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Aerospace Science and Technology
Journal Prestige (SJR): 0.796
Citation Impact (citeScore): 3
Number of Followers: 374  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 1270-9638
Published by Elsevier Homepage  [3183 journals]
  • Structural health monitoring for long-term aircraft storage tanks under
           cryogenic temperature
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(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.
       
  • Aerothermodynamic analysis for deformed membrane of inflatable aeroshell
           in orbital reentry mission
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Yusuke Takahashi, Taiki Koike, Nobuyuki Oshima, Kazuhiko Yamada An inflatable aerodynamic decelerator with a membrane aeroshell is a promising key technology in the reentry, descent, and landing phases of future space transportation. The membrane aeroshell is generally deformed by the in-flight aerodynamic force; however, the effects of the deformation on the aerodynamic heating are unclear. Here, we investigated aerodynamic heating for an inflatable reentry vehicle, Titans, in the hypersonic regime using flow field simulation coupled with structural analysis. Thermochemical nonequilibrium flows around the Titans with a deformed membrane aeroshell were reproduced numerically for an angle of attack (AoA) values between 0° and 40°. The maximum displacements of the membrane aeroshell by deformation at the AoAs of 0° and 40° were 6.7% and 6.6% of the diameter of the Titans, respectively. The difference in heat fluxes between the deformed and rigid shapes was a remarkable 188.8% for a 0° AoA owing to the considerable changes in the front shock wave shape. Meanwhile, it was indicated that membrane deformation at an AoA of 40° insignificantly affected the peak heat flux value on the inflatable torus because the considerable change in the shock wave shape observed for the case of 0° AoA did not occur. It was found that local wrinkles on the membrane aeroshell were formed by deformation, thus causing the heat flux to increase owing to an increase in local temperature gradient on the surface.
       
  • Resistojet thruster with supercapacitor power source – design and
           experimental research
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Jan Kindracki, Przemysław Paszkiewicz, Łukasz Mężyk The paper presents the design of and experimental research on a resistojet thruster with a dedicated power supply system based on supercapacitors. First, the description of two research stands used in different stages of the research is presented. The experimental approach divides the research into two phases. The first part focuses on optimization of the heating chamber, which has a laboratory power supply as a power source. Seven configurations of the chamber – differing in the number of coaxial channels and spiral heaters – are presented together with the results of the optimization process. The second stage describes research into thruster-like conditions with both power supply systems – supercapacitor based and with a laboratory power supply – with the geometry chosen based on the optimization process. A comparison is presented of propulsion parameters for various pre-heating times, power levels and power supply types, with a cold gas system as a reference.
       
  • Computational investigations into heat transfer over a double wedge in
           hypersonic flows
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Diego Expósito, Zeeshan A. Rana Recently developed OpenFOAM application hy2FOAM is employed to predict the aerodynamic heat transfer numerically and compared with the experimental data from the University of Illinois. Mach 7 nitrogen flow at 2.1 MJ/kg stagnation enthalpy, and Mach 7 nitrogen and air flows at 8 MJ/kg stagnation enthalpy over a double wedge geometry have been reproduced numerically assuming chemical and thermal non-equilibrium. Good agreement of mean heat transfer profiles has been observed, although none of the simulations achieved a steady-state. The reattachment heat transfer peak in the high enthalpy air case showed an improved agreement with the experimental data, which is due to the non-equilibrium in the flow field.
       
  • A method of 3D path planning for solar-powered UAV with fixed target and
           solar tracking
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Yu Huang, Jianguo Chen, Honglun Wang, Guofeng Su To harvesting the more net energy, a method of 3D path planning for solar-powered UAV with fixed target and solar tracking has been presented in this paper. However, how to deal with the coupling between UAV motion, mission constraints, energy production, and the energy consumption is the key to 3D path planning for solar-powered UAVs to continuously monitor fixed targets. Hence, in this study, the flight paths of SUAV will be planned on a virtual cylinder surface in 3D space, with the fixed target center as the origin. In order to realize the UAV path planning based on the virtual cylinder surface, firstly, the UAV motion is re-modeled, and the UAV's motion characteristics and force are analyzed based on this model. Then, based on the motion trajectory characteristics in relation to the force balance, the state variables at each waypoint of the UAV are parameterized in the form of the variables to be optimized and their first and second derivatives, and a spline interpolation function is introduced three times to obtain the first and second derivatives of the solution variables. Finally, with reference to the UAV's minimum power flight strategy, the horizontal plane component of the UAV flight speed is set to the minimum power level flight speed, while the yaw rate is solidified and the sun position is assumed to be unchanged for a short period of time to simplify the optimization process and thus obtain the final optimal solution. The simulation experiment shows that unlike the UAV that flies at a fixed height, the UAV herein will climb or descend to seek a more favorable attitude to maximize energy production.
       
  • Fuel injection location studies on pylon-cavity aided jet in supersonic
           crossflow
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Aryadutt Oamjee, Rajesh Sadanandan The current study numerically investigates the effect of fuel injection locations within a pylon-cavity aided Supersonic Combustion Ramjet (SCRAMJET) combustor on mixing enhancement, flame holding, fuel jet penetration and total pressure loss. RANS equations for compressed real gas are solved by coupled, implicit, second-order upwind solver. Two-equation SST model is used for turbulence modelling. The computational model is validated using experimental steady wall pressure data and 2D velocity field. The study uses seven distinct sonic fuel injection location cases of hydrogen fuel through a 1 mm diameter hole along the axis of the test section floor. All cases maintain crossflow of Mach number 2.2. The simulations show that the counter rotating vortex pair within the cavity plays a vital role in fuel dispersion and fuel jet penetration capability. The presence of pylon resulted in an increase of pressure loss by 7%, whereas the influence on total pressure loss due to transverse fuel injection is found to be insignificant. The injection locations within the cavity give around 55% (max) increase in fuel dispersion compared to location upstream of the pylon. Also the cavity floor locations give about 55% - 90% more flammable plume area than the injection from other locations.
       
  • A constrained reduced-order method for fast prediction of steady
           hypersonic flows
    • Abstract: Publication date: August 2019Source: Aerospace Science and Technology, Volume 91Author(s): Changqiang Cao, Chunsheng Nie, Shucheng Pan, Jinsheng Cai, Kun Qu A constrained reduced order model (ROM) based on proper orthogonal decomposition (POD) is proposed to achieve fast and accurate prediction of steady hypersonic flows. The proposed method addresses the convergence issue of the projection-based POD ROM which violates the boundary conditions by using a constrained Gauss-Newton iterative process. The constraints in the iteration are formulated by satisfying the physical boundary conditions. To achieve this, a weighting matrix constructed by an improved Gauss weighting function is adopted to determine the contribution of each cell to the constraint term. The proposed constrained reduced-order method is accelerated by a parallel algorithm based on message passing interface (MPI) and load balancing, which makes the method practical for prediction of complex flows with large memory cost. Fast predictions of hypersonic flows over the two-dimensional cylindrical blunt body and the three-dimensional reentry vehicle using the constrained reduced-order method show that the error is significantly smaller than that of the interpolation-based POD ROM and the projection-based POD ROM. Computing efficiency is increased by 2 ∼ 3 orders of magnitude compared to CFD.
       
  • Safe multi-cluster UAV continuum deformation coordination
    • Abstract: Publication date: August 2019Source: Aerospace Science and Technology, Volume 91Author(s): Hossein Rastgoftar, Ella M. Atkins This paper proposes a paradigm for coordination of multiple unmanned aerial vehicle (UAV) clusters in a shared motion space. UAVs are arranged in a finite number of teams each bounded by a leading triangle. Collective motion of each UAV cluster is managed by a continuum deformation defined by three leaders at the vertices of a leading triangle and followers contained within this triangle. Each triangular cluster can deform substantially to support maneuverability in constrained spaces. This paper specifies necessary conditions to guarantee obstacle avoidance as well as collision avoidance within and across all clusters operating in a shared motion space. Given initial and target configurations, an existing planner (A*) identifies the shortest coordinated leader UAV paths from initial to final configuration in a manner that satisfies safety constraints. An illustrative simulation case study is presented. Continuum deformation containment offers scalability in collision-free UAV motion planning not previously realized in the detect-and-avoid literature. The proposed multi-cluster coordination protocol also extends previous cooperative control to address detect-and-avoid (DAA) given multiple cooperative teams with different destinations.
       
  • Comprehensive assessment of newly-developed slip-jump boundary conditions
           in high-speed rarefied gas flow simulations
    • Abstract: Publication date: August 2019Source: Aerospace Science and Technology, Volume 91Author(s): Nam T.P. Le, Ehsan Roohi, Thoai N. Tran In this paper we numerically evaluate the recently developed Aoki et al. slip and jump conditions in high-speed rarefied gas flows for the first time. These slip and jump conditions are developed to be employed with the Navier–Stokes–Fourier equations. They were derived based on the Boltzmann equation with the first order Chapman–Enskog solution, and the analysis of the Knudsen layer. Four aerodynamic configurations are selected for a comprehensive evaluation of these conditions such as sharp-leading-edge flat plate, vertical plate, wedge and circular cylinder in cross-flow with the Knudsen number varying from 0.004 to 0.07, and argon as the working gas. The simulation results using the Aoki et al. boundary conditions show suitable agreement with the DSMC data for slip velocity and surface gas temperature. The accuracy of these boundary conditions is superior to the conventional Maxwell, Smoluchowski and Le boundary conditions.
       
  • Numerical study of cellular structure in detonation of a stoichiometric
           mixture of vapor JP-10 in air using a quasi-detailed chemical kinetic
           model
    • Abstract: Publication date: August 2019Source: Aerospace Science and Technology, Volume 91Author(s): Lijuan Liu, Qi Zhang Regularity of cellular structure that forms in the detonation of a stoichiometric gaseous JP-10/air mixture was investigated using a quasi-detailed chemical kinetic model consisting of 275 elementary reactions among 54 chemical species. Two-dimensional numerical simulations were conducted for the cases of a wave front structure that develops in a channel following strong initial perturbations. The effect that the grid resolution has on the detonation structure was examined. With a low grid resolution, relatively regular detonation structure was achieved, but an irregular structure developed at moderately high resolutions. The formation process of irregular structure can be divided into three states: overdriven state, regular state and irregular state. The detonation cell size obtained from numerical results about 50 mm is similar with the experimental results. Meanwhile, new modes appear earlier in wider domains than in narrower domains. This preliminary work provides references for our future research on detonation limits and fuel selection in cloud explosion.
       
  • Numerical and experimental investigation on hybrid rocket motor with
           two-hole segmented rotation grain
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Hui Tian, Lingfei He, Hao Zhu, Pengfei Wang, Xu Xu This paper investigates the effect of rotation angles on the fuel regression rate and the motor combustion efficiency in hybrid rocket motor with two-hole segmented grains. In this research, the propellant combination of 90% hydrogen peroxide (H2O2) and polyethylene (PE) is adopted, and two two-hole grains with the same configuration are used to conduct both numerical and experimental tests. A 3D simulation model is established to obtain combustion efficiency, fuel regression rate, temperature distribution and species mass fraction distribution. Meanwhile, in order to achieve the motor combustion efficiency and the average regression rate, 20 firing tests were conducted on a lab-scale hybrid rocket motor. The numerical and experimental results agree well and demonstrate that the rotation of after-section grain has no influence on the fore-section grain regression rate, while the after-section grain regression rate arises, compared with the no rotation case. Regression rates are fitted with the following empirical equation, r˙=aGoxn(1−e−D‾pm), in which the regression rate coefficients a and n of after-section grain changed with the rotation angle. The motor combustion efficiency is higher than the base operation condition. It increases gradually with the increase of rotation angle, and the increasing rate becomes slower. The combustion efficiency can reach to the maximum when the rotation angle is 45°, after which it drops slowly, and the rate of descent is slower than that of ascent.
       
  • Multidisciplinary design optimization of long-range slender guided rockets
           considering aeroelasticity and subsidiary loads
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Zhao Wei, Teng Long, Renhe Shi, Yifan Tang, Huaijian Li Due to the insufficient rigidity feature and long-range requirement, it is crucial to consider the aeroelasticity and subsidiary loads caused by the Earth rotation and fuel consumption when designing long-range slender guided rockets (LRSGRs). As a typical multidisciplinary design optimization (MDO) problem, the design optimization of LRSGRs confronts two critical challenges, i.e., accurate multidisciplinary modeling and efficient global optimization. To address the challenges, a novel MDO framework including MDO problem definition, multidisciplinary modeling, and metamodel-based optimizer is developed for LRSGR design. The LRSGR MDO problem is formulated to minimize the total mass subject to a number of practical engineering constraints such as bending mode frequencies, miss distance, and fall angle. Several disciplinary models including structure, aerodynamics, propulsion, mass, aeroelasticity, guidance control, and trajectory are established. To enhance the analysis accuracy, structural finite element analysis (FEA), three-channel autopilot, and high-fidelity trajectory models are adopted. In the aeroelasticity model, the unsteady aerodynamic loads are calculated by slender body theory and aerodynamic derivative method. The subsidiary loads including subsidiary Coriolis force, centrifugal inertial force, Coriolis force, and subsidiary Coriolis moment are incorporated in the trajectory model of LRSGRs. Since structural finite element, aeroelasticity, and trajectory models are computationally expensive (about 1.8 hours for one trial of system analysis on a well-equipped workstation), an adaptive radial basis function metamodel-based optimizer is integrated in the framework to solve the LRSGR MDO problem with moderate computational cost. The total mass of the studied LRSGR is decreased by 88 kg (i.e., 14% of the total mass) after optimization, which demonstrates the effectiveness and practicability of the proposed MDO framework for LRSGRs.
       
  • Performance analysis of dual-duct rotating detonation aero-turbine engine
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Zifei Ji, Huiqiang Zhang, Bing Wang A configuration for a dual-duct rotating detonation aero-turbine engine (DRDATE) is proposed. With the isolator and mixer arranged upstream and downstream from the rotating detonation combustor (RDC) respectively, the compatibility between turbomachinery and RDC can be realized. The conventional single annular RDC is replaced with a multi-annular RDC to expand the stable operation range of the RDC. A low-order analytical model of the rotating detonation process is presented, and comparisons between the results calculated by this model and the CFD solvers show reasonable agreement. Thereafter, a performance simulation model of the DRDATE is established, and further the variations in the overall performances with design parameters under three different flight conditions are investigated. The results demonstrate that, there exists an optimum compressor pressure ratio πopt that maximizes the specific thrust and an optimum pressure ratio πopt′ that maximizes the thermal efficiency for the DRDATE. With an increase in the compressor and turbine polytropic efficiency and turbine inlet temperature, both πopt and πopt′ increase monotonically. Comparisons between the DRDATE and conventional turbine engine reveal that, the former exhibits a major improvement in overall performance at low compressor pressure ratios, while the improvement decreases continuously with an increase in the pressure ratio. Moreover, as the turbine inlet temperature increases, the specific thrust improvement increases and the fuel consumption performance improvement decreases monotonically.
       
