Viscous flow and performance issues in a 6:1 supersonic mixed-flow
           compressor with a tandem diffuser
    • Abstract: Publication date: Available online 25 February 2019Source: Aerospace Science and TechnologyAuthor(s): Aravinth Sadagopan, Cengiz Camci The advancement of multi-dimensional and viscous computational tools has eased the accessibility and overall effort for thorough analysis of complex turbomachinery designs. In this paper, we computationally evaluate a high-pressure ratio supersonic mixed-flow compressor stage designed using an in-house mean-line code. Objective is to include three dimensionalities, viscous flow and compressibility effects including the shock wave systems into account. As mixed-flow compressors are advantageous especially for small jet engine applications we choose mass flow rate, stage total pressure ratio and maximum diameter as the main design constraints. This computational analysis is the second paper of a two-part series explaining strategy for designing a high-pressure ratio mixed-flow compressor stage. The high-pressure ratio and small diameter requirements push this compressor for a highly-loaded supersonic ‘shock-in rotor’ design with supersonic stator/diffuser.The used RANS based computational fluid dynamics model is thoroughly assessed for its ability to predict compressor performance using existing well-established experimental data. NASA Rotor 37 and RWTH Aachen supersonic tandem stator are chosen as the test cases for exhibiting similar flow characteristics to present design. The computational approach helps to shed light upon the mixed-rotor and supersonic-stator 3D shock structures and viscous/secondary flow. Stage performance map, pressure and velocity distribution of this high-pressure ratio mixed-flow compressor is obtained. Areas of design optimization are highlighted to further improve performance and efficiency. The in-house mean-line design code predicted a pressure ratio of 6.0 with 75.5% efficiency for a mass flow rate of 3.5 kg/s. The mean-line code obviously lacked to fully represent three-dimensionality effects due to its inherent over-simplifying assumptions thus, inclusion of RANS based computations improves the fidelity of mixed-flow compressor design performance calculations at a great rate. Comprehensive computational analysis of the stage shows that our design goal is met with a stage total pressure ratio of ΠTT = 5.83 with an efficiency of ηIS = 7% for a mass flow rate of m˙ = 3.03 kg/s. A total pressure ratio of 6.12 at 75.5% efficiency is reached with a 3.5% increase in design rotational speed.
  • Multi-objective optimisation of short nacelles for high bypass ratio
    • Abstract: Publication date: Available online 19 February 2019Source: Aerospace Science and TechnologyAuthor(s): Fernando Tejero, Matthew Robinson, David G. MacManus, Christopher Sheaf Future turbo-fan engines are expected to operate at low specific thrust with high bypass ratios to improve propulsive efficiency. Typically, this can result in an increase in fan diameter and nacelle size with the associated drag and weight penalties. Therefore, relative to current designs, there is a need to develop more compact, shorter nacelles to reduce drag and weight. These designs are inherently more challenging and a system is required to explore and define the viable design space. Due to the range of operating conditions, nacelle aerodynamic design poses a significant challenge. This work presents a multi-objective optimisation approach using an evolutionary genetic algorithm for the design of new aero-engine nacelles. The novel framework includes a set of geometry definitions using Class Shape Transformations, automated aerodynamic simulation and analysis, a genetic algorithm, evaluations at various nacelle operating conditions and the inclusion of additional aerodynamic constraints. This framework has been applied to investigate the design space of nacelles for high bypass ratio aero-engines. The multi-objective optimisation was successfully demonstrated for the new nacelle design challenge and the overall system was shown to enable the identification of the viable nacelle design space.
  • Experimental investigation on delay time phenomenon in rotating detonation
    • Abstract: Publication date: Available online 28 January 2019Source: Aerospace Science and TechnologyAuthor(s): John Z. Ma, Shujie Zhang, Mingyi Luan, Jianping Wang This study explored the phenomenon of the detonation formation delay time in a RDE with an annular combustion chamber with an array of holes fueled with hydrogen. It was determined that the delay time increased when the combustor prefilling time increased. Due to the explosion, the interaction of the shock wave and the pressure wave in the combustor leads to an increase in pressure near the inlet wall surface, thereby impeding the gas feeding. The wavelet transformation was used to analyze the pressure time signal associated with the detonation wave, which was found to be better than the fast Fourier transform typically used. Moreover, the phenomenon of single–double–single wave transformation was observed, of which the underlying mechanism can be explained by an interactive-adjusting process. Also, the experiment captured the reinitiation phenomenon of detonations after a long period of time (around 300 ms) of extinction.
  • CFD based criteria of stall onset in centrifugal compressors
    • Abstract: Publication date: Available online 4 January 2019Source: Aerospace Science and TechnologyAuthor(s): Adel Ghenaiet, Smail Khalfallah The determination of stall onset is very useful for the control and safe operation of centrifugal compressors at high pressure ratios. This is very delicate and intricate task since the details of the flow need to be acquired at several measuring planes. The state-of-the-art CFD tools have contributed substantially in the analysis of the flow structures and the design of centrifugal compressors. The present paper demonstrates the potentiality of numerical simulations to map the full flow field and assess some criteria to predict the stall onset, typically those relating the boundary layer growth and stability. Through the entire impeller and diffuser, the quantification of the blockage factor in addition to the modified loading criteria (diffusion rate) which combines the blade-to-blade and hub-to-shroud Richardson numbers seem more suitable in predicting the stall onset and stall cells positions. The well known models of centrifugal compressors operating at low-speed (LSCC) and at a high-speed (DLRCC) served for the validations. Besides, the present CFD based criteria are useful since they may easily integrate the design optimization chain.
  • Numerical simulation on thermal and mass diffusion of MMH–NTO
           bipropellant thruster plume flow using global kinetic reaction model
    • Abstract: Publication date: Available online 3 December 2018Source: Aerospace Science and TechnologyAuthor(s): Kyun Ho Lee A space propulsion system has a crucial role to perform several mission operations of a spacecraft successfully in an orbit. When a thruster is fired, the exhaust plume gas can have effects on the performance of a spacecraft because the expanded plume gas molecules directly collide with the spacecraft surfaces in the vacuum environment. Thus, the present study investigated more realistic plume flow behaviors using a global kinetic reaction model for an actual combustion process of a fuel and an oxidizer. To achieve this, the 4-step global combustion model of monomethylhydrazine and nitrogen tetroxide was incorporated for the first time in the plume flow analysis to reflect a more practical firing condition of a bipropellant thruster. Then, thermal and mass diffusion predictions of the plume flow were compared with the chemical equilibrium condition to examine the distinct differences between the proposed and conventional approach. For efficient numerical calculations, the Navier–Stokes equations and the DSMC method were combined to deal with a continuum flowfield inside the thruster and a rarefied plume gas flow outside the nozzle together. With the present analysis results, major differences in the thermal and mass behaviors of the plume gases were compared between the two reaction models, and their influences are also discussed.
  • Mechanism/structure/aerodynamic multidisciplinary optimization of flexible
           high-lift devices for transport aircraft
    • Abstract: Publication date: Available online 24 October 2018Source: Aerospace Science and TechnologyAuthor(s): Yun Tian, Jianchong Quan, Peiqing Liu, Doudou Li, Chuihuan Kong Mission adaptive variable camber wing in both chord-wise and span-wise directions that can improve the aerodynamic performance during takeoff, landing and cruising flight, will be the state-of-the-art high-lift system for next generation airliners. Based on NASA TrapWing model released on the 1st AIAA CFD High-lift Prediction Workshop, a smart high-lift system with “Flexible Droop Nose & Single Slotted Hinge Flap combined with spoiler deflection & Flexible Trailing Edge Flap” is proposed in this paper. The Flexible Droop Nose is actuated by kinematic chains mechanism, the Single Slotted Hinge Flap is actuated by simple hinge mechanism and the Flexible Trailing Edge Flap is actuated by link/track mechanism.A mechanism/structure/aerodynamic multidisciplinary optimization platform based on iSIGHT software is constructed for this smart high-lift system. This platform consists of stress analysis, high-lift configuration generation, high-lift configuration structure grid generation, computational fluid dynamics and optimization algorithm modules. The optimal takeoff and landing configurations with comprehensive performance of mechanism, aerodynamic and structure is then obtained after multidisciplinary optimization. Finally, the CFD results show that the aerodynamics performance of this smart high-lift system is more effective than the original NASA TrapWing model.