  • Implementation and verification of gust modeling in an open-source flow
           solver
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Mehdi Ghoreyshi, Adam Jirasek, Tyler Miller, Michael Nuzum, Roger Greenwood A gust simulation and modeling capability will support the aircraft certification process and therefore reduces the development cost. Two different gust modeling techniques are considered in this article: 1) a user-defined boundary condition available in the commercial flow solver of Cobalt and 2) a field-velocity method implemented in the open-source flow solver of FlowPsi. Specifically, this article describes implementation of gust models in the FlowPsi code including a new time-stepping scheme and a new grid-motion capability that could be used in the simulation of responding vehicles to wind gusts. In more detail, the three stage time-stepping scheme, as here developed, provides much better convergence properties than the Crank–Nicolson scheme available in the code. The grid motion using an external code, as here implemented, is a part of larger development plans to make FlowPsi communicate with external processes such as a finite element code. In addition, the FlowPsi code uses a field-velocity approach to simulate the wind gust which offers a reduced computational expense for simulating wind gust responses when compared to Cobalt. Instead of propagating the gust perturbation from the far inflow boundary as used in Cobalt, an artificial velocity profile is imposed on the grid cells to induce the gust effects on the vehicle. This article aims to determine the accuracy of FlowPsi in predicting gust responses when compared to Cobalt which has already been verified to produce accurate results. The two methods in different codes are compared for a number of gust profiles such as one-minus-cosine, step (sharp-edged), and a random gust profile. FlowPsi shows close convergence and agreement to Cobalt predictions for smooth gust profiles such as the one-minus-cosine, however, the code produces a numerical oscillation pattern at low free-stream Mach numbers when modeling gusts with a rapid change of velocity such as the step gust.
       
  • Use of Blended Blade and End Wall method in compressor cascades:
           Definition and mechanism comparisons
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Jiabin Li, Xin Li, Lucheng Ji, Weilin Yi, Ling Zhou Modern turbomachinery design uses the endwall treatment method to prevent the corner flow from separating. This study presents the Blended Blade and End Wall (BBEW) method, a passive endwall treatment method, and draws the distinction between the flow mechanisms of BBEW and fillet cascades.The first part of the study conducts an experimental investigation on the flow mechanism of a BBEW cascade. The results show that BBEW technology can stretch the spanwise area of vortices, thus reducing losses at a lower spanwise position, slightly increasing losses at a higher position, and improving the corner separation.In the second part of the study, comparisons of the flow mechanisms of the BBEW and fillet cascades are conducted using numerical methods. The results show that the fillet cascades induce a pressure difference at nearly the full range of the chord. BBEW cascade induces a pressure difference at a certain chordwise position, and the magnitude of the pressure difference is larger. The pressure difference tends to move the low-energy fluid away from the corner and reduce the cross flow accumulation, which reduces the endwall losses.
       
  • Application of deep learning based multi-fidelity surrogate model to
           robust aerodynamic design optimization
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Jun Tao, Gang Sun In the present work, a multi-fidelity surrogate-based optimization framework is proposed, and then applied to the robust optimizations for airfoil and wing under uncertainty of Mach number. DBN (deep belief network) is employed as the low-fidelity model, and the k-step contrastive divergence algorithm is used for training the network. By virtue of the well trained DBN model and high-fidelity data, a linear regression multi-fidelity surrogate model is established. Verification results indicate that the multi-fidelity surrogate model obtains more accurate predictions than the DBN model and is highly reliable as a prediction model. The multi-fidelity surrogate model is embedded into an improved PSO (particle swarm optimization) algorithm framework, and is updated in each iteration of the robust optimization processes for both airfoil and wing. Comparisons between multi-fidelity surrogate predictions and CFD results indicate that, the multi-fidelity surrogate predictions tend to approach the CFD results as the iteration number increases. The robust optimization results of airfoil and wing demonstrate that, the multi-fidelity surrogate model performs very well as a prediction model, and improves the optimization efficiency obviously.
       
  • A multivariable adaptive control scheme for automatic carrier landing of
           UAV
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Ziyang Zhen, Gang Tao, Chaojun Yu, Yixuan Xue This paper studies a multivariable model reference adaptive control (MRAC) scheme for the automatic carrier-landing control problem of unmanned aerial vehicles (UAVs) with system dynamics of nonlinearity, multivariable coupling and parametric uncertainty. A complete automatic carrier landing system (ACLS) for carrier-based UAVs is developed, which consists of a guidance subsystem and a flight control subsystem. The MRAC scheme is based on a state feedback for output tracking framework with relaxed design conditions, which guarantees the reference glide slope tracking. Simulation results of a nonlinear UAV model demonstrate that the multivariable MRAC based ACLS has a better carrier-landing performance than a fixed control based ACLS.
       
  • Combustion characteristics of a novel design of solid-fuel ramjet motor
           with swirl flow
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Omer Musa, Chen Xiong, Li Weixuan, Liao Wenhe The coupling of high speed to wide range is the key problem in the military applications. Solid-fuel ramjet (SFRJ) has drawn much interest since the beginning of the last century which could be used to extend the range and speed of the missiles as in ramjet-powered missiles. Nevertheless, solid-fuel ramjet suffers from low regression rate of the solid fuel. To enhance the regression rate of the solid-fuel ramjet, in this paper, a new design is proposed and numerically investigated. The new design uses two solid fuels with keeping the simplicity in the design of the classic solid-fuel ramjet. For the simulations, an in-house CFD code has been developed to solve Reynolds-averaged Navier–Stokes equations of turbulent, reacting, unsteady, and swirl flow. Simulations are carried out for the proposed and classic designs with and without swirl flow. The results are compared with the classic design for the same configuration and flow conditions. It is shown that the new design has improved the regression rate, reactants mixing degree, and performance of SFRJ. The proposed design offered two diffusion flames at which the new flame started from the inlet of the combustion chamber in the swirl flow case and near to the combustor's end for non-swirl case.
       
  • Multidisciplinary design optimization of aircraft wing using commercial
           software integration
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Abdelkader Benaouali, Stanisław Kachel In this paper, a fully automated framework dedicated to the high-fidelity multidisciplinary design optimization of aircraft wing is developed. This design framework integrates a set of popular commercial software using their programming/scripting capabilities. It goes through geometric modeling in SIEMENS NX, aerodynamic meshing in ICEM CFD, flow solution using ANSYS FLUENT, structural finite element modeling in MSC.PATRAN and structural sizing in MSC.NASTRAN. By adopting a parametric modeling methodology, the structural and aerodynamic metrics reflecting the wing performance can be evaluated given a description of its shape and dimensions. In order to overcome the high cost of simulation models and allow the efficient solution of high-fidelity optimization problems, a surrogate-based optimization strategy is adopted. The reliability of the proposed approach is investigated through its application to the design of a high-speed passenger aircraft wing. The optimization objective is to maximize the aircraft range, given by the Breguet equation, while maintaining the lift coefficient and the structural safety. The case study results in a 8.9% increase in the range by considering shape and structural design variables.
       
  • Observer-based linear parameter varying control design with unmeasurable
           varying parameters under sensor faults for quad-tilt rotor unmanned aerial
           vehicle
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Zhong Liu, Didier Theilliol, Liying Yang, Yuqing He, Jianda Han The tilt rotor unmanned aerial vehicle (TRUAV) has attracted considerable attention in recent years, and its control design, particularly during the transition procedure, is still a challenge. Aiming at this difficulty and by introducing virtual control inputs, a nonlinear model of the TRUAV as an equivalent linear parameter varying (LPV) system for controller design of these virtual values is considered in this study. With this idea, one of the main contributions of this paper is that the stability of TRUAV during the transition procedure could be ensured asymptotically in theory. An observer-based LPV control with fault-tolerant ability is synthesized considering the unmeasurable varying parameters of this LPV controlled plant and the underlying sensor faults. The second main contribution is to analyze the closed-loop stability with unmeasurable features and bias sensor faults using parameter-dependent Lyapunov matrices. The performances of proposed control method for TRUAV under sensor faults are illustrated with numerical results.
       
  • Axially functionally graded beams and panels in supersonic airflow and
           their excellent capability for passive flutter suppression
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Zhiguang Song, Yiyu Chen, Zhenyu Li, Jincheng Sha, Fengming Li Mode localization in the airflow direction can be observed in aeroelastic structures. On the contrary, can the aeroelastic stability of the structure be enhanced if the geometric sizes and material properties vary along the airflow direction' The present study mainly solves this problem. The mode localization phenomenon in panel flutter in supersonic airflow is displayed. Three arbitrary thickness functions according to the flutter mode are taken into account. Their panel flutter behaviors show that although the consistency between the variation of thickness and flutter mode can increase the flutter bound slightly, the increments are limited. Consequently, a novel strategy for passive control of the panel flutter is proposed by the optimal axially functionally graded (AFG) design of the panel. By investigating the sensitivity of each element in the aerodynamic stiffness matrix to the aeroelastic stability of the structure, the optimal thickness and Young's modulus functions are given out. Simulation results show that the optimal AFG design in this study can suppress the flutter essentially. It can increase the flutter bound of the structure by changing the flutter modes rather than only makes a slight extension on the original basis. Moreover, the designed thickness and Young's modulus are reasonable and applicable.
       
  • Robust optimization for a wing at drag divergence Mach number based on an
           improved PSO algorithm
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Jun Tao, Gang Sun, Xinyu Wang, Liqiang Guo In this study, an improved PSO (particle swarm optimization) algorithm is proposed and applied to the robust optimization of a wing at drag divergence Mach number. In order to reduce the number of design variables, a six-order CST (class/shape function transformation) method is employed for airfoil parameterization. For the purpose of improving the optimization efficiency, Delaunay graph mapping method is adopted for mesh deformation in each iteration of the airfoil optimization, and NURBS (non-uniform rational B-splines)-FFD (free-form deformation) method is employed for mesh deformation in each iteration of the wing optimization. For improving the standard PSO algorithm, CVTs (centroidal Voronoi tessellations) method is introduced to generate original positions of the particles more dispersedly, a second-order oscillating scheme is used and an FDR (fitness distance ratio) item is added for updating velocities and positions of the particles. By virtue of the improved PSO algorithm, single point optimization and robust optimization are conducted for both airfoil and wing. The results indicate that, comparing with the single point optimizations, the robust optimizations not only reduce drag coefficients of the airfoil and the wing at cruise Mach numbers, but also attenuate the drag increments as the Mach number increases up to drag divergence Mach numbers.
       
  • Initiation of oblique detonation waves induced by a blunt wedge in
           stoichiometric hydrogen-air mixtures
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Yishen Fang, Zijian Zhang, Zongmin Hu, Xi Deng Two-dimensional, oblique detonation waves (ODWs) in a stoichiometric hydrogen-air mixture are simulated with the reactive Euler equations using a detailed chemical reaction model. This study focuses on blunt wedge induced ODWs, which are not only influenced by inflow parameters but also the size of the blunt body. With the inflow parameters of flight altitude of 30 km and flight Mach number M0 of 8-10, the numerical results demonstrate that the blunt wedge is crucial to initiate the ODW. In the case of M0=10, the straight wedge without the blunt forebody can initiate the detonation. However, decreasing M0 causes the failure of initiation, which can be compensated by increasing the radius R0 of the blunt forebody. By adjusting R0, two initiation procedures are observed and distinguished: one is the wedge-induced initiation and the other is the blunt forebody-induced initiation. Although both have been independently studied before, in this study, their coexistence is demonstrated, and the mechanism is analyzed for the first time. A theoretical analysis based on the classic initiation theory is performed to elucidate the initiation mechanism, giving a good agreement between the critical radius with numerical results.
       
  • Investigation of pitch damping derivatives for the Standard Dynamic Model
           at high angles of attack using neural network
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Massoud Tatar, Mehran Masdari Numerical investigations of flow field around the Standard Dynamic Model (SDM) are performed using computational fluid dynamics approach. Initially, the static SDM is studied at various angles of attack up to 70∘ and an agreement between normal force and pitching moment predictions with experimental data is ensured, thanks to the polyhedral grids. Subsequently, the response of the SDM under single frequency sinusoidal pitching motions is computed and the associated pitching moment coefficient damping is obtained using two methods of classical Fourier coefficients and multilayer perceptron (MLP) artificial neural network. The results are compared to published experimental values. In the final stage, frequency sweep sinusoidal excitation in pitch axis is conducted with 30∘ amplitude in transonic flow and the MLP is exploited to calculate variable stability derivatives. It is observed that damping derivatives are highly dependent on both amplitude and frequency of oscillation. Also, an increase in the frequency of motion lowers the pitching moment damping. As the motion frequency rises, the pitching moment amplitude increment is seen to be greater than that of normal force. Polyhedral mesh as well as overset grid technique are adopted in flow field computations, leading to high fidelity of numerical simulations.
       
  • Station-keeping performance analysis for high altitude balloon with
           altitude control system
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Huafei Du, Mingyun Lv, Jun Li, Weiyu Zhu, Lanchuan Zhang, Yifei Wu Station-keeping endurance of high altitude balloon is the foundation of the assignments of environmental monitoring and communicational relaying. Exploiting the natural wind-field that varies with altitude by the altitude control system to extend the station-keeping endurance is proposed in this paper. A Matlab program is developed based on the theoretical model to simulate the station-keeping performance of the balloon in the real wind field. The trajectories of the balloon in different wind fields and the states of the balloon caused by the venting and pumping processes are discussed in detail. The results show that with the altitude control system it is possible to retain the balloon within the designated district for few days to a week. This can serve as a guideline for the design and initial flight tests of the serviceable high altitude balloon.
       
  • Mode multigrid - A novel convergence acceleration method
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Yilang Liu, Weiwei Zhang, Jiaqing Kou This paper proposes a mode multigrid (MMG) method, and applies it to accelerate the convergence of the steady state flow on unstructured grids. The dynamic mode decomposition (DMD) method is used to analyze the convergence process of steady flow field according to the solution vectors from the previous time steps. Unlike the traditional multigrid method, we project the flowfield solutions from the physical space into the modal space, and truncate all the high-frequency modes but only the first-order mode are retained based on the DMD analysis. The real solutions in the physical space can be obtained simply by the inverse transformation from the modal space. The developed MMG method ingeniously avoids the complicated process of coarsening computational mesh, and does not need to make any change for the grid in physical space. Therefore, it is very convenient to be applied to any numerical schemes with just a little change for the flow solver, which is also suitable for unstructured grids and easy for parallel computing. Several typical test cases have been used to verify the effectiveness of the proposed method, which demonstrates that the MMG can dramatically reduce the number of iterative steps for the different mesh types, different accuracy of spatial discretization and different time-marching schemes. The method is 3 to 6 times faster than the baseline method while ensuring the computational accuracy.
       
  • Stability and ground experiments of a spinning triangular tethered
           satellite formation on a low earth orbit
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): B.S. Yu, Z. Huang, L.L. Geng, D.P. Jin This paper studies the spinning stability of a triangular tethered satellite formation that flies on a low earth orbit. The spinning around the center of mass at a constant angular rate gives rise to a periodic motion in the non-inertial orbital frame. Floquet theory is used to analyze the stability of the periodic motion. A dynamic similarity between the on-orbit dynamics and ground experimental models is built to construct an equivalent ground experiment to verify the stability analysis. The analytical and experimental results show that a stable periodic motion can be guaranteed if the spinning angular rate of the system exceeds a critical value.
       
  • Fast task allocation for heterogeneous unmanned aerial vehicles through
           reinforcement learning
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Xinyi Zhao, Qun Zong, Bailing Tian, Boyuan Zhang, Ming You A task allocation problem for heterogeneous unmanned aerial vehicles (UAVs) in the presence of environment uncertainty is studied in this paper. Generally, the process of finding efficient allocation scheme can be computationally prohibitive. This work presents a Q-learning based fast task allocation (FTA) algorithm through neural network approximation and prioritized experience replay, which effectively offloads the online computation to an offline learning procedure. Specifically, the proposed approach develops a Q network that encodes the allocation rules. The Q network not only considers the effect of environment uncertainty, but also is capable of handling total different tasks. Comparison simulations are provided to show the efficiency of the proposed algorithm.
       
  • An efficient method combining adaptive Kriging and fuzzy simulation for
           estimating failure credibility
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Chunyan Ling, Zhenzhou Lu, Kaixuan Feng The failure credibility can be used to measure the safety level of the structure under the fuzzy inputs, but the computational efficiency for estimating the failure credibility is still a challenge. A novel method by combining the adaptive Kriging with fuzzy simulation (AK-FS) is proposed to efficiently estimate the failure credibility. The proposed method firstly employs the FS to transform the estimation of failure credibility into a classification problem, which can be viewed as a bi-level strategy. In the inner loop, a Kriging model for the actual complicated performance function is actively trained by U-learning function in the sample pool generated by FS until the convergent condition is satisfied, on which the samples are divided into failure group and safety group by the well-trained Kriging model instead of the actual performance function in the outer loop. Finally, the failure credibility is obtained by respectively searching the maximum joint membership degrees of the samples in these two groups. The proposed AK-FS method only evaluates the actual performance function in the process for constructing the Kriging model. Since the U-learning function can iteratively construct a sufficiently accurate Kriging model with model evaluations as little as possible, thus the failure credibility can be efficiently and accurately estimated by the proposed AK-FS method. The advantages of the proposed AK-FS method are demonstrated by several examples.
       
  • Preliminary design of multirotor UAVs with tilted-rotors for improved
           disturbance rejection capability
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Z.J. Chen, K.A. Stol, P.J. Richards The ability of multirotor UAVs to hold position accurately in a turbulent wind environment is important for a variety of commercial applications. In this work, a tilted-rotor configuration which offers the potential for improved disturbance rejection capability over traditional multirotor designs is presented, along with a design process to select optimal design parameters and components for a particular payload and hover time. The design process integrates models of UAV components and performance based off manufacturer data, and is applied to a range of payloads to explore resulting designs and trends. From the components considered used in this work, the tilted-rotor concept is found to provide the most benefit at payloads under 10 kg, and negligible benefit at payloads above 12 kg.
       
  • Aerodynamic and aeroelastic uncertainty quantification of NATO STO AVT-251
           unmanned combat aerial vehicle
    • Abstract: Publication date: August 2019Source: Aerospace Science and Technology, Volume 91Author(s): Andrea Da Ronch, Jernej Drofelnik, Michel P.C. van Rooij, Johan C. Kok, Marco Panzeri, Arne Voß Turbulence models based on Reynolds-averaged Navier-Stokes (RANS) equations remain the workhorse in the computation of high Reynolds-number wall-bounded flows. While these methods have been deployed to design the configuration developed within the NATO STO AVT-251 Task Group, their deficiencies in modelling complex flows are well-documented. However, an understanding of the sources of errors and uncertainties in RANS solvers, arising for example from different numerical schemes and flow modelling techniques, is missing to date. The aim of this work is to establish and quantify the impact that epistemic uncertainties within RANS solvers have on the aerodynamic and aeroelastic response of the combat aerial vehicle. This will produce a range of all possible values of interest due to the inherent uncertainty of RANS solvers, which is expected to be highly dependent on the flow conditions and geometry configuration. This information, in turn, is used to establish the robustness of the AVT-251 design and its performance metrics considering a high-g pull-up manoeuvre used for structural sizing. It is found that the static aeroelastic analysis without aerodynamic uncertainty (deterministic analysis) under predicted the largest generalised force, with an immediate consequence on the structural design.
       
  • Fractional order MIMO controllers for robust performance of airplane
           longitudinal motion
    • Abstract: Publication date: August 2019Source: Aerospace Science and Technology, Volume 91Author(s): Reza Mohsenipour, Mohsen Fathi Jegarkandi This paper presents fractional order multi-input multi-output (MIMO) controllers for the robust performance of airplane longitudinal motion. A novel necessary and sufficient criterion is offered by using the value set concept to analyze the robust performance of fractional order MIMO uncertain systems based on the location of the characteristic equation roots. The criterion is applicable to all linear time-invariant systems of commensurate and incommensurate orders with complex coefficients. The obtained results are applied to an uncertain linear model of a business airplane to improve the robust performance of its longitudinal motion by decentralized MIMO output feedback and MIMO state feedback controllers. Numerical simulations are conducted to confirm the effectiveness of the presented criterion and controllers.
       
  • Nonlinear reentry guidance law guaranteeing convergence before attainment
           of desired line-of-sight range
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Chunwang Jiang, Guofeng Zhou, Changsheng Gao, Biao Yang, Wuxing Jing Nonlinear guidance law guaranteeing convergence before attainment of desired line-of-sight (LOS) range in the background of reentry vehicle attacking a moving target on the ground is studied in this paper. Initially, a novel guidance model is established in which LOS range is treated as an independent variable, describing the relative motion between the vehicle and the target. The guidance model includes two differential equations that describe LOS's pitch and yaw motions in which the pitch motion is separately decoupled. Subsequently, guidance laws with disturbance suppression of the pitch and yaw motion of LOS are designed separately, and presented in the form of normal overload. Compared with traditional guidance laws, the proposed one guarantees that LOS angular rate converges to zero before LOS range decreases to the desired value. Finally, the correctness and validity of the guidance model and guidance law are verified by numerical simulation. The guidance model and guidance law proposed in this paper provide a new way for the design of fast convergent guidance law.
       
  • Global smooth sliding mode controller for flexible air-breathing
           hypersonic vehicle with actuator faults
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Yibo Ding, Xiaogang Wang, Yuliang Bai, Naigang Cui A global smooth sliding mode controller (GSSMC) is proposed for flexible air-breathing hypersonic vehicle (FAHV) under actuator faults and parametric uncertainties, consisting of global fast finite-time integral sliding surface (GFFIS), generalized smooth second-order sliding mode reaching law (GSRL) and smooth fixed-time observer. Firstly, nonlinear control-oriented model of FAHV is processed using input/output feedback linearization with flexible effects and actuator faults modeling as lumped matched disturbances. Secondly, a GFFIS is established to ensure finite-time convergence of states without singularity based on a newly proposed fast finite-time high-order regulator (FFR). The FFR is improved from standard finite-time high-order regulator via dilation rescaling, which can accelerate response speed avoiding complicated parameters selection. Meanwhile, GFFIS can eliminate initial reaching phase to enhance robustness of system due to characteristic of global convergence. Thirdly, a GSRL is presented to ensure finite-time convergence of sliding mode vector and its derivative without chattering based on a generalized smooth second-order sliding mode control algorithm, the stability and finite convergence time of which is analyzed via Lyapunov criteria in detail. Then, a smooth fixed-time observer is applied to estimate lumped disturbances in fixed time and avoid effects of parametric uncertainties. With the three components, GSSMC can drive FAHV subject to actuator faults and parametric uncertainties to follow desired values in finite time with smooth control signals. Ultimately, three sets of simulations are performed to verify the effectiveness of the methods proposed.
       
  • A velocity-free adaptive RISE-based trajectory tracking approach for
           quadrotors with desired model compensation
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Hongliang Wang, Xingling Shao, Jie Li, Jun Liu This paper presents a velocity-free desired compensation trajectory tracking strategy integrated with adaptive robust integral of the sign of the error (RISE) feedback mechanism for quadrotors concerning parametric uncertainties and external disturbances. The original cascaded dynamics of quadrotor is derived in a strict form with additive disturbances. Then, the adaptive RISE feedback controllers are respectively constructed in position and attitude loop, in which the control gains in RISE terms are adaptively updated online to ensure the robustness against uncertainties. In addition, to alleviate the measurement noise effect arising from the actual velocity signals, the velocity states in the model-based feedforward control are replaced with their desired values, then the desired compensation adaptive RISE controllers that depend on the desired trajectory and output tracking errors are synthesized, where the design conservatism on selecting the control gain in RISE is eliminated without knowing the prior bound of uncertainties, and enhanced performance robustness is also retained in the absence of velocity information. It is shown via Lyapunov analysis that the proposed method can guarantee the tracking errors to converge to the origin with asymptotic performance despite of bounded disturbances. The effectiveness and superiority of proposed method are validated through extensive simulations and comparisons.
       
  • Turbulent and transitional flow around the AVT-183 diamond wing
    • Abstract: Publication date: Available online 21 June 2019Source: Aerospace Science and TechnologyAuthor(s): Andrei Buzica, Christian Breitsamter For their low-radar signature, diamond-wing configurations with blunt leading edges recently attract attention for unmanned applications. For this purpose, the AVT-183 generic configuration is investigated both experimentally and numerically. The flow field is complex, exhibiting both vortical structures and attached flow regions. Despite numerous investigations of the fully turbulent case, little research focuses on the naturally transitioning boundary layer. This paper discusses the transitional and fully turbulent case comparatively. In addition, transient and steady flow-field measurements and oil-flow visualization complete the aerodynamic data set of the AVT-183 diamond-wing model. The separation phenomenon is discussed in detail for both laminar and turbulent leading-edge separation. Low and high frequency fluctuations correlate with the primary and the secondary separation processes, respectively, which are also analyzed.
       
  • Unscented Kalman filter state estimation for manipulating unmanned aerial
           vehicles
    • Abstract: Publication date: Available online 18 June 2019Source: Aerospace Science and TechnologyAuthor(s): H. Bonyan Khamseh, S. Ghorbani, F. Janabi-Sharifi Manipulating unmanned aerial vehicles (MUAVs) are aerial robots equipped with a mechanism to physically interact with the environment. State estimation of such robots is a challenging problem due to inherent couplings, nonlinearities and uncertainties of MUAV complex dynamics and, therefore, popular algorithms such as extended Kalman filter may not be applicable. With the above considerations, this paper formulates two variants of the unscented Kalman filter using (i) general and (ii) spherical unscented transform to address state estimation problem in MUAVs. In order to examine the effect of estimation quality on overall control performance, first the coupled dynamics of a quadcopter endowed with a robotic manipulator is presented. Next, a linear–quadratic–Gaussian (LQG) control is designed to achieve simultaneous control of the quadcopter and its manipulator. Then, the performance of each unscented Kalman filter algorithm is compared with that of extended Kalman filter in the context of estimation accuracy, overall control performance, and algorithm execution time. Additionally, sensitivity of the proposed approaches to increasing noise levels and total loss of sensory data are examined.
       
  • Investigation for electro-thermo-mechanical vibration of nanocomposite
           cylindrical shells with an internal fluid flow
    • Abstract: Publication date: Available online 18 June 2019Source: Aerospace Science and TechnologyAuthor(s): Dinh Gia Ninh, Nguyen Duc Tien In the present, based on geometrical nonlinearity in von Karman sense and classical thin shell theory (CLT), the vibrational analyses of conveying-fluid functionally graded carbon nanotube reinforced composite (FG-CNTRC) cylindrical shells with piezoelectric layers in thermal environment are taken into account by an analytical approach. The uniform and functionally graded CNTs are used to reinforce through the thickness of the shell. The fluid flow in the shell is mentioned as non-viscous, incompressible, isentropic and irrotational. Furthermore, piezoelectric layers are bonded onto the outer surface of the cylindrical shell to act as an actuator. The equations of motion are derived by using the CLT, von Karman nonlinearity theory, the fluid velocity potential and then solved by Galerkin's technique. Moreover, the fourth-order Runge-Kutta method is used to resolve the differential equations. The fundamental frequencies, responses of nonlinear vibration as time histories and bifurcation diagram are studied in this paper. Furthermore, the effects of CNT volume fraction, piezoelectric actuator, thermal environment, geometrical parameters, elastic medium and the fluid flow velocity are carefully analyzed. The obtained results that are validated with those of other studies can be used as benchmark solutions for an analytical approach serving in further research.
       
  • Analytical frequency response solution for composite plates embedding
           viscoelastic layers
    • Abstract: Publication date: Available online 18 June 2019Source: Aerospace Science and TechnologyAuthor(s): A. Alaimo, C. Orlando, S. Valvano In this work analytical damped free-vibration and frequency response solutions are obtained for the analysis of composite plates structures embedding viscoelastic layers. On the basis of the Principle of Virtual Displacements, Layer-Wise models related to linear up to fourth order variations of the unknown variables in the thickness direction are treated. Analytical solutions using the Navier procedure are presented for the analysis of isotropic, cross-ply composite and simply-supported plate structures. The modelization of multilayered structure materials takes into account the composite material properties and the frequency dependence of the viscoelastic material. Various external loads are considered: closed form solution for bi-sinusoidal pressure, constant distributed pressure and concentrated loads. Several analyses are carried out to validate and demonstrate the accuracy and efficiency of the present formulation for the study of viscoelastic plates, taking into account different lamination sequences and different plate aspect ratios.
       
  • Multi-failure probabilistic design for turbine bladed disks using neural
           network regression with distributed collaborative strategy
    • Abstract: Publication date: Available online 18 June 2019Source: Aerospace Science and TechnologyAuthor(s): Lu-Kai Song, Guang-Chen Bai, Cheng-Wei Fei The performance and reliability of aircraft engine are seriously affected by multiple failures induced by multi-physical loads. Multi-failure probabilistic design is an effective measure to estimate the multi-failure response traits and quantify the multi-failure risk for the improvement of component reliability. In this paper, we propose a neural network regression-distributed collaborative strategy (NNR-DCS) based on a developed two-step error control technique, to improve the efficiency and accuracy of multi-failure probabilistic analysis. We firstly mathematically model NNR-DCS and then introduce the corresponding multi-failure probabilistic framework. With respect to various failure modes such as deformation failure, stress failure and strain failure, the multi-failure probabilistic analysis of a turbine bladed disk is conducted to evaluate the proposed method. From this simulation, we gain the probabilistic distribution features, reliability degree and sensitivity degree of each failure mode and overall failure modes on turbine bladed disk, which provides a useful reference for improving the reliability and performance of aircraft engine. The comparison of methods (Monte Carlo method, RSM, DCRSM, DCFRM, NNR and NNR-DCS) shows that the proposed NNR-DCS holds high efficiency and accuracy for multi-failure probabilistic analysis. The efforts of this study offer an effective way for multi-failure evaluation from a probabilistic perspective and shed light on the multi-objective reliability-based design optimization of complex structures besides turbine bladed disk.
       
  • A novel three-axis attitude stabilization method using in-plane internal
           mass-shifting
    • Abstract: Publication date: Available online 14 June 2019Source: Aerospace Science and TechnologyAuthor(s): Liang He, Xiaoqian Chen, Krishna Dev Kumar, Tao Sheng, Chengfei Yue This paper investigates the feasibility of using four movable masses distributed in a plane to realize three-axis attitude stabilization. With the proposed distribution and moving the paired masses in the same or opposite direction simultaneously, aerodynamic torque or internal momentum exchange torque is acted on the spacecraft. Then the spacecraft is gradually stabilized under a three-phase control strategy. Each of the three control phases is designed based on optimal control, nonlinear and sliding mode control, respectively. The stability of the proposed controller is given. Furthermore, numerical simulation matches the analytical results presented in this paper. Results show that three-axis attitude stabilization is possible using only in-plane movable masses which weigh less than 10 percent of spacecraft mass. Finally, the proposed control system can be a viable alternative for three-axis attitude stabilization of spacecraft orbiting in low Earth orbit.
       