  • A sample-based approach to estimate the dynamic loads of components with
           nonlinear uncertain interfaces
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): E. Menga, M.J. Sánchez, I. Romero, S. Hernández Predicting aircraft dynamics and vibration loads at components' interfaces is a key task for ensuring a robust design and development of the product. Usually, whereas in the case of isolate components the dynamic behavior can be predicted quite accurately, when several components are assembled through discontinuous junctions the predictiveness of a model decreases. The junctions, whose mechanical properties are seldom well characterized experimentally, often introduce nonlinearities in the loads' path. Additionally, their behavior is intrinsically uncertain and as a consequence, the dynamic response of the connected structures becomes stochastic.We propose a sample-based approach which aims to cope with both aspects, nonlinearities and uncertainties, and can be split in two main tasks. First, the computational cost of each deterministic simulation is minimized considering that the global nonlinear behavior depends on localized sources of nonlinearities at the interfaces. Second, the uncertainties are propagated through the model by a non-intrusive method based on Sobol's low discrepancy design. Attention is paid to the global sensitivity indices, which are estimated by creating a meta-model based on Polynomial Chaos Expansion.An industrial application considering an aircraft component whose dynamic behavior is affected by uncertain free-plays at its interfaces is presented.
  • A comprehensive analysis on the structure of groove-induced shock waves in
           a linear turbine
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Ben Zhao, Mingxu Qi, Harold Sun, Xin Shi, Chaochen Ma A shock wave mostly occurs around a compression corner in supersonic flow and, when the corner is followed with another, two shock waves normally generate and, between them, an intersection is possible because of the increased shock angle of the downstream shock. A grooved-surface can generate multiple-shockwave structure in a local supersonic flow field as well. Different from the shockwave structure resulted from the compression corners, its structure may be approximately parallel or divergent when an additional expansion process occurs between two adjacent shock waves. This paper focuses exclusively on the shockwave structures based on both experimental and numerical methods. The physical mechanism for the generation of the approximately parallel and divergent shockwave structures is understood and then the approximately parallel shock wave structure is proved to belong to the same family with the divergent one. In addition, the relationship of the total oblique shock loss with the groove number is analyzed theoretically and numerically in entropy.
  • Nonlinear consensus strategies for multi-agent networks under switching
           topologies: Real-time receding horizon approach
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Fei Sun, Kamran Turkoglu In this paper, based on real-time nonlinear receding horizon control methodology, a novel scheme is developed for multi-agent nonlinear consensus problem under jointly connected switching topologies. The consensus problem is converted into a family of finite horizon optimization control process and is solved numerically to generate distributed control protocols in real-time. The stability is proved without the assumption that the topology is connected for all the time. Two benchmark examples on nonlinear chaotic systems provide validated results which demonstrate the significant outcomes of such methodology.
  • Numerical investigation on the forced oscillation of shock train in
           hypersonic inlet with translating cowl
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Wen Shi, Juntao Chang, Junlong Zhang, Jicheng Ma, Ziao Wang, Wen Bao To investigate the forced oscillations of shock train caused by sinusoidal backpressure perturbations with different amplitudes and frequencies in a hypersonic inlet equipped with translating cowl, numerical simulations have been conducted with the application of dynamic mesh method. The results reveal that under sinusoidal backpressure perturbations, the shock train oscillates and propagates upstream as the cowl moves downstream rather than crosses the shock-impact points abruptly with significant migration distance, compared to the result obtained under constant backpressure. Meanwhile, the amplitude of forced oscillation increases when the shock train leading edge oscillates around the adjacent shock-impact points. The larger amplitude of backpressure perturbation not only aggravates the forced oscillations but also increases the number of shock-impact points that the shock train crosses in one cycle, which leads to complicated changes in the shock train structures, even involving the separation mode transition. The frequency of forced oscillation equals to the one of backpressure perturbation invariably, but with certain phase lags due to the interference of background waves. Although the variable background waves do have the ability to affect the amplitude of forced oscillation, they are incapable of changing the frequency of forced oscillation.
  • Multi-blade shedding in turbines with different casing and blade tip
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Ibrahim Eryilmaz, Biläl Guenchi, Vassilios Pachidis A shaft failure in a gas turbine engine results in the decoupling of the turbine and the compressor. The turbine continues extracting work from the air flow causing the acceleration of the free-running turbine which can result in debris release or even disc burst. Post shaft failure the structural integrity of the engine must be guaranteed for product safety and certification purposes. To achieve this, speed limiting systems have to be integrated. One common method is friction between the rotor and stationary structures which occurs during unlocated failures where the bearing arrangement allows axial movement of the rotor under end load. Another possible mechanism is the destruction of the turbine rotor blades such that they cease to extract power from the incoming flow, decelerating progressively. Blade shedding involves rupture of blades which may result in their containment within the casing or rupture of the turbine casing. This research investigates the effects of excessive damage as blades rupture in the high-pressure turbine of a large civil engine. The research investigates different casing inclinations and shrouded/unshrouded blade configurations respectively. The nonlinear finite element software LS-DYNA is used to model two blade release scenarios which are; i) simultaneous release of all blades, and ii) simultaneous sectoral release of blades. The blades are released from firtrees considering the worst case scenario from a containment point of view. It is observed that a sector having a sufficient number of blades can result in the same effect caused by all blades impacting the casing. Containment requirements of shrouded and unshrouded rotors with different casing inclinations are compared as a function of the blade kinetic energy. Provided that the blade mass is kept constant, the effect of the casing inclination is found to be dominant when compared to the effect of blade tip geometry. Together with a rotor overspeed trajectory, the containment requirement of a simultaneous multi-blade shedding application for disk burst prevention is given. The research provides improved understanding of blade tip-to-casing interactions, to be used as an overspeed prevention mechanism, and contributes towards developing design guidelines for the next generation of aero engines in terms of fail-safe engine architectures.
  • Performance assessment of a closed-recuperative-Brayton-cycle based
           integrated system for power generation and engine cooling of hypersonic
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Kunlin Cheng, Jiang Qin, Hongchuang Sun, Chaolei Dang, Silong Zhang, Xiaoyong Liu, Wen Bao Hypersonic vehicle is a next generation aircraft/spacecraft with broad applications, but its development is limited by the high-power electricity supply and thermal protection of engine. This study presents an integrated system based on Closed-Recuperative-Brayton-Cycle (CRBC) for power generation and engine cooling, in which combustion heat dissipation is transferred by liquid metal and partly converted into electric power. An integrated system model which consists of a scramjet combustor, wall cooling channels and a CRBC power generator, is established to evaluate system performance. Results indicate that the integrated system can meet the demands of both high power generation and engine thermal protection for hypersonic vehicles. There is an optimal temperature of liquid metal at heater inlet for power generation performance. The peak electric power increases with fuel equivalence ratio. Liquid sodium exhibits excellent heat transfer performance for engine cooling, the mass flowrate of which becomes greater with smaller fuel equivalence ratio. Besides, the influence of power generation on propulsion is not significant. The specific impulse and specific thrust have maximum decrease of about 2% compared to the cases without heat dissipation. This research provides a novel scheme to achieve power generation and engine thermal protection for hypersonic vehicles.
  • A design strategy for a 6:1 supersonic mixed-flow compressor stage
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Aravinth Sadagopan, Cengiz Camci A surge in the small jet engine market due to aero-propulsion purposes generates requirement to develop compact and robust high-performance compressors. We address this necessity through the design of a single-stage high-pressure ratio mixed-flow compressor. Its compactness and reliability demonstrate its ability to replace a multistage axial design in the small aero-engine segment with high-performance envelope. We have perceived that though many design approaches are readily available for centrifugal and/or axial stages, mixed-flow compressor design systems are scarce.In this paper we intend to provide the designer a comprehensive background knowledge of a mixed-flow stage design. A brief historical development of these designs since the 1940s has been provided. It is observed that for a high-pressure ratio demand it necessitates a supersonic rotor exit flow. Hence, tandem stator configurations were investigated in the past to reduce blade loadings for efficient diffusion. However, most of the previous stage designs were inefficient due to inability of the stators to efficiently diffuse this supersonic flow. A tandem design based on Quishi et al. [1] has been implemented to solve this problem.A unique mean-line procedure based on isentropic equations is defined for mixed-flow stage. It is followed by a geometry construction technique based on Bezier curves. Furthermore, a rotor design evaluation study is conducted for 3.5 kg/s mass flow based on the mean-line code and additional computational analysis. Current computational results [2] have shown single-stage mixed-flow compressors designed using this method to generate reasonably high-pressure ratio up to 6:1 with 75.5% efficiency.