  • Data-driven Constraint Approach to Ensure Low-speed Performance in
           Transonic Aerodynamic Shape Optimization
    • Abstract: Publication date: Available online 13 June 2019Source: Aerospace Science and TechnologyAuthor(s): Jichao Li, Sicheng He, Joaquim R.R.A. Martins Aerodynamic shape optimization based on computational fluid dynamics has the potential to become more widely used in the industry; however, the optimized shapes are often criticized for not being practical. Techniques seeking more practical results, such as multipoint optimization and geometric constraints, are either ineffective or too time consuming because they require trial and error. We propose a data-driven constraint for the aerodynamic shape optimization of aircraft wings that ensures the overall practicality of the optimum shape, with a focus on achieving a good low-speed performance. The constraint is formulated by extracting the relevant features from an airfoil database via modal analysis, correlation analysis, and Gaussian mixture models. The optimization results demonstrated that this approach addresses the thin leading edge issue that had plagued previous optimization results, and further analysis demonstrated that this data-driven constraint ensures good low-speed off-design performance without sacrificing the transonic on-design performance. The proposed approach can use other airfoil databases and can even be generalized to other shape optimization and engineering design problems.
       
  • Modeling method of coating thickness random mistuning and its effect on
           the forced response of coated blisks
    • Abstract: Publication date: Available online 13 June 2019Source: Aerospace Science and TechnologyAuthor(s): Xianfei Yan, Junnan Gao, Yue Zhang, Kunpeng Xu, Wei Sun The mistuning modeling methods and mistuning effects on the vibration characteristics of uncoated blisks have been extensively investigated over the past years. Most of these studies mainly focus on single-factor mistuning. In this work, we study the effect of coating thickness random mistuning on the forced response, which contains mass, stiffness and damping mistuning simultaneously and has not yet been investigated in previous studies. Furthermore, to satisfy the requirement of variable coating thicknesses, a new finite element model (FEM) is presented to resolve the new stiffness and mass matrices when the coating thickness changes. This FEM only needs the element node coordinates of one sector for the uncoated blisk and is verified by ANSYS software. For the mistuned blisk, the Craig-Bampton (C-B) method and the subset of nominal system modes (SNM) are employed to statistical analysis. The results indicate that the maximum amplification factor of forced response grows monotonously with the increased mistuning value of coating thickness.
       
  • Observer-based fault tolerant control and experimental verification for
           rigid spacecraft
    • Abstract: Publication date: Available online 12 June 2019Source: Aerospace Science and TechnologyAuthor(s): Qinglei Hu, Xinxin Zhang, Guanglin Niu This work addresses the problem of observer-based closed-loop attitude stabilization in the presence of actuator faults, reaction wheel friction and external disturbances. As a stepping stone, an iterative learning disturbance observer (ILDO) is developed to estimate and compensate for the synthetic disturbances mentioned above. Specially, this ILDO does not need complex faults isolation operations or require knowledge related to these uncertainties. Furthermore, a typical proportional-derivative controller incorporating with the presented ILDO is employed to realize attitude stabilization for spacecraft subject to the actuator failures and uncertainties. The significant feature of the proposed strategy is that the uniformly ultimately bounded stability of the overall closed-loop system with observer-controller architecture could be guaranteed. The effectiveness and robustness of the developed scheme are investigated via a set of numerical simulations and analyses. Particularly, an experimental verification involving a hardware-in-loop platform is further provided to validate the engineering feasibility of the proposed strategy.
       
  • Magnetic sail-based displaced non-Keplerian orbits
    • Abstract: Publication date: Available online 12 June 2019Source: Aerospace Science and TechnologyAuthor(s): Marco Bassetto, Alessandro A. Quarta, Giovanni Mengali This paper deals with the problem of determining the requirements for the maintenance of circular, displaced, non-Keplerian orbits around the Sun by means of a magnetic sail-based spacecraft. The magnetic sail is an exotic propellantless propulsion system that gains thrust from the magnetostatic interaction between the solar wind and an artificial magnetic field, which is generated on board by an electrical current flowing through a loop of conducting material. The propulsive requirements are given in terms of characteristic acceleration and thrust (cone) angle. The analysis is performed using a recent mathematical model in which the magnetic sail thrust vector is expressed as a function of the Sun-spacecraft distance and the sail attitude. Moreover, a linear stability analysis is carried out to identify the (marginally) stable displaced orbits when an error in the orbital insertion is assumed.
       
  • RISE based active vibration control for the flexible refueling hose
    • Abstract: Publication date: Available online 12 June 2019Source: Aerospace Science and TechnologyAuthor(s): Zikang Su, Mingyang Xie, Chuntao Li During the pivotal docking process of autonomous aerial refueling (AAR), the flexible refueling hose's transient vibration resulted from the receiver aircraft's excessive closure speed must be immediately suppressed once it appears, due to its disastrous refueling accident possibility. This paper proposes a permanent magnet synchronous motor (PMSM) driven active vibration control scheme for the flexible refueling hose, via the robust integral of the sign of the error (RISE) feedback control and extended state observer (ESO). The active vibration control scheme can suppress the transient hose vibration under multiple disturbances, without angular velocity measurement. The unmeasured load torque is estimated and compensated via the ESO, together with other uncertainties. The proposed method takes advantages from both the disturbance observer-based control and RISE based disturbance suppression control, which tactfully bridge the gap between these two different anti-disturbance control methods. Extensive simulations results demonstrate that the proposed method can rapidly suppress the hose vibration with satisfactory accurately, even in the presence of the probe position perturbation and parametric uncertainties.
       
  • Appointed-time fault-tolerant attitude tracking control of spacecraft with
           double-level guaranteed performance bounds
    • Abstract: Publication date: Available online 12 June 2019Source: Aerospace Science and TechnologyAuthor(s): Mingmin Liu, Xiaodong Shao, Guangfu Ma This paper investigates the issue of appointed-time fault-tolerant control for spacecraft attitude tracking in the presence of external disturbances and actuator faults. By “appointed-time”, it is meant that the maneuver completion time can be preassigned offline according to mission-oriented demands. Firstly, appointed-time performance functions are tactfully developed, which seek to impose a priori desired performance metrics on both the attitude and angular velocity errors (double-level). After that, an adaptive fault-tolerant controller is derived using structurally simple error transformations in combination of asymmetric barrier Lyapunov functions. Based on Lyapunov synthesis, it is then shown that the derived controller is capable of guaranteeing the boundedness of all the signals in the closed-loop system, and of achieving double-level guaranteed performance bounds for output tracking errors, despite the presence of external disturbances and actuator faults. In particular, the attitude tracking can be accomplished in a user-appointed time without resorting to judicious control parameters selection. Finally, simulation results are presented to illustrate the efficacy of the proposed control scheme.
       
  • Non-uniform stator loss reduction design strategy in a transonic
           axial-flow compressor stage under inflow distortion
    • Abstract: Publication date: Available online 12 June 2019Source: Aerospace Science and TechnologyAuthor(s): Hanan Lu, Zhe Yang, Tianyu Pan, Qiushi Li In a boundary layer ingesting (BLI) propulsion system, the fans/compressors continuously operate at a highly distorted inflow condition due to ingesting the boundary layers on the aircraft fuselage. The BLI inlet distortion induces non-uniform whirl and axial velocities at the rotor outlet, leading to a circumferential and radial non-uniform incidence distribution for the downstream stator. One of the main challenges for the compressor performance is a bad match between the rotor outflow and the stator inlet geometry, which would cause flow separations in blade passages. This paper has employed a non-uniform stator loss reduction design strategy to make a redistribution of stator incidence at the compressor off-design working point so as to reduce the potential separation loss and enhance the stage aerodynamic performance. Firstly, three-dimensional full-annulus unsteady numerical simulations are carried out to reveal the specific flow features of the non-uniform incidence distribution at the stator inlet. Then, according to the circumferential and radial non-uniform incidence distribution, a spatial non-uniform stator loss reduction design strategy based on adjusting the local inlet metal angles of the blade airfoil sections is utilized to make it a good match between the rotor outflow and the stator inlet geometry. Finally, the inner flow fields are analyzed to test the effectiveness of the non-uniform loss reduction design strategy. The computational results indicate that the high-incidence regions at the stator inlet are notably reduced and a more uniform incidence distribution has been achieved along both the circumferential and radial directions. Moreover, the non-uniform loss reduction design has successfully suppressed the flow separations at both the hub and tip regions. As a result, the stator has achieved an 18.2% reduction of the aerodynamic loss at the re-designed point and the stage efficiency has been improved by 0.83%. In the meanwhile, the re-designed stator has decreased the aerodynamic losses over the whole compressor operating range, and the re-designed stage has also achieved a higher efficiency than the baseline one while maintaining the total pressure ratio.
       
  • Numerical investigation on detonation initiation using toroidal shock wave
           focusing
    • Abstract: Publication date: Available online 12 June 2019Source: Aerospace Science and TechnologyAuthor(s): Xiang Chen, Ningbo Zhao, Xiongbin Jia, Shizheng Liu, Hongtao Zheng, Zhiming Li Shock wave focusing was considered to be a potential alternative method to initiate detonation wave in detonation-based propulsion system. A two-dimensional numerical investigation was carried out for hydrogen-air mixtures. Toroidal shock wave formation and detonation initiation by shock wave focusing were analyzed. Besides, considering the flame accelerating zone width and length are potential geometric factors affecting detonation initiation. Effects of different structural parameters on jet intensity and shock wave focusing detonation initiation process were compared. The numerical results indicated that flame acceleration is the critical factor determining toroidal shock wave formation. One possible mechanism of rapid detonation initiation is that shock waves produced by two explosions can compress premixed gas rapidly, and accelerate the deflagration to detonation transition. In addition, with the increase of flame acceleration zone width, two explosion pressures decrease, as a result, the detonation initiation time and distance increase. When the flame accelerating zone width is 1 mm, the time and distance are 380 μs and 28.85 mm, which are the shortest. With the increase of flame acceleration zone length, two explosion pressures increase, the detonation initiation time and distance decrease. When the flame acceleration zone length is 140 mm, the shortest initiation time and distance are 30 mm and 382 μs respectively.
       
  • A nonlinear analytical formula for forced vibration analysis of the
           hard-coating cylindrical shell based on the strain energy density
           principle
    • Abstract: Publication date: Available online 11 June 2019Source: Aerospace Science and TechnologyAuthor(s): Yue Zhang, Jian Yang, Wei Sun, Hua Song In this paper, a nonlinear analytical formulation of the forced vibration of the hard-coating cylindrical shell is developed to investigate the nonlinear resonant characteristics of the shell. The strain dependence of hard coating and the elastic constraint with continuous variable stiffness are considered in the formulation. In order to fully introduce the effects of the strain dependence of hard coating on the resonant characteristics with base excitation, a novel analytical method is presented to determine the equivalent strain of hard coating according to the principle of equal strain energy density. The nonlinear governing equations of motion and the admissible displacement function are derived based on the Love's first approximation theory and the Gram-Schmidt orthogonalization process. A unified Newton-Raphson iterative solution method is employed to solve the nonlinear resonant frequency and response of the shell. As an example to demonstrate the feasibility of the developed analytical model, the forced vibration of the cylindrical shell coated with NiCoCrAlY + YSZ hard coating is implemented numerically and experimentally. Moreover, the influences of the storage modulus, loss modulus and thickness of hard coating on the forced vibration characteristics of the hard-coating cylindrical shell are analyzed in detail.
       
  • Robust trajectory optimization using polynomial chaos and convex
           optimization
    • Abstract: Publication date: Available online 10 June 2019Source: Aerospace Science and TechnologyAuthor(s): Fenggang Wang, Shuxing Yang, FenFen Xiong, Qizhang Lin, Jianmei Song The polynomial chaos (PC) theory and direct collocation method have been integrated to solve a lot of robust trajectory optimization problems. However, the computational cost and memory consumption increase significantly with the increase of the dimension of uncertain factors, and the nonlinearity of dynamic equations. To address this issue, a novel robust trajectory optimization procedure combining PC with the convex optimization technique is proposed in this paper. With the proposed procedure, the trajectory optimization can be implemented with high accuracy and efficiency, by taking advantage of the high accuracy of PC in addressing UP for highly nonlinear dynamics and the high efficiency of convex optimization in solving optimal control. The proposed robust trajectory optimization procedure is applied to two examples and compared with the existing method employing PC and the pseudospectral method. The results show that the proposed procedure can obtain highly accurate results similar to the existing method. Moreover, with the increase of random dimension, the optimal trajectory can still be generated very efficiently without significant increase of computational cost. These results demonstrates the effectiveness of the proposed procedure.
       
  • Biomimetic skeleton structure of morphing nose cone for aerospace vehicle
           inspired by variable geometry mechanism of honeybee abdomen
    • Abstract: Publication date: Available online 10 June 2019Source: Aerospace Science and TechnologyAuthor(s): Yuling Zhang, Jieliang Zhao, Weihua Chen, Xiaodong Guo, Shaoze Yan, Guotun Hu, Yuan Yuan, Pengfei Guo, Qiaoyan Cai Aerospace vehicle containing aeronautics and astronautics function realizes completely reusable, which is the key weapon for the future world countries to fight for air and space control power. However, the current nose cone of aerospace plane cannot respond optimally to the changes in the external environment. The constant structure of nose cone also greatly reduces the range of motion and flexibility. In this paper, a biomimetic skeleton structure of morphing nose cone inspired by variable geometry mechanism of honeybee abdomen is designed, which can achieve stretching and bending continuously. Based on the screw theory, the degree of freedom of morphing nose cone is calculated. Then a prototype of biomimetic skeleton structure with six segments for the morphing nose cone has been manufactured. Further, the simulation analysis and experimental test of the morphing nose cone are conducted respectively to evaluate the deformation ability and dynamic performance. The results show that deformation ability of the morphing nose cone meets the design requirements. The flexibility of drive mechanism in the morphing nose cone causes high-frequency components, which greatly reduces its dynamic performance. This research provides the way to develop a theoretical analysis and experimental support for the design of morphing nose cone for aerospace vehicle.
       
  • Rarefied airfoil aerodynamics based on the generalized hydrodynamic model
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): H. Xiao, D.H. Wang Numerical studies on rarefied gas flows around an airfoil are conducted by an implicit non-Navier-Stokes-Fourier (NSF) framework, namely, the generalized hydrodynamic (GH) model. A detailed study of the rarefied gas flow around an airfoil is performed following validations. Investigations show that all the characteristics of the rarefied effect on an airfoil observed by the particle method in a previous study are also observed in the present GH study. In addition, the constitutive relationships of the GH model provide a clear explanation regarding the distinctive features of the rarefied effect on the aerodynamic characteristics of an airfoil. Also, a recompression region is found at the trailing edge in a state involving subsonic flow and a high Knudsen (Kn) number. With increasing Kn number, the drag increases rapidly compared to the lift, and this effect results in a sharp decrease in the lift-drag ratio. Moreover, the NSF framework overpredicts the lift-drag ratio for subsonic cases and underpredicts the ratio for supersonic cases in rarefied gas flows with different angles of attack considered. Credible explanations are provided for the limitations of the NSF framework in a numerical study of the rarefied effects, including the near equilibrium state, considering the present constitutive relationships of the GH model. We show that the present GH model provides a new numerical tool for the investigation of the rarefied effect on aerodynamic characteristics.
       