  • Aircraft dynamics simulation using a novel physics-based learning method
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Yang Yu, Houpu Yao, Yongming Liu The fast and accurate prediction of flight trajectories is crucial for the real-time prognostics of the air transportation system. However, the computation costs of simulating aircraft dynamics can be expensive or even prohibitive especially for a large number of aircrafts in the airspace system. This study presents a novel physics-based learning method as a model order reduction (MOR) method for the simulation of aircraft dynamics. The idea of physics-based learning method is to integrate the underlying physics of aircraft dynamical systems into machine learning models to reduce training costs and enhance simulation performances. A recently proposed recurrent neural network (RNN) known as the deep residual RNN (DR-RNN) is used as a tool of physics-based learning. The application of the physics-based learning method is demonstrated on simulating the dynamics of a Boeing 747-100 aircraft. The results show that the DR-RNN can accurately predict aircraft responses and shows excellent extrapolation performances. Furthermore, a purely data-driven approach using the long short-term memory (LSTM) network is also used for the simulation. The comparison demonstrates that incorporating the physics of aircraft dynamics into the learning model can significantly improve prediction performances and effectively reduce training costs compared with using purely data-driven methods. Finally, it is found that the physics-based learning method exhibits superior computation efficiency compared with a classical numerical method since the physics-based learning method can use large time step sizes that violate the numerical stability condition while being explicit in time.
  • Fault-tolerant control for over-actuated hypersonic reentry vehicle
           subject to multiple disturbances and actuator faults
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Yue Yu, Honglun Wang, Na Li In this paper, a fault-tolerant control scheme is proposed for attitude tracking problem of hypersonic reentry vehicle subjected to multiple disturbances and time-varying actuator faults. For convenience of fault-tolerant controller design, conventional hypersonic reentry vehicle model is transformed into control-oriented model by incorporating actuator faults into lumped disturbance. Based on the lumped disturbance estimate provided by high order sliding mode observer, fault-tolerant controllers are developed within the frame of active disturbance rejection control in attitude loop and angular rate loop according to time-scale separation and singular perturbation principle. Then, the desired control moment is allocated to actuators based on recurrent neural network. With the well-designed fault-tolerant controllers and control allocation, a novel meta-heuristic dynamic adaptation salp swarm algorithm is employed to optimize control parameters to achieve minimum attitude tracking error. Comparative simulations are conducted to verify the effectiveness of the developed dynamic adaptation salp swarm algorithm and the investigated fault-tolerant control scheme.
  • Breakup prediction under uncertainty: Application to upper stage
           controlled reentries from GTO orbit
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Francois Sanson, Charles Bertorello, Jean-Marc Bouilly, Pietro M. Congedo More and more human-made space objects re-enter the atmosphere, and yet the risk for human populations remains often unknown because predicting their reentry trajectories is formidably complex. While falling back on Earth, the space object absorbs large amounts of thermal energy that affects its structural integrity. It undergoes strong aerodynamic forces and heating that lead to one or several breakups. Breakup events have a critical influence on the rest of the trajectory but are extremely challenging to predict and subject to uncertainties. In this work, we present an original model for robustly predicting the breakup of a reentering space object. This model is composed of a set of individual solvers that are coupled together such as each solver resolves a specific aspect of this multiphysics problem. This paper deals with two levels of uncertainties. The first level is the stochastic modeling of the breakup while the second level is the statistical characterization of the model input uncertainties. The framework provides robust estimates of the quantities of interest and quantitative sensitivity analysis. The objective is twofold: first to compute a robust estimate of the breakup distribution and secondly to identify the main uncertainties in the quantities of interest. Due to the significant computational cost, we use an efficient framework particularly suited to multiple solver predictions for the uncertainty quantification analysis. Then, we illustrate the breakup model for the controlled reentry of an upper stage deorbited from a Geosynchronous Transfer Orbit (GTO), which is a classical Ariane mission.
  • Optimal trajectories and normal load analysis of hypersonic glide vehicles
           via convex optimization
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Zhenbo Wang Hypersonic trajectory optimization has been intensively investigated through different approaches; however, the normal-load-optimal entry problems were barely studied and reported in the literature. Finding the optimal trajectories with maximum or minimum peak normal load is essential to evaluate the maneuverability and structural strength of the vehicle. In this paper, both the maximum and minimum peak-normal-load entry trajectories are explored using convex optimization. Based on the previous work, the maximum-peak-normal-load entry problem is firstly addressed by a Big-M method and a line-search approach. Through successive relaxations, the nonconvex discrete-event optimal control problem associated with maximum-peak-normal-load entry is transformed into a sequence of mixed-integer convex optimization problems. Then, a line-search technique is introduced to improve the convergence of the proposed method. Additionally, a sequential convex programming method is designed to solve the minimum-peak-normal-load entry problem to comprehensively analyze the normal load during the entry flight. There are efficient solvers that can solve each relaxed convex subproblem with a global optimum if the feasible set of the subproblem is nonempty. The convergence and accuracy of the proposed methodologies are demonstrated by numerical simulations, and the feasibility of the converged solutions is discussed based on an entry-corridor approach.
  • Optimal Staging of Serially Staged Rockets with Velocity Losses and
           Fairing Separation
    • Abstract: Publication date: Available online 9 March 2019Source: Aerospace Science and TechnologyAuthor(s): Aaron D. Koch Currently, algorithms exist for the optimal staging of serially staged rockets under loss-free conditions. These algorithms are based on the Tsiolkovsky rocket equation. Here, one variant is extended to include velocity losses and/or fairing separation. Instead of simply adding the velocity losses to the required loss-free Δv and freely distributing the total amount among all stages, a two-step process is implemented. First, the loss-free solution is obtained to determine the optimal velocity gain for each stage. Then, the rocket's stages are iteratively scaled up, starting with the uppermost stage and continuing downwards. The size of each stage is enlarged so that it generates Δv equal to the optimal velocity gain plus the losses occurring during its flight. Also, each stage accounts for the increased mass of the stages on top of it. Adding the velocity losses after the optimization step ensures that they are allocated to the correct stages. Ariane 40 is used as an example for a three-stage rocket. In this case, the proposed method produced realistic payload ratios, in contrast to the old idea of adding the velocity losses directly to the required loss-free Δv. As the method requires inputs that stem from a trajectory analysis, it works best when iteratively coupled to a trajectory optimizer. In doing so, Ariane 40's total payload ratio was increased, while taking both velocity losses and fairing separation into account.
  • Research of transition criterion for semi-empirical prediction method at
           specified transonic regime
    • Abstract: Publication date: Available online 8 March 2019Source: Aerospace Science and TechnologyAuthor(s): Yayun Shi, Tihao Yang, Junqiang Bai, Lei Lu, Hui Wang A transonic wind tunnel transition test is implemented on a fuselage-wing configuration with the sweep angle 35∘ for commercial aircraft. The wide range of angle of attack from −3.69∘ to 3.07∘ assures that with increasing angle of attack, the laminar to turbulent transition dominant factor varies from cross-flow (CF) vortices to Tollmien-Schlichting (TS) waves. With linear stability theory, the limiting N-factors are calibrated based on the pressure distribution by experiment or the Reynolds Averaged Navier-Stokes (RANS) solver using the fixed experimental transition location. The pressure distribution of the RANS solver agrees well with the experiment in general except some small discrepancies, which causes deviation by 0.6 for the limiting TS N-factor. The RANS solver and the stability analysis provide the limiting N-factors of 7.0 and 8.7 for CF-vortices and TS-waves at the two sides, respectively. In the between, the TS value decays with the CF value due to their interaction. Thus, the transition criterion for limiting N-factors is established for the laminar prediction tool of eN method at similar transonic wind tunnel. With the transition criterion, the transition location difference for 95% cases between the simulation and the experimental data is lower than 5% chord. The good match illustrates that the transition tool is accurate and robust for engineering applications, and also verifies the reasonability of the limiting N-factors. Therefore, the transition criteria at similar transonic conditions and well-performed eN transition tool can be applied for the future laminar wing design.
  • Experimental investigation of the effect of extended cowl on the flow
           field of planar plug nozzles
    • Abstract: Publication date: Available online 8 March 2019Source: Aerospace Science and TechnologyAuthor(s): Aqib Khan, Rohit Panthi, Rakesh Kumar, S. Mohammad Ibrahim Plug and ramp nozzles offer many advantages over the conventional converging-diverging counterpart and have received much attention in the recent decades. Variants of plug nozzles have been investigated for a wide range of operating conditions. The performance of these nozzles is dependent on the flow development on the plug or ramp surface, which in turn is greatly influenced by the cowl geometry. Experiments are conducted to study the planar plug nozzle flowfield for Mach 1.8 and 2.2 using half nozzle geometry. The work primarily focuses on the influence of the cowl length on the flow evolution on the plug surface at different nozzle pressure ratios (NPRs). The cowl of the outer nozzle is extended to 10, 30, 50 and 100% of the full plug length. This allows the supersonic flow to partially expand internally ahead of the throat section. Schlieren images are used to visualize the wave structure at different pressure ratios for different cowl lengths. For low NPRs, the nozzle with extended cowl behaves more like a conventional planar nozzle with strong shock waves. It is observed that the cowl length influences the pressure distribution on the plug surface only for low NPRs. The effect of side walls on the flow field of planar plug nozzles is also studied.