  • Geometry effect on the airfoil-gust interaction noise in transonic flows
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): Siyang Zhong, Xin Zhang, James Gill, Ryu Fattah, Yuhao Sun Numerical simulations are conducted to investigate the impacts of airfoil thickness, the angle of attack and camber on the airfoil-gust interaction noise in transonic flows where locally supersonic regimes and terminating shocks are present. The conclusions about the geometry effects based on the extensively studied subsonic cases are revisited. With the increase of airfoil thickness, the sound generation is reduced in the downstream direction as in subsonic flows. More sound is produced in the upstream direction for thicker airfoils due to the non-uniform mean flow and shocks in the near field. The compensative effect makes the overall sound reduction by the airfoil thickness less than the subsonic cases despite the significant difference in the radiation patterns. The acoustic responses to the single frequency gusts are sensitive to the airfoil angle of attack in transonic flows. However, the overall differences are reduced when multiple wavenumber components are superposed in isotropic turbulence, and the sound pressure levels are therefore close as in subsonic flows. Similarly, the significant variations in single frequency acoustic responses by airfoil camber are averaged by the superposition of various wavenumber components. However, apparent variations are still found in the upstream direction, especially for the turbulences with small integral length scales.
       
  • Non-modal stability analysis of low-Re separated flow around a NACA 4415
           airfoil in ground effect
    • Abstract: Publication date: Available online 6 June 2019Source: Aerospace Science and TechnologyAuthor(s): Wei He, José Miguel Pérez, Peng Yu, Larry K.B. Li In this numerical–theoretical study, we perform a linear non-modal stability analysis of the separated flow around a NACA 4415 airfoil over a no-slip ground at low Reynolds numbers (300⩽Re⩽500) and high angles of attack (12∘⩽α⩽20∘). We find that: (i) the strength of the recirculation zone behind the airfoil is a key parameter controlling the absolute/convective nature of the instability in the boundary layer downstream; (ii) when Re, α or the ground clearance increases, the energy gain also increases, with the optimal perturbations switching from being three dimensional to two dimensional; and (iii) classical hairpin vortices, or Klebanoff modes, can be produced by three-dimensional optimal perturbations on a two-dimensional steady base flow containing a laminar separation bubble. Knowledge of the spatiotemporal features of the optimal mode could aid the design of advanced strategies for flow control. This study offers new insight into the transient growth behavior of airfoil–ground flow systems at low Re and high α, contributing to a better understanding of the ground-effect aerodynamics of small insects and micro aerial vehicles.
       
  • Acoustic and radar integrated stealth design for ducted tail rotor based
           on comprehensive optimization method
    • Abstract: Publication date: Available online 5 June 2019Source: Aerospace Science and TechnologyAuthor(s): Zeyang Zhou, Jun Huang, Nana Wu For the integrated stealth issue of ducted tail rotor noise radiation and radar scattering, a comprehensive optimization method based on high frequency electromagnetic calculation theory and boundary element method is introduced. The radar cross section of the rotor, radiated noise and radar cross section of the ducted tail rotor are designed as optimization goals under the constraints of geometric parameters and aerodynamic force. The model of the ducted tail rotor is established by using the full factorial design and the flow field is constructed by high-precision unstructured grid technology. The aerodynamic characteristics of the ducted tail rotor is simulated by the computational fluid dynamics method based on Navier–Stokes equations and k–ε standard viscous model. The noise radiation is solved by boundary element method and the radar cross section value is calculated by physical optics method and physical theory of diffraction. On the basis of these calculation results, the optimization model of the ducted tail rotor is obtained and generated by comprehensive optimization method based on Pareto solution. The final scheme has been satisfactorily improved in terms of noise suppression, radar cross section reduction and aerodynamic lifting. The proposed approach is very effective and efficient for the acoustic/radar comprehensive stealth design of the ducted tail rotor.
       
  • Aerodynamic Optimal Design for a glider with the Supersonic Airfoil based
           on the Hybrid MIGA-SA Method
    • Abstract: Publication date: Available online 5 June 2019Source: Aerospace Science and TechnologyAuthor(s): F.P. Wang, Y. Xu, G.Q. Zhang, K. Zhang The aerodynamics of the glider with supersonic airfoil have been optimized numerically by combining the multi-island genetic algorithm (MIGA) with the simulated annealing (SA) methods. The corresponding results show that the improved hybrid MIGA-SA method can effectively solve the glider optimal design issues such as nonlinear, discontinuous, and multi-dimensional multimodal function etc. The glider optimal aerodynamic geometry can be quickly obtained by using of the hybrid MIGA-SA method under the conditions of huge design space and low calculating resource requirements. The optimized results have also indicated that the aerodynamic characteristics for the double curved airfoil are always superior to the hexagonal airfoil. The total length of wing span located closer to the constraint value can greatly increase the wing area, decrease wing load and benefit in gliding. The final optimal geometry can greatly extend the flight distance for the glider. These results could provide the reference data for designing the gliders with supersonic airfoil in aerodynamic geometry, control system as well as the structural optimal.
       
  • Experimental and numerical investigations of trailing edge injection in a
           transonic turbine cascade
    • Abstract: Publication date: Available online 4 June 2019Source: Aerospace Science and TechnologyAuthor(s): Jie Gao, Ming Wei, Weiliang Fu, Qun Zheng, Guoqiang Yue Trailing edge mixing flows associated with coolant injection are complex, in particular at transonic flows, and result in significant aerodynamics losses. A series of tests and calculations on a high-pressure turbine vane cascade with trailing edge injection were conducted to investigate the influence of trailing edge injection on the loss characteristics of the transonic profile. Wake traverses with a five-hole probe and measurements of the pressure distribution on the profile were taken for injection mass flow ratios of 0 and 4.1% under a test Mach number of 0.7. In the meantime, two-dimensional numerical predictions were carried out for exit isentropic Mach numbers of 0.7, 0.85 and 1.1 and injection mass flow ratios of 0, 4.1%, 4.7% and 5.2%. Numerical predictions show a reasonable agreement with the experimental data, and wake total pressure losses and flow angles as well as pressure distributions on the profile were compared to calculations without trailing edge injection, showing a significant effect of trailing edge injection on the wake development and its blockage effect on the flow in the vane cascade passage. The outlet total pressure loss and flow angle are increased with increasing exit Mach number. The trailing edge injection of 4.1% increases the static pressure in the rear part of vane suction side, and then reduces the vane loading, thus causing decreased wake losses at different exit Mach numbers. At transonic flows, the injection also reduces the strength of the shock wave at the trailing edge pressure side. In addition, the injection decreases the outlet pressure circumferential-unevenness at an exit isentropic Mach number of 0.7, while it increases the pressure circumferential-unevenness at higher exit Mach numbers.
       
  • Numerical investigation on flow control effects of dynamic hump for
           turbine cascade at different reynolds number and hump oscillating
           frequency
    • Abstract: Publication date: Available online 4 June 2019Source: Aerospace Science and TechnologyAuthor(s): Rongfei Yang, Dongdong Zhong, Ning Ge From takeoff to cruise, the operating Reynolds number (Re) of the low-pressure turbine (LPT) in an aero-engine decreases significantly, making the boundary layer near the aft portion of the blade surface susceptible to separation. This paper discusses the feasibility of using local dynamic surfaces as an active flow control method to suppress the laminar flow separation on the suction surface in PAKB LPT cascade at low Re. A local dynamic surface, shaped as a half-sinusoidal hump with an oscillation amplitude of 1 mm, was positioned just upstream of the peak velocity point. Unsteady Reynolds-averaged Navier–Stokes (RANS) simulations were performed to investigate the aerodynamic performance of the cascades with and without dynamic hump. At first, the effect of Re was studied with the hump oscillating frequency at 200 Hz. Compared with uncontrolled cascade at various Re from 25,000 to 150,000, the profile losses of the cascade controlled by dynamic hump reduced significantly at low Re condition where the separation bubble was very large, and increased slightly at high Re condition where the separation bubble was small. Then the effect of the hump oscillation frequency was investigated under the Re of 25,000. It was found that the controlled cascade loss reached a minimum in a certain hump oscillating frequency range. At low Re of 25,000, the loss of controlled cascade was mainly composed of low-loss continuous laminar vortices that were attached to the suction surface. Continuous vortices with a certain spacing confined the low-energy fluid near the suction surface by energy exchange between the free flow and near-wall flow, which suppressed the large-scale flow separation.
       
  • Numerical predictions of low Reynolds number compressible aerodynamics
    • Abstract: Publication date: Available online 3 June 2019Source: Aerospace Science and TechnologyAuthor(s): T. Désert, T. Jardin, H. Bézard, J.M. Moschetta Interest in low Reynolds number compressible flows is emerging due to prospective applications like flight on Mars and in the stratosphere. However, very little knowledge is available, both regarding the flow physics underlying this unique regime and the accuracy of numerical methods for its prediction. In this paper, low and high fidelity numerical approaches are compared with experimental measurements on both airfoils and rotors in the low Reynolds number compressible flow regime. It is shown that low fidelity approaches are suited to aerodynamic optimization despite high viscous and compressible effects. In addition, high fidelity approaches help reveal unique flow features of this regime.
       
  • Adaptive super-twisting sliding mode control of variable sweep morphing
           aircraft
    • Abstract: Publication date: Available online 3 June 2019Source: Aerospace Science and TechnologyAuthor(s): Binbin Yan, Pei Dai, Ruifan Liu, Muzeng Xing, Shuangxi Liu In this paper, a novel and more explicit longitudinal model of a wing-sweep morphing aircraft is proposed, accounting for the variations in aerodynamics, mass and inertial properties. In addition, open loop dynamic responses with different morphing rates are simulated, and a small morphing rate leads to big magnitudes of velocity and altitude variations. During morphing process, both aerodynamic parameters and mass distribution change dramatically, and it causes a difficulty in modeling the morphing aircraft precisely. Therefore, an adaptive super twisting algorithm sliding mode controller is proposed to track reference trajectory after input/output linearization of the longitudinal model. Simulation results show the designed controller works better with good tracking performance and smaller chattering compared with normal sliding mode controller, and the robustness of the designed controller is verified by simulations at the presence of aerodynamic perturbations.
       
  • Free vibration analysis of variable angle-tow composite wing structures
    • Abstract: Publication date: Available online 3 June 2019Source: Aerospace Science and TechnologyAuthor(s): A. Viglietti, E. Zappino, E. Carrera This paper investigates the possibility to improve the dynamic response of complex aeronautical structures using variable angle-tow composites. The study has been performed using an innovative numerical approach developed in the framework of the Carrera Unified Formulation able to study laminates with curvilinear fibres whose trajectories can be arbitrarily defined. Refined kinematic structural models have been used to deal with the complex behaviour of such structures. Several cases have been investigated in order to validate this approach and the results have been compared with those from classical modelling approaches. Simple beam models and complex wing structures, have been considered. The effects of different fibres-paths have also been studied and compared. The results confirm that an appropriate tow lay-up can be used to improve the performances of wing structures, i.e. innovative design solutions can be achieved.
       
  • A gradient-based aero-stealth optimization design method for flying wing
           aircraft
    • Abstract: Publication date: Available online 3 June 2019Source: Aerospace Science and TechnologyAuthor(s): Ming Li, Junqiang Bai, Li Li, Xiaoxuan Meng, Qian Liu, Bao Chen Flying wing layout has both favorable aerodynamic performance and outstanding radar scattering characteristics, which is regarded as an ideal stealth aircraft layout. In this paper, a gradient-based aerodynamic and stealth optimization design method is established by coupling the Free-Form Deformation approach (FFD), Radial Basis Function algorithm (RBF), Computational Fluid Dynamics (CFD), Physical Optics (PO) and the Sequential Quadratic Programming algorithm (SQP). The gradient of the aerodynamic characteristic parameters could be obtained by solving the discrete adjoint equations. With PO code differentiated by the automatic differentiation tool, the Radar Cross Section (RCS) and its gradient would be collected. Based on the aero-stealth optimization design method, three kinds of optimization strategies are applied to design a certain flying wing layout aircraft: 1) The drag coefficient is chosen as the objective function while the RCS is set as a constraint condition; 2) The RCS is selected as the objective function while the drag coefficient is set as a constraint condition; 3) The objective function is composed of the drag coefficient and the RCS through the Weighted Sum method. The optimization results show that when choosing the first two strategies, the optimization algorithm would push single discipline performance to the edge. Alternatively, it would obtain several trade-off configurations when selecting the last strategy. The results demonstrate that the gradient-based aero-stealth optimization design method can deal with the multidisciplinary optimization problems which require a large number of design variables. Without a large number of function evaluations, this method could rapidly and efficiently obtain the shape that meets the requirements of aerodynamic and stealth performances.
       
  • Effect of rear-fuselage design on the aerodynamics of a helicopter at
           side-slip angle
    • Abstract: Publication date: Available online 31 May 2019Source: Aerospace Science and TechnologyAuthor(s): Drew Gingras, Felix Weis, Bryan Godbolt, Sina Ghaemi The aerodynamic design of a helicopter fuselage encounters many complexities that arise from the large variation in side-slip angle during different flight maneuvers and gusty conditions. A wind tunnel investigation was performed on a scaled-down helicopter model at Reynolds number of Rew=9.6×105 to analyze its aerodynamic performance up to a 40° side-slip angle. Four rear-fuselage configurations were investigated including a removable motor geometry, as well as a cusped and a round enclosure for the motor. A six-axis load cell was used to determine three components of force and moment acting on the fuselage. Planar PIV measurements were performed in the fuselage wake to analyze effect of side-slip angle (β) and rear-fuselage on the wake flow. The drag force coefficients displayed a parabolic increase with rising β. The pitch moment strongly depended on the aft-body design. The cusped rear-body had the largest pitch moment (Cm) at zero-side slip while the round case had the smallest magnitude. For all the rear-fuselage configurations, the slope of the pitching moment (dCm/dβ) changed sign from negative to positive at β=20°. The largest change in dCm/dβ was observed for the baseline case while the round enclose had the smallest change in dCm/dβ. The increase in negative Cm with increase of β from zero to 20° was associated with the displacement of the low-pressure core of the wake toward the rear-fuselage. When β increased from 20° to 40°, the wake became skewed and the low-pressure core moved away from the rear-fuselage, causing reduction of the negative Cm toward zero.
       