  • Rotor aerodynamic shape design for improving performance of an unmanned
    • Abstract: Publication date: Available online 7 March 2019Source: Aerospace Science and TechnologyAuthor(s): Qing Wang, Qijun Zhao In order to improve the effective load of an unmanned helicopter, a new blade shape with a new airfoil section, non-linear negative twist, varied chord length and double-swept blade tip was designed by employing a CFD method coupled with an optimization method. Firstly, a new airfoil is designed with improved aerodynamic characteristics. By comparing the numerical data, the new airfoil has better lift-drag ratio, i.e., the maximum lift-drag ratio increases from 63.03 to 66.29 at Mach number of 0.3 and from 70.17 to 72.13 at Mach number of 0.4, compared with the original NACA8H12 airfoil. In addition, a new blade shape is designed based on the new airfoil. By comparing the numerical data, the CT of the design rotor increases from 4.79×10−3 to 5.07×10−3, and the maximum FM increases from 0.67 to 0.72 at the design state. After that, a verification test is performed in hovering flight. The test data indicated the a maxumu thrust increase of 3.18% at the design state. Meanwhile, the FM of design rotor is improved about 3.41% compared with the original rotor.
  • Resonance avoidance for variable speed rotor blades using an applied
           compressive load
    • Abstract: Publication date: Available online 7 March 2019Source: Aerospace Science and TechnologyAuthor(s): Robert Dibble, Vaclav Ondra, Branislav Titurus Varying the rotational speed of the main rotor is one method being considered to improve the performance of future rotorcraft. However, changes in rotor speeds often lead to resonant interactions between rotor blade modes and the rotor's excitation frequencies which increase the vibratory loads in the rotor. This research investigates the use of a compressive load to reduce a blade's natural frequencies and its potential to be used as a resonance avoidance technique by improving separation between the natural and excitation frequencies of a blade. The research presented herein describes and validates a model of a pretwisted rotating beam with non-coincident mass and elastic axes with an applied compressive load. The compressive load is applied at the elastic axis at the tip of the beam and is orientated towards the root of the beam. The beam model is then used in a case study to represent the rotor blade of a typical mid-sized civilian helicopter. The case study is performed to calculate the natural frequencies of a compressed blade for a reduction in rotor speed of up to 40% and evaluate the performance of the compressive load resonance avoidance technique. The results of the case study show that the compressive load improves the separation between natural and excitation frequencies over the full range of rotor speeds evaluated. The improved separation allows the rotor to operate safely with a reduction in rotor speed of up to 19%.
  • FV-MP model to predict lean blowout limits for multi-point lean direct
           injection combustors
    • Abstract: Publication date: Available online 7 March 2019Source: Aerospace Science and TechnologyAuthor(s): Lei Sun, Yong Huang, Ruixiang Wang, Xiang Feng, Zhilin Liu, Jiaming Wu The lean blowout (LBO) limit is a crucial performance for aircraft engine combustors. It is essential to obtain the LBO limit during the design stage of the aircraft engine combustors. The semi-empirical correlation is an important tool for quick prediction of the LBO limits. Among all the semi-empirical correlations for the prediction of the LBO limits, Lefebvre's LBO model is widely used for the swirl stabilized combustors. The Flame Volume (FV) model was proposed based on Lefebvre's LBO model to accommodate the effects of the geometry of the flame tube on the LBO. Meanwhile, the multi-point lean direct injection (MPLDI) combustor whose geometry of the dome is different from the traditional combustors is a promising low NOx emission combustor. Up to now, there are few existing semi-empirical correlations to predict the LBO limit for the MPLDI combustors although the prediction of the LBO limit is critical for them. Based on the FV concept and new physics-based analysis, the FV-MP (Flame Volume for the Multi-Point) model is derived to predict the LBO limits for the MPLDI combustors. The FV-MP model could accommodate the effects of the fuel staging and recessed pilot stage, in addition to the operating conditions, on the LBO limits of the MPLDI combustors and achieve better prediction accuracy than both the FV and Lefebvre's LBO models within the range of corresponding validation experiments. Compared with Lefebvre's LBO model, the FV model could double the prediction accuracy. Compared with the FV model, the FV-MP model could further double the prediction accuracy.
  • Effect of Adverse Pressure Gradient on Supersonic Compressible Boundary
           Layer Combustion
    • Abstract: Publication date: Available online 7 March 2019Source: Aerospace Science and TechnologyAuthor(s): Pu Zhang, Jinglei Xu, Yang Yu, Wei Cui Influences of adverse pressure gradient (APG) on supersonic turbulent boundary layers are numerically studied using Reynolds-averaged Navier-Stokes (RANS) equations. The RANS methodology is validated by comparing the numerical results with the existing experimental data. Although the flame is restricted in the boundary layer, the heat flux is reduced rather than increased for the suppressed turbulence momentum transportation ability. Moreover, APG contributes more to the heat flux than combustion heat release does. Compared with no-injection case, a large skin friction reduction can be obtained by boundary layer combustion, and further reduction can be achieved in APG state. Additionally, the effects of combustion in APG on the velocity laws of the wall show that White's velocity law is close to the numerical results in the outer region of boundary layer, and Nichols's velocity law is appropriate in the whole boundary layer unless the APG is too strong. Turbulence intensity influences are analyzed in the end. Results show that the additional reduction of skin friction due to induced combustion cannot offset the skin friction increase caused by high turbulence intensity.
  • Studies on the effect of imaging parameters on dynamic mode decomposition
           of time-resolved schlieren flow images
    • Abstract: Publication date: Available online 6 March 2019Source: Aerospace Science and TechnologyAuthor(s): M.V. Srisha Rao, S.K. Karthick Dynamic modal analysis enables new insights into the spatio-temporal dynaimcs of complex flow scenarios. Time resolved schlieren imaging provides significant information in compressible flow scenarios on flow structures and their evolution. We conduct a systematic study using synthetic images and experimental schlieren images on the effect of image acquisition parameters on the modal analysis by dynamic mode decomposition (DMD). We consider the effect of two important capture parameters - the capture rate (fs) and the exposure time (texp). Analysis is carried out on two sets of synthetic images, SI-I, an unsteady wavy interface created using a linear combination of sinusoids, and SI-II - hypothetical shock oscillations. Finally, a flapping supersonic jet is observed using high-speed schlieren with a nano-pulsed laser light source with three different imaging parameters. We find that among the two parameters the effect of exposure time on modal analysis and its interpretation is more pronounced than capture rate. An exposure time of 5% of maximum exposure produces 8% reduction in mode amplitude, and in case of long exposure the dynamic significance of modes undergoes complete change. If the flow images are instantaneous, then the spatial mode shapes of dominant modes remain the same irrespective of the capture rate. Aliasing has to be considered in sub-Nyquist capture rates, however, the actual frequencies can be suitably resolved.
  • A method for optimizing the aerodynamic layout of a helicopter that
           reduces the effects of aerodynamic interaction
    • Abstract: Publication date: Available online 6 March 2019Source: Aerospace Science and TechnologyAuthor(s): Yang Lu, Taoyong Su, Renliang Chen, Pan Li, Yu Wang The aerodynamic environment during the flight of a helicopter is complex because of the severe aerodynamic interaction between various components. To fully consider the effect of aerodynamic interaction in the initial stages of helicopter design and to eliminate or reduce its adverse effects, a comprehensive design optimization method for the aerodynamic layout of a helicopter that is capable of reducing the adverse effects of aerodynamic interaction is developed in this paper. To satisfy the requirements for precision and efficiency in the calculation model, an aerodynamic interaction analysis model of various helicopter components was established based on a viscous vortex particle and the unsteady panel hybrid method. To simultaneously consider the influences of the position and shape of the aerodynamic components on the aerodynamic interaction during the optimization process, parameter modeling of the helicopter's shape was performed based on the class function/shape function transformation (CST) method. A Kriging surrogate model of the objective function was further developed and combined with a hybrid sequential quadratic algorithm and genetic algorithm optimization strategy to establish a comprehensive optimization flow for the aerodynamic layout of a helicopter that reduces the adverse effects of aerodynamic interaction. Verification was carried out based on a fuselage shape derived from UH-60 helicopter. The optimization results showed that the use of the comprehensive optimization method for the aerodynamic layout of a helicopter can effectively reduce the adverse effects of aerodynamic interaction. Based on the optimization objectives, the efficiency of hovering increased by 4.7%, the hovering ceiling increased by 3.48%, the speed stability derivative increased by 264.7%, and the angle of attack stability derivative decreased by 26.4%.
  • The Potential of Helicopter Turboshaft Engines Incorporating Highly
           Effective Recuperators under Various Flight Conditions
    • Abstract: Publication date: Available online 6 March 2019Source: Aerospace Science and TechnologyAuthor(s): Chengyu Zhang, Volker Gümmer Incorporating recuperators into gas turbines shows considerable potential for lower emissions and fuel consumption. Nowadays the technology readiness of advanced compact heat exchanger has provided a solid foundation for the availability of lightweight, higher efficient recuperators which would find good acceptance on the rotorcraft without penalizing the operational capabilities. To understand the impact of recuperator on the whole system for further development of future recuperated helicopter, it is proposed to evaluate the potential of recuperated helicopter turboshaft engines with emphasis placed on highly effective primary surface recuperator. This paper presents a comprehensive multidisciplinary simulation framework, and the aircraft configuration selected is a generic helicopter, which is similar to the helicopter Bo105, equipped with two Allison 250-C20B turboshaft engine variants. The improved part-load performance against the reference non-recuperated cycle is discussed first, followed by the analysis and evaluation of two representative flight missions. The study is finally extended to quantify the flight time required to compensate for the additional recuperator weight under the flight condition of 0-250 km/h and 0-3000 m for different recuperator design effectiveness values. It is suggested that the selection of recuperator effectiveness should be dependent on the most commonly involved mission profile and flight duration, in order to offset the added parasitic weight of the recuperator. The established rotorcraft multidisciplinary framework proves to be an effective tool to conduct a comprehensive assessment for the recuperated helicopter under a wide range of flight conditions as well as at mission level.
  • Quadrotor Fault Tolerant Incremental Sliding Mode Control driven by
           Sliding Mode Disturbance Observers
    • Abstract: Publication date: Available online 6 March 2019Source: Aerospace Science and TechnologyAuthor(s): Xuerui Wang, Sihao Sun, Erik-Jan van Kampen, Qiping Chu This paper proposes an Incremental Sliding Mode Control driven by Sliding Mode Disturbance Observers (INDI-SMC/SMDO), with application to a quadrotor fault tolerant control problem. By designing the SMC/SMDO based on the control structure of the sensor-based Incremental Nonlinear Dynamic Inversion (INDI), instead of the model-based Nonlinear Dynamic Inversion (NDI) in the literature, the model dependency of the controller and the uncertainties in the closed-loop system are simultaneously reduced. This allows INDI-SMC/SMDO to passively resist a wider variety of faults and external disturbances using continuous control inputs with lower control and observer gains. When applied to a quadrotor, both numerical simulations and real-world flight tests demonstrate that INDI based SMC/SMDO has better performance and robustness over NDI based SMC/SMDO, in the presence of model uncertainties, wind disturbances, and sudden actuator faults. Moreover, the implementation process is simplified because of the reduced model dependency and smaller uncertainty variations of INDI-SMC/SMDO. Therefore, the proposed control method can be easily implemented to improve the performance and survivability of quadrotors in real life.
  • Aerothermodynamic analyses and redesign of GHIBLI Plasma Wind Tunnel
           hypersonic diffuser
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Pasquale Walter Agostinelli, Eduardo Trifoni, Raffaele Savino This article deals with the aerothermodynamic analyses of the hypersonic diffuser of GHIBLI Plasma Wind Tunnel at CIRA (Centro Italiano Ricerche Aerospaziali). This diffuser presented an unexpected behavior not ensuring nozzle under-expanded flow in test chamber, even in several free jet conditions. This experimental evidence was investigated following two approaches. First, a literature review pointed out that the design rules, generally recommended in supersonic diffusers design, were not observed in the case of GHIBLI diffuser. Second, a simplified numerical model was introduced and validated with experimental results, performing Computational Fluid Dynamics analyses of GHIBLI diffuser. Finally, in order to enhance the pressure recovery and to obtain the desired under-expanded conditions, two new candidate designs compliant with the literature recommended values have been proposed and analyzed via CFD.
  • Structural Health Monitoring for Long-Term Aircraft Storage Tanks under
           Cryogenic Temperature
    • Abstract: Publication date: Available online 4 March 2019Source: Aerospace Science and TechnologyAuthor(s): Dongyue Gao, Zhanjun Wu, Lei Yang, Yuebin Zheng, Wan Yin In order to monitor the structural conditions, an SHM technology is necessary for long-term aircraft storage tanks under cryogenic conditions. In this paper, a PZT- based Lamb waves SHM technology is developed for such storage tanks. In order to determine the survivability, durability of different PZT-epoxy sensor systems and functionality of the damage diagnosis method under cryogenic conditions of long-term storage tanks, a series of tests have been conducted. First, the durability of PZT-epoxy sensor systems under cryogenic environment was considered by cryogenic durability tests. Simultaneously, performance tests of different PZT-epoxy sensor systems were performed, include high strain performance test and Lamb waves propagation tests under different temperature environments. The high strain performance of different epoxy adhesives under cryogenic environments was investigated by lap shear strength tests. The functionality of different PZT-epoxy sensor systems was investigated by Lamb waves propagation tests. At last, the damage diagnosis ability of the SHM technology was evaluated in a composite damage diagnosis experiment under cryogenic temperature. Experimental results demonstrated that the developed SHM technology can withstand operational levels of high strain and long-term under cryogenic/room temperature on cryogenic storage tanks, and is functional in the cryogenic environment.
  • Leakage and rotordynamic performance of T type labyrinth seal
    • Abstract: Publication date: Available online 4 March 2019Source: Aerospace Science and TechnologyAuthor(s): Xingyun Jia, Qun Zheng, Yuting Jiang, Hai Zhang A new type of seal, called a T type labyrinth seal is proposed to reduce exciting force induced by seal region flow, thereby reducing rotor vibration in aero-engines. The whole circle computational model is performed, and the vibration characteristics parameters are defined at the boundary of the seal flow field to consider the effect of rotor vibration. The new type seal is analyzed along with the interlaced and straight through labyrinth seals. The enhanced Lomakin effect of seal flow field caused by the T-shaped tooth contributes to the great increase in the inward radial aerodynamic force. With the same axial length of the seal and the tip clearance, the leakage of T seal is approximately 23.6-25.3% less than the straight through labyrinth seal, and approximately 7.4-8.5% more than the interlaced labyrinth seal.
  • Orientation and size effect of a rectangle cutout on the buckling of
           composite cylinders
    • Abstract: Publication date: Available online 4 March 2019Source: Aerospace Science and TechnologyAuthor(s): Abolfazl Shirkavand, Fathollah Taheri-Behrooz, Milad Omidi In this article the effect of a rectangular cutout on the buckling behavior of a thin composite cylinder was investigated using numerical and experimental methods. To verify the finite element results, a limited number of tests was carried out on perforated and non-perforated glass/epoxy cylinders with [90/-23/23/90] layups. In the numerical analysis, linear and nonlinear approaches were employed to study the effect of initial imperfections on the buckling of the cylinders. Several key findings including the effects of cutout size and orientation, and the mutual effects of the cutout and initial imperfections on the buckling behavior were investigated in detail.In the presence of cutouts, the effect of initial imperfections on the buckling load is a function of the cutout size. In cylinders with rectangular cutouts, buckling analysis revealed that a rectangular cutout in the circumferential direction causes around 8% more reduction in the buckling load than the same cutout in the axial direction. Also, numerical findings illustrated that elastic stress concentration factors for the circumferential cutouts are much greater than those for the axial cutouts; thus premature failure around the cutout will trigger earlier buckling in the cylinder with circumferential cutouts.
  • An experimental study of the effects of different transverse trenches on
           depositing and temperature on a plate with film cooling holes
    • Abstract: Publication date: Available online 4 March 2019Source: Aerospace Science and TechnologyAuthor(s): Zhengang Liu, Zhenxia Liu, Fei Zhang, Yanan Liu Particles depositing is a severe damage to the aerodynamic and cooling performance of turbine guide vanes and a technique to decrease depositing is transverse trench. In this paper, the experiments are conducted to study the effects of different transverse trenches on the particles depositing and temperature on a plate. The particles are generated by atomizing molten wax. All the models are made based on a plate and have film cooling holes. Different transverse trenches with different widths and depths are installed along the row of the film cooling holes. The experimental results show that the trenches could alter the depositing mass and distribution on the pressure surface of the model. The narrower or deeper trench could result in the lighter depositing mass and more uniform depositing distribution on the vicinity downstream of the row of film cooling holes. However, the trenches have little effect on the depositing upstream of the row of film cooling holes on the pressure surface of the model. The more depositing under the experimental conditions leads to higher temperature on the pressure surface of the model. The temperature distribution on the vicinity downstream of the row of film cooling holes could also be altered by the trenches. The narrower or deeper trench could make the temperature distribution more uniform. This could be attributed to the effect of the trench on the depositing and the direct effect of itself on the temperature distribution.