  • Real-time simulation model for helicopter flight task analysis in
           turbulent atmospheric environment
    • Abstract: Publication date: Available online 31 May 2019Source: Aerospace Science and TechnologyAuthor(s): Honglei Ji, Renliang Chen, Pan Li This paper presents a real-time simulation model for the analysis of the helicopter flight tasks in turbulent atmospheric environment. First, recursive algorithms are developed for independent turbulence components by discretizing the high-order turbulence filters deduced from the von Kármán model. Then, the discrete filters are related according to the spatial correlation on transverse planes. The related filters are distributed in front of a helicopter and their turbulence components are transported along the longitudinal direction of airspeed to establish a distributed turbulence model. Next, the turbulence model is integrated into a flight dynamics model, and a real-time simulation model is formed and validated against flight test data. Finally, the model accuracy and computational speed with simulation step are investigated using numerical simulations. Results show that the model computational speed increases with the increment of simulation step. However, the lower-frequency helicopter motions may be distorted by the higher-frequency turbulence disturbance when a too large simulation step was used. Following the results, appropriate simulation steps are proposed. The model accuracy and computational speed meet the real-time requirements for the helicopter flight simulations in atmospheric turbulence.
       
  • Frozen orbit design and maintenance with an application to small body
           exploration
    • Abstract: Publication date: Available online 30 May 2019Source: Aerospace Science and TechnologyAuthor(s): Xiangyu Li, Dong Qiao, Peng Li Frozen orbits are ideal options for global mapping and observation. In this paper, a design and maintenance method of frozen orbit is presented and applied to the small body exploration. On the basis of the Legendre addition theorem, an analytic solution of the frozen orbit is developed, which can obtain frozen orbits with arbitrary orders of zonal harmonic coefficients in concise forms. Based on the proposed method, the distribution and properties of frozen orbits around Vesta are discussed. Then, an anti-chattering self-adaptive control law is proposed to maintain the frozen orbits in a high-fidelity environment. The control law is proved to be stable and robust against model uncertainties and unmodeled perturbations. Simulation results show that the proposed method can obtain the frozen orbit efficiently with desired performance, while the self-adaptive control law can keep the orbit stable with satisfactory fuel consumption. This study provides a new option for frozen orbit design in future exploration missions.
       
  • Experimental demonstration of an end-burning swirling flow hybrid rocket
           engine
    • Abstract: Publication date: September 2019Source: Aerospace Science and Technology, Volume 92Author(s): J.-Y. Lestrade, J. Anthoine, A.J. Musker, A. Lecossais Within the framework of the European H2020 HYPROGEO program, an innovative hybrid engine combustion chamber compatible with satellite requirements (constant thrust over very long burn-times) had to be developed. A first test rig was designed and tested in order to better understand the functioning of this innovative combustion chamber, and to help the design of a breadboard to demonstrate the efficiency of this new engine with respect to the mission requirements. Two test campaigns, the first on the test rig with 87.5% hydrogen peroxide, and the second on the hybrid engine breadboard with 98% hydrogen peroxide, were performed under various operating conditions to demonstrate the catalytic ignitability of this new hybrid engine, and the sustainment of a stable combustion over firing durations up to 180 s. The test campaigns also enabled the identification of the main influencing parameter on the fuel regression rate for this innovative combustion chamber.
       
  • High-precision control method for the satellite with large rotating
           components
    • Abstract: Publication date: Available online 30 May 2019Source: Aerospace Science and TechnologyAuthor(s): Feng Wang, Chi Wang, Xue-qin Chen, Cheng-fei Yue, Yi-fei Xie, Li-peng Chai Satellites with large rotating components may accomplish some specific scientific missions and have attracted a great deal of research interests. This paper investigates the control problem of such satellites. The model of the rotating component is carefully studied and both the static and dynamic unbalance are analyzed and demonstrated. To achieve a high precision control with unbalance rejection, a sliding mode control strategy is proposed. Practical simulations are conducted to verify the effectiveness of the proposed controller and simulation results illustrate that both the static and dynamic unbalance are perfectly compensated and a high-precision control is achieved.
       
  • Flatness-based finite-time leader–follower formation control of multiple
           quadrotors with external disturbances
    • Abstract: Publication date: Available online 29 May 2019Source: Aerospace Science and TechnologyAuthor(s): Xiaolin Ai, Jianqiao Yu This study considers the leader–follower formation control problem for multiple quadrotors in the presence of external disturbances. Based on the differential flatness theory, the underactuated quadrotor system is transformed into a fully actuated one with four degrees of freedom and four control inputs. Based on this model, a distributed finite-time observer is developed to reconstruct the leader's states for each follower. Then, an observer-based finite-time controller is proposed based on an adaptive disturbance rejection approach, which is independent of the upper bounds of the disturbances, to address the formation control problem for multiple quadrotors. The stability analysis indicates that the closed-loop system theoretically achieves the finite-time stability and robustness against the disturbances. Finally, a numerical simulation is provided to verify the performance of the proposed formation control scheme.
       
  • Design and implementation of MPC for turbofan engine control system
    • Abstract: Publication date: Available online 29 May 2019Source: Aerospace Science and TechnologyAuthor(s): Morteza Montazeri-Gh, Ali Rasti, Ali Jafari, Milad Ehteshami The objective of an aircraft engine control system is to provide required thrust as well as protection against the physical and operational limits. Model predictive control (MPC) technique is an attractive approach that incorporates input/output constraints in its optimization process to fulfill the control requirements of an aircraft engine. However, due to heavy computational burden of MPC, the real-time implementation of this algorithm is challenging and selection of MPC design parameters is crucial. This paper presents the design and hardware implementation of MPC algorithm as well as its HIL testing for turbofan engine control. In addition, a feedback correction technique is employed to compensate the effect of plant-model mismatch. For this purpose, a thermodynamic nonlinear engine model is firstly developed. A multivariable model predictive controller is then designed based on a discrete-time linearized state-space model where the horizons are obtained through a genetic algorithm optimization procedure. Moreover, this controller is implemented on an appropriate hardware taking the real-time aspects into account. Finally, an HIL platform is developed for testing the turbofan engine electronic control unit (ECU). For this purpose, a multi-level throttle command is applied to the ECU and the performance of the engine is evaluated. The results indicate that the controller satisfies all the engine constraints, and confirm the successful software and hardware implementation of the control algorithm in real-time.
       
  • Quasi-3D isogeometric buckling analysis method for advanced composite
           plates in thermal environments
    • Abstract: Publication date: Available online 28 May 2019Source: Aerospace Science and TechnologyAuthor(s): Vuong Nguyen Van Do, Chin-Hyung Lee A quasi-3D isogeometric thermal buckling analysis method for advanced composite plates such as functionally graded plates is presented. A new quasi-3D plate theory for the numerical analysis, which enables the normal deformations to be considered with only four unknowns, is introduced. A simple trigonometric normal shape function that is able to accurately express the through-thickness displacement while facilitating the numerical computation is developed. The refined quasi-3D theory is combined with the non-uniform rational B-spline (NURBS)-based isogeometric analysis (IGA) which satisfies the C1-continuity of the displacement field required by the proposed theory to resolve the thermal buckling issue. Several numerical examples which entail buckling analysis of diverse types of functionally graded plates including the one with a complex cutout under different temperature gradients through the thickness are simulated to validate the quasi-3D isogeometric approach, and to explore the thermal buckling response of functionally graded plates. The present quasi-3D IGA is concluded to be accurate and effective numerical method, and significance of including the thickness expansion effects in the thermal buckling assessment of functionally graded plates is confirmed.
       
  • Effects of fuel variation and inlet air temperature on combustion
           stability in a gas turbine model combustor
    • Abstract: Publication date: Available online 28 May 2019Source: Aerospace Science and TechnologyAuthor(s): Feier Chen, Can Ruan, Tao Yu, Weiwei Cai, Yebing Mao, Xingcai Lu In this study, influences of fuel variation and inlet air temperature on the combustion stability characteristics in a gas turbine model combustor were experimentally investigated. Test fuels involved three selected single component hydrocarbons, including one linear alkane (n-decane), one branched alkane (iso-octane) and one cyclic alkane (methylcyclohexane (MCH)). RP-3 jet fuel was also selected as a technical reference. For all the fuels, experiments were conducted at a fixed equivalence ratio of 0.86 and varying inlet air temperatures from 383 to 483 K. Results showed that RP-3 and n-decane exhibited similar stability behaviors. At low inlet air temperatures, when the combustor was fueled with RP-3 and n-decane, the flame was stabilized and anchored in the combustor. The combustor then shifted to thermo-acoustically unstable state when the inlet air temperature exceeded a threshold value, which was associated by large-scale flame shape variations. On the other hand, noticeable differences can be observed for MCH and iso-octane flames, which featured thermo-acoustically unstable combustion throughout all the tested conditions, and unique mode-shift phenomenon was observed when the inlet air temperature was raised from 403 K to 423 K. Additional flame dynamics was visualized by OH* chemiluminescence imaging. The underlying mechanisms that led to the differences in combustion stability and flame dynamics of the tested fuels were discussed with respect to their differences in physicochemical properties.
       
  • Morphing Control of a New Bionic Morphing UAV with Deep Reinforcement
           Learning
    • Abstract: Publication date: Available online 28 May 2019Source: Aerospace Science and TechnologyAuthor(s): Dan Xu, Zhe Hui, Yongqi Liu, Gang Chen With rapid development of aviation technology, materials science and artificial intelligence, aircraft design is pursuing higher requirements both in civil and military fields. The new generation of aircraft should has the autonomous capable of performing a variety of tasks (such as take-off and landing, cruising, maneuvering, hover, attack, etc.) under a highly variable flight environment (height, Mach number, etc.) and meanwhile maintaining good performance. Morphing aircraft can use smart materials and actuators to autonomously deform the shape according to the changes in flight environment and mission, and always maintain an optimal aerodynamic shape, therefore get flourished developments. Based on the ability of birds to stretch wings when flying at low speed and to constrict wings at high speed, a new bionic morphing UAV has been designed and developed as the study model by our team. In order to make this new aircraft be able to complete rapid autonomous morphing and aerodynamic performance optimization under different missions and flight conditions, we developed deep neural networks and reinforcement learning techniques as a control strategy. Considering the continuity of the state and action spaces for model, the Deep Deterministic Policy Gradient (DDPG) algorithm based on the actor-critic, model-free algorithm was adopted and verified on the classic nonlinear Pendulum model and Cart Pole game. After the feasibility was verified, morphing aircraft model was controlled to complete prescribed deformation using DDPG algorithm. Furthermore, on the condition that the DDPG algorithm can control morphing well, through training and testing on model using simulation data from wind tunnel tests and actual flight, the autonomous morphing control for the shape optimization of the bionic morphing UAV model could be realized.
       
  • Nanosatellite Attitude Estimation using Kalman-Type Filters with
           Non-Gaussian Noise
    • Abstract: Publication date: Available online 25 May 2019Source: Aerospace Science and TechnologyAuthor(s): Demet Cilden-Guler, Matti Raitoharju, Robert Piche, Chingiz Hajiyev In order to control the orientation of a satellite, it is important to estimate the attitude accurately. Time series estimation is especially important in micro and nanosatellites, whose sensors are usually low-cost and have higher noise levels than high end sensors. Also, the algorithms should be able to run on systems with very restricted computer power. In this work, we evaluate five Kalman-type filtering algorithms for attitude estimation with 3-axis magnetometer and sun sensor measurements. The Kalman-type filters are selected so that each of them is designed to mitigate one error source for the unscented Kalman filter that is used as baseline. We investigate the distribution of the magnetometer noises and show that the Student's t-distribution is a better model for them than the Gaussian distribution. We consider filter responses in four operation modes: steady state, recovery from incorrect initial state, short-term sensor noise increment, and long-term increment. We find that a Kalman-type filter designed for Student's t sensor noises has the best combination of accuracy and computational speed for these problems, which leads to a conclusion that one can achieve more improvements in estimation accuracy by using a filter that can work with heavy tailed noise than by using a nonlinearity minimizing filter that assumes Gaussian noise.
       
  • Inverse design methodology on a single expansion ramp nozzle for scramjets
    • Abstract: Publication date: Available online 24 May 2019Source: Aerospace Science and TechnologyAuthor(s): Kaikai Yu, Jinglei Xu, Zheng Lv, Guangtao Song An inverse design method on a single expansion ramp nozzle (SERN) is proposed in order to obtain good aerodynamic performance and full integration with airframes. By using this method, the flow parameters along the nozzle outlet and the shape of the outlet can be specified in advance. The details of the proposed method are introduced in this paper. Method of characteristics (MOC), known as a universal mathematical method, is the kernel of the design method and applied to generate the nozzle contour inversely. Also, the computational fluid dynamic (CFD) approaches are used to verify the effectiveness and accuracy of the proposed method. In the verification case, the relative error of the Mach number on the nozzle outlet in the main region is less than 0.2%. Then, an actual design case on SERN is performed using the proposed method. Two SERNs (SERN A1 and SERN A2) are designed using the conventional design method to examine the superiority of the proposed method. The comparison of the aerodynamic performance shows that the nozzle designed by the proposed method gains approximately 6.3% and 26.2% increments in axial thrust coefficient relative to SERN A1 and SERN A2 on the design condition, respectively. The improvement in the lift of the SERN designed by the proposed method is over 200% on the design condition. Besides that, findings also indicate that the proposed design method can offer a considerable improvement in the thrust coefficient and lift on the off-design conditions. In conclusion, the study can provide a powerful design method for SERNs.
       
  • Lag-twist coupling sensitivity and design for a composite blade
           cross-section with D-spar
    • Abstract: Publication date: Available online 24 May 2019Source: Aerospace Science and TechnologyAuthor(s): M.R. Amoozgar, A.D. Shaw, J. Zhang, C. Wang, M.I. Friswell In this paper, the effect of various parameters of a specific rotor blade cross-section on the effectiveness of a twist morphing concept is investigated. Then, by considering different constraints, a cross-section consistent with this morphing concept with high lag-twist coupling and low extension-twist, is developed. This lag bending-torsion coupling is used to change the twist of the blade during the flight, while the high values of extension-twist coupling is avoided. To this end, a concentrated mass is added to the blade, where its chordwise location varies in flight. When the mass moves in the chordwise direction, a local lag bending is introduced into the blade. This in-plane bending moment then changes the blade twist distribution through lag-twist coupling induced through stiffness tailoring in the blade cross-section. Therefore, this coupling plays an important role in this morphing concept. The one-dimensional dynamics of the blade is modelled by using the geometrically exact fully intrinsic bean equations while the 2D cross-sectional stiffness values are determined by using the VABS software. First, a blade which resembles the BO-105 main rotor blade in the fundamental frequencies is designed. Then, the effect of various parameters of the cross-section on the fundamental frequencies, the lag-twist coupling, and the extension-twist coupling are determined. It is found that the skin of the spar has the highest contribution to both the extension-twist and the lag-twist coupling. Finally, a cross-section compatible with the proposed morphing concept is designed and it is demonstrated that the twist of the blade may be changed significantly.
       