  • Application of the natural element method for the analysis of composite
           laminated plates
    • Abstract: Publication date: Available online 1 March 2019Source: Aerospace Science and TechnologyAuthor(s): Mohamed Amine Bennaceur, Yuang ming Xu Composite laminated plates are widely used in modern aerospace structures. Thus, pursuing an alternative and effective numerical approximate method for the analysis of composite laminated is always a demanding task for aerospace application. It is essential to determine the mechanical behaviors of those structures. This paper demonstrates the applicability of a novel meshless method in solving problems related to aeronautical engineering. The natural element method combines the advantages of meshless methods and finite element approaches. By its properties, it overcomes most of difficulties such as the imposition of Dirichlet boundary conditions or problems related to the size of the support domain comparing to other methods. The Stress Concentration Factor is investigated regarding its diminishing effect, the critical buckling loads and natural frequencies with their associated modes were determined and the results were compared to the analytical solutions. The proposed method is proved to have a good computational accuracy for composite plate analysis and also exhibit an automatic and optimal choice for the shape function and the imposition of essential boundary conditions, thus demonstrating the fitness and flexibility of this approach and a promising perspective for solving structural analysis problems of composite structures.
  • Dynamics and control of a tethered space-tug system using Takagi-Sugeno
           fuzzy methods
    • Abstract: Publication date: Available online 28 February 2019Source: Aerospace Science and TechnologyAuthor(s): Zhiping Zhang, Zhiwei Yu, Qianwen Zhang, Ming Zeng, Shunli Li Mitigation of space debris draws much concern today since its steadily increasing number and the consequent increased risk of space activities. Among all the active debris removal strategies, tethered space-tug is considered as a promising method due to its safeness of capture and dynamic stability of pull strategy. Although promising, challenges remain in its position and attitude control especially when the target is tumbling, which may cause damage to the appendages or result in collisions. This paper addresses the problem of position and attitude control of a tethered space-tug system with only tether tension available for the target. Based on Takagi-Sugeno fuzzy technology, a guaranteed cost guidance and control law is obtained by solving a set of linear matrix inequalities. The effectiveness of the guidance and control strategy is validated by numerical simulations.
  • Dynamics and control of spacecraft with a large misaligned rotational
    • Abstract: Publication date: Available online 26 February 2019Source: Aerospace Science and TechnologyAuthor(s): Fan Wu, Xibin Cao, Eric A. Butcher, Feng Wang A new type of earth observation mission using a spacecraft with rotational payload is introduced in this paper. The kinematic and dynamic equations for the nadir-aligned attitude tracking problem are derived in the presence of a misaligned rotational payload. A quaternion-based feedback control law is proposed and its stability is analyzed using the Lyapunov direct method. Comparisons are made between the simulation results of the proposed control law and that of a conventional PD feedback control law. The analytical solution of the spacecraft's residue angular momentum is obtained. Monte Carlo analyses are employed to demonstrate the angular rate accuracy of the system and to study the effects caused by the knowledge error of parallel misalignment. Simulation results show that the angular rate accuracy improves greatly by the proposed control law.
  • Novel quadrature element formulation for simultaneous local and global
           buckling analysis of eccentrically stiffened plates
    • Abstract: Publication date: Available online 22 February 2019Source: Aerospace Science and TechnologyAuthor(s): Jian Deng, Xinwei Wang, Zhangxian Yuan, Guangming Zhou Predicting both buckling load and mode shape of eccentrically stiffened panels correctly is of important and a rather challenging task. In this paper, a novel and efficient quadrature element formulation is developed to fulfill this challenging task. A high-order quadrature plate-stiffener element is proposed by assembling quadrature beam elements to the quadrature plate element via the displacement relations. Thus, the same plate mesh scheme can be used for buckling analysis of stiffened plate with different numbers of stiffeners located at arbitrary positions. Explicit formulations and solution procedures are given. Convergence studies are performed for stiffened plates with different shapes of cross-section of stiffeners and materials. A number of case studies are given. For validations, numerical results are compared with either existing solutions or finite element data. It is demonstrated that high accuracy can be achieved with relatively small number of nodes. Presented formulation is simple, straightforward, and reliable, which can allow a quick and accurate analysis of buckling behavior of eccentrically stiffened plates.
  • Elastic buckling response of rectangular GLARE fiber-metal laminates
           subjected to shearing stresses
    • Abstract: Publication date: Available online 22 February 2019Source: Aerospace Science and TechnologyAuthor(s): George S.E. Bikakis, Costas D. Kalfountzos, Efstathios E. Theotokoglou In this article, the elastic buckling response of rectangular GLARE fiber-metal laminates with three different support types subjected to shearing stresses is investigated using the finite element method and eigenvalue buckling analysis. Using validated FEM models, the buckling coefficient-aspect ratio diagrams of eight GLARE grades are obtained and studied along with the diagrams of three UD glass-epoxy composites and monolithic 2024-T3 aluminum. It is found that the critical average buckling stress and the buckling load of the materials increases for increasing metal volume fraction. The effect of the fiber-volume fraction on the shear buckling strength of simply supported and clamped fiber-metal laminates is also studied. It is found that the variation of the fiber-volume fraction has a small influence on the behavior of the buckling coefficient-aspect ratio diagrams of two specific GLARE grades. Furthermore, an approximate method is proposed in order to estimate the new shear buckling strength of a GLARE plate when the fiber-volume fraction changes.
  • A comparative study of blast resistance of cylindrical sandwich panels
           with aluminum foam and auxetic honeycomb cores
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Xuke Lan, Shunshan Feng, Qi Huang, Tong Zhou The dynamic response of cylindrical sandwich panels with aluminum foam core, hexagonal honeycomb core, and auxetic honeycomb core are compared numerically. A novel curved auxetic honeycomb core is designed, and the finite element models are built by employing ABAQUS–Explicit. To calibrate the numerical models, the experiments of sandwich panels with honeycomb core and aluminum foam cores are modeled. And the numerical results have a good agreement with the experiment date. The calibrated numerical models are used to simulate the dynamic response of cylindrical panels subject to external blast loadings. It is found that the cylindrical panels with auxetic honeycomb cores have a better performance than that with aluminum foam cores and hexagonal honeycomb cores in resisting blast loadings. A material concentration effect was observed in the auxetic honeycomb core due to the negative Poisson's ratio (NPR) effect. According to parameter studies, it is concluded that with the increase of curvature and face sheet thickness the blast-resistance of panels with both auxetic honeycomb core, hexagonal honeycomb core, and foam cores increased obviously, especially the panels with auxetic honeycomb cores. For the panels with auxetic honeycomb cores, increasing the back face sheet thickness can improve the blast-resistance performance more efficiently than increasing the thickness of front face sheet, which is opposite for the panels with foam cores and hexagonal honeycomb cores. Auxetic cores with a smaller unit cell aspect ratio and a smaller unit cell length ratio has a larger Poisson's ratio, and achieves better blast resistance performance. These simulation findings can guide well the theoretical study and optimal design of cylindrical sandwich structures subject to external blast loading.
  • Parametric reduced-order modeling of unsteady aerodynamics for hypersonic
    • Abstract: Publication date: April 2019Source: Aerospace Science and Technology, Volume 87Author(s): Zhiqiang Chen, Yonghui Zhao, Rui Huang A novel parametric reduced-order model (ROM) is proposed for efficiently predicting hypersonic unsteady aerodynamic responses under different flight conditions. The construction of the ROM is realized by computational fluid dynamics (CFD) simulations under the prescribed motion of the structure, while the results are processed via the proper orthogonal decomposition (POD) to obtain the predominant flow modes. Subsequently, to obtain the ROM valid for varying flight conditions, the method of interpolation in a tangent space to a Grassmann manifold is used to generate a new POD modal matrix for the arbitrary operating point in the considered parameter space. Finally, the least squares support vector machine (LS-SVM) is carried out to obtain the nonlinear relations between the applied excitations and the resulting POD coefficients. Once the parametric ROM is constructed, it can behave as a substitution of the full-order CFD flow solver in the considered parameter space. For demonstration purposes, the parametric ROM is used to predict hypersonic unsteady aerodynamic loads, flutter boundaries and limit-cycle oscillations of a double wedge airfoil over a wide range of the flight conditions, respectively. Numerical investigations show a good agreement between the results obtained by the ROM methodology in comparison to the full-order CFD solution over a wide range of the parameters. In addition, the ROM approach yields a significant speedup regarding unsteady aerodynamic calculations, which is beneficial for aeroelastic analysis, control and optimization applications.