  • Experimental and numerical study on thermodynamic performance in a
           designate pilot-ignition structure: Step
    • Abstract: Publication date: Available online 24 May 2019Source: Aerospace Science and TechnologyAuthor(s): Shilong Zhao, Yuxin Fan Reliable ignition and stable flame spreading were extremely important especially for the increasing expansion of flight envelope. A new pilot-ignition structure- step- was put forward with low resistance and developed to meet more critical challenges. To optimize structure parameters, thermodynamic performance with various step heights (30 mm, 40 mm, 50 mm, and 60 mm) was investigated in premixed combustion, such as ignition and blow-off loop measured by intelligent metal tube flowmeter, and flame development recorded by high-speed camera. Additionally, numerical simulation of flow field was performed to analyse the influence of flow pattern on thermodynamic characteristics. It was found that flow pattern occurred transition with the increase of step height, namely, couple-vortex recirculation in 30 mm-height step structure and single-vortex recirculation in the other three. Additionally, the four kinds of steps realised successful ignition and stability loop. Specially, ignition and blow-off performance of 30 mm-height step structure with particular couple-vortex recirculation was sensitive to temperature, and improved with the growing temperature; also its ignition limits were increasing with the growing velocity, especially up to 200 m/s, ignition limits were almost lower than the others. Once ignited, flame development rate had a convex-function relation to temperature and optimized temperature existed at an extreme point (at 800 K), it decreased progressively with growing velocity. Moreover, flame spreading height up to core flow increased first and then decreased with growing temperature, and it was monotone decreasing with the growing inlet velocity. The step structure conducted in this work was validated for its reliable ignition and broad stability loop. Low-resistance pilot ignition design was developed and employed for turbojet/ramjet.
       
  • AC Plasma Retarded Flame Spread over Thin Solid Fuels in a Simulated
           Microgravity Environment
    • Abstract: Publication date: Available online 24 May 2019Source: Aerospace Science and TechnologyAuthor(s): Zihan Wang, Qi Chen, Xingqian Mao The flame spread over thin solid fuels in a narrow channel by the action of AC plasma was first proposed and reported to show a trade-off between the flame instability produced by ionic wind and the kinetic enhancement stimulated by energetic and chemically active species. The 6mm height narrow channel apparatus has been demonstrated the ability to suppress buoyant flow in horizontally spreading flames, and can successfully simulate a microgravity environment. Plasma acted on flame spread over thin solid fuels led to a confined behavior not typically seen in previous researches, which was investigated by integrated studies of experiment and supporting analysis using BOLSIG+. Four primary variables including flow velocity, reduced electric field, oxygen concentration and electrode length were considered. Among them, the reduced electric field and the oxygen concentration were found the crucial factors affecting the flame spread and further the flammability limit. The experimental results showed the affected field of the flame spread matched the field acted by AC discharge, which suggested a one-to-one correspondence between AC plasma action and flame slowing down or extinguishing. The results also showed that the narrow channel apparatus could capture the essential features of the flame spread in a simulated microgravity, for example the flame-lets phenomenon. The dependence of ionic wind on the reduced electric field was additionally investigated by solving the electron energy deposition into different excited channels and the electron energy distribution function (EEDF), which was essentially in agreement with the experimental observations. This study would develop a new and successful tool with rapid control of spacecraft fire and response to future exploration vehicles.
       
  • Experimental study on the effects of simulated flight conditions on
           ignition and flame stabilization in a supersonic combustor
    • Abstract: Publication date: Available online 23 May 2019Source: Aerospace Science and TechnologyAuthor(s): Lang Li, Wenyou Qiao Effects of simulated flight conditions on ignition and flame stabilization in a kerosene fueled scramjet combustor were experimentally investigated in the present paper. Simulated flight conditions were Mach 4.5 and Mach 4.0, and the isolator entrance Mach number was 2.0. Wall pressure and flame emission were made during the experiments in an attempt to better understand the combustion characteristics, air throttling was used to enhance the flame stabilization. The results showed that, the kerosene ignition was achieved successfully under the Mach 4.5 flight condition, that of Mach 4.0 flight condition was blowout before the kerosene pressure reached the max pressure. The flame in the combustor without air throttling during Mach 4.0 flight condition only existed in the cavity and near the top wall before it was blown off, that of Mach 4.5 flight condition had spread into the isolator. The flame stabilization was also only achieved under the Mach 4.5 flight condition with the aid of air throttling, the flame in the combustor with air throttling during the Mach 4.0 flight condition was blowout before the throttling air was removed. Ignition and flame stabilization were easy to be achieved under the higher flight condition.
       
  • Identification of All the Inertial Parameters of a Non-cooperative Object
           in Orbit
    • Abstract: Publication date: Available online 23 May 2019Source: Aerospace Science and TechnologyAuthor(s): Qingliang Meng, Jianxun Liang, Ou Ma Knowing the dynamic properties of a non-cooperative spacecraft is critical for robotic capture and service in orbit. However, using visual observation alone is not sufficient to identify all the inertial parameters (the mass, mass center location, and inertia matrix) of an unknown target object. This paper presents a method to fully identify all the inertial parameters of a non-cooperative object in orbit with data from visual and force-moment sensors. We propose to use a flexible rod to change a target's movement, which is a prerequisite to identify the true values of the target's dynamic properties. A novel algorithm for processing the collected force and torque data is introduced, which reduces the effect of noise on the identification accuracy. Simulation results have shown that for a large target, we only need to apply a very small force to completely identify all the inertial parameters with an acceptable error in the presence of sensor measurement errors as well as a rough initial guess.
       
  • Integrated thermal protection system based on C/SiC composite corrugated
           core sandwich plane structure
    • Abstract: Publication date: Available online 23 May 2019Source: Aerospace Science and TechnologyAuthor(s): Ying Li, Lu Zhang, Rujie He, Yongbin Ma, Keqiang Zhang, Xuejian Bai, Baosheng Xu, Yanfei Chen A combined theoretical, numerical and experimental investigation of the integrated thermal protection system based on C/SiC composite corrugated core sandwich plane structure was conducted. Corrugated core sandwich plane structure was made from carbon fiber reinforced silicon carbide composite by a hot compression molding combined with precursor infiltration and pyrolysis method. The equivalent thermal conductivity prediction method for the C/SiC composite corrugated core sandwich plane was developed, and its heat transferring behavior was numerical analyzed. Based on the theory and numerical analysis, a novel integrated thermal protection system, which was consisted of three parts: C/SiC composite corrugated core sandwich plane, insulated aerogel filling in the core, and insulated aerogel adhering onto the down surface of the down facesheet, was designed, manufactured and examined by simulated atmospheric re-entry wind tunnel test. It was believe that this novel design of integrated thermal protection system is a potential thermal protection system for the next generation hypersonic flights.
       
  • On the performance of a body integrated diverterless supersonic inlet
    • Abstract: Publication date: Available online 23 May 2019Source: Aerospace Science and TechnologyAuthor(s): M.R. Soltani, R. Askari Extensive experimental investigations on a body integrated Diverterless Supersonic Inlet (DSI) were conducted. These inlets are implemented for both supersonic flow compression and boundary layer diversion using a three- dimensional bump in combination with a suitable cowl lip. Experiments were performed at a free stream Mach number of 1.65, the design Mach number, and at zero degrees angle of attack and zero degrees sideslip angle. To recreate the operational conditions more accurately and have a realistic performance characteristic, the intake was integrated with a fuselage and a forebody nose with an elliptical cross-section. Wind tunnel tests were conducted at critical, subcritical and supercritical operating conditions where DSI operates in its stable states. The results showed that the present DSI had acceptable performance characteristics in the stable operating conditions and provides the required mass flow and static pressure compression ratio. In addition, it has a wide stable subcritical operating condition. Moreover, a relatively different behavior in the DSI performance curve is evident due to its body integrated geometry and 3D configuration of the bump and cowl lip, especially in the supercritical conditions.
       
  • Nussbaum gain adaptive control scheme for moving mass reentry hypersonic
           vehicle with actuator saturation
    • Abstract: Publication date: Available online 22 May 2019Source: Aerospace Science and TechnologyAuthor(s): Haolan Chen, Jun Zhou, Min Zhou, Bin Zhao This paper addresses an actuator saturation problem of the moving mass hypersonic vehicles (HSVs) in the reentry phase. The saturation nonlinearity is modeled using a hyperbolic tangent function and the concomitant time-varying coefficients problem is handled via Nussbaum gain technique. Then backstepping technique is applied in control design and the explosion of complexity in traditional backstepping design is avoided by utilizing dynamic surface control. Subsequently, a Nussbaum gain adaptive controller is constructively framed to deal with the actuator saturation. Based on the disturbance observers, the robustness of the proposed controller is enhanced. The semiglobal stability of the closed-loop system is ensured via Lyapunov synthesis. Finally, numerical simulation results are presented to show the effectiveness of the designed control scheme.
       
  • A comparative study of multi-objective expected improvement for
           aerodynamic design
    • Abstract: Publication date: Available online 22 May 2019Source: Aerospace Science and TechnologyAuthor(s): Lavi Rizki Zuhal, Pramudita Satria Palar, Koji Shimoyama Multi-objective optimization in aerodynamics plays an important role in revealing trade-offs between conflicting objectives in order to discover important knowledge and insight for better future design. Of interest here is the use of Kriging surrogate models incorporated into a sequential Bayesian optimization (BO) strategy. In this paper, we studied four variants of multi-objective BO (MOBO) techniques that are based on expected improvement, that is, Euclidean-based EI (EEI), expected hypervolume improvement (EHVI), ParEGO, and expected inverted penalty boundary intersection improvement (EIPBII) to understand their capabilities on handling multi-objective aerodynamic optimization problems. Numerical tests were performed on a set consisting of six generalized Schaffer problems (GSP), five low-fidelity, and one high-fidelity airfoil design problems. Results suggest that EHVI is the only method which consistently performed well on artificial and aerodynamic problems. EEI yields the worst performance and is not suitable to deal with various problem complexities. ParEGO, although it performs modestly on GSP problem, surprisingly works well on the low- and high-fidelity problems. On the other hand, EIPBII encounters the opposite case, where it is one of the best performer on GSP but yields modest performance on the aerodynamic problems. In light of the results, we suggest that EHVI is a highly potential MOBO method to be applied for multi-objective aerodynamic design optimization.
       
  • Multi-objective design of optimal higher order sliding mode control for
           robust tracking of 2-DoF helicopter system based on metaheuristics
    • Abstract: Publication date: Available online 22 May 2019Source: Aerospace Science and TechnologyAuthor(s): Wafa Boukadida, Anouar Benamor, Hassani Messaoud, Patrick Siarry This paper deals with the trajectory tracking of a 2 Degrees of Freedom (DoF) helicopter system. The control strategy is designed by the combination of the robust control strategy (Higher Order-Sliding Mode Control (HO-SMC)) and the optimal control technique (Linear Quadratic Regulator (LQR)). Combining these two methods lies in the fact that the robust controllers tackle the uncertainties when the optimal controller performances are unaffected. As the performances of the Sliding Mode Control (SMC) greatly depends on the choice of the sliding surface, a novel method based on the solution of a Sylvester equation is proposed. Furthermore, the problem of deciding the optimal configuration of the LQR controller as well as the gain of the discontinuous control is considered as an optimization problem, which can be solved by the application of an efficient metaheuristic. The adequacy of the specific choice of the discontinuous gain is exhibited through general analysis. The main contribution of this paper is to consider a multi-objective optimization problem. For that, a novel dynamically aggregated objective function is proposed. As a result, a set of non-dominated optimal solutions are provided to the designer and then he selects the most preferable alternative. The proposed control strategy is applied for pitch and yaw axes control of the Quanser helicopter. Experimental results substantiate that the combination of the HO-SMC with the LQR method and metaheuristics results in not only reduced tracking error but also improved tracking response with reduced oscillations.
       
  • Modelling and control for the mode transition of a novel tilt-wing UAV
    • Abstract: Publication date: Available online 22 May 2019Source: Aerospace Science and TechnologyAuthor(s): Yongchao Wang, Yaoming Zhou, Chenghao Lin This paper mainly presents a multibody dynamics model and a novel control method for the mode transition of a new-style distributed propulsion tilt-wing UAV. Base on the technology of tensor flight dynamics, a multibody attitude dynamics model formulated in an invariant tensor form is developed for the mode transition such that the dynamics induced by the relative movement of the moving parts (wings and rotors) with respect to the fuselage could be formulated explicitly in the model. The control system is decoupled into two parts, namely the position subsystem and the attitude subsystem subject to input perturbation and external aerodynamic disturbances. A novel finite time altitude tracking controller is designed for position subsystem in terms of the existence of the external disturbances and the perturbations acting on the inputs such that the tilt-wing UAV can converge and move along the desired altitude trajectory in a finite time. Besides, a RISE-based attitude tracking controller is developed to control the attitude subsystem, which guaranteeing robustness to the external disturbances. Numerical simulations are carried out to illustrate the performance of the proposed controllers.
       
  • A passive approach for adjusting the diurnal temperature difference of the
           envelope of stratospheric light aerostat
    • Abstract: Publication date: Available online 21 May 2019Source: Aerospace Science and TechnologyAuthor(s): Jun Liao, Yi Jiang, He Liao, Di-e Xiao, Junjie Yuan, Zechuan Yang, Jun Li, Shibin Luo Stratospheric light aerostat flies relying on the air buoyancy, and has more advantages than other air vehicles. The diurnal temperature difference of the stratospheric light aerostat is important to the long-endurance regional station-keeping performance. In the present paper, a thermal model is proposed, which include the thermal model of the envelope, internal gas, computational model, and adjustment of the diurnal temperature difference. A comparison with related results in the literature is carried out to verify the model. The stratospheric light aerostat temperature distribution and Helium velocity field are simulated. The effects of the envelope radiation properties including emissivity, the absorptivity and the ratio of absorptivity to emissivity on the stratospheric light aerostat thermal performance are discussed. The results show that the envelope radiation properties have great influence on stratospheric light aerostat thermal performance and decreasing the ratio of absorptivity to emissivity of envelope can be a good way to improve thermal performance. The results are conducive to select envelope radiation parameters of the stratospheric aerostat.
       
  • Elastoplastic postbuckling analysis of moderately thick rectangular plates
           using the variational differential quadrature method
    • Abstract: Publication date: Available online 21 May 2019Source: Aerospace Science and TechnologyAuthor(s): E. Hasrati, R. Ansari, H. Rouhi In this research, the elastoplastic postbuckling response of moderately thick rectangular plates subjected to in-plane loadings is analyzed by a novel numerical approach. The influence of transverse shear deformation is taken into account via the first-order shear deformation theory (FSDT). Also, the elastoplastic behavior is captured based on two theories of plasticity including the incremental theory (IT) (with the Prandtl-Reuss constitutive relations) and the deformation theory (DT) (with the Hencky constitutive relation). Moreover, it is assumed that the material of plate obeys the Ramberg-Osgood (RO) elastoplastic stress-strain relation. First, the matrix formulations of strain rates and constitutive relations are derived. In the next step, according to Hamilton's principle, the weak form of governing equations is derived which is then directly discretized using the variational differential quadrature (VDQ) technique. The discretization process is performed by accurate matrix derivative and integral operators of VDQ. Plates with various boundary conditions under uniaxial and equibiaxial compressions are considered. It is first indicated that the present results are in excellent agreement with the analytical solutions existing in the open literature. Thereafter, the influences of geometrical properties, boundary conditions, elastic modulus-to-nominal yield stress ratio and value of power c in the RO relation on the elastoplastic postbuckling paths of plates are studied. Furthermore, several comparisons are made between the predictions of IT and DT.
       