  • Investigation of flame flashback phenomenon in a supersonic crossflow with
           ethylene injection upstream of cavity flameholder
    • Abstract: Publication date: Available online 21 February 2019Source: Aerospace Science and TechnologyAuthor(s): Guoyan Zhao, Mingbo Sun, Jinshui Wu, Xingda Cui, Hongbo Wang In the present study, a flame flashback phenomenon inside an ethylene-fueled scramjet combustor equipped with a cavity flameholder is investigated under flight Mach 5.5 condition. Experimental results exhibit the quasi-periodic combustion oscillation between the fuel injectors and the leading edge of the cavity under a specified condition. As an indispensable key sub-process of combustion oscillation, the flame flashback from the boundary layer downstream of the cavity is responsible for unsteady combustion process in scramjet combustors. Numerical simulation has been carried out in the specified condition. Attributed by (i) thicker boundary layer, (ii) closer thermal disturbance and (iii) improved local mixing degree, the flame flashback phenomenon can be induced by thermal throat which is generated by interaction between separated boundary layer and intense combustion. Quantitative analysis also indicates that the flame flashing is more sensitive to temperature fluctuation downstream of the cavity. In addition, a simplified combustion opening system model has been established to analyze combustion oscillation mechanisms, which theoretically demonstrates that the three factors mentioned above can destroy the balance of heat release and dissipation, causing the system cannot self-stabilize once certain temperature fluctuation thresholds in sensitive areas are exceeded. At the same time, the auto-ignition model excludes the possibility of flame flashback generated by auto-ignition effect.
  • Nonlinear analysis of FG-sandwich plates and shells
    • Abstract: Publication date: Available online 20 February 2019Source: Aerospace Science and TechnologyAuthor(s): M. Rezaiee-Pajand, E. Arabi, Amir R. Masoodi This study is dedicated to develop a mixed interpolated formulation for nonlinear analysis of plates and shells. Using Equivalent Single Layer (ESL) theory and the rule of the mixture scheme, the authors present a formulation for analyzing Functionally Graded (FG) sandwich structures. To incorporate large displacements and rotations, and alleviate the shear and membrane locking phenomena, a mixed interpolation of strain fields, including in-plane and shear strains, is utilized. Analyses are based on an isoparamertic 6-node triangular shell element. There are five degrees of freedom, including three displacements and two rotations at each node of this element. In order to validate the proposed formulation, some well-known benchmark problems are solved. Besides, the obtained responses are compared to the other available solutions, separately. Moreover, two other practical structures, including FG-sandwich plate and curved shallow panel, are considered to indicate the accuracy and high performance of the authors' approach.
  • Multi-objective Airfoil Shape Optimization Using an Adaptive Hybrid
           Evolutionary Algorithm
    • Abstract: Publication date: Available online 19 February 2019Source: Aerospace Science and TechnologyAuthor(s): HyeonWook Lim, Hyoungjin Kim In this study, multi-objective aerodynamic optimization problems were conducted with a hybrid evolutionary-adaptive directional local search method for convergence enhancement. The directional search operator includes selection of search direction and one-dimensional search. Probability for the directional operator is adaptively changed based on relative effectiveness of the directional search operator and evolutionary operators such as crossover and mutation. The adaptive directional operator is combined with a baseline evolutionary multi-objective algorithm (EMOA) such as NSGA-II or MOGA. Multi-objective airfoil shape optimization examples are defined as drag minimization/lift maximization and L/D maximization at high lift and cruise conditions in subsonic and transonic regimes. The CST method and the B-spline method were used for airfoil shape parameterization. Design examples with drag minimization and L/D maximization at cruise conditions are all found to have uni-modal design spaces, and the local search operator is effective for those examples. However, lift maximization and high angle of attack cases show multi-modality in the design spaces due to flow separations and thus the local search is not effective for those cases. Design results show that the adaptive directional search method significantly enhances convergence of problems in which the directional search is effective, and also minimizes unnecessary spending of computational budget for cases in which the directional search does not produce competitive solutions. The present method improves search performance for different airfoil parameterization methods and different baseline EMOAs. Statistical tests confirm that the adaptive hybrid method is superior to the baseline EMOA.
  • A proposed self-organizing radial basis function network for aero-engine
           thrust estimation
    • Abstract: Publication date: Available online 31 January 2019Source: Aerospace Science and TechnologyAuthor(s): Zhi-Qiang Li, Yong-Ping Zhao, Zhi-Yuan Cai, Peng-Peng Xi, Ying-Ting Pan, Gong Huang, Tian-Hong Zhang This paper proposes a new algorithm to construct self-organizing radial basis function neural networks (RBFNNs) for aero-engine thrust estimation. The algorithm can not only optimize centers and network size of the RBFNN but also automatically determine the connection weights. To reduce the dimensionality of particle and speed up the optimization process, spreads of an RBFNN are randomly initialized. Its weights are dynamically derived and adjusted by the product between the Moore–Penrose inverse of the hidden layer's outputs and the desired outputs. To optimize the centers and network size of the RBFNN, a strategy named multi-Gbest is adopted. Based on all these strategies, the proposed algorithm can effectively generate self-organizing RBFNNs with high accuracy. The successful application to aero-engine thrust estimation shows the practicability and effectiveness of the proposed algorithm.
  • Numerical investigation of the supersonic stabilizing parachute's heating
    • Abstract: Publication date: Available online 31 January 2019Source: Aerospace Science and TechnologyAuthor(s): Feng Qu, Di Sun, Kai Han, Junqiang Bai, Guang Zuo, Chao Yan In the design of the scale reduced model of the new generation reusable re-entry capsule, the supersonic stabilizing parachute should be applied to avoid the failure of the main parachute due to the uncertainty of the space capsule's attitude. Compared with the inflation technology during the subsonic process, special attention should be paid to the aerodynamic heating loads during the inflation process at supersonic speeds. In this manuscript, the inflation process of the supersonic stabilizing parachute is numerically studied by adopting the fluid–structure coupling method. The three-dimensional compressible Reynolds Averaged Navier–Stokes (RANS) equations are simulated and Menter's shear stress transport (SST) turbulence model is applied. Also, we study the aerodynamic heating loads of the stabilizing parachute by conducting numerical simulations at a typical reentry trajectory point. Results suggest that the wake of the capsule increases the temperature of the flow field remarkably. Also, the inner faces of the parachute encounter more sever heating loads than the outer faces, and the most sever heating load appears at the first horizontal parachute belt.
  • A pressure-controllable bump based on the pressure-ridge concept
    • Abstract: Publication date: Available online 18 February 2019Source: Aerospace Science and TechnologyAuthor(s): Zonghan Yu, Guoping Huang, Chen Xia, Joern Sesterhenn An innovative type of pressure distribution for the hypersonic aircraft forebody (bump) is presented, and this design is based on the newly established known as pressure ridge (PR) flow mechanism. Studies on the low-kinetic-energy fluid over the aircraft surface are summarized. Then, the challenges of the inlet–airframe integration at high speeds are discussed. The flow structure around the bump is analyzed in detail to eliminate the side-embedded shock (SES) effect at the inlet entrance. The concept of PR is proposed to improve the overall aerodynamic characteristics of the bump, namely, the boundary layer removal, the reduction of external drag, and the streamline direction at bump end-section. On the basis of the developed inverse method to generate the bump surface by the prescribed pressure distribution, the improved PR-derived bump is designed and numerically compared with the typical pressure-controllable bump (PCB) while the identical leading edge profile is imposed. Results demonstrate that the PR creates outward and inward pressure gradients. The removed amount of boundary layer increases with the increase in outward pressure gradient. Meanwhile, the inward pressure gradient determines the streamline pattern after the bump. The PR-derived bump is 29.2% lower in height than the typical one by using a proper outward pressure gradient. The uniform area of the new bump is 30% wider than that of the typical PCB. The near-wall streamlines of the new bump are adjusted from expanding to the parallel, thereby relieving the side-compression. Changing the location, width, and peak value of the PR can lead to great flexibility in the design and optimization of aircraft forebody subjected to hypersonic flow.
  • Numerical Investigation of the Aerodynamic Interaction between a Tiltrotor
           and a Tandem Rotor during Shipboard Operations
    • Abstract: Publication date: Available online 12 February 2019Source: Aerospace Science and TechnologyAuthor(s): Jian Feng Tan, Tian Yi Zhou, Yi Ming Sun, George N. Barakos Complex rotorcraft-to-rotorcraft interference problems occur during shipboard operations, and have a negative impact on safety. A vortex-based approach is used here to investigate the flow field and unsteady airloads of a tiltrotor affected by the wake of an upwind tandem rotor. In this work, the blade aerodynamics is modelled using a panel method, and the unsteady behavior of rotor wakes is modelled using a vortex particle method. The effects of the ship and sea-surfaces are accounted for via a viscous boundary model. The method is applied to a 1/48th scaled model of a CH-46 operating on a model-scale Landing Helicopter Assault ship. The predicted vertical velocities at the location of the downstream V-22 are compared with Computational Fluid Dynamics and experiments carried out at NASA Ames Research Center. The results show that the predicted vertical velocities compare reasonably well with experiments and Computational Fluid Dynamics. A V-22 tilt-rotor placed in the wake of the CH-46 is also simulated, and rolling moments of the V-22 are calculated to show the effect of the upstream CH-46 wake.
  • Parameterized nonlinear suboptimal control for tracking and rendezvous
           with a non-cooperative target
    • Abstract: Publication date: Available online 12 February 2019Source: Aerospace Science and TechnologyAuthor(s): Dengwei Gao, Jianjun Luo, Weihua Ma, Brendan Englot A specific parameterized nonlinear suboptimal control technique is proposed to control the relative position of a spacecraft in order to track a rotating target. The technique consists of using power series expansion to parameterize an SDRE (State-Dependent Riccati Equation) with an algebraic expression. One of the major contributions of this technique is the avoidance of online solution of algebraic Riccati and Lyapunov equations that will be much faster than the standard SDRE and θ−D control. Meanwhile, parameterized nonlinear suboptimal control is extended to adaptive form to verify robustness to unknown disturbances. Finally, we show two benchmark examples using this parameterized technique to construct controllers. Specifically, we also apply this technique to design the nonlinear control of a chaser spacecraft to track and rendezvous with a rotating non-cooperative target accompanied by an unknown translational maneuver. Numerical results demonstrate that the computational efficiency and tracking performance accuracy are superior to existing methods and an adaptive form is capable of offsetting unknown parameters.
  • Analysis of angular errors of the planar multi-closed-loop deployable
           mechanism with link deviations and revolute joint clearances
    • Abstract: Publication date: Available online 11 February 2019Source: Aerospace Science and TechnologyAuthor(s): Qiangqiang Zhao, Junkang Guo, Jun Hong, Zhigang Liu Accuracy is of great importance for deployable mechanisms because it greatly determines the performance of the antenna. The main purpose of this paper is to propose a method to analyze the angular errors of the multi-closed-loop deployable mechanism of the planar synthetic-aperture radar antenna considering link deviations and joint clearances. First, a two-link unit is explored to calculate the position errors of the joints due to the manufacturing imperfection only in the assembly process. On this basis, conducting a geometrical analysis, the formulations of the studied angular errors of two planar antenna panels after the introduction of clearances are established. Utilizing the rotatability law of the single-loop linkage, the worst-case scenario associated with a chaotic clearance-induced uncertainty is reduced to a purely geometrical problem. Accordingly, the maximum angular errors caused by these two error sources are obtained in an intuitive manner. Finally, the proposed method is demonstrated by a numerical example of the deployable mechanism and verified by experimental measurements.
  • Robust aerodynamic shape optimization—From a circle to an airfoil
    • Abstract: Publication date: Available online 11 February 2019Source: Aerospace Science and TechnologyAuthor(s): Xiaolong He, Jichao Li, Charles A. Mader, Anil Yildirim, Joaquim R.R.A. Martins Aerodynamic design optimization currently lacks robustness with respect to the starting design and requires trial and error in the flow solver and optimization algorithm settings to get a converged optimal design. We address this issue by developing ways to overcome robustness issues arising from shape parametrization, mesh deformation, and flow solver convergence. Our approach is demonstrated on the Aerodynamic Design Optimization Discussion Group (ADODG) airfoil optimization benchmarks to show the factors that dominate the robustness and efficiency. In the ADODG NACA 0012 benchmark, we address the additional issue of non-unique solutions. In the ADODG RAE 2822 case, we address solver failure due to shock waves, separation, and gradient accuracies due to the frozen turbulence model. Finally, we create a new, challenging aerodynamic shape optimization case that starts with a circle to test the robustness of our aerodynamic shape optimization framework. We use both fixed and adaptive parametrization methods to tackle this problem and show how we can exploit the advantages of adaptive parametrization methods to improve both robustness and efficiency. The combination of flow solver robustness, precision of gradient information, robust mesh deformation, and adaptive parametrization brings us closer to a “push-button” solution for airfoil design.
  • Effects of porosity and thermomechanical loading on free vibration and
           nonlinear dynamic response of functionally graded sandwich shells with
           double curvature
    • Abstract: Publication date: Available online 11 February 2019Source: Aerospace Science and TechnologyAuthor(s): Minh-Chien Trinh, Dinh-Duc Nguyen, Seung-Eock KimRésuméThe fundamental frequencies and nonlinear dynamic responses of functionally graded sandwich shells with double curvature under the influence of thermomechanical loadings and porosities are investigated in this study. Two material models are considered. The continuity requirement of material properties throughout layers are fulfilled by newly introducing refined effects of two porosity types regarding the average of constituent properties weighted by the porosity volume fraction. The first-order shear deformation theory taking the out-of-plane shear deformation into account is employed to obtain the Lagrange equation of motions. The number of primary variables reduces from five to three after introducing the Airy stress function. The system of dynamic governing equations is obtained by utilizing the Bubnov-Galerkin procedure. The natural frequencies are analytically computed by solving eigenvalue problems, and the fundamental frequencies are acquired by further assumptions about the inertial force caused by the shell rotation variables. The nonlinear dynamic responses of the functionally graded spherical, cylindrical, and hyperbolic paraboloid shells under the influence of different geometry configurations, loading conditions, and porosity types and degrees are obtained by applying the fourth-order Runge–Kutta method. The numerical results are presented and verified with available studies in the literature. Although porosities are usually considered material defects weakening the structure performance, this study has proved clearly that porosities stiffen the shell structures to some extent.
  • Investigation of a sliding alula for control augmentation of lifting
           surfaces at high angles of attack
    • Abstract: Publication date: Available online 8 February 2019Source: Aerospace Science and TechnologyAuthor(s): Thomas Linehan, Kamran Mohseni The ability to generate useful control forces on lifting surfaces at high angles of attack is particularly challenging due to boundary layer separation. A miniature collection of feathers on birds termed the alula, appears an intriguing solution to this control problem. Using surface-oil visualizations and direct force and moment measurements, we experimentally investigate the aerodynamics of a model alula(e) affixed to a thin, flat-plate, Image 1 rectangular wing. A critical parameter of the deflected alula considered, is not its orientation relative to the incoming flow, but rather its spanwise distance from the wing tip to which it is oriented. Control forces (lift and rolling moment) are proportional to this distance over a wide range of angles of attack. When centered on the stalled wing, a single alula generates a rolling moment of magnitude comparable to that produced by a conventional trailing-edge flap aileron in an attached-flow condition. Importantly, the wetted area of the alula is one order of magnitude less than the reference flap aileron. The uncharacteristically large control force of the alula stems from its ability to induce and stabilize a vortex that sweeps outboard across the span of the wing towards the wing tip. Changing the distance of the alula from the wing tip, varies the length of this ‘sweeping vortex’ and its associated interactions with the wing. A novel high-angle-of-attack control solution is proposed, the sliding alula, which entails coordinated shifting of two alulae to manipulate the length and asymmetry of stabilized ‘sweeping vortices’ on stalled wings. Results regarding control authority in cross flow and the gust mitigation potential of the sliding alula are also discussed.
  • High-resolution monitoring of aerospace structure using the bifurcation of
           a bistable nonlinear circuit with tunable potential-well depth
    • Abstract: Publication date: Available online 7 February 2019Source: Aerospace Science and TechnologyAuthor(s): Kai Yang, Zhen Zhang, Yanmin Zhang, Hao Huang Monitoring structural parameter and strain of a flight vehicle is important for safety inspection. To detect the nuanced variation of the structural parameter and strain, this study proposes a novel bistable nonlinear circuit with tunable potential-well depth (TPWD), and integrates it with electromechanical transducer (e.g. piezoelectric transducer and strain gauge) to form a high-resolution structural sensor. The TPWD bistable circuit's bifurcation feature can significantly magnify the response's difference before and after the variation of the structural parameter and strain. It is beneficial to detection of the tiny structural variation. The theory of the TPWD bistable circuit and its application in high-resolution structural monitoring is presented. Then, experimental studies of the TPWD bistable circuit and its performance coupled with a structure are performed, which quantitatively and qualitatively verifies the theoretical predictions and the circuit's feasibility for structural detection. It is seen that the circuit's potential well can be tuned to manipulate the bifurcation point, which can be adapted to different excitation levels. Finally, this paper performs investigations of the TPWD bistable circuit for detecting the variation of a wing structure's parameter and strain, respectively. Results show that by means of the TPWD bistable circuit's bifurcation, nuanced variations of both the structural parameter and strain can be evidently detected, leading to high-resolution structural monitoring.
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