  • Simulation of liquid jet primary breakup in a supersonic crossflow under
           adaptive mesh refinement framework
    • Abstract: Publication date: Available online 20 May 2019Source: Aerospace Science and TechnologyAuthor(s): Nan Liu, Zhenguo Wang, Mingbo Sun, Ralf Deiterding, Hongbo Wang Compressible two-phase flows were simulated based on the five-equation model under the Adaptive Mesh Refinement (AMR) framework to balance the requirements between space resolution and computational cost. And the simulation system was established in an open source software AMROC (Adaptive Mesh Refinement Object-oriented C++). A combination of Godunov method and wave propagation method was introduced to integrate numerical methods with the AMR algorithm. High speed and high liquid-gas density ratio are two main challenges in the simulation of liquid jet in a supersonic crossflow. To enhance the robustness of the simulation system, a MOON-type positivity preserving method was adopted in the development of the codes. Based on the system mentioned above, a liquid jet in a Mach 1.5 supersonic crossflow was simulated as the standard case to study the primary breakup process in the near field. The simulation captured the column and surface breakup which were the results of the development of the unstable waves in two directions respectively. The instabilities causing the surface breakup were found to be generated in the transonic region initially. Crossflow of a higher Mach number (Ma 1.8) was found being able to augment the instable waves along the injection direction and increase the number of instabilities responsible for the surface breakup. While there was no obvious enhancement of the penetration in the condition of periodic injection, extra unstable waves were imposed on both of windward and leeward liquid surface. The introduced unstable waves had an improvement on the column and surface breakup.
       
  • Rapid design approach for U-bend of a turbine serpentine cooling passage
    • Abstract: Publication date: Available online 16 May 2019Source: Aerospace Science and TechnologyAuthor(s): Changhee Kim, Changmin Son The goal of this study is to propose a rapid design approach for designing a minimum pressure loss U-bend in a turbine internal cooling passage using topology optimization. The total pressure loss for the flow in a bend region is a critical design parameter, as it augments the pressure required at the inlet of the cooling passage, resulting in a lower thermal efficiency. However, no design rules exist for generating a U-bend. The proposed rapid design approach can be applied as a 3D U-bend geometry creation tool. The minimization of the total pressure loss is achieved by means of topology optimization method that uses a continuous adjoint approach with steepest-descent method. The design space is considered as a porous medium with variable porosity. An incompressible steady-state flow (low-fidelity simulation) is considered at a Reynolds number of 100,000 based on the bulk inlet velocity at the domain inlet. 2D and 3D U-bend configurations are produced by the proposed rapid design approach in a few hours. Optimized geometries are automatically achieved around the bend region, leading to a reduced pressure loss.High-fidelity computational simulations are carried out on rapidly designed 2D and 3D U-bend configurations to demonstrate the proposed rapid design approach. The computational studies are performed by solving the compressible, turbulent steady-state Reynolds-averaged Navier-Stokes (RANS) equations with two turbulence models: Spalart-Allmaras model and Shear Stress Transport (SST) model. The high-fidelity simulations predict a relative improvement of 26.8% in total pressure drop with respect to the baseline configuration, mainly due to the reduction of the flow separation region along the inner side of the bend. The results indicate that rapid design approach can come up with design concepts for minimizing pressure loss around U-bend passages in the incompressible flow regime. This approach could be also applied to other fluid systems.
       
  • Robust tracking control of aero-engine rotor speed based on switched LPV
           model
    • Abstract: Publication date: Available online 16 May 2019Source: Aerospace Science and TechnologyAuthor(s): Tong-Jian Liu, Xian Du, Xi-Ming Sun, Hanz Richter, Fei Zhu Thrust control in aero-engines is achieved indirectly due to the lack of thrust sensing technologies. A related variable must be chosen for the control, typically high or low pressure rotor speeds. In this paper, an H∞ controller is designed for the rotor speed tracking in aero-engines. First, a small deviation linear model at steady state is obtained from the experimental data. Then, sets of small-deviation linear models are used to construct a linear parameter-varying (LPV) model with gain-scheduled parameters that capture the nonlinearity of the aero-engine dynamics. The LPV model is then converted to a switched convex polytopic form with hysteresis switching logic. Next, a theoretical sufficiency criterion is provided to guarantee H∞ performance based on linear matrix inequalities (LMIs). Relevant theoretical results are applied to prove stability when switching between subsystems. Simulation results are given to show the validity of the proposed design method, where the proposed the strategy with hysteresis switching logic can reduce the computational cost and avoid false switching due to disturbances.
       
  • Robust Adaptive Fuzzy Sliding Mode Control of Nonlinear Uncertain MIMO
           Fluttering FGP Plate Based on Feedback Linearization
    • Abstract: Publication date: Available online 15 May 2019Source: Aerospace Science and TechnologyAuthor(s): Mousa Rezaee, Reza Jahangiri, Rasoul Shabani In this study using an adaptive fuzzy sliding mode control (AFSMC) scheme, the robust stabilization of multi-input-multi-output (MIMO) nonlinear aero-elastic fluttering of the Functionally Graded Piezoelectric (FGP) plate in the presence of mismatched time-varying uncertainties have been investigated. It is assumed that the aerodynamic load is modeled by the first order piston theory and the piezoelectric patches are assumed to be bonded to the top and bottom surfaces of the plate in order to produce the controlling bending moment excitations. Using the airy stress function and applying the Hamilton's principle the governing coupled partial differential equations of motion are derived. Then considering the immovable simply supported edges boundary conditions and employing the aero-elastic multi-mode interactions and applying the Galerkin's method, the nonlinear coupled partial differential equations of motion are reduced to nonlinear ordinary differential equations in time. Then using the full state input-output feedback linearization technique, the nonlinear dynamics of the model is linearized and transformed into the multiple decupled single input-single output uncertain subsystems. In order to overcome the chattering phenomenon arises due to the sliding mode control (SMC) discontinuous inputs, a hybrid adaptive fuzzy sliding mode control technique is utilized to approximate the discontinuous synthetic control inputs. It is showed that considering the physical input limitations, the designed AFSMC control system, effectively suppress the fluttering motions in presence of the bounded external inaccuracies and it prevents the unwanted chattering of the subsystems inputs.
       
  • Adaptive constrained backstepping controller with prescribed performance
           methodology for carrier-based UAV
    • Abstract: Publication date: Available online 15 May 2019Source: Aerospace Science and TechnologyAuthor(s): Yang Zhang, Sheng-hai Wang, Bin Chang, Wen-hai Wu A novel adaptive constrained backstepping control with prescribed performance methodology for carrier-based unmanned aerial vehicle (UAV) in the presence of uncertainties, input constraints and unknown external disturbances is presented in this paper. This controller can guarantee the compensation tracking errors of the carrier-based UAV with prescribed performance including the steady and transient performance. Firstly, A new transformed system based on the compensation tracking errors, not the traditional tracking errors, is designed. Secondly, to deal with the UAV input constraints, the constrained command filters is introduced and the auxiliary dynamic is designed to eliminate the effect of input saturation. Thirdly, the prescribed performance methodology is introduced and the transient performance of the compensation tracking error is analyzed. It is proved that the proposed controller guarantees that all the signals of the closed-loop system are bounded by using Lyapunov method. Finally, the 6-DOF nonlinear carrier-based UAV model is used to demonstrate the effectiveness of the proposed control law. The simulation results show that the proposed controller is able to provide accurate tracking and satisfy the prescribed performance with unknown aerodynamic parameters, input constraints and external disturbances environment.
       
  • Reconstruction and analysis of non-premixed turbulent swirl flames based
           on kHz-rate multi-angular endoscopic volumetric tomography
    • Abstract: Publication date: Available online 14 May 2019Source: Aerospace Science and TechnologyAuthor(s): Hecong Liu, Jianan Zhao, Chongyuan Shui, Weiwei Cai The development of laser and sensor technologies have provided unprecedented opportunities for the extended applications of volumetric tomography. The recent progresses in computed tomography of chemiluminescence (CTC) have facilitated the understanding of turbulent flows and combustion instability. However, the current demonstrations of CTC can only provide either an instantaneous measurement with a good number of projections to achieve a good spatial resolution or time-resolved measurements (kHz-rate) but with a reduced number of projections which may cause a failure in resolving small details of the flames. In this work, we aim to develop a time-resolved endoscopic CTC system with 17 projections to achieve both good spatial and temporal resolutions. A new method was proposed here to calibrate projections that cover a field of view larger than 180 degrees. The system was then applied to a non-premixed turbulent swirl flame to reconstruct its time-resolved 3D structures. The experimental studies have shown that when only nine projections were used, parts of the flame structures would be lost. To fully recover the flame structures, a minimum of 16 projections should be used. Proper orthogonal decomposition and dynamical mode decomposition were then applied to analyze the time serious of 3D structures of a turbulent swirl flame.
       
  • Surrogate models for the prediction of the aerodynamic performance of
           exhaust systems
    • Abstract: Publication date: Available online 14 May 2019Source: Aerospace Science and TechnologyAuthor(s): Giorgio Giangaspero, David MacManus, Ioannis Goulos The aerodynamic performance of the exhaust system is becoming more important in the design of engines for civil aircraft applications. To increase propulsive efficiency and reduce specific fuel consumption, it is expected that future engines will operate with higher bypass ratios, lower fan pressure ratios and lower specific thrust. At these operating conditions, the net thrust and the specific fuel consumption are more sensitive to losses in the exhaust. Thus the performance of the exhaust needs to be accurately assessed as early as possible during the design process. This research investigates low-order models for the prediction of the performance of separate-jet exhaust systems, as a function of the free-stream Mach number, the fan nozzle pressure ratio and the extraction ratio (fan to core pressure ratio). In the current practice the two nozzles are typically considered in isolation and the performance is modelled as a function of their pressure ratio. It is shown that the additional degrees of freedom have a substantial impact on the metrics describing the performance of the exhaust system. These models can be employed at a preliminary design stage coupled with engine performance models, which require as input the characteristics of the exhaust system. Two engines, which are representative of current and future large turbofan architectures are studied. The low-order models investigated, generalized Kriging and radial basis functions, are constructed based on data obtained with computational fluid dynamics simulations. The data represents the characteristics of the exhaust of each engine, and they are provided for the first time for a wide operational envelope. The influence on accuracy of the type of surragate model and its settings have been quantified. Furthermore, the trade-off between the accuracy of the model and the number of samples has been identified. It is found that the exhaust performance metrics can be modelled using a low-order model with sufficient accuracy. Recommendations on the best settings of the model are also provided.
       
  • Isogeometric analysis of in-plane functionally graded porous microplates
           using modified couple stress theory
    • Abstract: Publication date: Available online 13 May 2019Source: Aerospace Science and TechnologyAuthor(s): Amir Farzam, Behrooz Hassani This paper examines bending, buckling and free vibration behaviors of in-plane functionally graded (FG) porous microplates by means of isogeometric analysis (IGA) and modified couple stress theory (MCST). A hyperbolic shear deformation theory is used, which does not need a shear correction factor. To take into account size-dependent effect, the MCST is employed to analyze functionally graded porous microplates. The IGA meets continuous requirement by using B-Spline or Non-Uniform Rational B-Spline (NURBS) functions. Various types of material distributions are assumed not only through plate thickness, but also in-plane material distributions. The effect of porosity on results is studied, while this parameter has not been paid attention yet for the analysis of in-plane and through-thickness functionally graded (FG) microplates. Furthermore, the effect of other parameters on the behaviors of microplates is investigated by several numerical problems. These parameters include boundary conditions, FG power index and material length scale parameter l.
       
  • Enhanced bondline thickness analysis for non-rigid airframe structural
           assemblies
    • Abstract: Publication date: Available online 13 May 2019Source: Aerospace Science and TechnologyAuthor(s): Pablo Coladas Mato, Philip Webb, Yigeng Xu, Daniel Graham, Andrew Portsmore, Edward Preston Adhesive bonding is a proven alternative to mechanical fasteners for structural assembly, offering lighter and thus more fuel efficient aircraft and cost-effective manufacturing processes. The effective application of bonded structural assemblies is however limited by the tight fit-up requirement, which is with sub-mm tolerance and can be a challenge for the industry to meet considering the variability of current part manufacturing methods and the conservative nature of the conventional tolerance stack-up analysis method. Such a challenge can discourage effective exploitation of bonding technologies, or lead to development of overengineered solutions for assurance. This paper addresses this challenge by presenting an enhanced bondline thickness variation analysis accounting for part deflection of a bonded skin-stringer assembly representing a typical non-rigid airframe structure. A semi-analytical model accounting for unilateral contact and simplified 1D adhesive flow has been developed to predict bondline thickness variation of the assembly under two typical curing conditions: namely autoclave curing and out-of-autoclave curing. The effects of component stiffness and manufacturing variations on bondline thickness are investigated by incorporating stringers of different stiffness, as well as shims of different thicknesses in-between the skin and stringer, in the stringer-skin assembly. A small-scale bonding demonstrator has been built and the physical results are in good agreement with the model prediction. It has been demonstrated that the part deflections need to be accounted for regarding fit-up requirement of bonded non-rigid structural assembly. The semi-analytical model offers more reliable and realistic prediction of bondline thickness when compared to a rigid tolerance stack-up. The analysis method presented can be a major technology enabler for faster, more economical development of the aircraft of the future, as well as of any analogue structures with high aspect ratios where weight savings and fatigue performance may be key objectives.
       
  • Effect of non-spherical particles on nozzle two-phase flow loss in
           nano-iron powder metal fuel motor
    • Abstract: Publication date: Available online 13 May 2019Source: Aerospace Science and TechnologyAuthor(s): Jin-yun Wang, Zai-lin Yang Metal iron powder is a promising new energy source that is of significant practical and research interest for future automotive power systems. However, the shapes of the particles are typically assumed to be spherical or an equivalent sphere when estimating the specific impulse of motors. Such an assumption lacks objectivity and can result in unreasonable estimations of two-phase flow losses. In order to better optimize the design of an engine, this study focuses on the influence of non-spherical particles (such as ellipsoidal and cuboid particles) on the characteristics of nozzle two-phase flow. Models for the governing equations of nozzle two-phase flow are developed to conduct a theoretical study to analyze the combustion properties of iron oxide particles and flow in the nozzle. In addition, experimental studies involving nanometer iron-powder particle combustion and engine thrust measurements are conducted to validate the results obtained from numerical calculations that are conducted using a fourth order Runge–Kutta–Gill method. The results indicate that particle morphology, size, and coagulation content play a significant role in the motor performance and the two-phase flow losses. Specifically, the hypothesis of the ellipsoidal model is in better agreement with the experimental findings compared to the other particle models.
       
  • A proposed design method for supersonic inlet to improve performance
           parameters
    • Abstract: Publication date: Available online 9 May 2019Source: Aerospace Science and TechnologyAuthor(s): M. Farahani, M.M. Mahdavi A new structure for the compression surfaces of a supersonic inlet is proposed which has improved the target performance parameter i.e. total pressure recovery ratio. This idea resulted in development of a new type of supersonic inlet, utilized with four ramps and a cone as the simultaneously compression surfaces. A prototype of the proposed inlet has been designed for a free stream Mach number of 3 and its performance has been evaluated via numerical simulation for both design and off-design conditions. The performance of the newly designed inlet has been compared to the existing experimental data of the inlets equipped with two double-cones. The acquired data are further compared with analytical calculations used for conventional supersonic inlets. The results confirmed the effectiveness of the main idea of the proposed design methodology. The total pressure recovery ratio of the newly designed inlet at its design condition, M∞=3, is calculated to be 76.8%.
       
  • Multi-objective optimisation of short nacelles for high bypass ratio
           engines
    • 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.
       
 
 
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